Proceedings of NASA Conference on Supersonic-Transport Feasibility Studies and Supporting Research

oi SCAT 4 - WING LEADING EDGE FLAPS TYPICAL SECTION

Figure 24

SCAT 4 - FINAL GENERAL ARRANGEMENT

WING AREA-5,150 SQ. FT.

ASPECT RATIO - 1.39 II Figure 25

3 - - * . . e - * - v e e r r '

SCAT 4 - L / D SUMMARY SUBSONIC WIND TUNNEL 0 1 0 2 0 L I D I '?

7.5 i MACH 3 . 0 CRUISE CONFIG.

I I I I I I O 2 0 L/D Figure 26 SCAT 4 - WEIGHT BREAKDOWN WEIGHT(LBS.) % T . O.G.W.

WING 40, 10.6 EMPENNAGE 8,000 2. I 2 5 , 9 0 4 6 . 7 FUSELAGE LANDING GEAR 2 6 , 8 0 0 7 . 0 4 6 , 0 5 0 12.0 PROPULS I 0 N FIXED EQUIPMENT - 2 5 , 2 4 5 6 . 6 WEIGHT EMPTY- 45.0 OPERATING EQUIPMENT - 4 , 7 5 6 I . 2 PAY LOAD 2 6 , I 2 5 6 . 8 ZERO FUEL WEIGHT 2 0 3 , 5 6 0 5 3.0 FUEL/WATER 47.0 GROSS WEIGHT 100.0

Figure 27

SCAT 4 - SIZE OPTIMIZATION

3,200 N. MI. RANGE 2 PSF CLIMB

390 r

I T/ W T.0.G.W.- 1000 LBS.

W / 6 =75 PSF

340 t

330 b, I I I I

4000 4500 5000 5500 6000 WING AREA - SQ. FT.

Figure 28 SCAT 4 - CHARACTERISTICS SUMMARY TAKE - OFF GROSS WEIGHT (LBS.)

384,O 00 FUEL (LBS.)

180,440 RANGE /PAYLOAD (N. MI./LBS.)

3200/26,125 ENGINE THRUST/AIRFLOW (LBS./LBS./SEC.)

44,000/508 SONIC BOOM, CLlMB/CRUISE (PSF.)

2.01 1 . 4 7 TAKE - OFF FIELD LENGTHISPEED (FT./KNOTS) 9500/165 AIQPORT NOISE - TAKE-OFFILAND (PNDB) 107/110 LANDING FIELD LENGTH/SPEED (FT.1 KNOTS) 746 0 / I 3 0 Figure 29 SCAT 4 STAB1 LlTY CHARACTER1 STlCS LONGITUDINAL LATERAL - DIRECTIONAL LOW SPEED APPROACH I .o SAT1 SFACTORY .8 DAMPING DAMPERS OFF CL .6

- I 2

c1/2 i .4 NSATISFACTORY

/*

I I .04 -.04 .4 .8 1.2 I .4

/A\ ROLL T O YAW PARAMETER

CmC.G.

\ve I

Figure 30 SCAT 15 - NASA CONFIGURATION GENERAL ARRANGEMENT SCAT 15 - NASA CONFIGURATION GENERAL ARRANGEMENT \ I \ I 'J Figure 31 i SCAT 1 5 - WING PLANFORM ORIGINAL TRAILING EDGE PROPOSED TRAILING EDGE

- - - - - - - - -

Figure 32 SCAT 15 - TAIL POSITION PROPOSED -----__ - - - - - - - Figure 33 .68 h 271'0 " rn Figure 34 SCAT 15 - L/D SUMMARY 15.1 SUBSONIC 15.0 CONFIGURATION

I

I

0 1 0 20 L/D 7.9 MACH 3.0 CRUISE 7 . 9 W.T.

CONFIG.

I I I 1 0 I O 20 L I D SCAT 15 - W E I G H T BREAKDOWN WEIGHT (LBS.) % T O.G.W.

,375 21.7 WING ,4 55 I .5 EMPENNAGE FUSELAGE 27,894 5. I LANDING GEAR 38,000 6.9 , 60,600 I I .o PROPULSION FIXED EQUIPMENT 27,955 5 . I WEIGHT EMPTY ,279 51.3 OPERATING EQUIPMENT - 4,756 0.9 PAYLOAD 26,I25 4.7 ZERO FUEL WEIGHT 313,160 56.9 FUEL / WATER 43. I GROSS WEIGHT S50,OOU 100.0 Figure 36 SCAT 15 - WlNG DESIGN LOADS MOVABLE FIXED PANEL AFT WING NOT SUBSONIC MANEUVER CRITICAL NOT II SUPERSONIC MANEUVER - CRITICAL II I1 LANDING i II TAXI SCAT 15 - CHARACTERISTICS SUMMARY TAKE - OFF GROSS WEIGHT (LBS.)

FUEL (LBS.1 236,840 RANGEIPAYLOAD (N. MI./LBS.)

126,125 ENGINE THRUST/AIRFLOW (LBS. / LBS./SEC.)

60,000 / 693 SONIC BOOM, CLIMB/CRUISE (PSF.)

2.0/ TAKE - OFF FIELD LENGTH/SPEED (FT./KNOTS) 7000/ I65 AIRPORT NOISE - T. O./LAND (PNDB) 90/95 LANDING FIELD LENGTH /SPEED (FT./KNOTS) 7260/130 SCAT 16 - NASA CONFIGURATION GENERAL ARRANGEMENT Figure .39 SCAT 16 - WING CHANGES A.R.= 1 3 FINAL A. R. = 6.0 Figure 40 SCAT 16 - ENGINE INSTALLATION (PROPOSED) --- ENGINE PIVOTED ENGINE SCAT 16 - FINAL ENGINE ARRANGEMENT i

REMOVAL RANCE ~-i FOR /T

Figure 42 SCAT 16 - HORIZONTAL TAIL LOCATION i ORIGINAL FINAL L O W TAIL T-TAIL 8 COMPENSATING IN ENGINE WAKE NEGATIVE WING DIHEDRAL

Figure 43

SCAT 1 6 - GENERAL ARRANGEMENT WING AREA - 3000 SQ. FT 33'10 Figure 44 SCAT 16-L/D SUMMARY 15.2 SUBSONIC 22.0 WIND TUNNEL CONFIGURATION I I I I I 0 IO 20 LID MACH 3.0 CRUISE I I I I I 0 I O 20 L I D SCAT 16 - WEIGHT BREAKDOWN WING EMPE FUSELAGE 2 8 , 9 5 0 6.8 LANDING GEAR 22,550 5.3 PROPULSION 4 3 , 3 9 5 10.2 FIXED EQUIPMENT - 30, I 0 5 7. I WEIGHT EMPTY 42.9 OPERATING EQUIPMENT - 4 , 7 5 6 1 . 1 6. I PAY LOAD 26, I 2 5 ZERO FUEL WEIGHT 50. I FUEL/ WATER 49.9 GROSS WEIGHT 100.0 Figure 46 SCAT 16- SIZE OPTIMIZATION 3,200 N. MI. RANGE 2 PSF CLIMB - - 7: 0.G.W.- 1000 LBS.

TAKE - OFF REQ. 8 - 420 AIRPORT SPEED REQ \---MIN. G . W .

- L LI I I I I I I I 2200 2400 2600 2800 3000 3200 3400 3600 WING AREA - SQ. FT.

Figure 47

SCAT 16 - CHARACTERISTICS SUMMARY TAKE - OFF GROSS WEIGHT (LBS.) 425,000 FUEL (LBS.) 212,124 RANGE/PAYLOAD (N. MI./LBS.) 3200/ 26,125 ENGINE THRUST/AIRFLOW (LBS./LBS./SEC.) 39,4001455 2.0/ I .36 SONIC BOOM, CLIMB/CRUISE (PSF.)

TAKE - OFF FIELD LENGTH/SPEED (FT./KNOTS) 9 5 0 0 / 1 6 5 AIRPORT NOISE - T 0. /LAND (PNDB) 95/90 LANDING FIELD LENGTHISPEED (FT./KNOTS) 7100/128 Figure 48 SCAT I 7 - NASA CONFIGURATION GENERAL ARRANGEMENT Figure 49 SCAT 17 - WING & ENGINE ARRANGEMENT \

x

Figure 50 SCAT 17 - FOLDING CANARD FINAL FIXED WING TIP Figure 51 Y O SCAT 1 7 AERODYNAMIC CENTER CONTROL AERODYNAMIC CENTER -%C I L I T.O. CRUISE LAND

r

I m -tL -9- 1 5 AERODYNAMIC CENTER 20 25

CENTER OF GRAVITY 1 1 8 1 1 4 1 1 0

STATIC MARGIN I * / " I 5

Figure 52 SCAT 17 - FINAL GENERAL ARRANGEMENT WING AREA - 5.500 SQ. FT.

ASPECT RATIO - 2.16 I I SCAT 17-L/D SUMk !RY 1.7 I I I I I 0 I O LID IO 20 LID Figure 54 S C A T 17 - W E I G H T BREAKDOWN WEIGHT (LBS.) % T. O.G.W.

9 . 2 WING 2.5 EMPENNAGE FU S EL AGE 2 6 , 6 0 0 7.0 LANDING GEAR 19, I 5 4 5.0 4 1 , 3 6 5 10.9 PROPULSION FIXED EQUIPMENT 2 9 , I 5 5 7.7 WEIGHT EMPTY 42.3 OPERATING EQUIPMENT - 4,756 I . 2 26, I 2 5 6.9 PAYLOAD ZERO FUEL WEIGHT 50.4 FUEL/ WATER 49.6 GROSS WEIGHT 100.0

SCAT 1 7 - SIZE OPTIMIZATION

3,200 N. MI. RANGE 2 PSF CLIMB

440 r

T.O.G.W. -

1000 LBS.

400 t

TAKE - OFF REQ.

AIRPORT SPEED REQ.

380[ 360 LMIN. G . W .

340 k, I I I I I

3000 4000 5000 6000 7000 8000 WING AREA - SQ. FT.

Figure 56 SCAT 17 - CHARACTERISTICS SUMMARY TAKE - OFF GROSS WEIGHT (LBS.) 380,000 FUEL (LBS.) 188,209 3200/26,125 RANGE /PAY LOAD (N.M I./LBS.)

ENGINE THRUST/AIRFLOW (LBS./LBS./SEC.) 39,4001455 SONIC BOOM,CLIMB /CRUISE (PSE) 2.011.35 TAKE - OFF FIELD LENGTH /SPEED (FT./KNOTS) 9500/ I65 AIRPORT NOISE -T.O./LAND (PNDB) 107/105 LANDING FIELD LENGTH /SPEED (FTJKNOTS) 72130/ 127 J

SCAT 1 7 - ALUMINUM

ALUMINUM - M 1.9 - 2.4 'I E " & 'I F I' ENGINE ENGINE 35,000 - 50.000 LBS. THRUST Figure 58 SCAT 1 7 AL - SPEED SELECTION 2 .o I .8 1.6 ALUMINUM WEIGHT 1.4 RATIO 1.2 I .o \-TITANIUM 8AL - IMO - IV .8 I I I I I I I 2 .o 2.2 2.4 2.6 2 8 3.0 3.2 CRUISE MACH NO.

SCAT V A L - ENGINE SELECTION

AFTERBURNING DUCT BURNING ENGINE TYPE TURBOJET TURBOFAN NASA E F DESIGNATION TURBINE INLET

I , 2,400 1 2,400

TEMP. , OR

SLS THRUST 1 3 4 , 2 0 0 1 34,100

6.0 5.3 CRUISE SFC

I 1.62 I 1.52

MACH 2.2 I I SFC HOLDING

I 1.29 I .87

AT 1500 FT.

I Figure 60

SCAT I7 AL - ENGINE SIZING

T.O.G.W. = 380,000 LBS.

I 39,400 LBS. I

TAKE OFF TRANS 0 NIC ACCELERATION

i

CRUISE AT 65,000' 00 LBS. ENG "Fit IN V A L Figure 61

SCAT 17AL - CRUISE THRUST

65,000 FT.

P = .0643 LBS/FT3 A I O 100% THRUST INCREASE- THRUST - IOOOLBS. 6 I I I I I 1 2.0 2.2 2.4 2.6 2.8 3.0 MACH NUMBER SCAT 17AL- CRUISE FUEL CONSUMPTION 65,000 FT.

2 . 0 r' 1.8 ' 1.6 SFC

-

1.4 sFc a

t

1 . 0 k, 1 I I

I I I I I 0 4 0 12 16 20 24 28 32 ~ ~ ._

THRUST - 1000 LBS

Figure 63 SCAT 17AL- ENGINE AIRFRAME MATCHING SFC 1.56 MIN. SFC MAX. L I D M = 3.0 ENGINE I' D " SFC 1.30 MIN. SFC MAX. L/D M ~ 2 . 2 ENGINE " F " SONIC BOOM SFC 1.63 I I I I I 30 40 50 60 70 ALT. 1,000 FT.

Figure 64 MACH 2.2 MISSION PROFILES ENGINE " F 'I ALTITUDE- 1000 FT IO ' . 0 RANGE - N. MILES Figure 63 SCAT 17 - RANGE PARAMETERS TOGW = 380,000 7 . 0

-

SFC LID SFC + 4 % + I I Yo 3.0

171,400 LBS i,GHT

- 10.0% SPEED - 265 O / o

A RANGE = - 21%

I)_._

=3.0 k2.2 ~ rlACH NO.

EMPTY WT.

Figure 66

SCAT 17AL - RANGE CAPABILITY

W/S= 69.3 350 I, I I , . I ' I I I I ; L 2400 2600 2800 3000 3200 RANGE - N. MILES Figure 67

SCAT 1 7 AL - GENERAL ARRANGEMENT

A 7 WING AREA - 6.350 SQ. FT ASPECT RATIO - 2.17 Figure 68

SCAT 17AL - CHARACTERISTICS SUMMARY

TAKE - OFF GROSS WEIGHT (LBS.) 440,000 500,000 FUEL (LBS.) 217,690 2 5 7,O 00 RANGE/PAYLOAD (N.MIJLBS.1 26,125 26,125 ENGINE THRUST/AIRFLOW (LBS./LBS./SEC.) 45,600/600 SONIC BOOM, CLIMB/CRUISE (PSF.) 2.0 TAKE - OFF FIELD LENGTH/SPEED (FT./KNOTS) 9500/165 9500/165 AIRPORT NOISE - T. 0. / LAND (PNDB) 107/105 107/105 LANDING FIELD LENGTHISPEED (FT./KNOTS) 7200/127 7200/ I27 Figure 69 STATIC LONGITUDINAL STABILITY lg - LEVEL FLIGHT SCAT 4 SCAT 1 5 TAKE -OFF SCAT 1 6 CRUISE M83.0 I SCAT 4 SCAT 1 7 I I I I 0 - 10

- 20 - 30

STATIC MARGIN - yo MAC

Figure 70 PITCH STABILITY IN CRUISE M = 3.0 SCAT I6 SCAT 1 7 2.5g MANEUVER ANGLE OF

ATTACK -

LEVEL DEGREES FLIGHT lg LEVEL FLIGHT L .08 .04 0 - . 0 4 -08 -04 0 m.04 -.08 '' C.G.

'' C.G.

Figure

LATERAL - DIRECTIONAL HANDLING QUALITIES

MIL F - 8 7 8 5 (ASG) REQUIREMENTS APPROACH CRUISE A T I SFACTO RY

I I I SATISFACTORY

DAMPING 3 I c1/2 2 I 0 .4 .8 1.2 1.4 0/V, ROLL TOYAW PARAMETER T DAMPER ON .I DAMPER OFF Figure 72 LONGITUDINAL HANDLING QUALlTl ES MIL F - 8 7 8 5 (ASG) REQUIREMENTS APPROACH CRUISE I

L ELECTRA-I

.I .2 .4 1 . 0

n DAMPER ON JSHORT PERIOD DAMPING RATIO

a DAMPER OFF

i M = 3.0 CRUISE ENGINE FAILURE SCAT 1 6 SCAT 1 7 ;CRITICAL ENGINE CRITICAL ENGINE

-

- c-

DAMPER! )AMPER5 DAMPERS D AMPERS OFF ON

OFF I ON

A02 CHANGE IN ANGLE OF ATTACK 3.4 - 5 I .o .2 AFTER 3 SEC.

(DEG.)

'J SIDE SLIP I .6 11.2 1.3 4.4 B i b l A X ANGLE (DEG.)

BANK ANGLE - 48

- 3 I - I I - 2

p , AFTER 3 SEC. (DEG.

Figure 74 SPEED STABILITY SCAT 1 6 .2c SCAT 1 7

-y

.I5

-rr/

T/ W .IO 'APP .Of c

-

1 1 1 1 1 ?O 140 160 180 200 20 140 160 180 200

APPROACH SPEED - KNOTS

Figure 75 MISSION COMPARISON SCAT 1 6 SCAT 1 7 3 2 0 0 N. MI. @ M=3,0

- I - - 182,550 158,900 BLOCK FUEL

RESERVE FUEL 2 9 , 6 0 0 I 2 9 , 3 0 0

BLOCK FUEL/MACH NO. 31,800/0.8 32,650/0.9 RESERVE FUEL 18,300 2 0 , 7 5 0 Figure 76 DIRECT OPERATING COST SUPERSONIC

r SCAT 1 6

I 3200 N. MI.

r SCAT ~17 I 1 2 o / o

I SCAT 1 6 I

N. MI.

IO00 SCAT 1 7 I O 0 1 0 SUBSONIC 600 N. MI.

SCAT 1 7 1 3 O / o SCAT 16 200 N. MI.

SCAT 1 7 Figure 77 OFF - DESIGN OPERATION 3200 N. MI. MISSION LAND WITH FULL RESERVES

SCAT 1 6 I SCAT 1 7

HOT - DAY RANGE 2670 RANGE TO ALTERNATE AFTER: MID POINT CABIN DECOMPRESSION I200 I060 MID POINT ENGINE FAILURE I280 I270

Figure 7 8

SONIC BOOM SENSITIVITY 3200 N.MI. RANGE 44c 42C C 10% TRANSIENT THRUST AUG.

40C D GROUND REFLECTIVITY

T0.G.W. -

FACTOR = 1 . 7 1000 LBS 38C

-

I I I I I I I I .6 1 . 8 2.0 2.2 2.4 2.6 2.8 CLIMB AP-PSF Figure 79 , I' , P , , I , RESEARCH AREAS AERODYNAMICS PITCH - UP OF SWEPT WING PLANFORMS SUBSONIC L/D, MAXIMUM LIFT, AND LATERAL CONTROL TRANSONIC L/D CAMBER AND TWIST ON LOW ASPECT RNIO WINGS DYNAMIC STABILITY DERIVATIVES 0 HANDLING QUALITY CRITERIA BY SIMULATOR STUDIES Figure 80 RESEARCH AREAS STRUCTURES 0 INTERFERENCE EFFECTS ERODYNAMIC HEATING 0 S IMULATOR STUDIES TO ESTABLISH STRUCTURAL CRITERIA 0 METHODS FOR ACCELERATED PR O OF AND FATIGUE TEST 0 CORROSION PROTECTION - FOR TITANIUM PROPULSION 0 LONG LIFE HIGH TURBINE TEMPERATURE ENGINES r 0 TURBOFAN DUCT AND AFTERBURNER COMBUSTION 0 HIGH STABILITY INLETS AND LIGHNVEIGHT EXHAUST 0 TAKEOFF AND LANDING NOISE Figure 8 1 4. SUMMARY OF BOEING,S,CAT FEASIBILITY STUDIES

* * - -' 8 4 * * . *

* - *

_ -

c By Lloyd T. Goodmanson, Wiliimi*IT. Hamilton, /

-

and Maynard L. Pennell J-JY

~ C P S

Boeing Coppany, Airplane Division 3

1 09

2 w

/ SUMMARY A f e a s i b i l i t y study of four configuration concepts f o r the super - sonic commercial air transport (SCAT) has been conducted by The Boeing Company under contract t o the Langley Research Center of the National Aeronautics and Space Administration.

The four concepts, i d e n t i f i e d as SCAT' s 4, 15, 16, and 17, were developed by t h e NASA.

The airplane configurations which have been developed t o meet spe - c i f i c design objectives are described and comparative performance is presented.

The technical aspects of the study i n t h e areas of struc - t u r a l design, propulsion, and aerodynamics are b r i e f l y reviewed.

The major conclusions a r e a s follows: The SCAT 16 and SCAT 17 concepts a r e feasible approaches t o the design of a supersonic transport.

The concept of variable - sweep wings, exemplified by SCAT 16, offers significant improvements i n operational f l e x i b i l i t y as compared with the conventional fixed - wing concept of SCAT 17.

SCAT'S 4 and 15 exhibit superior adrodynamic efficiency.

However, major deficiencies must be corrected through additional research i n order t o make these concepts feasible.

Titanium structure and advanced engine technology must be employed i n t h e supersonic transport i n order t o achieve reasonable economic per - Continued research i n materials, structures, and propulsion, formance.

a s well as i n aerodynamics, i s necessary i n order t o develop an airplane competitive with present subsonic jets.

A n aluminum airframe, designed Ih t o cruise a t Mach numbers near 2.0 and according t o t h e SCAT guidelines, has a gross weight too large t o be economically feasible.

I n addition, it could not meet the sonic - boom overpressure limitations.

Allowable sonic - boom overpressures w i l l have a major impact on the performance capability of t h e supersonic transport.

Based on the tech - n i c a l s t a t e of t h e art employed i n the SCAT study, an allowable over - pressure a r b i t r a r i l y selected too low could cause any configuration con - cept t o be not feasible.

% - 4 d ~ - qoe

i I .

INTRODUCTION The Boeing Coqany under N A S A Contract NAS~-2580 conducted a study of four airplane concepts developed by N A S A aeronautical research t o determine the f e a s i b i l i t y of t h e i r use as supersonic commercial air transports. The contract i n i t i a l l y required a cruise speed of Mach 3.0.

Later it w a s amended t o investigate one of these f o u r configurations a t a cruise speed i n the Mach 2.0 range. The SCAT study began on February 4, 1963, and has continued f o r 7 - months.

This paper i s divided i n t o three p a r t s t o summarize the study findings.

Part I defines t h e changes and improvements made i n the four p r i - mary SCAT configurations during t h e program. This p a r t a l s o summarizes the comparative performance of the SCAT models including the aluminum SCAT.

P a r t I1 summarizes the c r i t e r i a which governed the SCAT program.

This p a r t includes a review of specific technologies a s they were applied t o the various SCAT models.

Part I11 summarizes t h e findings and various trade - offs developed from the SCAT program, including economic considerations, and provides general conclusions which a r e intended t o help define future research i n t h e f i e l d of supersonic transports.

GUIDELINES The SCAT study guidelines on which a l l configurations were based a r e given i n t a b l e I. A l l four of t h e principal SCAT configurations were designed f o r Mach 3.0 cruise. The aluminum SCAT (SCAT 16AL) w a s t o be designed f o r cruise a t o r above Mach 2.0. A range of 3,200 nau - t i c a l miles with a payload of 26,125 pounds was specified with take - off, landing, and approach characteristics comparable t o those of present - day subsonic jets. L i m i t s on sonic - boom overpressures were established at 2 pounds per square foot ( p s f ) i n climb and at 1 . 3 psf f o r cruise.

f Aircraft noise a t the a i r p o r t and i n t h e neighboring community w a s t o be equivalent t o t h a t of present - day j e t s .

Figure 1 indicates the design features t h a t a r e common t o a l l of the Boeing configurations. The design speed f o r t h e four o r i g i n a l con - cepts w a s Mach 3.0.

The cruise speed selection of Mach 2.0 f o r SCAT 16~1, w i l l . be discussed later. All of the Mach 5.0 designs a r e based on titanium t o minimize s t r u c t u r a l weight. Engine - cycle studies indicated I that engine D, an advanced technology turbofan, provided the best per - a l l four Mach 3.0 SCAT'S. Engine F, which w a s chosen f o r formance f o r S C A T 16m, i s also a turbofan based on a s l i g h t increase beyond current A l l four original S C A T concepts and the later alumi - engine technology.

num SCAT a r e based on the S C A T guidelines.

SYMBOLS AND ABBREVIATIONS l i f t coefficient maximum l i f t coefficient pitching - moment coefficient directional - stability parameter compression yield strength ultimate t e n s i l e strength lift - drag r a t i o Mach number cruise speed dive speed m a x i m operating speed sonic - boom overpressure, psf maximum sonic - boom overpressure, psf i T thrust, l b thickness - chord r a t i o airspeed, knots

v

airplane weight, l b

w

angle of attack, deg a A change sweepback angle, deg * L E direct operating costs, cent/seat-mile Doc gross w e i g h t GW OWE operating, weight empty payload PL specific fuel consllmption SFC TO take - off I. CONFIGURATION D E T E L O P M E N T AND P E R F O R M A N C E

SCAT 4

S C A T 4 is a fixed arrow - wing concept based on the blending of the

engine nacelles with the wing t o achieve favorable aerodynamic inter - ference and t o minimize the s t r u c t u r a l weight through wing bending r e l i e f . (See f i g s . 2 and 3 . )

The configuration a t the end of 4 months of sCudy differed from

the original definition as a r e s u l t of the following changes: (1) To achieve l i g h t e r w e i g h t , it w a s decided t o pass the f u l l depth of the wing box below the f l o o r rather than carry the wing bending moment through the fuselage frames. This made it possible t o reduce the body diameter from 117 t o 113 inches. Also, the body length w a s increased t o s h i f t the crew and passenger compartments farward relative J t o the wing, thereby improving the airplane balance.

The sweepback ( 2 ) The wing planform experienced several changes.

angle w a s reduced from 750 t o 72.8O, as recommended by the NASA early i n the program, t o provide a b e t t e r compromise f o r Mach 3.0 cruise speeds. The wing t r a i l i n g edge a t the root was shifted aft with reduced This change reduced sweep t o allow the structural box t o be moved aft.

the wing weight as a result of the shorter struchral span.

( 3 ) Wing - tip elevons were t e s t e d and recommended by the N A S A when design studies revealed t h e d i f f i c u l t i e s involved i n providing adequate strength and s t i f f n e s s f o r an all - moving wing - tip control.

( 4 ) The trapezoidal, two - dimensional engine i n l e t s were replaced with lighter axisymmetric i n l e t s .

( 5 ) A delta planform w a s used f o r the horizontal t a i l instead of the higher aspect r a t i o , arrow configuration. This change provided more separation from t h e engine efflux and noise f i e l d s and increased t h e s t i f f n e s s f o r a given weight.

S C A T 15

For SCAT 15 ( f i g s . 4 and 5 ) a 7 5 O arrow wing blends i n t o the fuse -

lage t o achieve high aerodynamic efficiency. The engines a r e mounted as close as possible t o the wing undersurface t o develop favorable aero - dynamic interference and t o provide bending r e l i e f . Vertical f i n s a r e located on the wing t i p s i n a good flow f i e l d throughout the f l i g h t envelope. The t a i l location also yields an improvement i n t h e wing span loading by means of t h e end - plate e f f e c t s . An auxi$iary wing, pivoted from the main wing structure, produces higher subsoni,c l i f t - d r a g ratios and C L , ~ than can be obtained from the basic planform.

During the 4 months devoted t o studying SCAT 13 many possible

changes were investigated, but only a few were incorporatedA The major changes were a s follows: (1) The fuselage was lengthened t o provide volume f o r approximately 50 percent of the t o t a l required f u e l load and t o improve airplane balance.

( 2 ) A body - mounted horizontal t a i l was added t o improve s t a b i l i t y and control. However, it w a s found i n subsequent t e s t s t h a t t h e s i z e chosen was inadequate t o eliminate high - speed i n s t a b i l i t i e s of the f l e x i b l e airplane. The correct t a i l size w a s not incorporated because of termination of t h e SCAT 15 study.

After consideration of t h e data presented i n the midterm report a t

t h e end of 4 months of study, SCAT'S 4 and 15 were discontinued. The

r e s u l t s presented therefore r e f l e c t 4 month's work only.

-

The SCAT 16 concept ( f i g s . 6 and 7) i s based on a variable - sweep The objective of this approach i s t o y i e l d maximum off - design wing.

performance without s a c r i f i c i n g cruise capabilities. In contrast with SCAT'S 4, 15, and 17, the SCAT 16 configuration w a s i n i t i a l l y proposed with only three engines.

The SCAT 16 airframe definition changed i n t h e following areas during the study: (1) The midwing arrangement w a s changed t o a low - wing design because of the w e i g h t penalties associated with carrying wing bending moments through t h e fiselage frames.

( 2 ) The body cross section w a s increased because of the volumetric requirements of the main landing gear and fuselage f u e l tanks.

( 3 ) The wing pivot was shifted t o the fuselage side because of expected savings i n wing weight and t h e p o s s i b i l i t y f o r a l l e v i a t i n g wings - forward pitch - up.

( 4 ) The wing twist and camber were modified t o improve the aero - dynamic efficiency.

( 5 ) The wings - forward aspect r a t i o was reduced from 13 t o 7.5 t o improve wing weight.

(6) Four engines w e r e used instead of three t o provide the same degree of multiengine capability a s i n the other configurations.

(7) The horizontal t a i l was changed from an arrow t o a d e l t a plan - form t o obtain more s t i f f n e s s f o r a given weight and t o improve clear - ance f o r engine efflux and noise f i e l d s .

Figure 8 shows the general arrangement of t h e SCAT 16 f u e l tanks, principal spars and bulkheads, pivot and landing - gear locations, and the f l l g h t control surfaces. Some of the f u e l i s carried below t h e main cabin floor. This i s dictated by the limited volume available i n \ the wings and the airplane balance requirements. Wing f u e l i s carried between t h e front and rear spars i n t h e swinging section. The wing pivot i s located w e l l inboard a t the side of the body t o provide the maximum possible wing - span increase from the swep - tbackto the swept - forward position. stowed i n the lower lobe of The main landing gear is the body. A s the gear i s extended, it swings f m a r d and outboard t o the down position. Both leading - edge and inboard trailing - edge f l a p s a r e employed, along with spoilers and outboard ailerons. The horizontal A s t a b i l i z e r i s a variable - incidence type with trailing - edge flaps.

two - segment s p l i t rudder and spoiler system provides the yaw control.

S C A T 17 The SCAT 17 configuration ( f i g s . 9 and 10) incor$orates a wing having 5 9 O swept leading edge and zero - taper - ratio d e l t a planform.

axisymmetric power - plant i n s t a l l a t i o n s a r e mounted from the lower Four wing surface. A s i n i t i a l l y conceived, SCAT 17 had a canard,'fold-down wing t i p s , and a fold - down horizontal t a i l which w a s t o provide reduced drag and increased directional s t a b i l i t y during cruise.

t r i m During the study program the following changes were incorporated: (1) The fuselage w a s shortened, and the cross section was reduced This change resulted i n reduced t o accommodate four - abreast seating.

body wetted area and weight.

( 2 ) The wing thickness - chord r a t i o w a s increased from 2 t o 9 per-

cent and t h e folding wing t i p s were eliminated.

( 3 ) The canard and folding t a i l were replaced with a conventional high - aspect - ratio t a i l because of a longitudinal balance problem, inad - equate canard control power, and excessive drag.

(4) The power plants were moved forward on the wing t o improve overall balance and t o increase the t a i l moment arm.

Figure 1 1 i l l u s t r a t e s the general arrangement of t h e SCAT 17. Fuel i s carried both i n the exposed wing area between t h e front and r e a r spars and i n t h e inboard center section. A dry bay i s provided i n the wing t o accommodate the engine - driven a i r c r a f t accessories. The main gear i s stowed i n the wing adjacent t o the body. It swings a f t and The landing - gear s t r u t l i e s below the outboard t o the down position.

wing f u e l tank and i s covered by an aerodynamic f a i r i n g when retracted.

Full - span leading - edge flaps, inboard trailing - edge flaps, spoilers, The horizontal ")and an outboard flaperon are employed on t h e main wing.

s t a b i l i z e r i s a variable - incidence type with trailing - edge flaps. A two - segment s p l i t rudder and spoiler system provides the yaw control.

After the midterm contract review, t h e basic contract w a s amended t o include the study of an aluminum variable - sweep - wing configuration ( f i g . 12). The aluminum SCAT 16 w a s based on t h e original SCAT 16, - chord with appropriate modifications i n wing sweepback and thickness t / c = 4.75 percent r a t i o permitted by the lower cruise speed A m = TO0,

(

a t the pivot and 3 percent a t the wing tip) . The body cross section pro -

vides f o r four-abqeast seating with a large lower lobe f o r containment of f u e l and the landing gear.

The airplane has a much higher g r o s s weight than i t s titanium counterpart because of the lower strength - weight r a t i o of aluminum. The

SCAT 1 6 ~ ~ received less detailed study than SCAT'S 16 and 17 because of

the shorter t i m e available f o r refinement t o optimize t h e configuration.

Comparative Performance b

Data from the f i v e SCAT configurations a r e compared i n figure 13.

SCAT'S 4 and 15 are shown i n the shaded area t o indicate t h a t studies

Although SCAT 4

on them were discontinued a f t e r the midterm review.

s t a b i l i t y and control problems, par - showed the l i g h t e s t gross weight, t i c u l a r l y a t low speeds, l e d t o termination of the study. SCAT 15 showed the highest weight a t 490,000 pounds, and t h i s high weight was i t s prin - c i p a l drawback. Although the SCAT 15 exhibited excellent aerodynamic t h e weight of the dual overlapping wings l e d t o discon - characteristics, tinuance of the study. When b u i l t of titanium, SCAT'S 16 and 17 have i s about approximately the same weights. The aluminum SCAT, however, 90,000 pounds heavier.

The other f a c t o r s presented i n figure 13 - weight r a t i o , L/D,

- contribute

engine specific f u e l consumption and sonic - boom overpressure significantly t o the differences i n g r o s s weight f o r the f i v e airplanes.

The delta - wing SCAT 17 has t h e most favorable weight r a t i o . SCAT 16, on s l i g h t l y b e t t e r the other hand, has a somewhat higher l i f t - d r a g r a t i o , specific fuel consumption because of the engine match, and b e t t e r f u e l reserves. Both airplanes meet t h e sonic - boom overpressure l i m i t of 2 psf i n climb.

SCAT 1 6 ~ ~ has a f a i r l y high L/D primarily because of i t s lower SCAT 1 6 ~ ~ a l s o shows f a i r l y good specific f u e l consump- cruise speed.

t i o n and i t s weight r a t i o benefits from i t s large size.

The weight breakdown f o r four airplanes i s shown as t o t a l weights i n figure 1 4 and as percentages of t o t a l weight i n figure 15.

I n each chart, t h e weights of the three f i n a l airplanes a r e compared with the weight of the long - range Boeing 707 Intercontinental. A s charted i n figure 15, the supersonic transports show only about one - half of the payload - gross - weight r a t i o of present - day subsonic j e t s . Other sig - n i f i c a n t f a c t s shown i n the weight breakdowns are as follows: (1) O f the three f i n a l SCAT configurations, SCAT 17 has t h e l i g h t e s t s t r u c t u r a l weight.

(2) SCAT 17 uses somewhat more f u e l than e i t h e r SCAT 16 o r SCAT 1 6 ~ ~ because of i t s lower range factor.

( 3 ) A l l supersonic transports have a greater percentage of t h e i r gross weight i n f u e l than has t h e Boeing 707 Intercontinental.

Sonic - boom overpressure limitations have a fundamental e f f e c t on the s i z i n g of large supersonic airplanes. Figure 16 r e l a t e s airplane gross weight t o m a x i m u m sonic - boom overpressure i n climb f o r three of t h e SCAT configurations. Each curve on these figures represents a family of airplanes whose wing areas and power - plant sizes have been properly matched f o r minimum gross weight a t t h e sonic - boom overpres - sures shown. Sizes of t h e f i n a l SCAT 16 and SCAT 17, chosen t o meet t h e required climb sonic - boom overpressure of 2.0 psf, are indicated by the symbols. SCAT 1 6 ~ ~ designed f o r a 3,200 - nautical - mile range cannot meet the 2.0 psf requirement a t any gross weight; therefore, the air - plane was sized a t t h e knee of t h e curve.

Climb schedules f o r SCAT'S 16 and 17 sized f o r 2.0 psf overpres -

sure are shown i n figure 17. Note t h a t SCAT 17 must accelerate through

low supersonic speeds a t a higher a l t i t u d e than SCAT 16 t o meet the overpressure l i m i t . Both airplanes designed without regard t o sonic - boom overpressure would use the lower a l t i t u d e , maximum range - climb I n this case, e i t h e r would have a greater range f o r the same schedule.

gross weight o r could be smaller i f matched t o the same range. A s i s evident i n figure 16, both SCAT'S 16 and 17 a r e about 10 percent larger than would be necessary i f a sonic - boom overpressure of 2.5 psf were acceptable.

The mission p r o f i l e s f o r SCAT'S 16 and 17 a r e plotted on the upper portion of figure 18. Resulting sonic - boom overpressures throughout t h e 3,200 - nautical - mile mission are shown on t h e lower portion of the f i g - ure. Because of t h e higher required acceleration a l t i t u d e and smaller excess t h r u s t , SCAT 17 creates a 2.0 psf boom over a greater distance than does SCAT 16. The i n i t i a l sonic - boom overpressure i n cruise i s 1 . 5 psf or below f o r both models. However, SCAT 17 mast climb t o an a l t i t u d e above i t s optimum cruise a l t i t u d e f o r maximum range before it can start cruise with a 1.5 psf boom. Consequently, SCAT 17 will f l y a t a constant a l t i t u d e during the first portion of t h e cruise during which time the sonic - boom overpressure w i l l gradually decrease as gross weight decreases with f u e l burnoff. A t t h e end of cruise, SCAT 17 has a sonic - boom overpressure of about 1.2 psf.

Cruise boom overpressures f o r SCAT 16 start at about 1.5 psf and reduce t o about 1.1 psf a t t h e end of a climbing cruise.

During the early portion of t h e deceleration and descent, the boom overpressure w i l l increase s l i g h t l y .

Descent 1 0 1 overpressures a r e held below 1.5 psf as shown by decelerating t o Mach 1.0 before the airplane has descended t o 50,000 f e e t .

Sonic - boom overpressure r e s t r i c t i o n s have been shown i n t h e SCAT study t o have a major impact on t h e performance capability of t h e super - sonic transport. Since presently available data do not indicate a sig - n i f i c a n t difference i n public reaction t o sonic booms a t overpressures near 2 psf, it i s apparent that extreme caution should be used i f arbi - t r a r y r e s t r i c t i o n s are imposed. It a l s o should be recognized t h a t the tolerable overpressures probably w i l l be higher on over - water flights than on domestic routes. Therefore, f l e x i b i l i t y i n flight planning can r e s u l t i n significant decreases i n operating costs.

Well - controlled f l i g h t tests t o determine more accurately public

)

reaction t o sonic booms are necessary i n order t o resolve the con - f l i c t i n g requirements between performance and public acceptance of the supersonic transport. The United States possesses a unique capability t o conduct sonic - boom f l i g h t t e s t s , and it i s urged t h a t such a program be completed before the supersonic transport i s committed t o production.

11. CRITERIA AND TECHNOLOGY FGVIEW Structures The operational speed - altitude envelope of SCAT vehicles i s formed by a composite of boundaries representing aerodynamic, propulsion, and s t r u c t u r a l design limitations. This envelope varies somewhat with each SCAT configuration but t h e l i m i t s presented i n figure 19 a r e generally representative of those f o r supersonic transports. The maximum oper - a t i n g speed M M O i s selected t o give maximum performance and t o meet sonic - boom requirements. Margin i s provided f o r unintentional over - speeding t o dive speed MD a s required i n C i v i l A i r Regulations (CAR)

*

4b. The upper boundary i s the thrust - limited airplane ceiling. Tern- perature usually r e s t r i c t s the maximum Mach number, although f o r the basic SCAT study t h e maxim speed w a s specified t o be Mach 3.0. Other limitations may be imposed i n varying degrees by buffet, f l u t t e r , gust, aerodynamic s t a b i l i t y , and the engine. Sound s t r u c t u r a l design requires examination of the complete f l i g h t envelope f o r c r i t i c a l conditions.

The s t r u c t u r a l design i s affected by a l l f l i g h t regimes as w e l l a s landing and ground handling loads. Low - speed maneuvers and transonic gusts and maneuvers usually give the highest mechanical loadings. Cer - t a i n supersonic cruise and overspeed maneuvers have the added e f f e c t s

*

Anon.: Airplane Airworthiness; Transport Categories. C i v i l Air Regulations P a r t 4b, FAA, Jan. 7, 1963.

> of high temperature and thermal s t r e s s e s which make these maneuvers s l i g h t l y more c r i t i c a l i n certain areas.

One of the first major decisions i n developing a s t r u c t u r a l design i s t h e selection of the material t o be used. A group of seven of t h e most promising alloys including t i t a n i u m , s t a i n l e s s s t e e l s , and super - alloys w e r e considered i n the basic Mach 3.0 studies. Several heat - r e s i s t a n t aluminum alloys were a l s o considered f o r the SCAT 1 6 ~ ~ inves - tigations. Figures 20 and 21 graphically compare the r e l a t i v e weight efficiency of the prime candidates AM 335 s t a i n l e s s s t e e l , Ti-8AL-No-lV titanium alloy, and 202LT81 aluminum a l l o y a t room and cruise tempera - tures. I n t h i s chart the longer the bar, t h e b e t t e r the material i s f o r The titanium a l l o y i s the most e f f i c i e n t the p a r t i c u l a r application.

! material f o r components designed by s t a t i c load requirements, a s i n d i - cated by the tension and compression efficiencies ( f i g . 20). For the design of j o i n t s and components, such as t h e fuselage pressure cabin where fracture toughness and fatigue resistance a r e important, the t i t a - nium a l l o y again i s superior f o r t h e Mach 3.0 application ( f i g . 21).

Based on these considerations a s w e l l as corrosion resistance, ease of fabrication, and metallurgical s t a b i l i t y , Ti-8AL-No-lV was selected a s the primary s t r u c t u r a l material f o r the basic Mach 3.0 SCAT studies.

The heat - resistant 202bT81 aluminum alloy was judged best f o r most p r i - mary applications on the SCAT 1 6 ~ ~ .

studies were made t o deter - After selection of the basic material, mine t h e optimum cover - panel s t r u c t u r a l arrangements f o r the Mach 3.0 wing. Several arrangements t h a t appeared f e a s i b l e were evaluated on t h e basis of both w e i g h t and i n s t a l l e d cost, with consideration given t o j o i n t s and i n t e r n a l support structure.

Since a l l wings studied during the contract contained f u e l within the wing s t r u c t u r a l box, the insulating characteristics of each surface were considered i n d e t a i l .

Figure 22 shows f o u r of the panel types studied f o r the wing surface covers. A l l four types meet the f u e l insulation requirements. The single skin s t r i n g e r cover panel with backing board and bladder f u e l c e l l walls proved t o be the least expensive on an i n s t a l l e d cost basis.

The t i t a n i u m honeycomb sandwich panel was t h e most expensive and con - siderable fabrication development would be required t o bring it t o a production s t a t u s . The weight comparison of these four panels i s shown 4 1 i n figure 23 f o r t h e SCAT 16 wing structure. Based on a compromise between w e i g h t and i n s t a l l e d cost, the double skin stringer with a non - s t r u c t u r a l inner radiation shield w a s chosen f o r the upper surface and t h e double s t r u c t u r a l skin was selected f o r the lower surface of SCAT 16, as indicated by t h e shaded bars. For SCAT 17 the honeycomb struc - ture r e s u l t s i n the l i g h t e s t upper surface because of the lower wing loads ( f i g . 24).

However, the double skin s t r i n g e r arrangements were again chosen f o r the f i n a l design because of the high costs and devel - opment required f o r t h e honeycomb cover panel. After t h e midterm review when the contract was amended t o c a l l f o r an evaluation of an i aluminum airplane, a s t r u c t u r a l weight study w a s conducted on an alumi - num SCAT 16 sized t o accomplish the SCAT payload - range mission a t a Mach number of 2.0. Figure 25 shows the weight penalty associated with higher cruise Mach numbers. The s t r u c t u r a l weight penalty i s due t o the degradation of material properties by extended exposure t o the higher design temperatures. The optimum cruise speed was not precisely deter - mined because of the abbreviated contract period; however, it appears t o be near a Mach number of 2.2 when both t h e combined weight and range factor e f f e c t s a r e considered. Structural weight i s influenced strongly by configuration and t h e basic design guidelines and, therefore, the optimum speed may be expected t o vary somewhat a s these factors change.

Propulsion The operating conditions which significantly a f f e c t the engine- airplane matched performance a r e supersonic cruise, subsonic cruise, transonic acceleration, take - off, and holding, a s indicated i n f i g - ure 26. The optimum engine cycle c h a r a c t e r i s t i c s and size, i n s t a l l e d weight, and drag f o r any one of the operating conditions i s generally not optimum f o r any other condition, nor f o r t h e airplane a s a whole.

Therefore, the r e l a t i v e merits of given engines or engine design vari - ables cannot be assessed by a comparison of the propulsion - system data alone. Instead the t o t a l performance capability of p a r t i c u l a r airplane - engine configurations must be compared.

Both turbofan and turbojet engines were selected and data were supplied by the N A S A f o r t h i s study. Various degrees of augmentation were considered, from nonafterburning t o f u l l afterburning f o r the turbojets and from duct burning t o mixed - flow burning f o r the turbofans ( t a b l e 11). Low - flow versions of the turbofans also were studied because of the reduction i n transonic spillage drag and the improved engine - airflow nozzle matching a t cruise speed.

Two levels of engine design technology were included f o r the Mach 3.0 application. Engines A and B are representative of a small advancement over current m i l i t a r y engine weight and turbine i n l e t tem - perature designs. Engines C and D represent a significant advance i n these technology areas, p a r t i c u l a r l y i n the turbine i n l e t temperature. < i For the SCAT 1 6 ~ ~ , engines E and F represent the same general design l e v e l a s engines A and B.

The engine e f f e c t s on overall airplane capability f o r SCAT'S 16

and 17 are presented i n figures 27 and 28, respectively. The take - off

gross weight required t o meet the design payload a d range with a m a x i - mum sonic - boom overpressure of 2.0 psf i s significantly l e s s with t h e advanced turbofan engine D. Additional benefits a r e indicated i n t h e A s used i n t h i s evaluation it has a low - flow version of engine D.

reduced airflow schedule a t Mach numbers between 2.5 and 3.0. Engine D (low flow) w a s t h e only engine providing t h e capability of meeting the sonic - boom overpressure objective of 2.0 psf f o r SCAT 17 ( f i g . 28).

The superiority of engine D i s basically due t o t h e high turbine tem - perature and the accompanying higher cycle pressure r a t i o . These fac - t o r s give t h i s turbofan engine a thrust - weight r a t i o comparable t o t h a t of t h e high - temperature turbojet (engine C ) with good cycle efficiencies a t a l l Mach numbers. I n p a r t i c u l a r , the specific f u e l consumption a t t h e required thrusts f o r subsonic cruise and holding a r e significantly less t h m those f o r t h e other engines and t h e reserve requirements are, therefore, lower. These advantages, together with lower f u e l consump - t i o n during acceleration, more than o f f s e t t h e s l i g h t supersonic cruise SFC disadvantage of engine D a s compared with engine C .

Airplane sizing studies c a l l f o r the integration of a number of The performance trade i n factors a s detailed i n figures 29 t o 31.

matching engine D t o t h e SCAT 16 airframe i s shown i n figure 29 as con - tours of constant range capability a t a gross weight of 430,000 pounds.

These contours are p l o t t e d as a function of t h e t o t a l i n s t a l l e d engine airflow (sea - level s t a t i c r a t i n g ) and airplane wing area. Performance analyses such as t h i s one must be accomplished f o r a number of gross I n a l l cases, the airplanes weights when actually s i z i n g an airplane.

a r e accelerated along a speed - altitude schedule t o cruise which r e s u l t s The variations i n i n a maximum sonic - boom overpressure of 2.0 psf.

range on t h i s p l o t are caused by t h e trades between wing area and wing weight and between matched engine performance and engine weight.

Increasing the wing area improves t h e l i f t - d r a g r a t i o but r e s u l t s i n a greater weight empty, with l e s s f u e l available within t h e fixed gross weight. Similarly, l a r g e r engines provide greater acceleration capa - b i l i t y during the c r i t i c a l transonic acceleration condition and match Such an the airplane a t a lower cruising specific f u e l consumption.

engine offers improved f u e l u t i l i z a t i o n but a l s o increases the weight The variation of these i t e m s causes the gradients i n range empty.

capability shown. If range were the only consideration, a s l i g h t l y smaller airplane would f u l f i l l t h e SCAT requirements because the peak range capability i s approximately 3,250 nautical miles.

Take - off limitations have been added t o t h e range contours i n f i g - The CAR second - segment climb gradient of 0.03 f i x e s a lower ure 30.

l i m i t t o allowable engine t h r u s t since the l i f t - d r a g r a t i o i n the climb configuration i s r e l a t i v e l y insensitive t o wing area. The requirement f o r a take - off velocity no greater than 165 knots fixes a lower l i m i t t o wing area. For a given airplane configuration, such as SCAT 16, these two limits are actually determined by the h i g h - l i f t devices incorporated i n t o the wing. The balance between take - off l i f t coeffi - cient and climbout l i f t - d r a g r a t i o i s achieved by optimizing t h e f l a p setting. The variable - sweep - wing geometry of SCAT 16 eliminates t h e need f o r a penalizing compromise from the optimum combination of engine s i z e and wing area f o r range t o achieve good low - speed characteristics.

The requirement f o r a take - off f i e l d length no greater thqn 9,500 f e e t i s a l s o shown, although the take - off speed limitation i s more severe Inasmuch a s these values, 9,500 f e e t than t h e field - length requirement.

and 165 knots, t y p i c a l l y occur simultaneously f o r long - range subsonic j e t s , t h e greater thrust - weight r a t i o available with maximum t h r u s t augmentation on SCAT 16 provides the much shorter f i e l d length a t com - parable take - off speeds.

The landing requirements and t h e transonic t h r u s t margin have been The requirement f o r a touchdown speed added t o the data i n figure 31.

of 130 knots i s more r e s t r i c t i v e than the requirement f o r a 7,000 - foot landing - field length. The sizing of SCAT 16 was determined by t h e some - ) what a r b i t r a r y choice of a minimum transonic t h r u s t margin equal t o 0.3 a reasonable of the drag on a standard day. This margin was selected a s A somewhat compromise between standard - day and hot - day performance.

smaller airplane would r e s u l t i f l e s s t h r u s t margin were used (approxi - mately 0.2 corresponds t o the peak range); however, a greater range l o s s would occur on a hot day.

A similar sizing chart i s shown f o r SCAT 17 i n figure 32. The SCAT 17 characteristics which a r e fundamentally different from those of SCAT 16 r e s u l t i n more c r i t i c a l low - speed performance requirements.

The take - off performance required a significant compromise i n t h e air - plane sizing. The lower l i f t - d r a g r a t i o i n climbout, primarily due t o the lower span and l e s s effective f l a p s , sized the engines t o meet second - segment climb requirements. The lower l i f t capability of the wing and f l a p geometry has caused t h e take - off speed requirement t o increase t h e wing area considerably over the optimum area f o r range.

A l e s s conservative value of 0.22 f o r transonic t h r u s t margin was The hot - day range loss i s accepted t o meet t h e take - off requirements.

I n figures 31 accordingly somewhat greater than t h a t f o r SCAT 16.

and 32 the shaded area represents t h e available combinations of wing area and engine airflow which meet the other limitations.

Engine noise a t the a i r p o r t and i n the surrounding community rep - resents another l i m i t on propulsion - system selection. SCAT 16 alone meets the community noise goal of 112 PNdb maximum a t take - off on a a hot day. This low noise l e v e l i s due primarily t o the higher a l t i t u d e achieved over t h e community and the lower power s e t t i n g required t o The curves i n maintain the selected climb path a f t e r powen cutback.

figure 33 a r e f o r a hot day, since t h e bulk of c o m i t y noise com - p l a i n t s occur under these conditions. A point i s also shown on the p l o t f o r SCAT 16 on a standard day, which indicates a reduction of about 6 PNdb. This reduction i s t y p i c a l f o r the other airplanes as well.

, A Take - off along the p r o f i l e s p l o t t e d i n figure 33 i s made with full augmentation. For SCAT 17, f u l l augmentation i s required t o m e e t t h e second - segment climb gradient with one engine out. For SCAT 16 and SCAT L ~ A L , the take - off and climbout requirements could be met with lower power s e t t i n g s but the cornunity noise l e v e l would increase as a r e s u l t of the f l a t t e r climb gradient.

Noise i n t h e community under the flight path during landing, as shown i n figure 34, ranges from 7 t o 10 PNdb higher than t h e take - off noise shown i n f i g u r e 33. The higher landing noise i s due primarily t o However, the low a l t i t u d e over t h e commmity on the 3 O glide slope.

because of t h i s low a l t i t u d e , t h e high noise region does not spread t o t h e side as much as i n take - off. The take - off and landing noise levels would be equal a t a lateral distance of 1,000 t o 2,000 feet t o e i t h e r side of the f l i g h t path, and the landing noise would be less beyond this distance.

The fan whine contributes significantly t o t h e landing noise. If the fan whine contribution could be deleted through compressor, duct burner l i n e r , wd/or inlet design and operation, the landing noise f o r SCAT 16 would be reduced about 6 PNdb, as shown by the point a t t h e boundary of the c o m i t y . Reductions of t h e same magnitude would occur f o r t h e other airplanes as w e l l . Any improvements i n the airplane l i f t - drag r a t i o i n t h e approach configuration would a l s o reduce t h e landing A 10 - percent reduc - noise by reducing the approach power requirements.

t i o n i n approach power w o u l d reduce the noise l e v e l approximately 3 PNdb.

The continuing model noise test program at The Boeing Conpany began before the introduction of subsonic j e t s , and the data i n f i g - ures 33 and 34 are based on t h e r e s u l t s obtained from t h i s program.

During this study, models duplicating the SCAT engine nozzle velocities, temperatures, and stream - tube area r a t i o s were t e s t e d and these r e s u l t s were correlated with full - scale - engine t e s t r e s u l t s , Among the models tested w a s a coannular design t o simulate t h e duct burning fan e x i t nozzle.

Handling Qualities The Boeing SCAT studies included extensive analyses of s t a b i l i t y and handling q u a l i t i e s . The airplane configurations - have been modified t o ensure t h a t a l l basic reqgirements have been m e t .

Figures 35 t o 39 have been included t o i l l u s t r a t e the highlights of this work.

Investigations show t h a t both SCAT'S 16 and 17 possess s a t i s f a c - t o r y s t a b i l i t y characteristics throughout the f l i g h t envelope., For each airplane t a i l sizing, wing planform selection, and center - of - gravity location provide a configuration without pitch - up problems and with adequate s t a t i c and maneuver margins a t a l l f l i g h t conditions.

The e f f e c t s of a e r o e l a s t i c i t y on longitudinal s t a b i l i t y dCm/dCL a r e larger f o r SCAT 16 than f o r SCAT 17 as shown i n figure 35. The combined aerodynamic and e l a s t i c characteristics of the wing, fuselage, and t a i l cause the aerodynamic center t o move 23 percent forward f o r SCAT 16 ( t o t a l airplane) at supersonic A s a r e s u l t , t a i l sizing MD.

f o r minimum s t a b i l i t y margins i s as c r i t i c a l a t the supersonic condi - t i o n s (wings a f t ) as it i s a t the subsonic conditions (wings forward).

For SCAT 17, the aeroelastic aerodynamic - center movement i s only about

one - fourth t h e magnitude of t h a t f o r SCAT 16; consequently, the low - speed conditions d i c t a t e t a i l sizing and aerodynamic balance (a.c.

r e l a t i v e t o aft c.g. l i m i t ) .

The degree of s t a t i c directional s t a b i l i t y required f o r t h e SCAT'S

16 and 17 and the most c r i t i c a l f l i g h t condition which must be accounted

f o r have not been completely resolved a t t h i s t i m e . The t y p i c a l dete - ' r i o r a t i o n i n d i r e c t i o n a l s t a b i l i t y with angle of a t t a c k a t supersonic Mach numbers between 2.0 and 3.0 requires t h a t the uppermost boundary of the f l i g h t envelope (maximum a condition) be accurately determined and t h a t t h e minimum s t a b i l i t y l e v e l acceptable a t these conditions be established. Once an upper f l i g h t boundary i s defined, rigorous dynam - i c s studies w i l l be required t o establish v e r t i c a l - t a i l and ventral - area requirements t o assure f l i g h t safety during any conceivable maneuver.

Figure 36 r e l a t e s the deterioration i n directional s t a b i l i t y with angle of a t t a c k and s t r u c t u r a l a e r o e l a s t i c i t y f o r SCAT'S 16 and 17 a t t h e Mach 3.0 cruise condition. For reference, t h e angle of a t t a c k a t 2.5g i s shown f o r each airplane a t cruise conditions. T a i l s and ven - trals were a r b i t r a r i l y sized i n t h i s study t o provide a minimum CnP l e v e l of 0.0005 a t t h i s condition. Since higher angles of a t t a c k may be encountered a t off - design conditions, f'urther study i n t h i s area appears necessary.

Analytical studies and f l i g h t simulator evaluations have shown t h a t SCAT'S 16 and 17 possess acceptable low - speed longitudinal dynamic characteristics without s t a b i l i t y augmentation. Figure 37 indicates t h a t , i n t h e landing - approach condition, t h e longitudinal short - period o s c i l l a t i o n i s w e l l damped and s l i g h t l y longer i n period than t h a t of current j e t transports.

A s a r e s u l t , only a small amount of damping augmentation i s required t o optimize low - speed haadling qualities.

A t supersonic cruise, however, t h e longitudinal shQrt-period o s c i l l a t i o n i s very poorly damped f o r both configurations. Handling q u a l i t i e s f o r this f l i g h t condition appear t o be unacceptable without damping augmen - t a t i o n . Further f l i g h t simulator research w i l l be needed i n t h i s area.

The characteristic nature of the lateral - directional dynamic oscil - l a t i o n o r Dutch roll i s different for the SCAT 16 and SCAT 17 config - urations, as shown graphically i n figure 38, which includes the military A t low speeds, the SCAT 16 Dutch requirements specified i n MIL-F-8785.* roll oscillation i s b e t t e r damped with less r o l l i n g than t h a t of SCAT 17.

The b e t t e r damping of SCAT 16 is due t o t h e unsweeping of the wing. A t supersonic cruise where the S C A T 16 wing i s swept back 750, the Dutch roll oscillation i s less damped and with more rolling than f o r SCAT 17.

The augmentation requirements appear t o be moderate a t Mach 3.0 cruise and small at low speed f o r e i t h e r airplane. Present transports gen - e r a l l y f a l l i n the area near the SCAT 16 low - speed point. The f i n a l evaluation of handling qualities and augmentation f o r t h e supersonic transport w i l l require inclusion of the p i l o t i n the loop.

\ Positive speed s t a b i l i t y during landing approach w a s specified as The approach speeds a desirable characteristic f o r the S C A T airplanes.

which correspond t o a touchdown speed of 130 knots a t design landing Scat 16 has satisfactory speed s t a b i l i t y , weight are shown i n figure 39.

S C A T 17 i s s l i g h t l y which corresponds t o t h a t of present - day airplanes.

unstable but is well within the criterion of d(T/W) = - 0.0012 per knot dv which i s considered acceptable by NASA p i l o t s . Perhaps a more s i g n i f i - For the conditions shown, SCAT 16 cant factor i s approach attitude.

approaches a t a body a t t i t u d e of 4.30 with a speed of 135 knots and SCAT 17 approaches a t an a t t i t u d e of loo with a speed of 136 knots.

The relatively high a t t i t u d e of SCAT 17 a t touchdown probably would be considered more undesirable by p i l o t s than the s l i g h t speed i n s t a b i l i t y .

111. RESEARCH POTENTIAL AND CONCLUSIONS The Boeing SCAT f e a s i b i l i t y studies were very much l i k e the usual preliminary design investigations i n the early stages of the develop - The process is roughly described i n f i g - ment of any new airplane.

ure 4 0 . So many complex variables are interrelated t h a t both the i n i - t i a l choice of configuration and continued refinement require many rep - * This design cycle approach has been typ - e t i t i o n s of the design cycle.

i c a l of airplane development programs such a s those f o r the Boeing B - 47, the design cycle w a s During the SCAT studies, B - p , 707, and 727.

repeated several t i m e s but without the benefit of adequate wind - tunnel The design t e s t i n g of improvements each t i m e through t h e cycle.

*Anon. : Flying Qualities of Piloted Airplanes. Military Specifi - cation mL-F-8785( ASG) , Sept . 1, 1954; Amendment - 2, Oct . 17, 1955.

evolution does not stop once the o r i g i n a l configuration i s selected.

The design cycle i s repeated over and over t o optimize the t o t a l program.

Operating Cost Comparisons The effects of this continuing design refinement are c l e a r l y evi - dent i n t h e seat - mile economics of the Boeing 707, as shown i n f i g - The band of seat - mile costs f o r t h e 707 shows t h e improvement ure 41.

which has been achieved over a period of several years. The top of the band represents t h e d i r e c t operating costs (DOC) of the i n i t i a l offerings t o t h e a i r l i n e s . The bottom of t h e band represents the b e t t e r economies of t h e advanced 707 Intercontinental available today. A s shown i n f i g -

ure 41 on a r e l a t i v e scale, the DOC of SCAT'S 16 and 17 a r e not as good

as those of the i n i t i a l 707. However, neither of these SCAT designs has been optimized t o the same extent as was t h e 707 a t the time it was f i r s t offered t o the a i r l i n e s . The r e l a t i v e seat - mile costs of SCAT

1 6 ~ ~ shown f o r Mach 2.0 cruise a r e even higher than those for SCAT'S 16

and 17.

If additional e f f o r t had been applied and the design cycle repeated more often than t h e abbreviated period f o r the study of SCAT 1 6 ~ ~ allowed, seat - mile costs could probably have been reduced somewhat.

However, since the seat - mile costs of the SCAT 1 6 ~ ~ are so much higher

than those of SCAT'S 16 and 17, it i s doubtful i f any significant

improvement i n seat - mile costs would be achieved. It i s probable t h a t the DOC of the aluminum airplane w i l l never be competitive with those of the Mach 3.0 airplanes. The studies made t o date indicate t h e need f o r substantial improvements i n the supersonic transport i f it i s t o be com - p e t i t i v e from an economic standpoint with subsonic j e t s already avail - able. There appears t o be q l e opportunity f o r improvement of the supersonic transport i n the months and years ahead by repeating the design cycle.

However, improved designs do not evolve simply a s a r e s u l t of the passage of time. Intensive e f f o r t i s required and the research achievements i n a number of technology areas must be integrated i n t o an optimized configuration. Both basic research and the engineering application of many new ideas must be pursued vigorously i f desired r e s u l t s are t o be achieved.

I Aerodynamic Trades A s a p a r t of the SCAT studies, Boeing determined t h e impact of p o t e n t i a l improvements i n several important technology areas on the factors t h a t a f f e c t operating costs.

Figure 42 charts the influence of aerodynamic improvements on t h e gross weight of a supersonic transport * % 3 s t o perform a given job. For example, an"im&fefient. i n cruise L/D f o r SCAT 16 from 6.6 t o 7.0 would permit a reduction i n gross weight of around 30,000 pounds or approximately 7 percent. This i s a technology area where intensive e f f o r t i s j u s t i f i e d because of t h e impact of both gross weight and aerodynamic efficiency on operating costs. Improving aerodynamic efficiency w i l l require both t i m e and e f f o r t , but an L/D approaching 8 appears possible by 1970 t o 1975. If the L/D f o r SCAT 16 could be improved t o 8, the DOC would be reduced 13 t o 20 percent from the l e v e l shown i n figure 41.

Weight Trades Another obvious area f o r continued research i s i n the reduction of the weight empty of the airplane. Applying e f f o r t t o a t t a i n i n g minimum s t r u c t u r a l weight has been t r a d i t i o n a l i n airplane design, but the job becomes more important as it becomes more d i f f i c u l t . Structural - weight minimization has been a d i f f i c u l t technical problem i n t h e SCAT study, but the payoff i n lower gross weights makes further progress very desirable (see f i g . 43). For airplanes performing the same mission, 1,000 pounds saved i n operating weight empty can reduce the take - off gross weight by 5,000 t o 6,000 pounds. Alternatively, weight empty reductions can be converted d i r e c t l y i n t o allowable payload with accom - panying dramatic reductions i n seat - mile costs.

Propulsion System Trades The operating efficiency of an airplane i s always very sensitive t o t h e s t a t e of the a r t of the propulsion system. T h i s c r i t i c a l r e l a -

tionship i s p a r t i c u l a r l y t r u e f o r the supersonic transport. Figure 44

shows t h e e f f e c t of an improvement i n engine design achieved principally through the use of higher turbine i n l e t temperature. Instead of a SCAT 16 with a gross weight of 430,000 pounds with engine D (turbine i n l e t temperature of 2700° R ) the gross weight would have been about 490,000 pounds with engine B (turbine temperature of 2400° R ) . A similar e f f e c t i s apparent f o r SCAT 17.

e Figure 45 shows t h e e f f e c t on gross weight of small variations i n i three factors affecting propulsion - system i n s t a l l a t i o n efficiency. The r e l a t i v e l y large e f f e c t s point up t h e need t o achieve an optimum i n s t a l - l a t i o n as well as an optimum engine. Looking ahead, the guarantees which an airplane manufacturer must eventually make t o the a i r l i n e cus - tomer will permit only very minor tolerance variations i n factors such as i n s t a l l a t i o n weight, nozzle efficiency, and i n l e t recovery. Com - pounding of a number of individual tolerances could have a tremendous effect on the f i n a l gross weight of the supersonic transport.

- -

a D J i * 3.

Sonic - Boom Effects Other important factors r e l a t e d t o t h e freedom of choice which the airframe designer w i l l have i n the eventual selection of the supersonic transport configuration were brought out by the SCAT studies. For example, a i r c r a f t designers have depended f o r many years on increasing payloads f o r improving t h e economics of new airplanes. I f t h e same sonic - boom overpressures a r e assumed as those which were used i n the SCAT studies, the amount of exploitation of s i z e e f f e c t s f o r obtaining improved economics will be severely limited. The left - hand chart i n figure 46 indicates the e f f e c t of range and payload variations on DOC when t h e 2 psf sonic - boom - overpressure constraint i s observed. If the sonic - boom overpressure were increased t o 2.3 psf, reductions i n seat - mile costs, such a s those shown i n the right - hand chart, a r e immediately

!

available. Thus, i f it were possible, f o r example, t o operate a t higher sonic - boom overpressures over ocean areas, improved economics on i n t e r - continental routes should r e s u l t . The reduced width of the shaded area on the right - hand chart i s a l s o significant. Direct operating costs show a substantially smaller increase with increasing design range a t an overpressure of 2.5 psf than a t 2 psf.

C onel u s i on s The p r a c t i c a b i l i t y of a supersonic transport has been confirmed by the SCAT studies. This conclusion must, of course, be confirmed by construction and t e s t i n g of f u l l - s c a l e hardware before t h e t r u e answer i s known. However, a t present it appears t h a t a t l e a s t two of the f i v e configurations studied a r e quite feasible when judged by the standards established f o r the SCAT studies. There i s no clear - cut choice between these two configurations although the variable - sweep - wing airplane appears t o have b e t t e r " off - design point " characteristics than t h e delta - wing configuration. Furthermore, it i s e n t i r e l y possible t h a t one of the other configurations can be made more a t t r a c t i v e i f s u f f i c i e n t inventiveness i s applied.

Certainly attempts t o develop configurations which a r e basically superior t o e i t h e r SCAT 16 or SCAT 17 a r e t o be encouraged.

The economic f e a s i b i l i t y of the SCAT configurations i s not as c l e a r , cut as the technical f e a s i b i l i t y . The present configurations with air - frame and engine costs generally considered probable f o r t h i s type of airplane w i l l not have as good DOC as t h e present - day highly refined subsonic j e t s . Further research and optimization of specific designs should do much t o reduce the gap i n d i r e c t operating costs which pres - ently e x i s t s . During t h e i n i t i a l operation of the supersonic transport, t h e higher u n i t costs would undoubtedly be more'khan o f f s e t by higher load factors and/or higher fares. However, i n order t o be economically sound f o r future operation, the ultimate product must be improved.

I n view of the marginal acceptability of any of t h e configurations studied, an attempt must be made t o s e l e c t f o r further intensive work those types of configurations with t h e m a x i m u m improvement potential.

For example, t h e extensive use of titanium i n t h e structure of t h e air - plane, although requiring additional development, appears t o be neces - sary f o r t h i s airplane whether it i s t o be operated a t Mach 2 or a t higher speeds.

Opportunities f o r further aerodynamic refinement and design opti - mization of SCAT 16 appear greater than f o r the other configurations studied and further investments i n understanding and improving this con - figuration are p a r t i c u l a r l y desirable.

f Problems such as t h e sonic boom can likewise have a major influence on the ultimate usef'ulness of t h e supersonic transport. Although the sonic boom i s a very real problem from the standpoint of public reaction, it i s a t l e a s t as great a problem from the standpoint of the designer.

Careful study and research of t h i s phenomenon i s s t i l l very much i n order.

Perhaps t h e most important thing t o bear i n mind i s t h a t September 1963 i s l i t e r a l l y only t h e threshold of the supersonic transport era.

The research and development t h a t has been expended toward the develop - ment of the supersonic transport must be continued and expanded. Aero - dynamics, structures, and propulsion have always been important t o the designer. They a r e even more important today. Industry must work on these fundamentals t o the m a x i m u m extent that i t s resources will permit.

Since these resources a r e limited and since substantial m i l i t a r y devel - opments i n the supersonic airplane area are lacking, the supersonic transport requires d i r e c t government funding f o r continuing and inten - sified research. The supersonic transport i s an important subject, one i n which the stakes will be extremely large. Both the quality and the extent of the government research program must be consistent with t h e scope of the problem. These research e f f o r t s must be extended f o r many years i f the kind of a program which the American public has come t o expect of j o i n t government - industry e f f o r t s i s t o be achieved. Boeing enters the supersonic era confident t h a t a superior U.S. a i r l i n e r can be achieved but fully cognizant of the extent of the e f f o r t which lies ' ahead.

! C A B L E I . - SCAT GUIDELINES

Cruise speed. . . . . . . . . . . . . . . . . . . . . . . . . . Mach 3

(Mach 2 for 16AL) Range . . . . . . . . . . . . . . . . . . . . . . . . . . 3,200 n.mi.

Payload . . . . . . . . . . . . . . . . . . 125 passengers (26,125 lb)

Take - off:

Speed . . . . . . . . . . . . . . . . . . . . . . . . . . 165 knots

Distance . . . . . . . . . . . . . . . . . . . . . . . . . . 9,500 ft

Landing :

Speed . . . . . . . . . . . . . . . . . . . . . . . . . . l3Oknots

Distance . . . . . . . . . . . . . . . . . . . . . . . . . . 7,000 ft

Approach (speed stability desired) . . . . . . . . . . . . . 145 knots

Sonic - boom overpressure:

Climb . . . . . . . . . . . . . . . . . . . . . . . . 2 psf or less

Cruise . . . . . . . . . . . . . . . . . . . . . . . 1 . 5 psf or less

Airport noise . . . . . . . . . . . . . Comparable to present - day jets

TABLE 11.- ENGINE TYPES Mach 3 . 0 Mach 2 . 0 Turbine Turbine Turbine

r Engine

temperature, temperature, temperature ,

2 , 4000 R 2,700° R

2,400' R a (4 ( a 1 A C E Nonafterburning Turbojet Afterburning Afterburning N afterburning B D F ~ Duct burning Duct burning Turbofan Mixed burning Duct burning Low flow Low flow

I

aPresent-day technology.

bAdvanced technology.

DESIGN MACH 3

1 MACH 2

I SPEED 1

BASIC TITANIUM ALUMINUM STRUCTURAL (8AL- lMO -1V) (2024 -T81) MATERIAL POWER ADVANCED TURBO - FAN PLANT (ENGINE D) (ENGINE F) ALL STUDIES BASED ON SCAT GUIDELINES Figure 1 SCAT 4 ORIGINAL CONCEPT Figure 2 SCAT 4 FINAL CONFIGURATION - JUNE AXISYMMETRIC INLETS 96 F 'T / TIP ELEVONS GROSS WEIGHT 380,000 LB WING AREA 5150 FT2

1 - 233FT - 1

Figure 3 SCAT 1 5 ORIGINAL CONCEPT Figure 4 SCAT 1 5 FINAL CONFIGURATION - JUNE BODY LENGTHENED D GROSS WEIGHT 490,000 LB WING AREA 5200 SQ FT Figure 5 SCAT 1 6 ORIGINAL CONCEPT, Figure 6 c SCAT 1 6 , r - 7

WING PLANFORM -/

I 1 TAILCHANGED1

CHANGED i !

GROSS WEIGHT 430,000 LB WING AREA 4 0 0 0 S Q F T - 2 1 2 FT Figure 7 SCAT 1 6 FUEL AND PRIMARY STRUCTURE : Figure 8 1 1 8

S C A T 17 ORIGINAL CONCEPT

n

Z C R O S S SECTION Figure 9 S C A T 17 F I N A L CONF I GURAT I ON NO FOLD

NG TA1L7

9 NON-FOLD IN^^ , BODY CROSS SECT1 ON POWER PLANT MOVED FWD.

NO CANARD \ GROSS WEIGHT 445,000 L B

WING AREA 5,600 F T *

T H I CKNESS I NCR 1 - 2 2 0 ' , - 1 Figure 10 FUEL AND PRIMARY STRUCTURE Figure 11 SCAT 16AL

t ,r-7

WING AREA 4750 FT2 Figure 12 Figure 13 WEIGHT BREAKDOWN A

500 1

G R O S S W E I G H 707 S C A T S C A T SCAT - 320 8 16 1 6 A L Figure 14 ,

WE I GHT BREAKDOWN - PERCENT

A

WEIGHT 4 0 2 0 SCAT SCAT 707 SCAT - 320 B 17 16 I 6 A L Figure 15

SCAT SIZING - SONIC BOOM

I

5 0 0 ’\ S C A T 16AL -\ GROSS WT, \ I 1000 LB 450 ~

- --

-----_____ SCAT 17

4 00 SCAT 16 SCAT 16 3501 I I I I I I I I I I I i I I I 1.5 2.0 2.5 A PMAX 9 PSF Figure 1 6 - A L T ITU DE, 1000 FT \ Figure 17 MISSION PROFILE SONIC BOOM OVERPRESSURES 80 - X I O 3 ALTITUDE, FT 50- RANGE, N.MI.

Figure 1 8 SCAT OPERATIONAL LIMITS ALTITUDE, 1000 FEET 0 1.0 2.0 3.0 MACH NUMBER Figure 19 MATERIAL SELECTION TENSILE COMPRESSION

-

IIII( 500 1000 500 1000 FCY ttu DENSITY DENSITY AM355 STEEL 8AL-IMO-1V TITANIUM 2024-T81 ALUMINUM Figure 20 FAIL - SAFE STRUCTURE JOINTS 500°F TEMP 200°F 70" F 70°F 500 1000 500 1000 FATIGUE STRENGTH FRACTURE STRENGTH DENSITY DENSITY AM355 STEEL 8AL-1 MO-IV TITAN I UM 2024-T8I ALUMINUM Figure 21 WING CONSTRUCTION MACH 3 AIRPLANE RELATIVE UNIT COST BACKING BOARD 8 BLADDER CELL NON - STRUCTURAL

4 INNER SKIN

DOUBLE STRUCTURAL SKIN HONEYCOMB SANDWICH 2 . 0 1 . 0 0 Figure 22

JilEl GH? kOMPARI

SCAT 1 6 WING UPPER LOWER SURFACE SURFACE BACKING BOARD 8 BLADDER CELL NON - STRUCTURAL INNER SKIN DOUBLE STRUCTURAL SKIN HONEYCOMB

1 - 1

SANDWICH

-

0 0.5 1 . 0 0 0.5 1 . 0 1 . 5 Figure 23 WEIGHT COMPARISON SCAT 1 7 WING UPPER LOWER SURFACE SURFACE BACKING BOARD 81 BLADDER CELL NON - STRUCTURAL INNER SKIN DOUBLE STRUCTURAL SKIN HONEYCOMB SANDWICH

-

0 0.5 1 . 0 0 0.5 1.0 1 . 5 Figure 2 4 MACH NUMBER EFFECTS SCAT 16AL STRUCTURAL WEIGHT G W = 520,000 POUNDS - O W E 190- LESS ENGINES, - 1000 I8O POUNDS J I70

- "L MATCHED

AIRPLANE - 2.0 2.2 2.4 CRUISE MACH NUMBER Figure 27 CRITICAL ENGINE DESlGN CONDITIONS ALTITUDE, SUPERSONIC F T CRUISE S F C T 70,000 60,000 50'000 CRUISE TO 40,000 ALTERNATE

1 SFC-

THRUST MARGIN 0 1.0 2 .o 3 . O MACH NUMBER Figure 26 600- - 5 5 0 GROSS - WEIGHT, ENGINE 1000 LB - 4 5 0 B (LOW FLOW)

'L

MATCHED -C (AFTERBURNING) - AIRPLANE -D (BASIC) D (LOW FLOW) I I Figure 2 7 ENGINE S E L E C T I O N SCAT 17 M=3.0 TITANIUM - RANGE=3200 N.MI.

PL =26,125 LB 5 5 0

- ;:; K! F i DE q J & ENGINE

5 0 0 GROSS WEIGHT, B ( L O W FLOW) IOOOLB 450 - C (AFTERBURN I NG) D (HIGH FLOW) - D (LOW FLOW) I I I I 2.0 2 5 3.0 MAX. SONIC BOOM OVERPRESSURE IN CLIMB APhnnx 7 PSF Figure 28 AIRPLANE SIZE ANALYSIS SCAT 1 6 G W = 430,000 LB AP =2.0 PSF 3 5 0 0 ~ ENGINE A'R 3000 FLOW, POUNDS 2 5 0 0 PER SECOND 2 0 0 0

I 5 0 0 t

L I I I I 2000 3000 4000 5 0 0 0 WING AREA, SQUARE FEET Figure 29 AIRPLANE SIZE ANALYSIS SCAT 1 6 G W = 430,000 LB A P = 2.0 P S F

f :605KTS

ENGINE T O 9 5 0 0 F T ~ O O F I O O O F T

I

FLOW, POUNDS2500 PER SECOND

t

2 o o o ~ I 5 0 0 LI I I I 2 000 3000 4000 5000 WING AREA, SQUARE FEET Figure 30 AIRPLANE S I Z E ANALYSIS G W = 430,000 LB S C A T 16

AP= 2.0 P S F

7000 FT LANDING

c

TOUCHDOWN

ENGINE I

- MIN. THRUST - POUNDS 25001 - 9 5 0 0 FT SECOND

PER 2oool

------- -

1500 1

L l I I 1 3000 4000 5000 WING AREA, SQUARE F E E T F i g u r e 31 AIRPLANE SIZE ANALYSIS G W = 445,000 L B SCAT 1 7 A P M A ) ( = ~ . O PSF

3500 t

ENGINE 3000 AIR FLOW, POUNDS PER SECOND 2000 i

1500 c

L I I I I I 3000 4000 5000 6000 7000 WING AREA, SQUARE FEET F i g u r e 32 2 - ALTITUDE, - - - - 1OOOFEET I - NOISE SCAT 1 7 UNDER

FLIGHT SCAT 1 6 - - - - -

PATH, 1 1 0 PNdb STANDARD DAY - I I I I I I I .o 1.5 2.0 2.5 3.0 3.5 4.0 DISTANCE FROM BRAKE RELEASE, MILES LANDING NOISE I30 SCAT 1 7 NOISE UNDER FLIGHT 120 PATH, NO FAN NOISE\ PNdb a END OF RUNWAY SCAT 1 6 V I I I I I I 0 .5 I .o 1 . 5 2 .o 2.5 3.0 DISTANCE FROM RUNWAY THRESHOLD, MILES Figure 34 I AFT C.G. LIMIT SCAT 16

so5 r

.5 PITCHING - . I , .2 -3 .4 I I I I

MOMENT 0 - - - - - -

---.

I \ COEFF ~ F L E X I B L E

I I

- . I O LIFT COEFFICIENT Figure 35 STATIC DIRECTJONAL STABILITY MACH = 3.0 CRUISE SCAT 1 6 SCAT 1 7 A L E = 75 " .003 .002 cnP .OOl

0 5 I O 1 5 20 5 IO ' 1 5 20

BODY ANGLE OF AT ACK, DEGREES Figure 36 L .6 A SCAT 17 SUBSONIC J E T S

-'-

CRUISE - .4 UN DAM P E D.

- N A T U R A L A ' FREQUENCY, C P S 2 - - . I I I I I I I 1 I I Figure 37 LATERAL - DIRECTIONAL DAMPING CHARACTERISTICS NO AUGMENTATION MI L - F - 8785 M l NlMUMS 4 r

I LANDING r AUGMENTED

RECIPROCAL X L I C n n n m n n n u 17 I OF CYCLES TO HALF- AMPLITUDE 4 . 6 8 1.0 1.2 DEG BANK ANGLE J - EQUIVALENT LATERAL VELOCITY FPS Figure 38 SCAT 177 - .4 - .3 THRUST REQUIRED (T/W) - .2 APPROACH SPEED - .I

0 I I I - t I I I 1

100 120 140 160 180 200 220 EQUIVALENT AIRSPEED, KNOTS DESIGN EVOLUTION CONFIGURATION

/GG&i+ ( )-a

DESIGN "'",",",I

u t s i G N

u

I MECHANICAL TUNNEL DESIGN TESTS WEIGHTS TUNNEL 3, ANALYSIS MODELS Figure 40 OPERATI NG COST COMPAR I SON .6Oc L I I I 1 I IO00 2000 3000 RANGE, N. MI.

AERODYNAMIC TRADES CRUISE L/D 6.5 8.0 7 1 0 75 - & T SCAT 1 6 1 7 Figure 42 WEIGHT TRADES 10,000

t

/

t

L I I I I 350 400 450 500 GROSS WEIGHT, 1000 LB Figure 43 PROPULSION SYSTEM TRADES ENGINE ENGINE B ENGINE B 2400 " R - 500,000 GROSS WEIGHT, POUNDS - 450,000 ENGINE D i 400,000 SCAT 1 6 SCAT 1 7

Figure 44

f I ? .

PROPULSION SYSTEM TRADES INSTALLATION 60,000 40,000 SCAT 1 7 A GROSS WEIGHT, LB 20,000 A INSTALLATION A NOZZLE A INLET WEIGHT=IO% EFFICIENCY=I% RECOVERY =2% Figure 45 E C O N O M I C T R A D E S A P = 2 . 0 PSF AP=2.5 PSF

\*

DIRECT OPERATING COST . - 300 400 500 600 300 400 500 600 GROSS WEIGHT, 1000 LB Figure 46 N A S A SUMMARY AND ASSESSMEXT O F FEASIBILITY - STUDY FESULTS

By John G. Lowry d *

- - f = - - - ?

P NASA.Langley Research Center S U M M A R Y The studies have indicated t h a t through the use of an advanced engine technology and the use of an advanced material, titanium, e i t h e r t h e variable - sweep approach o r the fixed - wing approach can meet the , , i required supersonic commercial a i r transport (SCAT) design objectives.

Based on the same design objectives, an aluminum bhch 2.2 airplane did not have the same range, payload, o r sonic - boom capability even with prohibitive increases i n gross weight. Significant research w i l l be required i n the areas of aerodynamic performance, handling qualities, sonic boom, propulsion, and s t r u c t u r a l fabrication before the super - sonic transport w i l l be a success.

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INTROIXJCTION !The supersonic commercial a i r transport f e a s i b i l i t y studies were designed t o supplement and complement related government and industry programs. The N A S A program directed toward establishing a general l e v e l of supersonic transport efficiency studied many different SCAT concepts. Four of the more promising concepts evolved from t h i s program were selected f o r the f e a s i b i l i t y studies.

Before discussing the r e s u l t s of these studies, it would be wcitll t o review t h e purpose of the f e a s i b i l i t y studies and see t h e types of r e s u l t s t h a t were desired.

The main objective of t h e SCAT f e a s i b i l i t y studies was t o have the contractor evaluate and compare the promising design concepts r e l a t i v e t o a l l facets of a successful commercial design. It w a s not expected, nor desired, t h a t tihe contractors a r r i v e a t a f i n a l design of a trans - ' port f o r the free-dorld a i r l i n e s .

A s p a r t of the evaluation and com - parison, the contractors were t o c l a r i f y t h e present technical s t a t e of the art f o r guidance of tlie national program and t o define require - ments f o r future N A S A research. The studies were t o be conducted i n s u f f i c i e n t depth t o assure an adequate degree of confidence i n gross weight and payload - range characteristics.

An assessment of the overall study will be made t o examine the f e a s i b i l i t y of the various concepts, and t h e results of trade studies a r e discussed t o assess the importance of variations i n ground rules on t h e o v e r a l l c h a r a c t e r i s t i c s of a supersonic transport. The dis - cussion i s general i n nature since a discussion of the details is presented i n the contractors' summary reports.

CONCEPT FEASIBILITY

During t h e first 4 months of t h e contract e f f o r t , t h e contractor

studied a l l four of t h e configurations furnished by the NASA, evaluating each with the four d i f f e r e n t engine types furnished with t h e data.

A f t e r a review by t h e NASA, t h e contractors were instructed t o focus t h e i r a t t e n t i o n on SCAT 16 and SCAT 17 and use engines "B," r'C,rt and "D" i $ during t h e remaining p a r t of the study e f f o r t . SCAT 4 and SCAT 15 were dropped from t h e study e f f o r t because there w a s not s u f f i c i e n t t i m e t o correct t h e deficiencies found i n these concepts. A t the t i m e of the review, two aluminum SCAT designs were added t o the study e f f o r t .

Boeing was t o study SCAT 16 and Lockheed was t o study SCAT 17 with a h - minum alxoy as t h e basic s t r u c t u r a l material. The contractors were t o choose t h e cruise Mach number which they f e l t would be compatible with t h e aluminum structure.

Before discussing the f e a s i b i l i t y of t h e four configurations, it would appear desirable t o compare the f i n a l configurations of SCAT 16 and SCAT 17. (See f i g s . 1 and 2.) There are two very noticeable dif - ferences i n the f i n a l configurations f o r SCAT 16 from the o r i g i n a l con - (See f i g . 1.) The f i r s t i s t h e reduction i n wing aspect r a t i o cept.

t h a t was made t o improve the s t r u c t u r a l weight characteristics. The second is t h a t both contractors used a four - engine arrangement with two engines mounted below t h e wing and two above and t o the r e a r of the wing instead of t h e three - engine arrangement of t h e o r i g i n a l concept.

The contractors' f i n a l configurations of SCAT 17 ( f i g . 2) differ considerably from each other as w e l l as from t h e o r i g i n a l concept.

Boeing, f o r example, used an a f t - t a i l configuration whereas Lockheed used a canard configuration. The contractors changed t h e engine loca - t i o n both spanwise and vertically. Although it i s not apparent from t h i s figure, both contractors moved the engines up closer t o t h e wing $ and Lockheed actually integrated t h e outboard engines with the wing structure, the engines being mounted on the wing - chord plane. Compari - sons of t h e gross weights, payload - range characteristics, and other mis- sion performance data f o r SCAT 16 and SCAT 17 were presented i n paper no. 3 by R. Richard Heppe and J i m Hong and i n paper no. 4 by Lloyd T.

Goodmanson, W i l l i a m T. Hamilton, and Maynard L. Pennell.

Several of t h e more important factors t h a t could determine the f e a s i b i l i t y of t h e concepts are rated i n a broad or gross manner i n t a b l e I. Figure 3 is presented so t h a t the SCAT numbers can be associ - The r a t i n g system used i s shown at the ated with configuration concepts.

and P being used t o delineate the bottom of t a b l e I, with ratings A, M, d i f f i c u l t y of obtaining s a t i s f a c t o r y c h a r a c t e r i s t i c s f o r the p a r t i c u l a r item. The ratings are based on an N A S A assessment of t h e contractors' Even results and show the strong and weak points of t h e various concepts.

a quick glance at t h i s t a b l e shows that there are no outstanding configu - rations. I n comparing the configurations, it should be realized that t h e r e are d i f f e r e n t degrees of sophistication of design incorporated i n

t h e various SCAT configurations since SCAT 4 and SCAT 13 were studied f o r

only 4 months. The ratings f o r a l l the configurations indicate t h a t much

additional research i s needed t o a r r i v e at a s a t i s f a c t o r y transport.

The poor ratings of SCAT 4 and SCAT 15 indicate the areas of deficiency found i n t h e early p a r t of t h e study which were responsible

f o r their being dropped from the f e a s i b i l i t y study. SCAT 4 had severe

s t a b i l i t y problems with no immediate solution apparent and SCAT 15 was plagued with very high wing weights resulting from the two overlapping wings and t h e i r attendant problems, A comparison of SCAT 16 and SCAT 17 shows t h a t although both are capable of performing t h e mission requirements, each has many problems t h a t must be solved. A s can be seen from the ratings, the seriousness of these problems varies between the two configurations. When t h e over - a l l f e a s i b i l i t y of t h e concepts i s assessed, it appears t h a t there i s not much choice between SCAT 16 and SCAT 17 as supersonic transports.

O f t h e two t h a t were dropped, SCAT 4 and SCAT 15, it should be pointed

out t h a t e i t h e r would be a very good configuration i f the deficiencies could be corrected. Work i s now underway t o provide a means of a l l e -

viating t h e s t a b i l i t y problems associated with SCAT 4 and a study i s

underway t o a r r i v e a t a configuration of the SCAT 13 concept t h a t has b e t t e r s t r u c t u r a l characteristics.

This work i s discussed i n paper no. 14 by A. Warner Robins and Richard T. Whitcomb.

TRClDE STUDIES Several trade studies were made by the contractors t o determine importance of some of t h e guideline r e s t r a i n t s on the overall char - a c t e r i s t i c s of a supersonic transport. A f e w of t h e more important ones w i l l be discussed t o point out t h e sensitive areas.

The contractors were given the choice of the materials they used f o r basic structure, except f o r t h e aluminum SCAT'S, t h e only r e s t r a i n t being t h a t t h e structure would be s a t i s f a c t o r y f o r Mach 3.0 cruise.

Fig -

ure 4 shows a structural - material comparison of t h e r a t i o of the operating

weight empty t o the operating w e i g h t empty f o r an aluminum structure designed f o r room temperatures as a function of cruise Mach number. It is quite obvious that the use of titanium will r e s u l t i n an a i r c r a f t l i g h t e r than one using e i t h e r aluminum o r stainless steel. This figure was pre - pared from independent studies made by both Boeing and Lockkeed and the r e s u l t s compare favorably with the SCAT configurations weighed f o r the various materials. It i s quite obvious, of course, that aluminum w o u l d not be a satisfactory material f o r Mach 3 . 0 cruise. I n f a c t , f o r Mach numbers of about 2.0, aluminum structure would r e s u l t i n a considerably heavier a i r c r a f t than an aluminum a i r c r a f t designed at room temperatures and much heavier than a titanium a i r c r a f t which would have material capa - b i l i t i e s of operating a t higher Mach numbers. Since l o w gross weight i s necessary from a sonic - boom point of view, every e f f o r t should be made t o use titanium as a structural material f o r a supersonic transport regardless of i t s cruise Mach number.

'$ A comparison of the aluminum SCAT configurations studied by the two contractors with the corresponding titanium SCAT configuration Table I I ( a ) pre - designed f o r a Mach number of 3.0 is shown i n t a b l e 11.

sents the r e s u l t s of the study by The Boeing Company t o determine the characteristics of a Mach 2.0 aluminum S C A T 16 sized t o f l y the same mission as the Mach 3.0 SCAT 16. It i s evident t h a t the aluminum a i r - c r a f t weighs more than the titanium airplane and does not meet the The sonic - boom overpressures are 2.23 lb/ft2 sonic - boom requirements.

i n climb and 2.0 l b / f t 2 at start of cruise f o r the aluminum SCAT 16 com - pared with 2.0 l b / f t 2 and 1.5 l b / f t 2 f o r the titanium aircraft capable of operating at Mach 3.0.

Table I I ( b ) shows a similar comparison f o r SCAT 17 prepared by the Here again the take - off gross weights are Lockheed - California Company.

higher f o r the aluminum SCAT and the a i r c r a f t w i l l not meet the sonic - boom restrictions, the sonic - boom overpressure being 1.8 l b / f t 2 at s t a r t of cruise instead of the specified 1.5 lb/ft2. For these studies Lockheed used a cruising speed of M = 2.2. Thus, it would appear t h a t an aluminum SCAT cruising at Mach numbers from 2.0 t o 2.2 would not, w i t h the engines used i n these studies, meet all the mission requirements. If t h e aluminum SCAT had been powered with an advanced engine of engine "D" technology, the performance would have been b e t t e r and would have compared more favorably with the titanium a i r c r a f t . However, it i s not believed t h a t t h i s increase i n engine performance would be sufficient t o compen - sate f o r the greater s t r u c t u r a l weight of an aluminum SCAT and the ben - e f i c i a l effects on the airplane characteristics, including sonic boom, of flying a t a Mach number of 3.0 instead of 2.0 o r 2.2.

The very powerful effect of placing r e s t r i c t i o n s on the sonic - boom overpressures during transonic acceleration on airplane gross weight i s shown i n figure 5 . In t h i s figure are plotted the sonic - boom overpres - sures as a function of gross weight. The large band represents the outer l i m i t s of the variation of gross w e i g h t with sonic - boom overpres - sure f o r the two S C A T configurations studied by the two contractors.

It i s interesting t o note tg@:thare is.Dply about a 60,000-p0und dif - ference i n gross w e i g h t between the.figh&st and heaviest of the con - figurations studied by the contractors. The s e n s i t i v i t y of gross w e i g h t t o sonic - boom overpressure f o r the best of these configurations a t an overpressure of 2.0 indicates the importance of the selection of the design overpressure. I f , f o r example, the design boom overpressure had been 2 . 1 i n s t e a d of 2.0, a difference probably imperceptible t o the general public, the gross weights would be from 20,000 t o 40,000 pounds t less than the weights based on an overpressure of 2.0 lb/ft2.

O n the other hand, if the overpressure had been a r b i t r a r i l y selected at 1 . 9 lb/sq f t , there would have been a 20,000 - t o p,OOO-pound increase i n gross weight and the heavier of the SCAT configurations would have r' been incapable of meeting the sonic - boom requirements and the 3,200 - nautical - mile range. Since the sonic - boom problem i s one of utmost importance t o the design of the transport and can very well determine whether the a i r c r a f t w i l l be capable of economically per - forming the design missions, extreme care should be exercised i n picking the m a x i m u m overpressure during transonic acceleration t o assure that undue r e s t r i c t i o n s are not placed on the design of the a i r c r a f t .

These r e s u l t s indicate the extreme importance of continuing the studies t o determine the intensity of sonic booms t h a t will be acceptable t o the public f o r commercial transport operation.

Figure 6, a comparison of the variation of take - off gross w e i g h t with the transonic sonic - boom overpressures f o r three of the four study engines, shows t h e importance of the sonic - boom - overpressure r e s t r i c - t i o n on engine performance requirements.

If engine "C" with a f t e r - burning had been used instead of engine rtD" (low flow), a considerable increase i n gross weight would have resulted. In fact, i f engine "B" had been used, t h e a i r c r a f t would be incapable of meeting the sonic - boom requirement as set down i n the " Guidelines. " These r e s u l t s clearly show t h e need for an improved state - of - the - art engine of a t l e a s t engine "D" performance t o assure a satisfactory supersonic transport, particularly with the current sonic - boom restrictions.

For the S C A T studies the current F A A International Fuel Reserves were used, with an alternate a i r p o r t distance of 250 nautical miles.

' Since fuel reserves are regulated from a safety point of view and are based, i n part, on t h e current air t r a f f i c control system and weather minimums, variations i n f u e l reserves were studied t o determine t h e i r importance on a i r c r a f t gross w e i g h t and/or range. Figure 7 shows the take - off gross w e i g h t as a function of range f o r a typical configura - t i o n with variations i n f u e l reserves of *lO,OOO pounds from the SCAT values. Constant payload volume w a s assumed i n tietermining these vari - ations. The center curve i s the variation of take - off gross weight with range with the use of the SCAT f u e l reserves. The lower curve shows the improvements i n gross weight or range t h a t would be obtained i f f u e l reserves could be reduced by 10,000 pounds and indicates t h e importance of improving the a i r t r a f f i c coh%rol system, and so forth, so t h a t the reserves can be reduced without compromising the safety of commercial operations. If,, f o r some reason, however, it were necessary t o increase the fuel reserves, the a i r c r a f t would become very heavy or have decreased range.

Figure 8 has been prepared t o show the effects of range and pay - load on the take - off gross weight of the airplane. For these r e s u l t s the fuselage w a s resized t o accommodate the additional passengers f o r increases i n payload and f o r fewer passengers i n the case of decreased payload. The r e s u l t s i n t h i s figure are f o r an a i r c r a f t meeting t h e sonic - boom - overpressure r e s t r i c t i o n 4 of 2.0 l b / f t 2 during transonic The center curve again represents the variation of gross acceleration.

weight with range f o r a typical S C A T configuration with i t s payload of 26,125 pounds. Increases i n range above t h e design value of 3,200 nau - In f a c t , t i c a l m i l e s r e s u l t i n very large increases i n gross weight.

a range of about 3,500 nautical miles would appear t o be the m a x i m u m range f o r t h i s configuration and would require increases i n gross weight, of about 100,000 pounds. Increasing the payload t o 35,000 pounds, the a t a l l top curve, would also require large increases i n gross weight ranges. With 35,000 pounds payload t h i s particular S C A T would not meet the mission range requirement, since it has a maximum range of about The very rapid increase i n gross weight with increase i n 3,000 miles.

range f o r take - off gross w e i g h t s above 400,000 pounds is a result of t h e sonic - boom - overpressure limitation of 2,O lb/ft2 during accelera - t i o n and 1.5 lb/ft* during cruise. Another indication of the impor - tance of sonic - boom r e s t r i c t i o n s i s the f a c t that take - off gross weights of between 550,000 and 650,000 pounds appear t o give the maxi - mum range f o r t h i s configuration regardless of its payload. A s would be expected, there would be large decreases i n take - off gross weight if the payload were reduced t o values lower o r large increases i n range than 26,125 pounds as indicated by the large difference between the SCAT payload and a payload of 15,000 pounds. Because of the very pro - nounced e f f e c t t h a t the sonic - boom r e s t r i c t i o n s have on the airplane's gross,weight and range, it is evident that t h i s picture would be con - siderably changed if t h e reductions i n gross weight indicated by The i Boeing Company f o r improved structures and advances i n aerodynamic state of the art could be realized. Such reductions would lower the design point and would r e s u l t i n much b e t t e r payload range character - i s t i c s because the adverse effects of sonic boom become less pro - nounced f o r gross w e i g h t s less than 400,000 pounds. If, however, an advanced engine of "D" performance were not available, an airplane using current state - of - the - art engines, engine "'B," f o r example, would have even more adverse characteristics than are shown i n the figure.

a - 9 . 1 * e * e a I n looking a t the mission requiremeGts ?o"r commercial transport a i r c r a f t , it appears t h a t certain mission f l e x i b i l i t i e s may be required because of sonic - boom sensitive areas or operational procedures. Fig - ure 9 has been prepared t o show the varia$ion i n range f o r two flight profiles using the g r o s s w e i g h t for t h e b$sfepmi,c3s$o? (3,200 nautical miles a t a Mach number of 3.0). Mission A is a " fle#@t pkofile where the point of departure i s i n a sonic - boom sensitive area and the air - c r a f t i s required t o f l y subsonically f o r some distance p r i o r t o accel - erating t o supersonic speeds, i n t h i s case 800 nautical miles. Mission B is a subsonic mission a t a Mach number of about 0.85. I n t h e mixed subsonic - supersonic mission A, SCAT 16 has a range greater than f o r the normal mission and SCAlcl7 has a range of about 3,200 nautical miles.

s flight plan (mission A ) could be used by SCAT 16, with i t s basic sion range of 3,200 nautical m i l e s , f o r nonstop Paris - New York flights i n winter with the headwinds since it would have a range of about 4,000 s t a t u t e miles or close t o the value t h a t i s needed f o r day - i n and hy-out operation. This plan would, of course, result i n a block t i m e about halfway between the supersonic time and the t i m e f o r the current subsonic j e t s . Mission B i s an indication of the subsonic aerodynamic performance of SCAT 16 and SCAT 17 and r e f l e c t s the good subsonic performance of SCAT 16.

CONCLUSIONS The conclusions r e f l e c t the very powerful influence of sonic boom on a l l phases of the study and indicate i t s importance as a design parameter f o r the supersonic transport.

Either the fixed - wing or the variable - sweep concept will meet t h e SCAT mission requirements with very l i t t l e choice, for example, between

S C A T 16 and SCAT 17. If mission f l e x i b i l i t y i s required, SCAT 16

appears t o offer t h e better compromise; however, i f structural simplic - i t y and w e i g h t are of primary importance, S C A T 17 might offer t h e better compromise.

The results indicate t h a t i f SCAT i s t o be a competitive transport, "priced propulsion systems are required. This requirement means t h a t t d engine mst have high i n l e t turbine temperatures and be an advance over the current state of t h e art with performance characteristics a t least equal t o those of engine " D. " The r e s u l t s a l s o indicate that i f the airplane i s built with t i t a - nium as the basic s t r u c t u r a l material, there would be a significant reduction i n gross w e i g h t over a stainless - steel airplane.

The r e s u l t s indicate th& t h e aluminum SCAT configurations studied would not meet the mission requirements.

The aluminum SCAT configura - t i o n s w i t h cruise speeds of Mach numbers from 2.0 t o 2.2 would weigh more than the titanium Mach 3 . 0 a i r c r a f t ; would, of course, f l y slower; and would not m e e t the sonic - boom r e s t r i c t i o n s if it had a range of 3,200 nautical m i l e s and carriea " the 26,123-po~nd payload.

The studies a l s o emphasize t h a t any substantial increase i n f u e l reserves, payload, or range will be very costly insofar as gross weight i s concernbd.

For the transport t o have satisfactory characteristics, research i n several areas i s urgently needed. The a i r c r a f t needs t o have b e t t e r % aerodynamic performance than w a s incorporated i n t o these studies, and $1 configurations need t o be devised t h a t can obtain the high aerodynamic performance while s t i l l maintaining l o w s t r u c t u r a l weight. Research i s needed i n the areas of s t a b i l i t y and control t o improve t h e flying qual - i t i e s of the a i r c r a f t . Research and development work i n the engine f i e l d i s required t o have available an engine of a t l e a s t class "D" capability by the t i m e the a i r c r a f t i s ready t o enter i n t o service.

Since titanium i s by far the l i g h t e r material and offers the greatest potential, considerable work i s needed t o establish the l e v e l of con - fidence i n titanium t h a t now e x i s t s f o r aluminum. Since sonic boom i s so important t o the operational procedures of the a i r c r a f t and t o i t s gross weight and performance, research i s needed t o see whether there a r e possible means of reducing the booms a s well as research i n the f i e l d of public acceptance t o provide documented evidence from which t h e boom r e s t r i c t i o n s can be s e t a t the highest acceptable level.

1 4 6 / G - GOOD A - ACCEPTABLE M - MARGINAL P - POOR TABLE I1 COMPARISON OF ALUMINUM AND TITANIUM AIRPLANES (a) SCAT 1 6 ALUMINUM TITANIUM TAKE - OFF GROSS WEIGHT, LB 520,000 430,000 RANGE, N. MI. 3,200 3,200 CRUISE SPEED M=2.0 M = 3.0 WING AREA, F T ~ 4,750 4,000 ENGINES F D (LOW FLOW) SONIC - BOOM OVERPRESSURE: I N CLIMB , L B / F T ~ 2,23 2.00 A T START OF CRUISE , L B I F T ~ 2.00 1845 ( b ) SCAT 1 7 ALUMINUM TITANIUM TAKE - OFF GROSS WEIGHT, LB 500,000 380,000 c RANGE, N. MI, 3,200 3,200 CRUISE SPEED M= 2.2 M = 3.0 W I N G AREA, F T ~ 7,200 5,500 ENGINES F D SONIC - BOOM OVERPRESSURE: I N C L I M B , L B I F T ~ 2 .o 2 .o AT START O F CRUISE,LB/FT~ I .8 I .5 : ' ;: ORIGINAL CONCEPT II BOEING LOCKHEED FINAL CONFIGURATIONS Figure 1 SCAT 1 7 ORIGINAL CONC E PT I

} - - 2 19' - 6" ,-------d 253'-3"

BOEING LOCKHEED FINAL CONFIGURATIONS Figure 2 ORIGINAL STUDY CONCEPTS SCAT 4 SCAT 1 7 SCAT 1 5 SCAT 1 6 i l-i Figure 3 STRUCTURAL MATERIALS I 2 3 DESIGN CRUISE MACH NUMBER

Figure 4

SON I C - BOOM COMPARISON

TRANSONIC ACCELERATION 500rx103 I GROSS WEIGHT, LE -

I I

0 1 . 5 2 .o 2.5 SONIC - BOOM OVERPRESSURE, W F T 2 Figure 5 ENGINE COMPARISON SONIC BOOM I GROSS

WE LB I GHT, '*'~

ENGINE B I C D (LOW FLOW) 4 0 0 [ 360

L L

0 1.5 2.0 2.5 SON IC - B O O M OVERPRESSURE, LB/FT* Figure 6 RANGE FOR VARIOUS FUEL RESERVES CONSTANT PAYLOAD VOLUME 500 /\-SCAT + 10,000 L B GROSS SCAT WE I GHT, DESIGN POINT L B 10,000 L B i t l 3000 4000 0 2000 RANGE, N. MI.

Figure 7 6 0 0 - X 103 PAYLOAD - - GROSS WE I GHT, LB -

-

I I I I I I L I 4000 5000 0 2000 3000 RANGE, N. MI.

Figure 8 f M I SS ION FLEX IBlLlTY BASIC MISSION: Mz3.0; RANGE = 3,200 N. MI.

ALTITUDE MISSION B \ \ RANGE RANGE, N. MI.

SCAT 16 SCAT 17 MISSION A (SUBSONIC FOR 800 N. MI.). * * 3 , 4 5 0 3,200 L MISSION 6.. .. .. . . . . . . . . .. .. . .. ... . . . . . . . . .

4 , 5 0 0 3,500 Figure 9 . .

.,. .

. I .i- ,; 2 : @ANALYSIS AND CORRELATION OF AIRCKAET WAVE DRAG By Roy V. Harris, Jr.

I NASAJangley Research Center S U M M A R Y Two approaches t o t h e analysis and correlation of a i r c r a f t wave The f i r s t approach i s t o compare t h e a i r c r a f t area drag a r e discussed.

distribution with t h e slender - body - theory optimum and t h e second approach employs the supersonic area rule t o determine the equivalent - body area distribution.

Although the optimum - body wave - drag variations can be used t o indi - cate t h e gross trends i n a i r c r a f t wave drag w i t h equivalent - body fine - ness r a t i o and t h e degree o f afterbody closure, the r e l a t i v e l y large amount of s c a t t e r i n the correlating data, due t o deviations i n a i r - c r a f t area d i s t r i b u t i o n s from t h e optimum, indicate t h a t more precise computing techniques a r e required i n order t o estimate t h e a i r c r a f t wave - drag levels.

Several computer programs have been developed which apply the slender - body theory i n combination with t h e supersonic area r u l e t o t h e solution of a i r c r a f t wave drag. One such program, developed by The Boeing Company f o r use on the IBM 70% electronic data processing and t h e system, has been studied a t t h e Langley Research Center, r e s u l t s of t h i s study indicate t h a t , i n addition t o providing reason - ably accurate supersonic wave - drag estimates, the computer program provides a useful t o o l which can be used i n design studies and f o r

configuration optimization. A dTgdde

INTRODUCTION * Since t h e r u l e w a s formulated, and v e r i f i e d experimentally, t h a t r t h e transonic wave drag of an a i r c r a f t i s essentially t h e same as t h e wave drag of an equivalent body of revolution having t h e same cross - sectional - area d i s t r i b u t i o n a s the a i r c r a f t ( r e f . l), attempts have been made t o estimate a i r c r a f t wave drag by examining t h e equivalent - These attempts have led t o two approaches body area distributions.

for analyzing t h e equivalent - body, and hence t h e a i r c r a f t , wave drag.

The first and simplest approach i s t o compare the a i r c r a f t area dis - t r i b u t i o n with t h e slender - body - theory optimum. The second approach i s more complex and employs t h e supersonic area rule (ref. 2 ) t o

153 - 1 6 - 3

r I .

.I I S determine the equivalent - body area distributions. This paper discusses briefly the first approach, and then presents in more detail some recent developments in the application of the second approach to the analysis and correlation of wave drag.

SYMBOLS A cross - sectional area C D , , J ~ ~ wave - drag coefficient D drag 2 length M Mach number

v velocity

x,y,z coordinates along X, Y, and Z axes X,Y,Z axis system of airplane

e azimuth angle

1-L Mach angle P density Subscripts: MAX maximum BASE base DISCUSSION In the design of supersonic aircraft, the Sears - Haack (ref. 3 ) or Haack-Adams (ref. 4) optimum - body area distributions are often used as the basis for establishing the cross - sectional - area distribution of the aircraft. However, as can be seen in figure 1, attenpts to correlate aircraft wave drag with that of the optimum body indicate a consider - able degree of scatter in the data. This figure shows the variation * e 3 : e : , * p u . .

of wave - drag coefficient based on maximb" cro&sectional area with equivalent - body fineness r a t i o . The two curves show t h e optimum - body wave - drag variations. For the purposes of t h i s paper; the optimum body i s considered t o be the body of revolution which, according t o slender - body theory, has the minimum wave drag f o r a given length, volume, and base area. The upper curve is the wave - drag variation f o r t h e Sears - Haack optimum body which has 100 - percent afterbody closure (ref. 3 ) . The lower curve represents the optimum body which has a base area equal t o 50 percent of the m a x i m u m area. The symbols indicate wave - drag coefficients determined by wind - tunnel t e s t s of models of a wide v a r i e t y of complete a i r c r a f t configurations. This comparison between a i r c r a f t wave drag and t h e theoretical optimum i s made a t the near - sonic Mach number of 1.2 because the slender - body theory does not ) consider the optimum - body wave - drag variations with Mach number. The Mach number effects, as presented subsequently, can be significant, p a r t i c u l a r l y a t the lower fineness r a t i o s .

The deviations i n the a i r c r a f t wave - drag levels from the optimum and t h e rather large amount of s c a t t e r i n the data can be a t t r i b u t e d t o t w o effects. F i r s t , due t o t h e techniques f o r simulating engine i n s t a l - l a t i o n s and t h e sting - mounting system used on t h e models from which these data were obtained, the various configurations have d i f f e r e n t amounts of afterbody closure ranging from about 60 t o 90 percent. A s can be seen by t h e difference i n l e v e l s of t h e t w o optimum - body curves, o f afterbody closure can have a powerful e f f e c t on the the degree equivalent - body wave drag. Second, as shown i n the sketch of the area d i s t r i b u t i o n of a t y p i c a l wind - tunnel model ( f i g . l ) , t h e a i r c r a f t area distributions are r a t h e r lumpy and depart considerably from the optimum.

Although the optimum - body wave - drag variations can be used t o indicate equivalent- the gross trends i n a i r c r a f t wave drag with variations i n body fineness r a t i o and the degree of afterbody closure, the large amount of s c a t t e r i n t h e data due t o deviations i n the a i r c r a f t area distributions from t h e optimum indicates that more precise computing techniques a r e required i n order t o estimate t h e a i r c r a f t wave - drag level.

Reasonably good wave - drag estimates can be made near a Mach number of 1 if the slender - body theory i s applied t o the actual a i r c r a f t area distributions. This procedure can be extended t o higher Mach numbers

'

by using the supersonic area rule t o determine the equivalent - body area distributions. A review of t h i s procedure i s given i n figure 2.

Each equivalent body of revolution i s determined by passing a s e r i e s of p a r a l l e l cutting planes through the configuration. The cut - t i n g planes are inclined with respect t o t h e a i r c r a f t axis at the Mach angle p. The area of the equivalent body at each s t a t i o n i s the pro - jection onto a plane normal t o t h e a i r c r a f t axis of the area intercepted It is evident that the series of parallel cutting by the cutting plane.

can be oriented at various angles 9 around the aircraft axis; planes and in order to determi'ne the drag accurately, a family of equivalent - 9, must be consid - bodies, each corresponding to a particular value of Thus, at each Mach number, a series of equivalent bodies of revo - ered.

lution are generated. The wave drag of each equivalent body is deter - mined by the von Karman slender - body formula (ref. 5) which gives the drag as a function of the free - stream conditions and the equivalent - body area distribution. The wave drag of the aircraft at the given Mach num - ber is then taken to be the integrated average of the equivalent - body wave drags.

For most practical applications, the complexity Qf this procedure A s a requires that it be adapted to the high - speed electronic computer. \ j ,~ result, several digital - computer programs have been developed which apply this theoretical approach to the solution of aircraft wave drag. One such program, developed by The Boeing Company for use on the IBM 7090 electronic data processing system, has been under study at the Langley Research Center and details of this program will be made available by A simplified description of the program and some of the the hASA.

results of this study are shown in figures 3 to 6.

A major problem in adapting this procedure to machine computation is that of describing a rather complex aircraft to the computer in suf - The manner in which an aircraft is mathematically ficient detail.

described to the computer for this program is illustrated in figure 3 .

The lower right portion of the figure shows a typical aircraft for which the supersonic wave drag is to be computed. The upper left portion of the figure shows the aircraft as it is described to the computer.

The locations of all the aircraft components are referred to an X - , The fuse - Y - , Z - axis system with its origin at the nose of the fuselage.

lage is assumed to be sufficiently close to a body of revolution that it - sectional - area distribution. The can be described in terms of its cross variation in fuselage radius along the axis between stations is assumed to be linear.

The wing is described as a sequence of streamwise airfoils distrib - The contour of the wing is assumed to be linear uted along the span.

between successive ordinates. The horizontal and vertical tails are I described in a manner similar to that of the wing.

The engine nacelles are located by specifying the x, y, and z ordinates of the nacelle center line at the inlet face, and are described in a manner similar to that of the fuselage by giving the radii at suc - cessive stations. The discontinuities caused by.the inlet and exit faces are eliminated by assuming that infinitely long cylinders extend in both directions from t h e i n l e t and the e x i t . The effe‘cis of i n l e t spillage on the wave drag can be included by properly contouring the cylindrical extension near the i n l e t face.

Once the a i r c r a f t description has been stored i n t h e memory u n i t of the computer, the equivalent - body area distributions a r e determined by solving f o r the normal projection of the areas intercepted by t h e cutting planes.

I n addition t o t h e a i r c r a f t wave drag, which i s evaluated by applying the method of Eminton and b r d (ref. 6 ) t o the solution of the von K&m&n i n t e g r a l ( r e f . 5 ) , a check on the accuracy of the equivalent - , body area d i s t r i b u t i o n i s provided. F i r s t , the exact volume of the wing ‘ of the mathematical model i s computed, and then, independently, the vol - ume i s found by integrating t h e area d i s t r i b u t i o n of the wing equivalent body. If a s u f f i c i e n t number of cutting planes have been used t o define the wing equivalent body of revolution, then the t w o values of wing vol - ume should be e s s e n t i a l l y the same.

I n addition, the program l i s t s the wave drags of the a i r c r a f t equivalent bodies a t each Mach number, a s well as selected equivalent - body area distributions. This additional information i s p a r t i c u l a r l y useful i n t a i l o r i n g a configuration f o r minimum wave drag since, i n order t o optimize a configuration at some supersonic Mach number, it i s necessary t o examine the s e r i e s of equivalent bodies corresponding t o the p a r t i c u l a r Mach number. It should be noted t h a t the area distribu - t i o n s required i n the computation of sonic - boom overpressures a r e provided.

In order t o determine the accuracy of the wave - drag estimates obtained from t h i s procedure, the drag w a s first computed f o r the opti - mum body of revolution; the r e s u l t s were then compared with experiment and t h e more precise characteristics theory. Figure 4 shows a compari - son of t h e computed wave drag with experimental r e s u l t s f o r the optimum body having a base - to - maximum - area r a t i o of about 53 percent. The var - i a t i o n s i n wave drag with Mach number are shown f o r optimum - body fine - ness r a t i o s of 7, 10, and 13. The data points were obtained i n the Langley 8 - foot transonic pressure tunnel, t h e Langley 4- by &foot Br I supersonic pressure tunnel, and t h e Langley Unitary Plan wind tunnel by integrating t h e measured surface - pressure coefficients.

The characteristics theory, indicated by t h e solid l i n e , shows excellent agreement with the experimental results. The slender - body theory, shown by t h e short - dash l i n e , gives good agreement near a Mach number of 1. However, as t h e Mach number i s increased, the slender - body theory overestimates t h e optimum - body wave drag. It should a l s o be noted t h a t t h e e f f e c t s of Mach number a r e greater a t the lower * T P This greater depar - fineness ratios than at the higher fineness ratios.

ture from slender - body theory should be expected as the bodies become less slender. The long - dash line shows the results obtained from the machine program which uses the slender - body theory in combination with the supersonic area rule.

A s would be expected, the characteristics theory gives the best However, when the slender - body agreement with experimental results.

theory is applied to the proper equivalent bodies, as in the machine program, the Mach number effects on the optimum - body wave drag are pre - dicted with a fair degree of accuracy.

The most severe test of the theoretical approach used in this machine program lies in its application to the calculation of the drag of wings. Figure 5 shows a comparison of the machine-computed wave drag with experimental results for a series of semispan wings.

In the table at the right of figure 5 are shown sketches of the wings and the Mach numbers for which computations have been made. A l l The first the wings in the series had circular - arc airfoil sections.

wing had a trapezoidal planform and a linear spanwise thickness distri - bution. The second wing had a complex planform with a linear spanwise thickness distribution. The third wing had a complex planform as well The final wing had an as a complex spanwise thickness distribution.

The computations arrow planform with a linear thickness distribution.

were made for Mach numbers ranging from about 1 . 4 to 2.2.

At the left of figure 5 is shown a plot of experimentally deter - mined wave drag plotted against the computed values. The solid line is The the locus of perfect agreement between theory and experiment.

experimental wave - drag coefficients were found by subtracting the equi - valent flat - plate turbulent skin - friction drag coefficients from the As can be seen from the figure, the measured total drag coefficients.

program tends to underestimate the wave drag of the semispan wings.

This result for wings alone is not surprising, since a wing departs con - siderably from the equivalent body of revolution assumed by the theory.

A comparison of the machine - computed wave - drag coefficients with experimental results for complete airplane configurations is shown in figure 6. A s indicated in the table, the comparison is made over the

Mach number range from 1.4 to 3.2 for several of the SCAT configurations

and a typical supersonic fighter. Again, the plot on the left shows, as a solid line, the locus of perfect agreement between the experimen - tally determined wave - drag coefficients and the machine - computed values.

The experimental wave - drag coefficients were determined for each con - figuration by subtracting the equivalent flat - plate turbulent skin - friction drag and an estimated camber drag from the wind - tunnel results.

A s can be seen from figure 6, t h i s comparison indicates that the machine program, which uses slender - body theory i n combination with the super - can produce good estimates of a i r c r a f t wave drag a t sonic area rule, supersonic speeds.

CONCLUDING REMARKS I n conclusion, it appears t h a t although the optimum - body wave - drag variations can be used t o indicate the gross trends i n a i r c r a f t wave drag with equivalent - body fineness r a t i o and the degree of afterbody closure, the r e l a t i v e l y large amount of s c a t t e r i n the correlating data, i due t o deviations i n a i r c r a f t area distributions from the optimum, indi - cates t h a t more precise computing techniques a r e required i n order t o estimate the a i r c r a f t wave - drag levels.

Several computer programs have been developed which apply the slender - body theory i n combination with the supersonic area rule t o the solution of a i r c r a f t wave drag. One such program, developed by The Boeing Company f o r use on the IBM 7090 electronic data processing system, has been studied at the Langley Research Center, and the results of t h i s study indicate that, i n addition t o providing reasonably accu - r a t e supersonic wave - drag estimates, the computer program provides a useful t o o l which can be used i n design studies and f o r configuration optimization.

REFERENCES 1. Whitcomb, Richard T.: A Study of the Zero - Lift Drag - Rise Character - NACA i s t i c s of Wing - Body Combinations Near the Speed of Sound.

Rep. 1213, 1956. (Supersedes NACA RM ~ 5 2 ~ 0 8 . ) 2. Jones, Robert T.: Theory of Wing - Body Drag a t Supersonic Speeds.

NACA Rep. 1284, 1956.

3. Sears, W i l l i a m R.: O n Projectiles of M i n i m u m Wave Drag. Quarterly Appl. Math., vol. IV, no. 4, Jan. 1947, pp. 361-366.

Determination of Shapes of Boattail Bodies of Revo - 4. Adams, Mac C . : lution f o r Minimum Wave Drag. NACA TN 2550, 1951.

5. von K&rm&n, Th. : The Problem of Resistance i n Compressible Fluids.

R. Accad. d ' I t a l i a , C 1 . Sei. Fis., Mat. e N a t . , vol. X I I I , 1935, pp. 210-265.

6. Eminton, E., and Lord, W. T.: Note on the Numerical Evaluation of the Wave Drag of Smooth Slender Bodies Using Optimum Area D i s t r i - butions f o r Minimum Wave Drag. Jour. R.A. S., vol. 60, no. 541, Jan. 1956, pp. 61 - 63.

COMPARISON OF AIRPLANE WAVE DRAG> WITH

SLENDER - BODY - THEORY OPTIMUM

WAVE - DRAG M = 1,2 COEFFICIENT, EXPERIMENT CD, WAVE TYPlCAL WDE& OAIRPLANE MODELS WITH VARIOUS AFTERBODY

A w, CLOSURES

A M A X THEORY (OPTIMUM BODIES)

- - - 100% AFTERBODY

CLOSURE

- 50% AFTERBODY

.2 .. CLOSURE

I I I

0 4 '

7 9 II I3 FINENESS RATIO Figure 1 ILLUSTRATION OF WAVE - DRAG COMPUTING PROCEDURE a MATHEMATICAL REPRESENTATION OF ILLUSTRATIVE AIRPLANE FOR MACHINE - COMPUTING PROCEDURE MATHEMATICAL MODEL

i

Y ILLUSTRATIVE AIRPLANE Figure 3 COMPARISON OF COMPUTED WAVE DRAG WITH EXPERIMENTAL RESULTS OPTIMUM BODIES OF REVOLUTION ABASE/AMAX = 0.532 'D, WAVE

"~ AMAX

X

-- -- - - - -- - _ _ _ _

&? - - CHARACTERISTIC -------___ SLENDER BODY

----

SLERDE BODY WITH

SUPE SONIC AREA i

CD, WAVE RULE FINENESS RATIO = IO FINENESS RATIO = 1 3 .08 M M Figure 4 COMPARISON OF MACHINE - COMPUTED WAVE DRAG WITH EXPERIMENTAL RESULTS SEMISPAN WINGS EXPERIMENTAL cD, WAVE WING MACH NUMBERS 0 h1.4; 1.8; 2.0; 2.2

A 11.4; 1.8; 2.0; 2.2

0 A b . 4 ; 1.8; 2.0; 2.2

A A 1.6; 2.0; 2.2

Figure 5 .010- CONFIGURATION MACH NUMBERS 0 SCAT 1 5 2.3; 2.6; 3.0 0 SCAT 15 - A 2.3; 3.0 0 SCAT 15-6 I .6; 2.2 - .005 v SCAT 1 6 2.4; 2.6; 3.0 C7 SCAT 16 - A 2.4; 2.6; 3.0 a SCAT 4 2.3; 2.6; 3.0; 3.2 u FIGHTER 1.4; 2.2 Figure 6 APPLICATION O F WING WARP AND AERODYNAMIC 1 - C E TO IMPROVE SUPERSONIC PERFORMANCE By Francis E. McLean and Harry W. Carlson

- -

NASA,Langley Research Center A brief review i s made of current informtion on highly swept 'warped - wing arrangements obtained from experiments w i t h isolated wings, with wing - body configurations, and w-ith complete configurations. The use of analytical methods f o r the solution of arbitrary planform prob - l e m s i s discussed. )Q LfTtw4 INTRODUCTION It i s well known that linearized theory predicts high l e v e l s of supersonic performance for wings with leading edges which are mept

behind the Mach cone - t h a t is, subsonic leading edges. A prime ingre -

dient of t h i s performance potential i s the t h e o r e t i c a l premise that many wings of t h i s c l a s s can be warped t o produce l i f t more e f f i c i e n t l y than w i n g s with leading edges that l i e ahead of the Mach cone. Several experimental investigations (refs. 1 t o 3 ) have been conducted on wings which were warped, according t o the specifications of theory, t o provide these better l i f t i n g efficiencies. The r e s u l t s of these experiments were rather disappointing. Because of these disappointments, there i s some feeling t h a t the advantages of wing w a r p are questionable and t h a t l i t t l e i s t o be gained by the use of wings with subsonic leading edges.

I

However, the f a c t t h a t three of the NASA SCAT configurations make use of t h i s type of wing i s evidence of a continued i n t e r e s t i n the design concept and i s also evidence of the belief that research w i l l find the key which w i l l allow application of the concept t o a p r a c t i c a l airplane with improved supersonic performance.

T h e purpose of the present paper i s t o discuss some of the present information on highly w e p t warped - wing arrangements t h a t has been obtained from tests of isolated wing and wing - body configurations, from t e s t s of complete configurations, and from a n a l y t i c a l methods.

J SYMBOLS l o c a l chord pitching-moment coefficient drag coefficient zero - lift drag coefficient of symmetrical configuration drag - due - to - lift factor (fig. 1) l i f t coefficient drag l i f t Mach number Reynolds number maximum thickness of a i r f o i l section sweepback angle of wing leading edge Sub s c r i p t s : DES de sign MAX maximum OPT optimum TRIM trim value THEORY t h e o r e t i c a l i lSESULTS AND DISCUSSION ? - w - * ; 7 - 7r2- Design Considerations A t t h e outset, it would be i n order t o consider the design region of high l i f t i n g efficiencies as specified by theory (ref. 4) f o r the Mach 3 . 0 cruise condition of the supersonic transport. I n figure 1 the t h e o r e t i c a l drag - due - to - lift f a c t o r s &$CL2 are shown as a function of leading - edge sweep angle f o r a Mach number of 3.0. The v e r t i c a l l i n e (long dash, short dash) at 7O.>O separates t h e sweep - angle region i n t o a supersonic - leading - edge condition and a subsonic - leading - edge condition as indicated i n the figure. In the supersonic - leading - edge I region (leading - edge sweepback less than 7 0 . 5 O ) , theory indicates some minor drag - due - to - lift advantages from the use of highly notched arrow wings with wing w a r p , but, because of wave - drag disadvantages, t h i s region i s generally discounted as a suitable design region f o r t h i s type of wing (ref. 4). For all p r a c t i c a l purposes, then, the drag - due - t o - l i f t f a c t o r s f o r wings with supersonic leading edges can be repre - shown by the solid - line curve.

sented by the constant value From figure 1 it can be seen that as the leading - edge sweepback i s increased t o clear the zero - lift drag peak (symbolized by the long dashed curve on t h e upper p a r t of the fig.), the t h e o r e t i c a l drag - due - to - lift The upper band (shaded area) represents values diverge i n t o tW.0 bands.

the drag - due - to - lift f a c t o r s f o r flat l i f t i n g surfaces and i s bounded by the curve f o r the delta wing ( s o l i d curve) and that f o r highly notched The lower band arrow wing (dashed curve) as indicated i n t h e figure.

(cross - hatched area), which i s similarly bounded, represents the corre - sponding drag - due - to - lift f a c t o r s i f optimum wing w ~ r p i s applied t o the It i s apparent that an airplane designed i n this subsonic leading - wing.

edge region must obtain some of the theoretical drag - due - to - lift poten - t i a l predicted f o r optimum wing warp t o exhibit high l e v e l s of super - sonic perf ormance.

I n attempts t o t r a n s l a t e the drag - due - to - lift potential of warped wings that i s shown i n figure 1 i n t o an increased l i f t - d r a g r a t i o , o r an important parameter has been found t o be the design l i f t efficiency, a coefficient CL,DEs. Design l i f t coefficient C L , D E ~ i s the l i f t coefficient f o r which a l l the l i f t i s obtained by wing warp. A wing with a high design l i f t Coefficient i s extremely warped, whereas a design l i f t coefficient of zero designates a flat l i f t i n g surface w i t h no w a r p . There i s a certain design l i f t coefficient; that is, a certain degree of warp f o r which, theoretically, the highest possible l i f t - d r a g This design l i f t coefficient i s called the optimum r a t i o i s obtained.

design lift coefficient CL,om. A w i n g w i t h either more or less w a r p w i l l theoretically provide less l i f t - d r a g r a t i o .

For example, with the use of t h e o r e t i c a l inputs from figure 1, figure 2 indicates t h e variation of t h e o r e t i g a l l i f t - d r a g r a t i o with l i f t coefficient f o r three wings. The three w i n g s have the same plan - form, are at the same Mach number of 3.0, and have the same assumed value of C D , ~ of 0.0060. The solid curve represents the flat wing with zero design lift coefficient (CL,DES = 0)' a wing which depends on a t t i t u d e f o r i t s l i f t . The short - dashed curve represents a wing which has a design l i f t coefficient equal t o one - half' the optimum T h i s wing d e p n d s on both w a r p and a t t i - ( C L , ~ S = O . ~ C L , O ~ T = 0.062).

The long - dashed tude t o reach t h e l i f t coefficient of peak efficiency.

curve represents a wing which i s warped f o r the optimum design l i f t coefficient (CL,-,E~ = CL,O~T = 0.124). A t the design l i f t coefficient of t h i s optimum wing, t h e highest possible t h e o r e t i c a l l i f t - d r a g r a t i o The point t o be made from this i s obtained f o r the assumed conditions.

figure i s t h a t not much i s lost i n t h e o r e t i c a l l i f t - d r a g r a t i o i f l e s s A limited design l i f t coefficient, or than the optimum w a r p i s used.

reasonably good r e s u l t s as w i l l be shown i n degree of w a r p , can lead t o sub sequent figures .

Isolated - Wing and Wing - Body T e s t s Ekperimental r e s u l t s , t y p i c a l of those t h a t have been obtained on t h i n wings w-ith the r e l a t i v e l y extreme w a r p associated with near - optimum design l i f t coefficients, are shown i n figure 3. The comparisons i n the figure are shown as the incremental increase i n m a x i m u m l i f t - d r a g r a t i o of t h e w,arped wings over those of corresponding f l a t wings: %or t h e three configurations shown i n the figure, which were near - optimum (CL,DES = C L , O ~ T ) warped designs at Mach numbers 2.0, 2.5, and 3.0, the t h e o r e t i c a l l y predicted increase i n maximum l i f t - d r a g r a t i o over corresponding f l a t configurations was of t h e order 2.0 as represented A s can be seen from the symbols and dashed curve, by t h e solid curve.

' which indicate t h e experimentd r e s u l t s , these extremely warped surfaces were only s l i g h t l y b e t t e r than the f l a t wings.

Two reasons have been given f o r t h i s l a c k of agreement between F i r s t , the extremely warped optimum theory and experiment (ref e 3 ) .

surface v i o l a t e s t h e linearized theory from which it i s designed, and second, t h e induced transonic flow normal t o the leading edges causes a breakdown i n the careful balance between wing slope and pressure which has been designed i n t o the wings.

i A s suggested by research on two - dimensional cambered a i r f o i l s , i n the transonic flow regime b e t t e r l i T t i n g effi-ciency may be obtained by using less than i d e a l camber combined wifh a b l k z o f Sttack. This com - promise i s e s s e n t i a l l y represented by a limiting condition such as f o r the three - dimensional wing.

CL,DEs = 0.5CL,0pT Some r e s u l t s 0-btained with t h i s l i m i t a t i o n on the magnitude of l i f t carried by t h e warped surface i s shown i n figure 4. The wings of at Mach numbers the configurations shown i n t h i s figure were designed 2.0, 2.2, and 2.6 t o carry part of the l i f t by means of thq angle of attack as mentioned previously. It can be seen by the symbols and dashed l i n e , w"aich indicate t h e experimental results, that the agree - '\ ment w i t h theory i s much b e t t e r with these wings than with t h e extremely warped optimum designs shown i n figure 3, p a r t i c u l a r l y at t h e l o w Mach a numbers, It can be noted, however, by t h e solid diamond symbol a t Mach nmiber of 2 . 6 , t h a t there i s a strong nonlinear e f f e c t of thick - ness on the comparison of experiment w i t h theory. Increasing the thick - ness r a t i o from 0.025 t o 0.040 reduced the incremental l i f t i n g e f f i - ciency of the warped surface by more than 50 percent.

The design l i f t limitation i l l u s t r a t e d i n figure 4 appears t o have

an application i n the design of thin, low - aspect - ratio configurations.

that the r e s t r i c t i o n which has been It should be pointed out, however, employed ( C~-,,DES = O , 5 C ~ , o p r ) i s a r b i t r a r y and t h a t the nature of the limitation would depend on t h e p a r t i c u l a r configuration.

Investigations such as those discussed i n figures 2 t o 4 have given an indication of the r e s t r i c t i o n s which appear t o be necessary t o get reasonable r e s u l t s from l i n e a r theory design of isolated wings.

Research presently under way w i l l hopefully lead t o greater increases i n the l i f t i n g e f f i c i e n c i e s and some inputs i n t o the thickness problem.

Nevertheless, one m i g h t have t o go contrary t o these limitations, or r e s t r i c t i o n s , i n the design of a complete airplane as indicated by the following discussion.

Complete Configuration Tests # I I n t h e design of SCAT 13 the wings had t o be thick because of vari - able sweep, and longitudinal trim was a major problem. The f l a t version of the configuration had shown an i n a b i l i t y t o t r i m t h e moments due t o O n the b a s i s of t h e s h i f t of aerodynamic center w i t h Mach number.

available information, an extremely warped wing appeared t o be necessary t o provide the zero - lift pitching moment required for trim. Figure 5 indicates quite c l e a r l y t h a t the warped surface, which w a s a basic linear - theory design, provided the desired t r i m characteristics.

I n figure 5 t h e m a x i m u m trimmed l i f t - d r a g r a t i o of the S C A T 15 configuration a t M = 2.96 i s shown as a function of s t a t i c margin It can be - seen from the f i g u r e t h a t the warped version of the hC&L.

configuration, CL,DES CL,OW, was trimmed near m a x i m u m l i f t - d r a g r a t i o over a wide range of s t a t i c margins, whereas the f l a t version could not be trimmed t o the required l e v e l of aC,.$CL of - 0.15. Pos - s i b l y t r i m could have been obtained on the flat version with a larger t a i l but the l o s s i n l i f t - d r a g r a t i o would have been substantial.

A s pointed out i n paper no. 4 by Lloyd T. Goodmanson and h i s col -

leagues, the SCAT 15 airplane had the highest aerodynamic performance l e v e l of any of the SCAT configurations. SCAT 15 was able t o establish \r t h i s r e l a t i v e l y high performance l e v e l , i n s p i t e of high thickness r a t i o and extreme wing w a r p , because of favorable component interference effects. Other than t h e planform of the wing, the basic ingredients of these favorable interference e f f e c t s were a positive l i f t produced by the nacelle arrangement under t h e wing and an e f f i c i e n t l i f t produced by outboard t a i l s operating i n an upwash f i e l d . (See ref. 6.) The r e s u l t s of these favorable interference e f f e c t s are shown i n figure 6.

I n figure 6, variation of the r a t i o C D / C D , ~ H G - B O D ~ with l i f t coefficient i s shown f o r the SCAT 15 a t These curves show M = 2.96.

t h a t when the nacelles were added under the wing - body, the drag coef - f i c i e n t at zero l i f t increased by about 14 percent. A s a r e s u l t of the favorable l i f t produced by the nacelles near a l i f t coefficient of about Q.1, however, the drag coefficient of the wing - body - nacelle arrangement i s the same a s t h a t of the wtng-body configuration. Furthermore, when the horizontal and v e r t i c a l t a i l s (represented by the shaded region) were added t o make the complete configuration, the drag coefficient a t zero l i f t was about 28 percent greater than t h a t of the wing - body con - figuration. These r e s u l t s show t h a t the combined favorable interference e f f e c t s reduced the drag l e v e l of the complete configuration t o t h a t of the wing - body configuration a t a l i f t coefficient of about 0.1. The e f f e c t s i l l u s t r a t e d can be interpreted t o represent an increase i n Further research l i f t i n g efficiency due t o the added components.

appears necessary t o define more c l e a r l y the f a c t o r s which contribute t o these favorable interference e f f e c t s i n order t h a t they may be e f f i - c i e n t l y applied t o other configurations.

i From the r e s u l t s shown i n figures 5 and 6 f o r the SCAT 15, it can be concluded t h a t desirable t r i m characteristics and supersonic effi - ciency may be obtained by making use of a warped wing, derived basically from l i n e a r theory, i n combination with favorable component l i f t -

interference effects. A s reported i n a previous paper (paper no. 4

by Goodmanson e t al.), these aerodynamic advantages of the SCAT 15 could not overcome the s t r u c t u r a l penalties associated with the concept.

There i s reason t o hope, however, t h a t a wing planform and surface can J be found t o employ the aerodynamic ,@&ant@gw p i t h o u t severe structural penalties and without the low - speed Gitch-up problems found on highly swept arrow wings. Part of the present research e f f o r t i s directed toward the use of arbitrary planform methods t o find such a wing.

Analytical Methods Digital computer programs which are available f o r the solution of a r b i t r a r y planform problems are outlined i n figure 7. The programs w i l l solve a variety of problems as indicated i n the figure. M r . Ralph Carmichael of the N A S A Ames Research Center has formulated a more con - ' p l e t e program f o r use i n the design of an ogee wing planform f o r com - parison with the S C A T 17.

Although l i n e a r theory has shortcomings, it has been useful i n pointing out the design areas of high performance and, properly limited, provides warped surface shapes which can be worked with t o obtain high l i f t i n g efficiency.

The programs indicated i n figure 7 have been used t o design d e l t a and ogee configurations f o r comparison a t a Mach number of 2.2. Because low, - aspect - ratio wings which were considered f o r these con - of the thin, figurations, a moderate degree of wing w a r p (approaching the limitation CL,DES = O.?CL,O~T) was considered permissible and desirable. Figure 8 indicates the e f f e c t of wing warp on the performance of these two con - figurations a t a Mach number of 2.2 and indicates the variation of l i f t - drag r a t i o with l i f t coefficient as predicted by theory and determined by experiment. It can be seen from the figure that wing warp improved the performance of both configurations and that there w a s reasonably good agreement between theory and experiment f o r both f l a t and warped versions of the configurations. It should be pointed out t h a t the values of lift - drag r a t i o shown i n figure 8 do not represent trimmed values. The t r i m considerations f o r the d e l t a and ogee configurations are discussed i n paper no. 12 by Robert T. Taylor e t al.

\ CONCLUDING RFSIARKS I n summary, consideration of highly swept warped - wing arrangements a moderate warped surface at angle of indicates that, f o r t h i n wings, attack m y be desirable. Wing warp offers favorable t r i m character - i s t i c s and i n combination with component l i f t interference e f f e c t s can lead t o a high performance configuration. Analytical methods a r e being employed t o find a configuration which can use these aerodynamic advan - tages without severe structural penalties.

FBFERENCES 1. Carlson, Harry W.: Aerodynamic Characteristics a% Mach Number 2.05 of a Series of Highly %ept Arrow Wings Employing Various Degrees NASA T M X-332, 1960.

of Twist and Camber.

Aerodynamic 2. Hasson, Dennis F., Fichter, Ann B., and Wong, Norman: Characteristics at Mach Nunibers From 1.6 t o 2.8 of 74' Swept Arrow Wings With and Without Camber and Twist. N A S A T M X - 8, 1959.

30 Hasson, Dennis F., and Wong, Norman: Aerodynamic Characteristics at Mach Numbers From 2.29 t o 4.65 of 80° Swept Arrow Wings With and NASA TM X-175, 1960.

Without Camber and Twtst.

The Problem of Obtaining 4. Brown, Clinton E., and McLean, Francis E. : High Lift - Drag R a t i o s at Supersonic Speeds. Jour. Aero. Sci., VOL 26, no. 5, ~ a y 1959, pp. 298-302.

Theoretical.

5. Brown, Clinton E . , McLean, F . E., and K l u n k e r , E. B.: and Experimental Studies of Cambered and Twisted Wlngs Optimized Advances i n Aero. Sci., vol. 3, for Flight a t Supersonic Speeds.

Pergmon Press (New York) , 1961, pp. 415-430.

Aero - 6. Robins, A. Warner, Spearman, M. Leroy, and Harris, Roy V.: dynamic Characteristics a t Mach Numbers of 2.30, 2.60, and 2.96 of a Supersonic Transport Model With a Blended Wing - Body, Variable - %eep Auxiliary Wing Panels, Outboard T a i l Surfaces, and a Design N A S A TM X-815, 1963.

Mach Number of 2.6.

DRAG - DUE - TO - LI FT FACTORS

M = 3.0 I .6 - I .2

- dCD

acf . 8 OPTIMUM WING .4 I I 1 I I I I I I I 0 60 64 68 72 76 80 84 LEADING - EDGE SWEEP ANGLE, ALE, DEG Figure 1 EFFECT OF DESIGN LIFT COEFFICIENT ON THEORETICAL LIFT - DRAG RATIO = 0.062 LIFT COEFFICIENT, CL Figure 2 INCREASE IN MAXIMUM LIFT - DRAG RATIO DUE TO WING WARP CL, DES R CL,OPT ; R E: 4 X 1 0 6 ; CIRCULAR - ARC THICKNESS DISTRIBUTIONS, t/c =0.025 TO 0.030 I .ot I I I 0 1.5 2 .o 2.5 3.0 MACH NUMBER, M INCREASE IN MAXIMUM LIFT - DRAG RATIO DUE TO WING WARP CL,DES~ 0.5 CL,OpT ; R = 4 X106 ; CIRCULAR - ARC THICKNESS DISTRIBUTIONS, t / c =0.025 TO 0.030 I .o EXPERIMENT /

* ; = 0.04

I I

LI

I 0 1 . 5 2.0 2.5 3.0 MACH NUMBER, M Figure 4 a FAVORABLE TRIM CHARACTERISTICS DUE TO WING WARP M=2.96; R=3.7Ix1O6; NACA 65A-SERIES, t/c=0.045 AT ROOT TO 0.035 AT TIP I I I I T I I -.I2 -.I6 -.20 -.24 0 -.04 -.08 dC rn /dCL Figure 5 INTERFERENCE EFFECTS ON AIRPLANE CONFIGURATION WITH WARPED LIFTING SURFACES M = 2.96; R = 3.71 X IO6; CL, DES FJ CL,OPT CD CD, W I N G - BODY

z --

c

- -----'z-:----=

1 . 0 COMPLETE CONFIGURATION

1 LWING - BODY - NACELLES

.6 I I I I I I I 0 . 0 2 .04 .06 .08 . I O . I 2 .I4 LI F T COEFFl CI ENT, CL Figure 6 \ .

COMPUTER PROGRAMS AVAILABLE FOR SOLUTION OF ARBITRARY - PLANFORM PROBLEMS I. GIVEN: TOTAL LI FT 1. GIVEN: LIFT DISTRIBUTION SOLUTION : OPTIMUM LIFT SOLUTION: SURFACE SHAPE DISTRIBUTION DRAG - DUE - TO - LI FT OPTIMUM SURFACE FPlCTOR SHAPE 2. GIVEN: SURFACE SHAPE SOLUTION: LIFT DISTRIBUTION TO - LIFT DRAG-DUE- FACTOR DRAG - DUE - TO - LIFT FACTOR LI FT - CURVE SLOPE GIVEN: COMBINATION OF SOLUTIONS FROM PROGRAMS I AND I1 SOLUTION: OFF - DESIGN POLARS EFFECT OF WING WARP ON PERFORMANCE OF DELTA AND -0% EE CON F I G U R AT I ON S M = 2.2; R= 3 X lo6 ; CL,DES 0.5 CL,OPT; CIRCULAR - ARC THIEKNESS DISTRIBUTIONS , t i c = 0 ~ 0 2 TO 0.03 THEORY EXP .. - FLAT d ---_ I 7 WARPED L L - - D D I 0 .08 .I6 .24 .32 LIFT COEFFICIENT, CL Figure 8 i 8. STTJDDS OF SKIN F R I C T I d N ~ A p % U J & R ~ ~ C $€XEDS B y K. R. Czarnecki, Mary W. Jackson, and W i l l i a m J. Monta NASA Langley Research Center SUMMARY A b r i e f review has been made of the most significant r e s u l t s from recent research on turbulent boundary layers applicable t o the supersonic The review indicates that: t r ns o r t . High Reynolds number flat - plate

P P

boundary - layer skin f r i c t i o n with heat transfer w i l l tend t o agree with the predictions of the Sommer and Short T ' method and those theories or methods which are i n approximate agreement with this T ' method. Three - dimensional boundary - layer flow probably w i l l increase skin - friction drag over tha - t expected from two - dimensional flat - plate theories or experiment. I n general, l o c a l flow conditions and the proper correlating parameters w i l l have t o be used f o r the accurate prediction of drag due t o surface roughness.

k u/7-#8(

INTRODUCTION The successful design of a supersonic transport i s c r i t i c a l l y dependent upon making a satisfactory estimate of the airplane drag.

One of t h e major components of airplane drag i s skin f r i c t i o n . Conse - quently, it becomes mandatory t o have reliable methods f o r estimating skin - f r i c t i o n drag. Detailed examination of the available theories and experimental data, however, reveals many deficiencies and lack of proper verification. The purpose of t h i s paper i s t o present and discuss some of the highlights from recent research being made on turbulent boundary layers t o f i l l i n some of the voids i n the desired information. Specif - i c a l l y , a b r i e f review w i l l be made o f - t h e skin - friction l e v e l on smooth ,surfaces a t high Reynolds numbers, the e f f e c t s of some three - dimensional boundary - layer flows on skin f r i c t i o n , and t h e e f f e c t s of surface rough - ness on the skin - friction drag.

S Y M B O L S drag coefficient CD roughness drag coefficient based on integrated pressure ( ' ~ 9 qk) av distribGtions, roughness height, and average dynamic pres - sure within boundary layer over roughness height roughness drag coefficient based on integrated pressure

( cD, ¶&

distributions, roughness height, and free - stream dynamic pre s sure average skin - friction drag coefficient based on wetted sur - CF .

face area and free - stream dynamic pressure transformed average skin - friction coefficient CF" ') pressure coefficient cP k roughne s s height 6* boundary - layer displacement thickness k/6* r a t i o of roughness height t o boundary - layer displacement t h i ckne s s free - stream Mach nulllber M, r model radius measured normal t o body axis free - stream Reynolds number per foot RFT Reynolds number based on free - stream conditions and distance R, t o v i r t u a l origin or t o body nose transformed Reynolds n&er

sc"

X distance from v i r t u a l origin o r a x i a l distance from body nose adiabatic w a l l temperature Taw free - stream stagnation temperature T t average model wall temperature T W f ree - stream s t a t i c temperature T W temperature r a t i o

Tw / T W

T ' reference temperature a angle of a t t a c k \ RESULTS AND DISCUSSION Skin Friction at High Reynolds Numbers I n a recent report by Peterson (ref. 1) a large amount of turbulent skin - friction data f o r smooth f l a t p l a t e s were correlated and compared with several methods of predicting the skin f r i c t i o n . The analysis shows that, by a s m a l l margin, the Sommer and Short T I , o r reference tempera - ture, method may give the most r e l i a b l e estimate of the value of skin f r i c t i o n .

One of the major deficiencies existing i n t h i s correlation i s the lack of r e l i a b l e experimental skin - friction r e s u l t s a t the high Reynolds numbers t h a t w i l l be encountered i n f l i g h t by the supersonic transport. Figure 1, therefore, presents some f l a t - p l a t e skin - friction r e s u l t s with Reynolds numbers up t o about 150 X 10 6 recently obtained on the sidewall of the Langley 4- by 4-foot supersonic pressure tunnel by boundary - layer profile surveys and l o c a l skin - friction balances a t Mach numbers of 1.61 and 2.20. The data were obtained within the constant Mach number rhombus of the t e s t section a t s i x o r seven Reynolds numbers per foot a t zero heat transfer. The data were reduced t o coefficient form by a method u t i l i z i n g both momentum surveys and l o c a l surface shear.

(See ref. 1.) I n the figure, the ordinate CF i s the average skin -

f r i c t i o n coefficient, and % i s the Reynolds number based on free -

stream conditions and the distance t o the v i r t u a l origin. Included i n the figure are the curves determined by the Sommer and Short T ' method and several of the other more r e l i a b l e theories o r empirical methods t h a t can also handle heat transfer.

Inspection of the p l o t s i n figure 1 indicates t h a t these high Reynolds number f l a t - p l a t e skin - friction data are i n good general agree - ment with the l e v e l of the Sommer and Short T ' curve and t h e various other theories represented here, and that the variation of CF with Rx~ a t these high Reynolds numbers w i l l tend to follow the predictions of these theories.

I n figure 2 are presented some preliminary high Reynolds number skin - f r i c t i o n r e s u l t s obtained on an ogive - cylinder model by the usual tech - ",nique of force balance minus nose - and base - pressure drag i n the Langley * ' 9- by 6 - foot thermal structures tunnel a t These are three - M, = 2.98.

dimensional boundary - layer flow data and are t h e only high Reynolds num - ber r e s u l t s available which include the e f f e c t s of heat transfer. The data were obtained with fixed transition at two stagnation temperature conditions and a model preheat condition to approximately zero heat transfer. The a c t u a l Reynolds number range w a s from about 36 X 10 6 t o 200 X LO6 and the wall - to - stream temperature r a t i o Tw/Tm varied 9 , ) , J i i I , I 4 + i s I < from 1.4 t o 2.5. Because of the varying amounts of heat transfer involved, the experimental data are presented i n terms of the coeffi -

cients CF* and %* which are the transformed

equivalent incompressible - flow zero - heat - transfer skin - friction coeff i c i n t s and transformed Reynolds numbers, respectively. The data were reduced t o this form by the Sommer and Short T ' method. (See ref. 2.) For pur - poses of comparison, t h e standard Von Karman - Schoenherr incompressible - flow zero - heat - transfer skin - friction curve is included as a solid line.

Examination of the data i n figure 2 indicates t h a t the l e v e l of the experimental data is below the Von Kazman-Schoenherr curve. The f a c t that the experirnental r e s u l t s l i e approximately 10 percent below the Von Karman - Schoenherr curve, whereas, f o r reasons t o be i l l u s t r a t e d shortly, the data were expected t o l i e somewhat above the reference curve, i s ascribed t o insufficiently accurate knowledge of e i t h e r the model nose or base drag. T h i s error i n the experimental r e s u l t s w i l l have l i t t l e i f any e f f e c t on the slope of the curve, and the slope of the experimental variation of CF* with Rx* i s i n good agreement w i t h the theoretical prediction. Consequently, the conclusion can be made that these experimental three - dimensional boundary - layer r e s u l t s at b & , = 2.98 w i t h heat transfer (fig. 2) and the lower Mach number f l a t - p l a t e zero - heat - transfer data (fig. 1) indicate that t h e Sommer and Short T' method and other methods giving similar r e s u l t s can be used t o extrapolate low Reynolds number model data t o the high Reynolds nmbers and heat - transfer r a t e s of the supersonic transport.

Skin Friction i n Three - Dimensional Boundary Layer Airplane skin f r i c t i o n has often been estimated on the basis of two - dimensional f l a t - p l a t e theory o r experiment. In the discussion of figure 2, however, mention w a s made that the skin - friction drag of the ogive cylinder w a s expected t o be higher than that indicated f o r the body by f l a t - p l a t e theory. Bqerimental evidence as t o why this trend w a s expected i s shown i n figure 3 . These data were obtained by the standard force - balance and nose - base - pressure technique (with fixed t r a n s i t i o n ) on a minim-drag type of body w i t h a fineness r a t i o of 13 i n a number of Langley f a c i l i t i e s over a Mach number range from 0.60

t o 3.95, a true Reynolds nuniber range from 3 X 10 6 t o 13 X lo6, and at

zero heat transfer. In order t o simplify the comparisons, the data have been transformed t o the incompressible - flow form by the Sommer and Short T t method as i n figure 2 and are compared w i t h t h e Von K a m - Schoenherr reference curve.

The r e s u l t s presented i n figure 3 show that the drag of a three - dimensional axisymmetric body i s al'Ftays higher than that predicted by the Sommer and Short T ' method f o r a flat p l a t e of equal wetted surface area. Similar drag increases were noted f o r other axisymmetric bodies i n this p a r t i c u l a r investigation and on an ogive - cylinder model ( i n r e f . 3 ) Furthermore, preliminary examination of wake surveys recently made behind flat and/or cambered and twisted sweptback wings has revealed the p o s s i b i l i t y of increased overall skin f r i c t i o n f o r the case i n which small spanwise boundary - layer flow may e x i s t on a three - dimensional body such as a wing. In view of all these indications it becomes apparent t h a t it may be d i f f i c u l t , i f not impossible, t o achieve a skin - friction drag l e v e l on an airplane equal t o t h a t attainable on a f l a t p l a t e of equal wetted area at i d e n t i c a l free - stream conditions. E s t i m a t e s of 7st of these three - dimensional boundary - layer flow e f f e c t s can be made, -khou& not with the saae order of accuracy as f o r the basic flat - plate skin f r i c t i o n .

Drag of Surface Roughness Elements Roughness drag on t h e supersonic transport can be a problem and can lead t o a decrease i n performance. Considerable experimentation on the e f f e c t s of surface roughness on supersonic turbulent skin f r i c t i o n has been accomplished. Some of the work of NASA on t h i s subject has been The present s t a t u s of knowledge per - reported i n references 4 t o 7.

taining t o the e f f e c t s of surface roughness on turbulent skin f r i c t i o n

a t supersonic speeds i s i l l u s t r a t e d b r i e f l y i n figures 4 t o 7 relating

t o fomard- and rearward - facing steps and surface waves, which represent I n the basic shapes from which most roughness shapes can be derived.

figure 4 i s presented the drag coefficient f o r a two - dimensional forward -

facing step roughness and i n figure 5 i s presented t h e drag coefficient f o r a two - dimensional rearward - facing step roughness, both as a function of the r a t i o of step height t o boundary - layer displacement thickness k/6)e. These data were obtained at a Mach number of 2.20 over a range of Reynolds number per foot a t zero heat t r a n s f e r by means of pressure distributions determined on a variable - height block mounted i n the thick boundary layer on the sidewall of the Langley 4 - by 4 - foot supersonic pressure tunnel, and thus exclude chariges i n skin f r i c t i o n on streamwise aiigfaces ahead of and t o the rear of the steps.

For the forward - facing step roughness ( f i g . 4) the drag coefficient w a s constant when it w a s based on the roughness height and the average dynamic pressure within t h e boundary layer over t h e height of the step (the undisturbed boundary layer) except when t h e step w a s immersed within Within t h i s the lowest 1 percent of the t o t a l boundary - layer thickness.

Similar results p a r t of t h e boundary layer the flow becomes subsonic.

w e r e obtained at a Mach number of' 1.61 except t h a t the correlating drag coefficient value w a s higher, about 0,8, and indicated that an e f f e c t of Mach number is present.

For the rearward-facing step roughness (fig. 5 ) , the drag coeffi - cient was constant when based on the roughness height and t h e free - stream dynamic pressure, the relation again breaking down f o r the step immersed i n the lowest 1 or 2 percent of the t o t a l boundary - layer thick - Once more, decreasing the Mach number t o 1.61 increased the cor - ness.

relating drag coefficient t o 0 . 1 8 and again indicated an influence of the free - stream Mach number.

I n figure 6 i s depicted the pressure distribution (taken from ref. 6) over a sinusoidal - wave type of surface roughness.

The wave w a s one of a number b u i l t i n t o the cylindrical part of an ogive - cylinder model. The free - streamMach number i s 2.01; the Rm range i s from 1.25 x 106 t o 7.08 x 106. The measured surface p r o f i l e i s shown, with \) a greatly exaggerated ordinate scale, i n the lower part of the figure.

The dashed l i n e represents two - dimensional linearized - theory calcula - t i o n s using measured l o c a l flow conditions (experiment, smooth body) j u s t outside the boundary layer and the measured surface profile.

The most salient feature of figure 6 is t h a t at high values of Rm, when the r a t i o of boundary - layer thickness t o roughness height i s l e a s t , it i s possible t o predict with good accuracy the pressure distribution and, hence, the wave drag of t h i s type of roughness. A s Rm decreases, however, and the r a t i o of roughness height t o boundary - layer thickness increases, there i s an onset of flow separation within the recessed or trough p a r t s of the roughness elements as evidenced by the decreases i n this area, which r e s u l t s in a decreased drag and i n more d i f -

i n 5

f i c u l t y i n predicting the drag.

It i s apparent that the drag of a l l ty-pes of surface roughness i s dependent upon l o c a l stream conditions, but t h a t the controlling drag parameters w i l l vary with the type of roughness.

A s shown i n refer - ence 7 and verified even more vividly by recent pressure t e s t s of surface roughness i n the transonic Mach number area, most of the surface rough - ness drag at supersonic speeds i s wave drag with only small increases i n boundary - layer momentum losses.

I n the l i g h t of the knowledge derived from figures 4 t o 6 i n this series on surface roughness, it should be apparent from theoretical con - j siderations that the e f f e c t s of heat t r a n s f e r on roughness drag gener - a l l y w i l l be small. This deduction i s investigated by analysis of data such as those presented i n figure 7. The data were obtained f o r two roughness configurations on an ogive - cylinder model i n the previously mentioned Mach number 2.98 tests i n the Langley 9 - by &foot thermal structures tunnel. The ordinate CD i s the coefficient which repre - sents the combined drag of the smooth - body skin f r i c t i o n plus surface roughness based on wetted surface area and stream flow conditions. The abscissa i s the average model - wall - free - stream temperature ratio.

From the r e l a t i v e constancy of the increment i n drag between the smooth model and the roughness configurations over the temperature - r a t i o range, it i s readily established that the data confirm the pre - viously mentioned deduction that heat transfer w i l l have r e l a t i v e l y l i t t l e e f f e c t on the increment i n drag due t o surface roughness.

In order t o provide an interpretation of these roughness r e s u l t s as applied t o a supersonic transport, figure 8 was prepared. Flight at 60,000 feet at M, = 2.20 and a near 0 ' was assumed. The drag coef - f i c i e n t s f o r a forward - and rearnard - facing step roughness are plotted a s a function of step height. The coefficients are f o r 1 l i n e a l foot of step roughness located on the wing of the supersonic transport having a reference wing area of 2,300 square feet. The drag coefficients are based on t h i s wing area and free - stream dynamic pressure.

For the rearward - facing step roughness, the drag coefficient based on the airplane reference conditions increases l i n e a r l y with step height i n t h e range of k/6* f o r which t h e correlating drag coefficient i n figure 5 was constant. The step drag i s also independent of distance O f course, f o r the very s m a l l heights, the along the wing surface.

drag contribution tends toward zero somewhat more rapidly, but the trend can hardly be distinguished.

For the fomard-facing step roughness, the drag coefficient increases with height not only because of t h e increased f r o n t a l area but a l s o because the average dynamic pressure increases on the f r o n t face. The drag coefficient, therefore, increases w i t h step height at a f a s t e r r a t e than linearly. For t h i s roughness, there i s an e f f e c t of step location due t o the variation with surface distance of the average dynamic pressure i n the boundary layer over a given height.

As a further i l l u s t r a t i o n , it i s estimated that a thousand f e e t of each type of step roughness O.O>O inch high scattered over the wing w i l l T h i s drag contribution increase the airplane drag by nearly 4 percent.

can be readily decreased by requiring smaller surface mismatch and ~ lesser lengths of roughness. A surface waviness of height t o length

r a t i o of only - of 1 percent, but covering t h e complete airplane wetted

surface area between stringers and ribs, can increase airplane drag by 2 percent. T h i s type of roughness drag may be somewhat more d i f f i c u l t t o a l l e v i a t e t o acceptable values.

i , SUMMARY O F RESULTS From t h i s b r i e f review of recent turbulent boundary - layer research applicable t o the supersonic transport it may be summarized that: High Reynolds nuniber f l a t - p l a t e boundary - layer skin f r i c t i o n with heat t r a n s f e r w i l l tend t o agree with the predictions of the Sommer and Short T ' method and those theories o r methods which are in approximate agreement with t h i s T ' method.

Three - dimensional boundary - layer flow on the supersonic transport skin - friction drag over that e q e c t e d from two - probably w i l l increase

theories or experiment . )

dimensional. flat - plate In general, l o c a l flow conditions and t h e proper correlating param- roughness shape, w i l l have t o be used for t h e eters, which vary with accurate prediction of drag due t o surface roughness.

i . > ' A Comparison of Experimental and Theoretical 1. Peterson, John B., Jr. : Results f o r t h e Compre s sible- Turbulent -Boidary-Layer Skin Friction N A S A TN 0-1795, 1963.

With Zero Pressure Gradient.

2. Sommer, Simon C., and Short, Barbara J . : Free - Flight Measurements of Turbulent - Boundary - Layer Skin Friction i n t h e Presence of Severe NACA TN 3391, Aerodynamic Heating at Mach Numbers From 2.8 t o 7.0.

7 1 Czarnecki, K. R., and Monta, W i l l i a m J . : Boundary - Layer Velocity Pro - ' f i l e s and Skin Friction Due t o Surface Roughness on an Ogive Cylinder of 1.61 and 2.01. NASA TN D - 2048, 1963.

at Mach Numbers Inves - 4. Czarnecki, K. R., Robinson, Ross B., and Hilton, John H., Jr.: t i g a t i o n of Distributed Surface Roughness on a Body of Revolution a t a Mach Number of 1.61. NACA TN 3230, 1954.

3 . Sevier, John R., J r . , and Czarnecki, K. R,: Investigation of Effects of Distributed Surface Roughness on a Turbulent Boundary Layer Over a Body of Revolution at a Mach Number of 2.01, NACA TN 4183, 1958.

6. Czarnecki, K. R., Sevier, John R., Jr., and Camel, Melvin M.: Effects of Fabrication - Type Roughness on Turbulent Skin Friction at Supersonic Speeds. NACA TN 4299, 1958.

7. Czarnecki, K. R., and Monta, W i l l i a m J . : Pressure Distributions and Wave Drag Due t o Tw.o-Dimensional Fabrication - Type Surface Roughness on an Ogive Cylinder at Mach Numbers of 1.61 and 2.01. N A S A TN D - 835, 1961.

SKIN - FRICTION DRAG IN WALL BOUNDARY LAYER EXPERIMENT THEORY SOMMER AND,SHORT T'

_------ MONAGHAN T

.003

- --

AND VAN DRIEST CF .002

.003 r

CF .002 Figure 1 SKIN - FRICTION DRAG ON OGIVE CYLINDER Mm = 2.98 EXPERIMENT .005 Ttl O F 0 300 300 (TW a Taw) 0 600 VON KARMAN-SCHOENHERR CF* .002 i I I

i o 60 i o ' : o ' r ~ o x l o 6 .ool,o 20

RX* Figure 2 SKIN - FRICTION DRAG OF THREE - DIMENSIONAL BODIES FINENESS RATIO, 1 3 EXPERl MENT M, 0 0.60 0 2 . 0 1 0 2.50 A 2.96 b 3.95 .003 ON KARMAN -SCHOENH - O o 2 - .oo!7 RX* 20 x I 0 6 Figure 3 DRAG CORRELATION FOR FORWARD - FACING STEP

*9 P M m = 2.20

0 .2 .4 . 6 .8 1.0 1.2 1.4 1.6 1 . 8 2 . 0 2 . 2 k/8*

Figure 4

DRAG CORRELATION FOR REARWARD - FACING STEP M , = 2.20 EXPERIMENT RFT 0 0.7~10~ 0 1 . 1 0 2 . 3 A 3 . 4 AIRFLOY b 4.4 h 5.4 a 6.4

@@%kzz+-

I I I I I I I I I I I 0 . 2 . 4 . 6 .8 1 . 0 1 . 2 1 . 4 1 . 6 1 . 8 2 . 0 2.2 k/8* Figure 5 PRESSURE DISTRIBUTION OVER WAVE - TYPE ROUGHNESS -.lSy M,=2QI; kz0.053 IN.

EXPERIMENT a 5.97 -.08 0 4.7% A 3.62 b 2.41 MOOTH BODY .08 .I6 0 45 46 47 48 49 x,in.

Figure 6 EFFECT OF HEAT - TRANSFER RATIO ON ROUGHNESS DRAG Ma= 2.98; R,= 94 x IO 6 .003t EXPERIMENT 0 SMOOTH CD 0 COMBINATION S T E P u .002 ~ - - O WAVEM L.1 I I I I I I I 0 1 . 2 1 . 6 2.0 2.4 2.8 3 . 2 3 . 6 4 . 0 T./T,, Figure 7 DRAG OF I FOOT OF ROUGHNESS ON SUPERSONIC TRANSPORT Ma=2.20; ALTITUDE, 60,000 FT; WING AREA, 2,300 SQ FT; a=Oo CD 80 AIRFLOW 80 AIRFLOW

--

401 -'>

CD

-4 0. .02 .04 .06 k , IN. D8 .IO -12 -14

.02 .04 .06 D8 .IO -12 -14 0.

k , IN.

Figure. 8 9. SOME FACTORS AFFECTING T u R B m SKIN FRICTION w.& AT SWERSONIC SPEEDS By Albert L. Braslow, John B. Peterson, Jr., and Donald I. McRee . & NASA Langley Research Center .

SUMMARY Preliminary r e s u l t s are presented f o r two investigations of f a c t o r s affecting turbulent skin f r i c t i o n at Mach 3 : an investigation of the use of air injection i n t o the turbulent boundary layer through rearward - facing inclined flush and step s l o t s t o reduce skin f r i c t i o n , and an investiga - t i o n of the contribution t o skin - friction drag of three - dimensional rough - ness t r i p s used t o f i x boundary - layer t r a n s i t i o n .

A JGPird

SYMBOLS i model drag coefficient CD t o t a l airplane cruise drag coefficient with air CD,INJECTION injection t o t a l airplane cruise drag coefficient without air CD,NO INJECTION injection skin - friction drag coefficient based on wing area CD,F skin - friction drag coefficient based on wetted area cF cT thrust coefficient k . height of roughness mass rate of flow of injection air &

k o mass rate of flow of free - stream air, p , % S

M Mach number b 3 . , ' '.

roughness Reynolds number, = ukk -

Rk free - stream Reynolds number per foot S flat - plate reference area velocity at top of roughness uk

urn free - stream velocity

W width of t r a n s i t i o n t r i p X distance from leading edge kinematic v i s c o s i t y at top of roughness "k free - stream density Po3 Po3 free - stream viscosity Subs c r i p t s : 0 without injection at zero angle of a t t a c k a = O t h , theory A I R INJECTION Introduction The flow over t h e surfaces of the supersonic transport at super - A large percentage of sonic cruise w i l l be almost e n t i r e l y turbulent.

the t o t a l airplane drag w i l l then be due t o turbulent skin f r i c t i o n . It would be most desirable, of course, t o find a means f o r obtaining exten - sive regions of laminar f l o w t o reduce t h e skin - friction contribution t o the t o t a l drag. Althoughthis has been a goal ever since Prandtl formu - l a t e d the concept of t h e boundary layer, even today very l i t t l e laminar Research on laminar flow, how - flow i s attained on airplanes i n flight.

For example, the Nortbrop Corporation, under ever, i s still continuing.

contract t o the U.S. A i r Force, i s currently f l i g h t t e s t i n g the i .

laminar - flow control of the X - 2 1 airplane. Results of t h i s program should provide considerable information on the p r a c t i c a l aspects of using suction boundary - layer control at subsonic speeds. Research i s a l s o planned at Langley and at Northrop, under a combined USAF, FAA, and N A S A contract, on laminar - flow control at supersonic speeds. Such e f f e c t s as shock - boundary - layer interactions, wing sweep, noise, and vibrations on laminar - flow control w i l l be investigated.

Since sufficient p r a c t i c a l experience w i t h laminar - flow control i s not yet available at any speed and considerable fundamental research on i t s application at supersonic speeds i s s t i l l required, laminar-flow con - t r o l does not appear feasible f o r the first generation supersonic trans - Any p o s s i b i l i t y of reducing the l e v e l of the tmbulent skin f r i c - > p o r t .

ition, therefore, should not be overlooked.

Previous research has shown that injection of air i n t o the turbulent boundary layer through porous surfaces w i l l significantly reduce the skin f r i c t i o n . The use of extensive porous surfaces on an airplane, of course, i s not a very p r a c t i c a l arrangement. I n addition, such injection would result i n a complete loss of a l l the momentum available i n t h e injected air and a net drag increase rather than a decrease could occur. A pre - liminary investigation was made recently a t a Mach number of 3 t o deter - mine whether injection through rearward inclined s l o t s would reduce the skin f r i c t i o n and a t the same time provide some recovery of the momentum of the injected air.

Discussion Rearward - facing inclined f l u s h s l o t s and rearward - facing inclined step s l o t s of various widths were tested near t h e leading edge of a flat - Pertinent results f o r the most effective plate model. (See f i g . 1.)

s l o t configuration so far tested are presented i n figure 2 where f l a t - The p l a t e drag coefficient i s plotted against a mass - flow parameter.

lower curve presents t h e variation of the measured CD with rate of air This i n j e c t i o n through a 0.0085 - inch rearw'ard-facing inclined f l u s h s l o t .

drag coefficient includes the reduction i n skin f r i c t i o n as well as the The upper curve indicates the momentum recovered from the injected air.

that could be obtained i f the same air were ejected reduction i n CD The f a c t t h a t t h e overboard i n the usual manner through a sonic nozzle.

drag i s lower when the air i s injected i n t o the boundary layer indicates a reduction i n skin f r i c t i o n .

A possible application of s l o t injection t o one of the SCAT config - A four - engine S C A T 16 configuration urations i s presented i n figure 3 .

a t a Mach number of 3 and an a l t i t u d e of 65,000 f e e t was selected.

The abscissa i s the air injection f l o w i n pounds per second through inclined s l o t s on selected p a r t s of the airplane where the air used is p a r t of It w a s estimated that t h e i n l e t bleed air at the t h e inlet bleed air.

assumed f l i g h t condition amounts t o about 100 pounds per second. The ordinate i s the r a t i o of the t o t a l airplane drag coefficient w i t h injec - t i o n over t h e indicated p a r t s t o the t o t a l drag coefficient without injection. The denominator includes the reduction i n drag obtained from the expected thrust of all the i n l e t bleed air. U p t o the first v e r t i - c a l l i n e , it w a s assumed that p a r t of t h e bleed air, as indicated on the abscissa, i s injected through a s l o t near the leading edge of each of the four nacelles. The momentum recovered w i t h the s l o t s and the thrust l o s t from the decreased mass flow through t h e i n l e t bleed nozzles i s included i n the curve. The m a x i m injection flow rate over the nacelles

of about 2’3 pounds per second i s equivalent t o a flow parameter of 1.00 -

) the maximum plotted i n figure 2. A t t h i s injection f l o w rate, an over - If additional all reduction i n drag of about 1.1percent i s indicated.

bleed air can also be injected over the v e r t i c a l and horizontal tails up t o the same m a x i m u m flow rate of 1.00, the net drag reduction i s same rate of injection over the aft part of the fuse - 2.3 percent. The lage behind the passenger compartment increases the drag reduction t o about 3 percent. If f u r t h e r research indicates that the skin f r i c t i o n w i l l continue t o decrease w i t h increased flow rate (as f i g . 2 indicates may be possible), f u r t h e r gains will be possible by injecting a greater amount of the available boundary - layer bleed air. I n addition, perhaps other sources of air, such as part of the cabin cooling air, can be used f o r injection over other surfaces of t h e airplane. The f e a s i b i l i t y of doing this depends upon many considerations, one of these being the weight increase required f o r any additional ducting. The point t o be made, however, is t h a t t h e t o t a l drag reductions that a r e indicated on t h i s figure warrant f u r t h e r study of the use of air injection for reducing the turbulent skin f r i c t i o n .

BOUNDARY - LAYER TRIPS Introduction I n paper no. 8 by K. R. Czarnecki, Mary W. Jackson, and W i l l i a m J.

Monta, recent turbulent skin - friction data at high Reynolds numbers a r e presented t o assist i n the extrapolation of wind - tunnel model drag data To obtain fully turbulent flow over t h e wind - t o f l i g h t conditions.

tunnelmodels, it i s almost always necessary t o f i x t r a n s i t i o n i n the To do this, distributed leading - edge regions of the various components.

roughness t r i p s are most commonly used. The use of trips, however, requires a knowledge of t h e i r e f f e c t on the model drag. Ideally, the t r i p w i l l cause an increase i n drag due t o the forward movement of tran - s i t i o n only and w i l l have no additional drag itself. If t h e t r i p i s too 194 ,- \ / I ..

, > ' > i ..I :*: * large, i n e i t h e r width o r roughness height, it will an additional amount of drag t o the model. This drag increment i s very d i f f i c u l t t o and t h e use of such t r i p s should be avoided. I n estimate accurately, order t o provide some information on the drag contribution of various three - dimensional boundary - layer t r i p s , an experimental program i s under - way at Mach numbers of 3 and 4 on a f l a t - p l a t e model.

Discussion I n figure 4, photomicrographs of typical boundary - layer t r i p s are shown. An indication of the size of the t r i p s can be obtained by noting t h e 1/4-inch length shown. The airflow i s from l e f t t o right and t h e leading edge of the f l a t p l a t e can be seen i n each of the photographs.

The l e f t and center photographs show l/k-inch-wide bands of carborundum

grains - the l e f t one beginning at t h e leading edge and t h e center one

A narrow band of p a r t i c l e s beginning 1/4 inch behind t h e leading edge.

beginning S/k inch behind the leading edge i s sham i n the right photograph.

The drag r e s u l t s obtained with these three t r i p s a t a Mach number of 3 are presented i n figure 5. The drag coefficient of the flat p l a t e i s plotted against t h e mean height of the carborundum grains i n t h e t r i p . The drag coefficient w a s determined by a boundary - layer momentum survey 8- inches from the leading edge. The roughness height at which t h e t r i p s at t h e 1/4-inch position equal a roughness Reynolds number of 600 i s indicated. It i s apparent t h a t t h e t r i p s at 1/4 inch do not a f f e c t t r a n s i t i o n when they are smaller than t h i s height and do cause After t h e t r i p i s high t r a n s i t i o n t o move forward when they are larger.

enough t o cause transition, there i s a range of s i z e s where the drag i s constant; thus, there i s probably no additional drag due t o t h e t r i p s themselves over this range. A t larger sizes, the t r i p does cause addi - The wide t i o n a l drag as shown by the increase at the l a r g e r heights.

t r i p at t h e leading edge causes an additional drag which i s larger than t h a t associated with moving t r a n s i t i o n t o the leading edge. This addi - t i o n a l drag i s most l i k e l y due t o t h e f a c t that the roughness required t o fix t r a n s i t i o n very near t h e leading edge actually protrudes through * the boundary layer.

Ia order t o determine whether t h e drag measured with the narrow band of distributed carborundum includes any t r i p drag, comparison w a s made with a t r i p consisting of a single row of evenly spaced spherical p a r t i c l e s . The spherical p a r t i c l e s (fig. 6) were placed at 1 . 9 inches from the leading edge, s l i g h t l y ahead of natural transition, t o allow the The photograph on the use of larger, more e a s i l y handled, p a r t i c l e s .

r i g h t i s of a narrow band of distributed carborundum at the same position.

i The single row of spherical particles i s close t o the i d e a l t r i p , since it contains the l e a s t number of p a r t i c l e s t o cause drag and t h e height of the p a r t i c l e s can be accurately controlled.

. .

**e -

(. * 4 - *a * Figure 7 shows the r e s u l t s of t h i s comparison at Mach 3. The drag The narrow coefficient i s once again plotted against roughness height.

band of distributed roughness i s represented by the t r i a n g l e s and the single row of glass beads by the circles. All the data presented are The curves show that f o r roughness Reynolds numbers greater than 600.

a narrow band of distributed p a r t i c l e s produces the same drag r e s u l t s as a single row of g l a s s beads. The roughness height caa be two t o three times t h a t required t o f i x t r a n s i t i o n before either the single row o r narrow band of distributed roughness starts t o contribute roughness These r e s u l t s show t h a t t h e use of a correctly sized narrow band drag.

of sparsely distributed roughness on wind - tunnel models w i l l fix t r a n s i - t i o n as desired and w i l l not cause an additional t r i p drag.

A n example of the application of this type of t r i p t o a wind - tunnel model (fig. 8) i s shown i n figure 9. The configuration (fig. 8) w a s a clipped - tip delta - wing airplane model which w a s tested at M = 1.61.

The model w a s t e s t e d with t r a n s i t i o n free and with three different s i z e s of carborundum grains i n narrow t r i p s , (See f i g . 9.) The turbulent curve represents the model drag coefficient calculated with complete turbulent The gradual increase flow and w i t h a wave drag coefficient of 0.0092.

i n drag w i t h Reynolds number f o r the free - transition case indicates a slow forward movement i n transition. A s t h e roughness s i z e i s increased, the drag i s equivalent t o the value f o r t r a n s i t i o n at the t r i p at lower u n i t Reynolds numbers, as expected. I n fact, the roughness Reynolds num - ber f o r t r a n s i t i o n at t h e t r i p i s approximately 600 i n each case. The data points show that through a limited Reynolds number range, t r a n s i t i o n i s fixed at the roughness without additional g r i t drag. A t the higher Reynolds numbers, where the larger g r i t heights tested are considerably greater than the height required t o f i x t r a n s i t i o n , the g r i t begins t o cause additional drag. This result clearly indicates t h a t if the g r i t i s sized t o fix t r a n s i t i o n at a low Reynolds number, there i s a limit t o t h e range of Reynolds number through which tests should be made with only one g r i t size.

CONCLUDING R E M A R K S The r e s u l t s of the tests on three - dimensional boundary - layer t r i p s show that a narrow band of correctly sized sparsely distributed carbo- rundum located behind t h e leading edge w i l l f i x t r a n s i t i o n without addi - t i o n a l t r i p drag. The use of such t r i p s , then, w i l l allow an extrago- l a t i o n of wind - tunnel drag t o full - scale values with confidence. However, i f a wide t r i p , a leading - edge location, o r too large a roughness height i s used, excess t r i p drag w i l l : result which is d i f f i c u l t t o estimate.

The use of such t r i p s , therefore, should be avoided.

With regard t o air injection i n t o the turbulent boundary layer, t h e i n i t i a l r e s u l t s on injection through rearward - facing inclined s l o t s indicate that further study i s warranted.

INJECTION CONFIGURATIONS REARWARD - FACING INCLINED FLUSH SLOT REARWARD - FACING INCLINED STEP SLOT Figure 1 EFFECTIVENESS OF AIR INJECTION M =3.0 .0020 r 0 .2 . 4 .6 .8 I .o 2 m

- -

CEO ma3 Figure 2 c USE OF INLET BLEED AIR FOR BOUNDARY - LAYER INJECTION 4 - ENGINE SCAT 16; Mz3.0; 65,000 FT 1.00

\

\

CD~INJECTION CD,NO INJECTION.^^ NACELLES TAILS ~ FUS.

L I .92 I I I I I I 1 2 24 36 48 60 72 m , LB/SEC

Figure 4 L-2104-4

CARBORUNDUM TRANSITION TRIPS M= 3.0 x,lN. w,lN.

0 0 0.25 0.25 0.25 a 0.25 0.05 -Oo4[ .003 - Rk 600

* 0 ° ' I

I I 1 1 1 J 0 .002 .004 .006 .008 .OlO .012 k,lN.

Figure 5 Figure 6 L - 2104- 6 COMPARISON BETWEEN SINGLE ROW AND DISTRIBUTED ROUGHNESS TRIPS M = 3.0 0 SINGLE ROW, SPHERES; x=1.91N.

A DISTRIBUTED; W = 0 . 1 IN.; ~=1.91N.

.oo A .oo CD .oo Rk= 600 I I I I I I I I 0 .004 .008 ,012 ,016 .020 .024 .028 ,032 k, IN.

MODEL CONFIGURATION

/7

I

I Figure 8 i EFFECT OF BOUNDARY - LAYER TRIPS ON CD+=O M = 1.61 k, in.

0 0.012 0 .008 0 .005 i i

’ 1 0 ) CONFIGURATION EFFECTS ON TIlE CHARACTERISTICS

.-_.

OF DELTA - WING SUEERsOmC TRANSPORTS Adrien E. Anderson SUMMARY

// 9Xa-

This paper discusses a nmiber of i t e m s which are characteristic of or influence t h e c h a r a c t e r i s t i c s of delta - wing supersonic transports.

The topics discussed include: a comparison between the s t a b i l i t y char -

a c t e r i s t i c s of the basic SCAT 1 - 7 configuration and configurations simi -

lar t o the SCAT 17 designs proposed by the two f e a s i b i l i t y study con - t r a c t o r s ; the lateral s t a b i l i t y problem which arises when a design i s compromised t o achieve higher efficiency and directional s t a b i l i t y ; pos - sible configuration and control u t i l i z a t i o n changes which can be made t o improve efficiency; t h e e f f e c t of a i r f o i l section and flap deflection on transonic efficiency; low - speed high - lift studies; and f i n a l l y , aile - The ron and spoiler c h a r a c t e r i s t i c s obtained at high Mach numbers, material presented i s based on a portion of the studies made at the A m e s Research Center i n the course of evaluating t h e S C A T 17 configuration.

4 JTV0/2

INTRODUCTION The primary emphasis i n the SCAT 17 configuration i s toward s i m - p l i c i t y i n design, and at the same time meeting the rigorous low - speed requirements and the high efficiency needed at supersonic speeds. The aft s t a b i l i z i n g t a i l w a s incorporated i n t h e design t o allow good longi - t u d i n a l s t a b i l i t y combined with adequate control a t l o w speeds and t o Drooping the s t a b i l i z e r t o large serve a dual function at higher speeds.

anhedral angles permits some control over the natural rearward aerodynamic - center s h i f t which occurs with increase i n Mach number and allows improve - *$ ment of the directional s t a b i l i t y which tends t o decrease with increase I n t h e f e a s i b i l i t y study the two contractors submitted i n Mach nuniber.

configurations which employ e i t h e r the canard o r the aft s t a b i l i z e r , but not both. Thus, it i s of i n t e r e s t t o study the s t a b i l i t y characteristics

of t h e a l t e r n a t e concepts and compare them with t h e basic S C A T 17

configuratgon.

The development of high efficiency i n t h e transonic and low - speed because at flight zones i s as important as high efficiency at M = 3 transonic speeds, where the drag of the a i r c r a f t i s high, high e f f i - ciency allows operation of the a i r c r a f t at lower drag levels, and l e s s engine t h r u s t i s required f o r acceleration. At landing, high efficiency permits lower landing speeds. The importance of high efficiency has led t o an extensive study of the effectiveness of leading - edge f l a p s as a means for improving the characteristics of the configuration.

The p o s s i b i l i t y of engine f a i l u r e at high speeds places large demands on the aileron system, since aileron effectiveness has decreased considerably w i t h Mach number. Thus, an investigation w a s made t o deter - mine t h e a b i l i t y of wing spoilers t o provide roll. control at supersonic speeds.

The material presented i n t h i s report i s from studies made i n the Unitary Plan wind tunnels and the 40- by 80 - foot wind tunnel a t the Ames Research Center.

aerodynamic center, percent E a.c.

b wing span b' l o c a l span C wing chord

-

C wing mean aerodynamic chord cog* airplane center of gravity C canard l i f t coefficient

CL

rolling-moment coefficient

dihedral parameter, - &L

aa

yawing - moment coefficient Cn acn

d i r e c t i o n a l s t a b i l i t y , -

Cns

as

drag elevon spoiler height above surface of wing incidence angle lift lift - to - drag ratio Mach number thickness of wing section wing thickness to chord ratio angle of attack angle of sideslip dihedral angle deflection angle of control surface Subscripts: a aileron f wing trailing - edge flap m8x maximum n wing leading - edge flap S stabilizer i T wing tip W wing DISCUSSION Early in the feasibility studies the two study contractors took exception to the folding stabilizer of the basic SCAT 17 on the basis that the jet blast - impinging on the stabilizer in the down position might create stability problems as well as thermal and dynamic structural problems. Subsequently, the stability problem appears to be an insig - nificant one on the basis of large - scale wind - tunnel studies, and the structural situation appears to be one of whether you desire to pay the weight penalty of the heavier structure required to withstand the blast.

The study contractors suggested alternate designs which avoided the concept of the folding aft stabilizer. Thus, the thought arises as to the relative static stability characteristics of the three versions. In the NASA wind - tunnel tests, various studies were made to determine the aerodynamic contribution of a particular component of the basic config - uration and in some cases substitute components. From these data it is possible to compare alternate configurations. This information is being .

presented to show the potentiality of various components and configura - tions and should not be construed as reflecting on the contractors' pro - posals, for no attempt was made to optimize the geometric relationship among the components in the alternate cases.

Figure 1 presents the variation of aerodynamic center a.c. in percent of the mean aerodynamic chord E with Mach number for the three configurations indicated in the silhouettes. In the case of the alter - nate configurations, the fuselage length and the location of the wing on the fuselage are identical to those of the basic configuration (center silhouette in fig. 1 ) . At a Mach nmiber of 0.9, as noted in the fig- ure, the aft stabilizer dihedral angle is changed from 0 ' to -75' (down) on the basic SCAT 17. This change reduces the variation of aerodynamic center with Mach number. Had the stabilizer not been deflected, the movement of aerodynamic center would have been rearward, and at M = 3 would have been located at the value indicated by the " dot. " The aerodynamic - center variation for the canard configuration without sta - bilizer is very similar to that of the basic configuration; however, the aerodynamic - center location is slightly forward. Removal of the canard produced a sizable rearward shift of the aerodynarcic center and a larger variation of aerodynamic center with Mach number, as indicated by the top curve.

The large rearward aerodynamic - center shift produced by removing the canard indicates that the aerodynamic - center travel in the canard , configuration might be reduced through the use of either a folding tip, a variable sweep, or a servo - controlled canard. The canard, if servo - controlled at subsonic speeds and adjustable at supersonic speeds, would allow a low - speed aerodynamic center which is almost equal to the cruise

-

aerodynamic center, namely, 34 percent c versus 32.5 percent E . An

abrupt change in aerodynamic center would occur at the Mach number where the canard is locked to the control system and some form of stability augmentation undoubtedly would be required. In the case of the i - e . * ,

* * 0 . . . .(I*

canard - off configuration, deflecting t h e stabilizer up would provide some aerodynamic - center control.

If the low - speed s t a t i c margin i s a r b i t r a r i l y set at 2 percent (which i s equivrtlent t o around 18 inches on a 425,000 - pound a i r c r a f t ) and assuming no center - of - gravity change with Mach nmber, the s t a t i c f o r t h e basic configuration, margin would be 9.0 percent at M = 3 9 percent on the canard version, and around 13 percent on the canard - aft configuration. It should be borne i n mind t h a t t h i s anaLysis assumes a r i g i d a i r c r a f t and that aeroelastic e f f e c t s may a l t e r these results.

The directional s t a b i l i t y c h a r a c t e r i s t i c s of the t h r e e configura - t i o n s at a = 4 ' i s indicated i n figure 2. The abrupt increase i n ? ! Cnp at M = 0.9 f o r the basic configuration r e s u l t s from deflecting the stabilizer. If the s t a b i l i z e r i s not deflected, the M = 3 sta - b i l i t y l e v e l i s lower as indicated by the " dot. " This value i s very close t o that f o r the canard configuration. The directional s t a b i l i t y of the canard configuration i s less than that f o r the canard - off config - uration because the canard produces a destabilizing e f f e c t by increasing the pressures on the up - wind side of the fuselage. These data are f o r zero canard deflection, but w i t h the model s e t at cruise a t t i t u d e (a = bo). Thus, a f u r t h e r decrease i n directional s t a b i l i t y w o u l d be expected when the canard i s deflected or t h e angle of attack is increased.

In t h e canard configuration presented, the v e r t i c a l t a i l is aft of the wing t r a i l i n g edge and hence i s more effective than if it were located d i r e c t l y over the wing.

The dihedral parameter i s presented as a function of Mach number i n Positive s t a b i l i t y i s indicated by a negative value of figure 3.

6 % B o The s t a b i l i t y l e v e l decreases rapidly i n t h e supersonic range and i s lowest for the basic configuration. A t M = 3 , the value of Cz indi -

P

cates a s m s l l amount of negative s t a b i l i t y and, when compared w i t h the " dot, " shows t h a t deflecting a surface could lead t o a serious problem.

I n t h e course of evaluating the basic SCAT 17 configuration, the use of deflected wing t i p s was investigated as a means of controlling the aerodynamic center. The strong negative dihedral e f f e c t of deflec - t i n g the wing t i p s i s indicated i n figure 4. The s o l i d curve represents the case when both the s t a b i l i z e r and the wing t i p s are undeflected.

Deflecting the stabilizer at M = 0.9 reduces the s t a b i l i t y somewhat; at but when the wing t i p s are deflected from 0 ' at M = 0.9 t o -60' occurs and a negative sta - M = 1.4, an appreciable reduction i n c z B b i l i t y condition results. (The wind - tunnel studies indicated that the deflecting of t h e wing t i p should be programed o v e r t h e Mach number range from 0.9 t o 1.4 t o avoid abrupt variations i n aerodynamic - center travel.) It would appear then that using wing - tip deflection t o control i aerodynamic - center travel would be detrimental in the case of all three configurations shown in figure 3, since each has such low values of at the higher Mach numbers.

czB

In figure 5, the directional stability which could be achieved at M = 3 by deflecting the wing tips is presented and is found to be about twice that for the configuration with the wing tip and stabilizer at 0 ' and only half again as large as that for the configuration wherein the stabilizer had been deflected to a dihedral angle of -75O. Although the deflected wing tips increased Cnpy their detrimental effect,on C2

P

warranted the consideration of other configurations. The aerodynamic contribution of the alternate tail arrangements shown in figure 6 was determined in the course of the investigation. A side view of the low stabilizer deflected - 7 5 ' , which is the reference configuration, is shown at the left. In the central figure is a high stabilizer, at zero dihedral, in combination with a small ventral fin which allows ground clearance. Note the relationship of the tail surfaces to the wing. The square end on the fuselage is the shape of the wind - tunnel model which was sting mounted. At the right is a high stabilizer deflected to a dihedral angle of 37.5' and a large ventral fin. The large ventral fin provided approximately one - half as much with only one - tenth the

cnP

penalty of the deflected wing tips at Positive deflection M = 3 .

czP

of the stabilizer to 3 7 . 5 ' provided approximately one - half as much Cn P as deflecting the wing tips, but was beneficial in its effect on Cz

P

and aerodynamic - center shift.

Figure 7 is included to indicate how deflecting the wing tips and

the stabilizer affects aerodynamic - center travel on the basic SCAT 1 7 .

The smaller the variation in aerodynamic - center travel, the smaller will be the trim drag. How important is aerodynamic - center The question is: travel in affecting the trim drag of the basic configuration? Figure 8 presents the wind - tunnel model values of (L/D)-, at a Reynolds number

of 5 X lo6, as a function of Mach number. The lower curves are f o r the

three configurations in figure 7 when they are trimmed by deflecting

the canard. The assumption is that there is a 2 - percent static margin at M = 0.7 and that the center of gravity does not change with increase in Mach number. The trim curve for the configuration with the surfaces undeflected is quite far below the untrimmed curve. On the

other hand, deflecting the stabilizer to a dihedral angle of - 7 5 ' raises

the curve appreciably, and deflecting the wing tips brought small gains in the low supersonic range, but a slight loss at M = 3 .

Now there are other techniques for improving trim, two of which were investigated in the studies. One, shown in figure 9, was to ineor - porate wing incidence. The configuration considered here is the one with the s t a b i l i z e r deflected at -75O, the t i p s at Oo, and the canard i s used f o r trim. It w i l l be seen t h a t an incidence of - 1 ' (negative being t r a i l i n g edge up) produces an appreciable increase i n the trim ( L/D 1-9 whereas +lo caused a small loss a t the higher Mach numbers.

Another possibility, shown i n figure 1 0 , i s t o develop m a x i m u t i l i z a t i o n of the control systems. The curve f o r trim with the canard alone is a repeat of the s o l i d curve i n figure 9. If t r i m is by elevon alone, t h e trim L/D drops greatly. However, when a combination of both the elevon and canard i s used, it i s possible t o achieve an addi - amounts t o t i o n a l gain i n (L/D)-. In t h i s case the elevon s e t t i n g roughly 2O throughout the Mach number range.

1 Figure 1 1 presents a summary of what may be accomplished when the two concepts j u s t described a r e applied i n trimming t h e a i r c r a f t . The top curve i s for the - untrimmed configuration, whereas t h e bottom curve i s f o r t h e configuration with the stabilizer deflected, but t h i s t i m e the wing i s at -lo incidence and both the canard ( C ) and elevon (E) are used f o r optimum t r i m . This trimmed curve i s much closer t o the untrim - med curve and higher than the curve f o r the configuration where the wing t i p s w e r e deflected (refer t o f i g . 8). Thus, it appears t h a t high may be achieved without resorting t o deflec - values of t r i m (L/D)max t i n g the wing t i p s .

I n t h e SCAT 17 studies, the t h a t w a s realized i n the (L/D)- transonic speed range w a s rather disappointing. It w a s believed t h a t the low L/D values might be due t o the shape of the forward portion of the a i r f o i l section of the wing. The data of figure 12 a r e presented t o f a c i l i t a t e an understanding of the approach used t o improve the L/D characteristics.

The basic a i r f o i l section of the wing w a s a 30 - 70

hexagon, 2 percent thick. The 30- and 70 - percent points are indicated by the l i n e s i n the plan view and the 30 - percent chord s t a t i o n by t h e arrow i n the section view. Several years ago Menees of the Ames Research Center reported (ref. -1) t h a t a d e l t a planform wing incorporating the outboard 5 percent of t h e semispan of a wing having conical camber over t h e outboard 20 percent of the semispan had essentially the same effi - ciency as a wing with 20 percent of t h e semispan cambered* I n S C A T 17,

t h i s concept was simplified - the outboard 5 percent of the l o c a l semi -

'# span w a s deflected down approximately 3 . 1 ° i n t h e streamwise plane.

The The chord of t h i s portion of the wing increases toward the t i p .

wind - tunnel studies indicated t h a t t h i s type of leading edge would cause no penalty a t M = 3 when l e f t deflected. The increase, i n over the plane wing for three transonic Mach numbers, indi - (L/D)= cated t o the r i g h t i n the figure, w a s less than achieved i n Menees' investigation.

,, A flap with a constant chord equal to 1 0 percent of E was deflected to achieve further curvature of the leading edge. Gains in of around one unit were obtained at deflection angles of 1 2 ' (L/D)ma to l 5 O at speeds below M = 1 , but an insignificant gain was made at M = 1.2. The semispan of the flap was shortened to the 76 - percent sta - tion to see whether the gain in ( L / D ) - would warrant the complexity of incorporating a flap at the wing tip.

The increments in ( L / D ) , , were around 2/3 those indicated in the figure.

The airfoil section used in the study made by Menees was an NACA 0003-63. Since his results indicated a gain in ( L / D ) - of 3, 2, and 1.3 units for M = 0.70, 0.95, and 1.20, respectively, it was felt that the thickness as well as the curvature of the leading edge was

' i

playing an important part in the results. Tests have been made recently with the leading edge modified as indicated in the bottom section of fig - ure 12. In this case the camber line is that used by Menees and the top half of the airfoil section is an NACA 0002, which reaches maximum thickness at the 30 - percent chord station. The lower surface has been Taired straight back from the leading - edge radius to the 30 - percent chord point. The gains in (L/D)= were not significant; in fact, at M = 3 there was a loss in ( L / D ) - . These results are consequently considered to indicate that further study is required in this area.

The large - scale model used in the low - speed high - lift studies is presented in figure 13. In this model the airfoil section is an NACA 0003 - 03, which provides for a wing with a sharp leading edge, but with smooth curvature from the leading edge to the 30 - percent chord sta - tion. Figure 1 4 gives an indication of the gains to be obtained by for the wing - body combi - deflecting the leading - edge flap at M = 0.2 nation. In the plot of L/D as a function of CL the undeflected flap condition is indicated by the solid curve, the 30° flap deflection by the dash - double dot curve, and the 45' flap deflection by the dash - dot curve. In this case, appreciable gains were made, and when the canard indicated in the figure was added to trim the vehicle, only small losses occurred. When only the portion of the flap out to the 76 - percent semi - span station was deflected, the peak of the curve fell about mid - L/D way between the 0 ' and 45' deflection curve. The optimum leading - edge flap angle will vary with the deflection of the trailing - edge flap. i This information was taken from an investigation recently made by David G. Koenig of the Ames Research Center. This investigation also included information regarding midchord flaps and small wing planform variations.

Other work completed at Ames used the large - scale model to evaluate flap designs, effectiveness of various canard planf~rms, and canard ' The best way to sum up this work and indicate flaps (refs. 2 and 3 ) .

.the present state of the art insofar as the achievements of reasonable landing approach speeds is to use figure 15, which is from an IAS paper

by J. L . Jones, et al. (ref. 4 ) . Inddc?t?d i $ figure is a compari -

son of the thrust - to - gross weight rat& and c>*pCjh&ing flight veloc - ities to be expected with various nose and trailing - edge flap combina - tions. At the right is a canard configuration with no wing flap deflections and the center of gravity at the 18 - percent chord station.

This vehicle landing at a wing loading of 40 pounds per square foot and

a static margin of 3 percent would have to approach at 212 knots or higher to avoid the range of speed instability which is the condition where an increase in thrust is required to maintain the flight path with

decreasing speed. Deflecting the leading - edge flap 4 5 ' and the trailing -

edge flap loo (same static margin) reduces the speed to around 1 7 0 knots.

p e addition of the aft tail permits the trailing - edge flap angle to be increased to 200 and provides a more rearward center - of - gravity position,

-

namely, 2 6 percent c. The speed is then reduced to 1 4 5 knots. The fourth curve in the figure is for a larger canard which free floats.

This allows a landing speed of 135 knots with the center of gravity at

-

30 percent c.

Aileron effectiveness was found to decrease considerably at super - sonic speeds and additional roll capability appeared necessary. Consid - eration was given to the use of wing spoilers and the results of the study are given in figure 16. On the right is a plan view of one wing panel.

The aileron, designated as ailelon A had a chord equal to 15 percent of the local chord, and extended to the 76 - percent semispan station. Three sizes of spoilers located along the 80 - percent - chord line and mounted perpendicular to the wing surface were tested. The spoiler designated as spoiler B extended to the 34 - percent semispan sta - tion, while that designated spoiler C extended to the 76 - percent station.

In the left portion of the figure is a plot of rolling moment as a func - tion of Mach number in the supersonic range.

Airplane roll is indicated for two differential aileron deflections, k5' and +loo. Aileron effectiveness at M = 3 is 0 . 0 0 0 1 4 and is ade - quate to hold wings level to 5 ' of sideslip for deflections of +5O. At lower Mach numbers, however, higher differential deflections would be required. Tests made to determine the effectiveness of various segments of the aileron indicated that the outboard three panels provide nearly "11 of the roll. Spoiler B was not too effective for a height - to - chord ratio of 0.03.

M = 3 by going to twice A sizable gain was made at the height. The latter effectiveness was almost doubled by going to the full - span spoiler. More research is considered necessary in the area of aileron - spoiler effectiveness, because spoilers of this size, that is, over 6 feet high, at the inboard end essentially double the drag of the aircraft. Perhaps a combination of spoiler and aileron is the answer.

f * u * Additional material including engine - out and dynamic stability studies of a configuration similar to SCAT 1 7 will be found in refer- ences 5 through 8 .

CONCLUDING REMARKS The results presented herein on SCAT 17 indicate that: .

1 . The aerodynamic - center shift and the directional stability of the three types of configurations may be improved by deflecting surfaces, but there is the danger of getting into a dihedral - effect problem on all SCAT 17 configurations if the wing tips are deflected.

2. The trim drag may be improved by the use of negative wing inci - dence and judicious use of the canard and elevons so as to eliminate the need for deflecting the wing tips.

3. In the transonic and low - speed flight zones, improvements in have been gained by the use of leading - edge flaps, and further L/D improvements appear possible through the modification of the airfoil section, although additional studies are required.

4 . An evaluation has been made of simple high - lift devices which, when combined with the aft stabilizer, permit landing speeds comparable to present - day jet - transport aircraft.

5. Aileron and spoiler studies indicate deficiencies which indicate a need for further study.

"* REFERENCES 1 . Menees, Gene P.: L i f t , Drag, and Pitching Moment of an Aspect - Ratio - 2 Triangular Wing With Leading - Edge Flaps Designed t o Simu - late Conical Camber. NASA MEMO 10-5-58A, 1958.

2. Brady, James A,, Page, V. Robert, and Koenig, David G. : Large - Scale Low - Speed Wind - Tunnel Tests of a Delta Winged Supersonic Transport Model With a Delta Canard Control Surface. NASA TM X-643, 1962.

3. Koenig, David G., Brady, James A., and Page, V. Robert: Large - Scale Wind - Tunnel Tests a t Low Speed of a Delta Winged Supersonic Trans - port Model i n the Presence of the Ground. NASA TM X - 644, 1962.

4. Jones, J. Lloyd, Jr., Eunton, Lynn W., Gregory, Thomas J., and Nebs, Walter P., Jr.: A C r i t i c a l Study of Delta - Wing Configurations f o r the Supersonic Transport Application. Paper No. 63-5, Inst. Aero - space Sci., Jan. 1953.

5. Fletcher, LeRoy S.: S t a t i c S t a b i l i t y Characteristics of a Delta Winged Configuration With a Canard Control and Nacelles at Mach Numbers From 0.25 t o 3.50. NASA T M x - 651, 1962.

6. Gnos, A. Vernon, and Kurkowski, Richard L.: Effect of Off - Design I n l e t Mass Flow upon S t a t i c S t a b i l i t y of a Triangular Wing Configu - ration With a Canard Control and Pylon Mounted Nacelles f o r Mach Numbers From 0.65 t o 3.50. NASA TM X - 658, 1962.

7. Fletcher, LeRoy S.: S t a t i c S t a b i l i t y m a r a c t e r i s t i c s of a Delta Winged Airplane Configuration With Nacelles, a Trapezoidal Canard, and a Drooped T a i l at Mach Numbers From, 0.70 t o 3.52. NASA X-780, 1963.

8. Fletcher, LeRoy S.: Dynamic Rotary S t a b i l i t y Derivatives of a Delta - Winged Configuration With a Canard Control and Nacelles at Mach Numbers From 0.25 t o 3.50. NASA TM x - 781, 1963.

CONFIGURATION EFFECT ON AERODYNAMIC CENTER CL'0O 6 O r /----

- - - 4 4

50 c /'

2 0 ~ 1 0 0 I 2 3 MACH NUMBER Figure 1 CONFIGURATION EFFECT ON DIRECTIONAL STABILITY a=40 rS .008 r 0°&-750

.006 1

0 I 2 3 MACH NUMBER Figure 2 CONFIGURATION EFFECT ON DIHEDRAL PARAMETER a = 4 O

I+ -

--OO3 c

-.oo I PER DEG

. 0 0 1 t l

I I I I I I 2 3 MACH NUMBER Figure 3 DIHEDRAL EFFECT, BASIC SCAT 1 7 STAB I L I Z ER WING TIP -.002 - per deg - 00 00 0 . \ -00--750 0 0 \ -

--------_____ '. --

I I ,00--750 00-600 I I

.oo Io

I ; 3

MACH NUMBER

Figure 4

i I . ' !

DIRECTIONAL STABILITY, BASIC SCAT 1 7 - .008 a = 4 0 rT - 75O 0 " d - 6 0 ' ,006 0 0 .

I

c"s 0 0 - per .004 deg - .002 DEFLECTION DEFLECTION I I I I I 0 I 2 3 MACH NUMBER Figure 5 ALTERNATE TAIL ARRANGEMENTS HIGH STABILIZER

r =oo HIGH STABILIZER r =37.50

L O W STABILIZER rs=-750 +SMALL SENTRAL

+ LARGE VENT~AL I Figure 6 -"*

f 0 . ;" -- ii a *-.a

AERODYNAMIC - CENTER CONTROL, BASIC SCAT 1 7 50 r DEFLECTION 0 I 2 3 MACH NUMBER Figure 7 EFFECT OF SURFACE DEFLECTION ON L/D CANARD FOR TRIM MACH NUMBER Figure 8 EFFECT OF WING INCIDENCE ON TRIM L / D +- I O

rs 0 ' : - 750

rT = 00

- .. CANARD FOR TRIM

a - TRIMMED - 6 -

t

I I I I I I 4 L I 2 3 0 MACH NUMBER Figure 9 EFFECT O F CONTROL UTILIZATION ON TRIM L I D

r,= 0°,-750

I O

r,= 00

i w = o o - TRIMMED - -

ELEVON-- -----

- - I I I I I I C 1 I 2 3 MACH NUMBER Figure 10 TR M L/D POSSIBLE WITHOUT DEFLECTING WING TIPS I O TRIMMED I I I I I I I 2 3 MACH NUMBER Figure 11 EFFECT OF AIRFOIL SECTION AND LEADING EDGE FLAP ON TRANSONIC L/D 0.70 0.95 1.26 0 0 0 0.18 0.35 0.18 P.C.C. + FLAP DEFL 1.10 1 . 0 0 0.31 1.20 1 . 0 0 0 . 4 0

I Mo

Figure 12 LARGE SCALE SCAT 1 7 MODEL USED IN HIGH LIFT STUDIES A - 2 8 x 8 Figure 13 EFFECT OF WING LEADING EDGE F L A P DEFLECTION L I D / - - 0 .I .2 .3 A .5 .6 CL Figure 14 LANDING APPROACH CHARACTERISTICS - c . * . a 0 LANDING WING LOADINGs40psf STATIC MARGIN= 3% TRIMMED FLIGHT A l p 6 8 cg,o/oF 45" 20" 26 .30 r / n - .24 E

\

3-

# : I-Cn CnO - 3 E .I8 g"3 0 " 0" 1 8 I- L I 1 1 1 1 0 100 1 4 0 180 220 260 VE LOCITY , k t s Figure 15 AILERON AND SPOILER ROLL EFFECTIVENESS .007 .006 .005 .004 =1 .003 .002

t

I I 1 0 I 2 3 M Figure 16 STABILITY AND HIGH - LIFT STUDIES APF'LICABIX TO VARIABLE - -SWEEP SUPERSONIC - TRANSPORT - COI!ITIGURATIONS By William J. Alford, Jr., Vernard E. Lockwood, Linwood W. McKinney and Richard K. Greifl NASAsAmes R e s & & Cente3 SUMMARY A consideration of the stability and high - lift characteristics applicable to variable - sweep supersonic transport configurations indi - cates that the more outboard pivot locations can be used to reduce the aerodynamic - center variation with wing sweep angle, and this reduction results in higher supersonic trimmed lift - drag ratios. Of the several configuration concepts considered, the smallest maximum lift - drag - ratio variation with stability level or center - of - gravity location was obtained with the auxiliary wing concept.

The outboard pivot wings with fixed forward area have pitch instability problems at high lifts but these prob - .lems can be minimized by use of suitable forewing configurations and wing devices. The large spans provided by variable sweep allows the efficient attainment of high lift coefficients that provide very good take - off and landing characteristics. Representative lateral - directional stability and dynamic stability characteristics of several variable - sweep concepts are included to indicate the type of research information currently available. A bibliography of variable - sweep studies is also included.

INTRODUCTION The use of variable sweep offers potential not available to fixed wings and also poses several unique design problems.

The purpose of this paper is to review several types of variable - sweep schemes and to discuss some advantages and disadvantages of two of these with reference to the longitudinal - stability - supersonic - perfomance trade - offs and the variation of longitudinal stability with lift, and to discuss the high lift characteristics attainable with the low sweep angles and large spans offered by variable sweep.

The lateral - directional stability and dynamic stability characteristics of several variable - sweep concepts are also presented to indicate the type of research information avail - able. An extensive bibliography of variable - sweep studies is included for reference.

C O E F F I C I E N T S AND SYMBOLS Unless noted otherwise, the coefficients are based on the geometric characteristics of the sweptback wings.

A aspect ratio bf span of flaps b , span of wing C wing chord

-

C mean aerodynamic chord lift coefficient CL CD drag coefficient drag coefficient at zero lift ‘D, o pitching - moment coefficient Cm tail pitching - moment - coefficient contribution, (am) tail

(Cm)tail on - ( ‘ m l t a i l off

- & m

slope of pitching - moment curve through zero lift coefficient &L

Cms + C % damping - in - pitch parameter

C n , - CnB damping - in - yaw parameter wing efficiency factor maximum lift - drag ratio Mach number relative velocity stalling speed chord - plane horizontal - tail location H1 H2 mi8-high horizontal - tail location high horizontal - tail location H3 U angle of attack sweep angle of forewing A m sweep angle of leading edge ALE i 6f flap deflection angle

s, horizontal - tail deflection

l3EXKJLTS AND DISCUSSION Variable - Sweep Pivot Schemes Several of the many possible variable - sweep wing - pivot schemes are shown in figure 1 .

Shown at the top left is the single inboard type which has good low - speed pitch characteristics in the low sweep position because of its straight leading edge and small fixed forewing area. It has, however, a large longitudinal stability variation with increasing sweep angle because of the large movable area and small forewing and, if the advantages of low sweep angles are to be maintained, would probably require wing translation to provide reasonable handling qualities and minimum trim drag.

(See refs. 1 to 6. ) The next scheme is the " virtual pivot, " where the pivot does not physically lie on the wing surface.

This arrangement should have excellent all - around aerouynamic characteristics but would also require wing translation with the associated mechanical

complexity. (See refs. 7 and 8 . )

Shown at the bottom left of the figure is the double inboard pivot which provides both good pitching - moment char - acteristics throughout the lift range and minimum stability variations ',with sweep angle (ref. 9 ) ; however, its use would present some additional complexity and require a wide fuselage to be fully effective.

At the top right of figure 1 is shown the single outboard pivot scheme as utilized on the SCAT 1 6 concept.

Previous investigations (refs. 1 and 2) have shown that this general arrangement can provide minimum longitudinal - stability change with wing sweep angle but that it has pitch problems at high lifts when the fixed forward area (shown as the hatched area) is required to have an extremely high sweep angle.

The auxiliary wing type as utilized by the SCAT 15 concept has a large longitudinal - stability change with wing sweep angle but maintains good supersonic performance levels because the trim drag associated with its wing - tip controls is low and relatively insensitive to stability level.

The final scheme shown is the "M" wing type (See refs. 10 and 11.)

with outboard pivot. Data recently obtained indicate a reasonably small longitudinal - stability variation with wing sweep with a potentiality for having reasonable pitching - moment variations with lift throughout the sweep range.

Longitudinal - Stability - Supersonic - Performance Trade - offs For the SCAT 15 and SCAT 16 concepts, the variation of trimmed

as a function of the longitudinal stability parameter - & m

(L/D)ma &L Inasmuch as a given for a Mach number of 2.96 is presented in figure 2 .

level of longitudinal stability is associated with a particular center - of - gravity location, this figure presents, in essence, the variation of trimmed ( L / D ) , , , with center - of - gravity location. A comparison of the absolute levels of the lift - drag ratio is not pertinent to this discus - sion since this comparison has been covered in paper no. 2 of this com - pilation by Donald D. Baals. Of interest is the fact that for SCAT 15, which was trimmed with wing - tip tails operating in an upwash field, the highest trimmed lift - drag ratios were obtained with a center - of - gravity location giving a rather large stability level whereas SCAT 16, which was trimmed with a conventional aft tail operating in the wing - body downwash field, attained its maximum L/D at a considerably lower static margin, and exhibited a rather rapid decrease in L/D as the stability level was increased.

Before discussing the stability levels or center - of - gravity loca - tions that are required for these configurations and the implications with regard to performance, it is well to consider briefly the effects of aeroelasticity since this property can appreciably affect the sta - bility level. The effect of aeroelasticity on longitudinal stability and trim requirements is illustrated in figure 3 for a highly swept con - figuration at a Mach number of 3.0. For a given center - of - gravity loca - , ,) tion the longitudinal stability of the elastic aircraft is of course less than that of the rigid aircraft both with tail off and with tail on because of the increasing washout of the elastic wing with increasing lift. If the rigid aircraft is assumed to have an optimum aerodynamic shape for cruise at the design lift coefficient, the flexible aircraft will be optimum only if it is constructed in such a manner that when it is loaded to the design cruise condition, it will assume the same shape as the optimum rigid aircraft. The tail - off center of pressure, as well as the pitching moment of the elastic aircraft, will then be identical to that of the rigid aircraft at the design condition regardless bf the ch ge in stability. Therefore, at the design CL the tail load and incidence required to trim, and the resulting trimmed L/D will be unaffected by aeroelasticity.

The results of figure 2 are therefore applicable to

either flexible or rigid aircraft if the value used for - & m is that

&L - of the equivalent rigid configuration. The aircraft stability is, how - ever, affected by elasticity not only through the change in static mar - gin but also through changes in other stability derivatives and must be considered in handling - qualities evaluations.

In order to determine the stability levels that are required for I the SCAT 15 and SCAT 1 6 configurations, it is necessary to consider the variation of aerodynamic center with wing sweep angle and Mach number as phsented in figure 4. For reference, center - of - gravity locations are assumed that vary with sweep angle because of the weight of the wing and its internal components. These center - of - gravity variations are for illustrative purposes only and do not necessarily correspond to those used by the contractors in their feasibility studies.

For SCAT 15, a center - of - gravity position A can be selected which provides a small amount of stability for low speed while producing nearly maximum available trimmed ( L / D ) , , , in supersonic cruise (see fig. 2) even though a large static margin exists at Mach 3 . 0 . For SCAT 16, on the other hand, selection of a center - of - gravity position B which gives only a small trim - drag penalty in cruise requires a minimum

sweep angle of 3 5 ' to kOo to achieve low - speed stability. (See fig. 4.)

A lower minimum sweep angle would be desirable to improve the subsonic lift - drag ratio, pitch - up characteristics, performance of high - lift devices, and lateral - directional handling qualities. At a center - of - gravity position C which gives a reasonable level of static stability for low speeds at 16O of sweep, a noticeable penalty in trim drag must be paid in supersonic performance because of the large static margin.

(See fig. 2. ) This trim - drag penalty can be reduced somewhat (theoreti - cally at least) by a more optimum design of wing warp.

Since the highest cruise L/D was obtained with center - of - gravity 'llocation B, it is of interest to find a means of obtaining low - speed stability at lower sweep angles for this more rearward center - of - gravity location. An illustration of what can be done by adjustment of the wing pivot location, the importance of which was pointed out in reference 1, is indicated in figure 5 where the variation of the longitudinal stabil -

ity parameter - acm with wing sweep angle is presented for SCAT 1 6 at low

&L speed. The experimental points are from wind-tunnel tests made wi%h the horizontal tail off.

The calculated curves were obtained from a modified Because the calculation did lifting - line procedure for the wing alone.

not account for the fuselage contribution, the calculated curves have Am = 7 6 O been arbitrarily shifted vertically to bring the results for The calculation for pivot into agreement with the experimental results.

which corresponds to the experimental case, shows that the location 2, effect of sweep angle on the static margin can be predicted with fair Calculations were also made for pivot locations 1 and 3 , accuracy.

which lie on the locus of pivots and which provide identical planforms The open - wing posi - at maximum sweep and equal spans at minimum sweep.

tion shown corresponds to pivot location 2 . The results illustrate the large increase in stab5lity at'low sweep angles that can be gained by moving the pivot point outward. It is apparent, therefore, that with a more outboard pivot location a considerably lower minimum sweep angle can be utilized while retaining low - speed stability with the more rear - ward center - of - gravity location indicated by center - of - gravity location B, compatible with the desired supersonic performance. It should be pointed out that the original SCAT 1 6 design is characterized by an extremely high aspect ratio wing which makes it difficult to provide a pivot location and rotating wing - panel - geometry relationship compatible with the outboard pivot concept. It therefore appears desirable to reduce the aspect ratio and sacrifice some of the extremely high subsonic performance in order to provide the desired geometry.

Longitudinal Stability Characteristics The variation of pitching - moment coefficient with lift coefficient for the SCAT 15 and SCAT 16 configurations are presented for low speeds in figures 6 and 7 s t n d for transonic and supersonic speeds in figure 8 .

For low speed and low sweeps (fig. 6), both configurations exhibit varying degrees of instability at high lift; thus, each must be fitted with special devices to eliminate or minimize these undesirable trends.

For the high - sweep case at low speeds (fig. 7 ) , which is important for

an emergency landing associated with stuck - back wings, the SCAT 1 6 type appears reasonably satisfactory whereas the auxiliary - wing type of SCAT 13 has undesirable characteristics that would also require addi - tional devices. The effects of horizontal - tail deflection on the longi - tudinal stability characteristics are seen to be minor for both the low -

sweep case (fig. 6) and the high - sweep case (fig. 7) and imply that

major conclusions concerning the high - lift instabilities will not be appreciably affected by the horizontal - tail deflections required for trim.

The subsequent figures will therefore present data for only one tail set - ting. At transonic speeds (fig. 8 ) , the reverse is apparent in that the SCAT 1 6 type indicates an instability whereas the SCAT 15 type does not, at least for the lift coefficient range investigated. At M = 3.0 both types appear to be fairly satisfactory for the lift - coefficient range shown, although there is some indication that SCAT 15 is approaching a statically unstable condition at the higher lift coefficients.

The cause of the high - lift instabilities which characterize these configurations at low sweep angles can generally be traced to the fore - wing where lift effectiveness increases with angle of attack and pro - duces a strong vortex flow that induces high angularities on the out - board pwels and causes them to stall or separate. The effects of forewing geometry on the - low - speed pitch characteristics of an outboard - pivot configuration are presented in figure 9. As would be expected, - the degree of high - lift instability, or pitch - up, is reduced with reduc tions in forewing sweep angle and area for the tail - off configuration.

The bottom graphs of figure 9 present the same forewing curtfigurations For the 7 1 5 ' forewing case, although ' with a low horizontal tail added.

an improvement is noted in longitudinal stability at high lift, the curve is still unacceptable. For the 7 0 ' forewing, howevqr, a stable, although nonlinear, variation is obtained. I f the forewing is deflected down about the hinge line indicated on the sketch, a very desirable variation is obtained. It should be noted that this configuration utilized engines mounted within the fuselage so that it was possible to employ a horizontal tail in a low location and thereby provide desirable pitch characteristics. For the SCAT 1 6 type configuration, however, there are conflicting requirements between the podded - engine location and horizontal - tail location which result in the tail being forced above the wing - chord plane in order to avoid hot gas efflux effects on the tail structure. ' & e effects of horizontal-tail location on the low - speed tail pitch contributions of the SCAT 1 6 type configuration are For the.mid-high tail location a negligible tail presented in figure 1 0 .

contribution results from placing the tail in the extreme downwash flow field caused by the strong vortex flow emanating from the body nose and fortified by that shed from the forewing. The data further show the well - known desirability of positioning the tail as l o w as possible.

For she high location, the tail contribution is initially large but decreases rapidly with increasing angle of attack and then reverses as the tail moves into the aforementioned downwash field.

I f it is assumed that a moderately high tail must be used for practical reasons, it is apparent that the achievement of a pitch - up - % free configuration will require very careful tailoring of the detailed arrangement of the horizontal tail in combination with wing flow - control devices. Some selected results from such a detailed study at low speeds on a SCAT 1 6 type of configuration utilizing two nacelles side by side under each wing are presented in figure 1 1 . The best horizontal - tail arrangement found which would remain clear of the jet blast was at a mid - high location with 3 0 ' anhedral. With a forewing of 73.5O sweep and an outer wing panel at 25O sweep, a definite instability at high i l i f t i s evident. Addition of a leading - edge slat on the outer panel eliminated t h i s i n s t a b i l i t y with the exception of a minor break near maximum l i f t . The angle of a t t a c k a t t h i s point i s above 180 and it i s believed t h a t even t h i s minor i n s t a b i l i t y could be eliminated with more careful tuning of the slat. An a l t e r n a t e arrangement of wing flow - control devices consisting of a notch i n the outer panel leading edge combined with a 290 downward deflection of the forewing nose f l a p also produced f a i r l y s a t i s f a c t o r y pitch characteristics through the l i f t range.

The contribution of the notch alone w a s about one - half the difference shown i n figure 1 1 between the clean wing and the notch plus forewing deflection except f o r a s l i g h t pitch - up near maximum l i f t coefficient.

Recent M - Wing Variable - Sweep Investigation Another type of variable - sweep pivot scheme t h a t has been the sub - j e c t of recent preliminary studies i s the M - wing type with outboard pivot. The s t r u c t u r a l aspects of t h i s type of configuration were not included i n the f e a s i b i l i t y studies and therefore the p r a c t i c a l aspects have not as yet been evaluated. The low - speed longitudinal s t a b i l i t y characteristics of such an M - wing configuration are presented i n f i g - ure 12. This configuration employed a high horizontal - tail location.

The static - margin variation with wing sweep angle i s presented f o r both t a i l - o f f and tail - on conditions. For the tail - on case the s t a t i c m a r -

-

gin s h i f t s about 17 - percent c as the wing i s swept from 3 4 O t o 74'.

For the t a i l - o f f condition, at 34' sweep there are large variations i n the pitching - moment coefficient with pitch - up occurring a t CL = 0.5 followed by a very stable break a t CL = 1.0.

The i n i t i a l pitch - up i s presumed t o be due t o the s t a l l i n g of the outer wing panel. Fixes on the outer wing panel, as shown e a r l i e r , should a l l e v i a t e these undesir - able trends. Since the t a i l - o f f curve breaks stable at the high lifts, a s i t u a t i o n occurs t h a t should be compatible with the use of a high t a i l which breaks unstable when it enters t h e wing downwash f i e l d . The addi - t i o n of such a t a i l indicates t h a t the desired pitch trends are mostly achieved and the use of outer wing panel f i x e s should provide even more l i n e a r curves.

High - Li f t Characteristics The configuration that has been used t o obtain extensive high-lf data applicable t o variable - sweep transports both at the Ames and the Langley Research Centers i s shown i n figure 13.

The principal d i f f e r - ence between t h i s wing and conventional low - sweep wings i s the addition of the forewing. High - lift devices investigated included a leading - .edge slat and trailing - edge f l a p s of e i t h e r the single s l o t t e d o r double s l o t t e d types.

It should be noted t h a t t h i s configuration has the horizontal t a i l below the chord l i n e of the wing. Some selected r e s u l t s f o r a single s l o t t e d f l a p deflected 400 a r e presented i n f i g - ure 14.

The trimmed l i f t - c o e f f i c i e n t characteristics shown i n the upper left graph are for a flap semispan of 98 p & c & o f , t h e wing semispan.

Maximum trim lift coefficient values of 2 . 4 or higher based on the area

of the 1 3 . 5 ' swept wing are seen attainable with this type of configu - ration.

Increasing the wing sweep from 13.50 to 250 causes a loss in CL of approximately 0.25 or about 12 percent at an angle of attack The question of the lift that is available for landing in the of 12O.

case where the wings are stuck back'is illustrated in the lower part of The comparison shows that no increase in lift at landing the graph.

a, 2 loo, was obtained from deflection of these flaps at a

attitudes, wing sweep of 7 5 O and indicated that careful flap design will be required to obtain a lift increment for the higher sweep angles.

The upper right part of figure 14 shows the effect of flap span on ' the high - lift characteristics. A reduction of flap span from 0.98h/2 to 0.67bJ2 with the wings at 250 sweep reduced the lift coefficient by a value of 0.15 and resulted in a lift coefficient of 1.8 at an angle of attack of 1 2 ' . Consideration of ground effect should result in landing lift coefficients in excess of 2.0 even with the partial - span flap.

With regard to the pitch characteristics at high lift, the data

for 2 5 ' of wing sweep with partial - span flaps are presented in the lower

portion of figure 1 4 . The data show that the pitching - moment - coefficient

variation with lift coefficient is stable over most of the lift range and that the tail is capable of trimming the model over most of thelift range. It should be remembered, however, that this configuration has a relatively low horizontal tail which provides better variations of pitching moment with lift than other horizontal - tail locations. I f it is found that the use of a higher tail location leads to insoluble pitch - up problems, some reduction of forewing area by use of a more inboard pivot location may be required although, as shown earlier, this would probably require the use of higher minimum sweep angles to restrict the aerodynamic - center variation with wiqg sweep angle.

The effect of high - lift devices on the drag characteristics of the configuration shown in figure 13 is presented in figure 15. The pro - gressive improvement provided by more sophisticated devices is evident.

It is also interesting to note that an envelope drag polar constructed , by using the zero - lift drag coefficient and drag - due - to - lift factor of the basic wing is nearly tangent to each of the high - lift - device curves.

This condition implies that high - lift devices can be designed which will retain the span efficiency of the basic wing. The resulting high lift - drag ratios will lead to l o w values of thrust required in the landing approach and will thereby reduce the problems of community noise.

For reference, the lift coefficients corresponding to 1 . 3Vs and ( L / D ) , , for the double slotted flap configuration are shown. Since (L/D)- occurs at a higher coefficient and therefore at a lower speed than the assumed approach ?pee& bf 2. 3Vs, speed stability during the landing approach will be obtained with the efficient high - lift system shown.

Some results recently obtained at the Langley Research Center during a study to determine the effect of the addition of a forewing on the lift characteristics of an aspect - ratio - 10 wing having a very effi - cient high - lift system are presented in figures 1 6 and 17. This wing - body arrangement utilized an unswept wing having leading - edge slats in combination with a double slotted flap as shown in figure 1 6 . All coef - ficients are based on the area of the wing without the forewing. At

zero angle of attack a bo deflection of the flaps -provided an increment

in CL of 2.0. At an angle of attack of 1 0 ' a maximum lift coefficient of approximately 3 . 1 was obtained. The addition of the forewing caused an increase in the slope of the lift curve and increased the maximum lift coefficient by about 0.1. The effect of the forewing on the drag characteristics was to cause an increase at low lifts and a small reduc - tion at high lifts. (See fig. 17.) A comparison of the experimental data with the envelope drag polar constructed by using the basic - wing zero - lift drag and drag - due - to - lift factors again indicates that this high - lift system is very efficient.

Lateral - Directional Stability and the The variation of the directional - stability parameter CnP with angle of attack for low speed effective dihedral parameter C 2P and supersonic speeds are presented in figures 1 8 and 1 9 , respectively, for both the SCAT 15 and 16/ configurations. For the SCAT 16 configura - tion whose original body nose cross section was near circular but some - was invariant with a up to about loo, what flat on the bottom, CnP increased appreciably between loo and 20°, and decreased sharply above 200, for the A m = 1 6 0 condition. Additional tests made with a fuse - lage nose section with a circular cross section caused to decrease

CnP

rapidly above u = loo. It is, therefore, apparent that fuselage cross section can have a large effect on the directional - stability variation with a especially when long fuselage forebodies are used. For the

'

becomes exceedingly stable at the highest angles high - sweep case,

c " P

of attack. For SCAT 15, the low - sweep condition becomes unstable at an angle of attack of about 1 9 ' whereas the high - sweep case, although very nonlinear, generally becomes increasingly stable at the higher angles of attack. With regard to the effective dihedral parameter -CzP, the values for the SCAT 1 6 low - sweep wing are seen to reduce at the higher angles of The attack whereas for the high - sweep case an increase is observed.

reduction at high angles of attack for the low - sweep wing are associated with the st@ling of the outer panel men.t"ion"ed previously in connection with pitch-ap problems and, when wing fixes are employed, the reduction would be expected to be considerably less.

For SCAT 15 both in sweeps have large levels at the higher angles of attack.

For supersonic speed (fig. 19) both configurations have similar characteristics, cnP decreasing as M increases because of the reduction in vertical - tail side - force effectiveness, and become unstable at relatively low angles of attack. The level of effective dihedral also decreases.with Mach number. It is beyond the scope of this paper to assess the effect of the rather high values of effective dihedral on handling qualities since this assessment would require extensive simulator studies.

I Dynamic-Stability Derivatives Knowledge of the dynamic stability derivatives is required in order to predict handling - qualities and stability augmentation require - ments. Experimental measurement of these derivatives requires careful application of specialized techniques. Two different wind - tunnel tech - niques employing forced oscillation of the model have been employed at the NASA Ames and Langley Research mnters for measurement of the damping derivatives in pitch and yaw of models of a preliminary variable - sweep transport concept. Representative results at zero angle of attack are presented in figure 20 to illustrate the type of sweep and Mach number effects that might be expected for a variable - sweep trans - port. The data, regardless of wing - sweep position, are referred to the The geometric characteristics of the wing in the 7 5 ' sweep position.

was generally independent of wing damping - in - pitch parameter C

ms + c %

The data measured by the sweep for sweep angles of 25O, 30°, and 7 5 O .

two techniques show reasonably good agreement throughout the Mach number

range. The damping - in - yaw parameter C n , - Cni was also relatively

independent of wing sweep 'but, in this case, the agreement between the two techniques was good only at supersonic speeds. The results indicate - only minor effects of sweep, and Mach number effects reasonably compati ble with the well - known variations in the lift effectiveness of the tail ' surfaces with Mach number. The complete results for the supersonic part of the Langley investigation are reported in reference 12.

CONCLUDING RENARKS A consideration of the stability and high - lift characteristics applicable to variable - sweep supersonic transport configurations indicates that the more outboard pivot locations can be used to reduce the aerodynamic - center variation with wing sweep angle and resul.b&n higher supersmic trimmed lift - drag ratios. Of the several configura - tion concepts considered, the smallest m a x i m lift - drag-ratio variation with stability level or center - of - gravity location was obtained with the auxiliary - wing concept. The outboard pivot wings with fixed forward area have pitch instability problems at high lifts but these problems can be minimized by use of suitable forewing configurations and wing devices. The large spans provided by variable sweep allow the effi - cient attainment of high lift coefficients that provide very: good take - off and landing characteristics. Representative lateral - directional stability and dynamic stability characteristics of several variable- sweep concepts are included to indicate the type of research information currently available.

A bibliography of variable - sweep studies is also included.

J FU3FERENCES 1 . Alford, William J., Jr., and Henderson, William P . : An Fixploratory Investigation of the Low - Speed Aerodynamic Characteristics of Variable - Wing - Sweep Airplane Configurations. NASA TM X - 142, 1959.

2. Alford, William J., Jr., Luoma, Arvo A., and Henderson, William P . : Wind - Tunnel Studies at Subsonic and Transonic Speeds of a Multiple-Mission Variable - Wing - Sweep Airplane Configuration.

NASA TM X-206, 1959.

3. Spencer, Bernard, Jr.: Stability and Control Characteristics at *i Low Subsonic Speeds of an Airplane Configuration Having Two Types of Variable - Sweep Wings. NASA T M X-303, 1960.

4. Bielat, Ralph P., Robins, A . Warner, and Alford, William J., Jr.: The Transonic Aerodynamic Characteristics of Two Variable - Sweep Airplane Configurations Capable of Low - Level Supersonic Attack.

NASA TM X-304, 1960.

5. Bielat, Ralph P., and Pierpont, P . Kenneth: Transonic Aerodynamic Characteristics of a Variable - Sweep Airplane Configuration Having a 12 - Percent - Thick Wing and an Inboard Pivot Location.

NASA TM X-429, 1960.

6. Hammond, Alexander D., and Polhamus, Edward C. : The Effects of Korizontal-Tail Position on the Longitudinal Aerodynamic Charac - teristics at Transonic Speeds of a Variable - Sweep Aircraft Having an Inboard Pivot. NASA TM x - 612, 1961.

7 . Kemp, William B., Jr., Becht, Robert E., and Few, Albert G., Jr.: Stability and Control Characteristics at Low Speed of a - - Scale Bell X-5 Airplane Model. Longitudinal Stability and Control.

NACA RM ~ 9 ~ 0 8 , 1950.

8 . Kemp, William B . , Jr., and Becht, Robert E . : Stability and Control Characteristics at Low Speed of a - - Scale Bell X-5 Airplane Model. Lateral and Directional Stability and Control, NACA RM L5OC17a, 1950.

9 . Polhamus, Edward C . , Alford, William J., Jr., and Foster, Gerald V . : Subsonic and Supersonic Aerodynamic Characteristics of an Airplane Configuration Utilizing Double - Pivot Var&able-Sweep Wings.

NASA TM x-743, 1 9 6 2 .

10. Driver, Cornelius, Spearman, M. Leroy, and Corlett, William A.: Aerodynamic Characteristics at Mach Numbers from 1.61 to 2.86 of a Supersonic Transport Model With a Blended Wing - Body, Variable - Sweep Auxiliary Wing Panels, Outboard Tail Surfaces, and a Design Mach Number of 2.2. NASA TM X-817, 1963.

1 1 . Harris, Roy V., Jr., and Corlett, William A . : Longitudinal Aero - dynamic Characteristics at Mach Numbers From 0.50 to 1.20 of a Supersonic Transport Model With a Blended Wing - Body, Variable - Sweep Awriliary Wing Panels, and Outboard Tail Surfaces. NASA TM x - 837, 1 9 6 3 .

12. Delaney, Bobby R., and Thompson, Wilson E. : Dynamic Stability Characteristics in Pitch and in Yaw for a Model of a Variable - Sweep Supersonic Transport Configuration at Mach Numbers of 2.40, 2 . 9 8 , and 3.60.

NASA TM x - 761, 1963.

i ?

BIBLIOGRAPHY Campbell, John P., and Drake, Hubert M. : Investigation of S t a b i l i t y and Control Characteristics of an Airplane Model With Skwed Wing i n the Langley Free - Flight Tunnel. NACA TN 1208, 1947.

Donlan, Charles J., and Sleeman, W i l l i a m C., Jr.: Low - Speed Wind - Tunnel Investigation of the Longitudinal S t a b i l i t y Characteristics of a Model Equipped With a Variable-Sw'eep Wing. NACA RM ~ 9 ~ 1 8 , 1949.

Kemp, W i l l i a m B., J r . , Becht, Robert E., and Few, Albert G., J r . : .

S t a b i l i t y and Control Characteristics a t L m S p e d of a - Scale B e l l 1 r; X - 5 Airplane Model. Longitudinal S t a b i l i t y and Control. NACA RM ~ 9 ~ 0 8 , 1950.

Keq, W i l l i a m B., Jr., and Becht, Robert E. : S t a b i l i t y and Control Characteristics a t Low Speed of a L -Scale B e l l X-5 Airplane Model.

Lateral and Directional S t a b i l i t y and Control. NACA RM LTOC17a, 1950.

Silsby, N0rma.n S., Morris, Garland J., and Kennedy, Robert M.: Longi -

t u d i n a l Characteristics at Mach Number of 1.24 of - - Scale Semispan

Model of Bell X - 5 Variable - Sweep Airplane With Wing Swept Back 60° From Tests by NACA Wing - Flow Method. NACA RM L5OEO2a, 1950.

Morris, Garland J., Kennedy, Robert M., and Silsby, Norman S.: The Effect of seepback on the Longitudinal Characteristics a t a Mach

Number of 1.24 of a - - Scale Semispan Model of the B e l l X-5 Airplane

From Tests by the SACA Wing - Flow Method. NACA R M ~50128, 1950.

Sawyer, Richard H., Kennedy, Robert M., and Morris, Garland J.: Longitudinal - Control Effectiveness and Downwash Characteristics at a

Mach Nunber of 1.24 of a - - Scale Semispan Model of the B e l l X - 5

\ 30 Airplane as Determined by the NACA Wing - Flow. Method. NACA RM L50Kl5, 1951.

Kemp, W i l l i a m B., J r . , Becht, Robert E., and Few, Albert G., Jr.: Investigation of t h e Low - Speed Aerodynamic Characteristics of a Variable - Sweep Airplane Model With a Twisted and Cambered Wing.

N A C A RM ~ 5 ~ ~ 2 2 , 1952.

Kemp, W i l l i a m B., Jr. : A n Investigation of the Lw-Speed Longitudinal S t a b i l i t y Characteristics of a Wept - Wing Airplane Model With Two Modifications t o the Wing - Root Plan Form.

NACA RM L52EO7, 1952.

Becht, Robert E., and Byrnes, Andrew L., Jr.: Ilnrestigation of the Low - Speed Aerodynamic Characteristics of a Variable - Sweep Airplane Model With a Wing Having Partial - Span Canibered-Leading-Edge Modifications.

NACA RM L52GO8a, 1952.

Spreemann, Kenneth P., and Alford, W i l l i a m J., Jr.: Small - scale Transonic Investigation of the Effects of 'Pwist and Camber on the Aerodynaslic Characteristics of a 60°42' Weptback Wing of Aspect Ratio 1.94.

NACA RM ~51121, 1952.

>

Spreemann, Kenneth P., and Alford, W i l l i a m J., Jr.: Investigation of the Effects of Twist and Camber on the Aerodynamic Characteristics of a 5 0 ~ 3 8 ~ Weptback Wing of Aspect Ratio 2.98.

TransonickBump Method.

NACA RM ~ 5 1 ~ 1 6 , 1951.

Alford, W i l l i a m J., Jr., and Bymes, Andrew L., Jr.: Small - scale Tran - sonic Investigation of the Effects of Partial - Span Leading - Edge Camber on the Aerodynamic Characteristics of a 50038' Sweptback Wing of Aspect Ratio 2.98.

NACA RM ~ 5 2 ~ 0 8 a , 1952.

Kolnick, Joseph J., and Kennedy, Robert M.: The Effects of Sweepback on Longitudinal Characteristics of a - - Scale 1 Semispan Model of the B e l l X - 5 Airplane as Determined From NACA Wing - Flow Tests at Transonic Speeds. NACA RM L52I23, 1952.

Silsby, Norman S o , and Morris, Garland J.: Longitudinal - Control Effec - tiveness and Downwash Characteristics at Transonic Speeds of a -..) Scale Semispan Model of the B e l l X-5 Airplane as Determined by the NACA Wing - Flow Method.

NACA RM L52Kl.2, 1953.

Bielat, Ralph P., and Campbell, George S o : A Transonic Wind - Tunnel Investigation of the Longitudinal S t a b i l i t y and Control Character - i s t i c s of a 0.09 - Scale Model o f the B e l l X-5 Research Airplane and i Comparison With Flight.

NACA RM ~ 5 3 ~ ~ 8 , 1953.

Donely, Philip, and G i l l i s , Clarence L.: Some Design Considerations Pertinent t o the Rough - Air Behavior of Airplanes at Low' Altitude.

NACA RM L53JOlb, 1953.

English, Roland D.: Some Effects of Aeroelasticity and Sweepback on t h e Rolling Effectiveness and Drag of a l/ll-Scale Modekof t h e Bell NACA RM L53118b, X-5 Airplane Wing at Mach N W e r s From 0.6 t o 1 . 5 .

1953.

Becht, Robert E. : S t a b i l i t y Characteristics at L w Speed of a Variable - Sweep Airplane Model Having a P a r t i d l y Cambered Wing With Several NACA RM L53L14, 1954.

Chord - Extension Configurations .

Morris, Garland J., and Silsby, Norman S.: Distribution of L i f t and Pitching Moment Between Wing and Fuselage and Effects of Wing Flexi -

b i l i t y and Dive Brake on a 1/3O-Scale Semispan Model of the B e l l X-2

Airplane at Transonic Speeds a s Determined by the NACA Wing - Flow \+ / Method. NACA RM L54i~t.1, 1934.

S t a b i l i t y and Control Char - Spearman, M. Leroy, and Fo-mer, Gerald V.: a c t e r i s t i c s at a Mach Number of 2.01 of a Variable - Wing - Sweep Config - NASA TM X-32, 1959.

uration With Outboard Wing Panels Swept Back 7 5 O .

An Exploratory Alford, W i l l i a m J., Jr., and Henderson, W i l l i a m P.: Investigation of the Low - Speed Aerodynamic Characteristics of Variable - Wing-Sw'eep Airplane Configurations. NASA T M X-142, 1959.

Alford, W i l l i a m J., Jr., Luoma, Arvo A . , and Henderson, W i l l i a m P.: Wind - !Tunnel Studies at Subsonic and Transonic Speeds of a Multiple - Mission Variable-Wing-Sw'eep Airplane Configuration. NASA '151 x - 206, Effects of Various Modifi - Spearman, M. Leroy, and Foster, Gerald V.: cations on the Supersonic S t a b i l i t y Characteristics of a Variable - NASA TM x - 260, Wing - Sweep Configuration at a Mach Number of 2.01.

1960 .

Fogter, Gerald V.: S t a b i l i t y and Control Characteristics at Mach Numbers of 2.50, 5-00, and 3.77- of a Variable - Wing - Sweep Configuration With NASA TM X - 267, 1960.

Outboard Wing Panels Swept Back 7 5 O .

foster, Gerald V.: Effects of Spoiler - Slot - Deflector Control on the Aerodynamic Characteristics at a Mach Number of 2.01 of a Variable - , Wing - Sweep Configuration With t h e Outer Wing Panels Wept Back 7 5 O .

N A S A T M X-273, 1960.

Spencer, Bernard, Jr .: S t a b i l i t y and Control Characteristics a t Low

Subsonic Speeds of an Airplane Configuration Having Wo Types of N A S A '151 X-303, 1-960.

Variable - Sweep Wings .

.x. 'I Foster, Gerald V., and Morris, Odell A.: S t a b i l i t y and Control Charac - t e r i s t i c s at a Mach Number of 1.97 of an Airplane Configuration Having m o .Ty-pes of Variable - Sweep Wings. NASA T M X-323, 1960.

The B i e l a t , Ralph P., Robins, A. Warner, and Alford, W i l l i a m J., Jr.: Transonic Aerodynamic Characteristics of Two Variable - Sweep Airplane Configurations Capable of Low - Level Supersonic Attack. NASA T M X-304, 1960 .

and Robinson, Ross B.: Spearman, M. Leroy, S t a b i l i t y and Control C h a r - a c t e r i s t i c s at a Mach Number of 2.01 of a Variable - Sweep Airplane

Configuration Capable of Low - Level Supersonic Attack - Outer Wing

Ehept 7 5 O . NASA T M X-310, 1960.

'B

Robinson, Ross B., and Howard, Paul W. : S t a b i l i t y and Control Charac - t e r i s t i c s at a Mach Number of 1.41 of a Variable - Sweep Airplane Con -

figuration Capable of Low - Level Supersonic Attack - Outer Wing h e p t

75' and 108~. N A S A T M X - 320, 1960.

Luoma, Arvo A., and Alford, W i l l i a m J . , Jr.: Performance, Stability, and Control Characteristics at Transonic Speeds of Three V/STOL Air - plane Configurations With Wings of Variable Sweep. NASA T M X-321, 1960.

Foster, Gerald V., and Morris, Odell A.: Aerodynamic Characteristics i n Pitch a t a Mach Number of 1.97 of Two Variable - Wing - Sweep V/STOL Configurations With Outboard Wing Panels Swept Back 7 5 O . NASA TM x - 322, 1960.

Foster, Gerald V., and Morris, Odell A.: S t a t i c Longitudinal and Later.& Aerodynamic Characteristics a t a Mach Number of 2.20 of a Variable - Wing - Sweep STOL Configuration. N A S A T M X-329, 1960.

Robtnson, Ross B., and Spearman, M. Leroy: S t a b i l i t y and Control Char - a c t e r i s t i c s a t a Mach Nuniber of 2.2 of a Variable - Sweep Airplane Con -

figuration Capable of Low - Level Supersonic Attack - Outer Wing

Swept 75'. N A S A TM X-330, 1960.

Luoma, Arvo A.: S t a b i l i t y and Control Characteristics at Transonic i b Speeds of a Variable - Wing - Sweep Airplane Configuration With Wing Out - NASA T M X-342, 1960.

board Panels Swept 113.24' and 75'.

Staff of the NASA: Compilation of Papers Summarizing Some Recent N A S A Research on Manned Military Aircraft. NASA T M X - 420, 1960.

B i e l a t , Ralph P., and Pier-pont, P. Kenneth: Transonic Aerodynamic Characteristics of a Vari&le-Sweep Airplane Configuration Having a 12 - Percent - Thick Wing and an Inboard Pivot Location. NASA T M X-429, Trescot, Charles D., Jr., and Spencer, Bernwd, Jr.: Effect of Reynolds Number on the Low - Speed b n g i t u d i n a l Aerodynamic Characteristics of NASA T M X - 434, 1961.

TW.0 Variable - Wing - Sweep Airplane Configurations, S t a b i l i t y and Control Charac - Robinson, Ross B., and Howard, Paul W.: t e r i s t i c s at a Mach Number of 2.2 of a Variable - Sweep Airplane Con - figuration Having a 12 - Percent - Thick Wing Wept 7 5 O and an Inboard Pivot Location. NASA T M X - 435, 1960.

Ward, Robert J.: Transonic Aerodynamic Characteristics at a Wing Sweep of 1 0 4 ' of a Variable - Sweep Airplane Configuration With an Over - and - Under Engine Installation. NASA T M X - 444, 1961.

Cornette, Elden S.: Wind - Tunnel Investigation of the Wing Buffet Response of a Variable - Wing - Sweep Model a t Subsonic and Transonic Speeds. N A S A T M X - 443, 1961.

Hayes, W i l l i a m C., J r . , and Thompson, Wilson E.: Wind - Tunnel Measure - ments a t Subsonic and Transonic Speeds of the Rolling S t a b i l i t y

Derivatives of a Variable - Sweep Airplane Configuration - Outer Wing

Panels Swept 1 0 8 ' . N A S A TM X-465, 1961.

Longitudinal Aerodynamic Characteristics at Transonic quoma, Arvo A.: Speeds of Two V/STOL Airplane Configurations With Skewed and Variable - Sweep Wings. N A S A T M X-527, 1 9 6 1 .

S t a t i c Longitudinal and Lateral Morris, Ode11 A,, and Foster, Gerald V.: Aerodynamic Characteristics at a Mach Number of 2.20 of a V/STOL A i r - plane Configuration With a Variable - Sweep Wing and With a Skewed Wing Design. NASA TM X - 521, 1961.

A Simplified Method f o r Estimating Subsonic Spencer, Bernard, Jr.: Lift - Curve Slope a t Low. Angles of Attack f o r Irregular Planform Wings.

N'ASA TM x - 525, 1961.

Henderson, W i l l i a m P., and Hammond, Alexander D.: Low - Speed Investiga - t i o n of High - Lift and Lateral Control Devices on a Semispan Variable - NASA T M X-542, 1961.

Sweep Wing Having an Outboard Pivot Location.

Bielat, Ralph P., and Robins, A. Warner: S t a b i l i t y and Control Charac - teristics at Transonic Speeds of TW.0 Variable - Sweep Airplane Config - NASA TM X - 559, 1961. urations Differing i n Wing - Pivot Locations.

Spearman, M. Leroy, and Robinson, Ross B.: Effects of Wing Sweep and Horizontal - Tail Position on the Aerodynamic Characteristics a t a Mach Nuniber of 2 . 2 of a Variable - Sweep Airplane Configuration Having an N A S A T M X-95, 1961.

Inboard Wing Pivot.

Vogler, Raymond D., and Turner, Thomas R.: Explo.ratory Lm-Speed Wind - Tunnel S t a b i l i t y Investigation of a Supersonic Transport Configuration With VariabLe-Sweep Wings. NASA T M X - 597, 1961.

Stonesifer, John C., and Goetz, Robert; C.: Transonic and Supersonic F l u t t e r Trend Investigation of a Variable-%eep Wing. NASA 5% X-598, 1961.

Hayes, W i l l i a m C., Jr., Kemp, W i l l i a m B., Jr., and Thompson, Wilson E.: Wind - Tunnel Measurements and Estimated Values of the Rolling S t a b i l i t y Derivatives of a Variable - Sweep Airplane Configuration a t Subsonic and Transonic Speeds. N A S A T M X-600, 1961.

Hammond, Alexander D., and Polhamus, Edward C.: The Ef'fects of

a

Horizontal - Tail Position on the Longitudinal Aerodynamic Character - i s t i c s at Transonic Speeds of a Variable - Sweep Aircraft Having an Inboard Pivot. NASA T M x - 612, 1961.

Sleeman, W i l l i a m C., J r . , and Robins, A. Warner: Low - Speed Investiga - t i o n of the Aerodynamic Characteristics of a Variable-Seep Supersonic Transport Configuration Having a Blended Wing and Body. NASA TIM x - 619, 1962.

Ward, Robert J., and McKee, John W.: Low - Speed Aerodynamic S t a b i l i t y and Control Charaeterist i c s of a Cambered Fuselage, Variable - Sweep Supersonic Transport Configuration.

NASA T M x - 632, 1962.

Sleeman, W i l l i a m C., Jr.: Low - Speed Investigation of t h e Effects of and Extension of the Wing - Root Leading - Edge Horizontal T a i l H e i g h t Sections on a Variable - Sweep Supersonic Transport Configuration.

NASA TM x - 681, 1962.

Effect of Model Modifications on the Transonic Drag Bielat, Ralph P.: C'haracteri stics of a Variable - Wing - Sweep Airplane Configuration Having an Outboard Wing Pivot. N A S A TIM X-695, 1962.

H i l l j e , Ernest R., and Wiley, Harleth G.: Transonic Dynamic S t a b i l i t y Characteristics i n Pitch and i n Yaw f o r a Model of a Variable - Sweep Airplane Configuration Capable of Low - Level Supersonic Attack. NASA TM x - 618, 1962. i Spencer, Bernard, Jr.: Low - Speed Longitudinal Aerodynamic Character - i s t i c s Associated With Variations i n the Geometry of the Fixed Portion of a Variable - Wing - Sweep Airplane Configuration Having an Outboard Pivot. NASA TM x - 625, 1962.

Luoma, Arvo A.: Longitudinal Aerodynamic Characteristics at Transonic Speeds of a V/STOL Airplaae Configuration With a Fixed Delta Wing Having Auxiliary Variable - Sweep Outboard Panels. NASA TM x - 661, 1961.

Land, Norman S., Wood, John H., and Fougbner, Jerome T., Jr.: A n Inves - t i g a t i o n of the Structural Characteristics of a Simplified Model of a Variable - Sweep Wing. N A S A T M x - 662, 1962.

Re, Richard J., and Cassetti, Marlowe D.: Transonic Longitudinal Aero - dynamic Characteristics of' a Model of a Tactical Fighter Airplane With a Fixed Delta Wing and Variable - Sweep Wing Panels.

NASA TM X-674, 1962.

Capone, Francis J., and Lee, Edwin E., Jr.: Transonic Aerodynamic Characteristics of Three V/STOL Fighter Models With Variable-Sweep or Skewed Wings and Different Engine Installations. N A S A TM X-706, 1962.

Vogler, Raymond D.: Low - Speed Wind-'Punnel S t a b i l i t y Investigation of a Supersonic Transport Model With Variable - Sweep Wings Equipped With H i g h - L i f t Devices. N A S A T M x - 728, 1962.

Ruhlin, Charles L., and Gurley, John R., Jr.: Transonic F l u t t e r Inves - t i g a t i o n of Models of a Proposed Variable - Sweep W i n g .

N A S A TM X-739, 1962.

Henderson, W i l l i a m P.: Low - Speed Longitudinal S t a b i l i t y Characteristics of a Supersonic Transport Configuration With Variable - Sweep Wings Employing a Double Inboard Pivot. N A S A TM X - 744, 1962.

Shaw, David S., and Henderson, W i l l i a m P.: Wind - Tunnel Investigation at Mach Numbers From 1.60 t o 2.86 of the S t a t i c Aerodynamic Charac - t e r i s t i c s of a Supersonic Transport Configuration With Variable - Weep Wings Employing a Double Inboard Pivot.

NASA lllvl X - 745, 1962.

Foster, Gerald V.: S t a t i c Longitudinal and Lateral Aerodynamic Charac - teristics at a Mach Number of 2.20 of a Model of an 8 2 O Delta - Wing NASA T M X-707, Airplane Having Auxiliary Variable - Sweep Wing Panels.

1963 0 ' Spearman, M. Leroy: S t a t i c Longitudinal and Lateral Aerodynanic Charac - t e r i s t i c s at Mach Numbers of 1.41 and 2.20 of a Model of a Low - Aspect - Ratio 83.5' Delta - Wing Airplane Having Auxiliary Variable-&eep Wing Panels. NASA T M x - 708, 1963.

Spearman, M. Leroy: Longitudinal and Lateral Aerodynamic Character= i s t i c s at Mach Numbers From 0.60 t o 2.20 of a Variable - Sweep Fighter Model With Wing Sweep Angles From 2 5 O t o 7 5 O . NASA T M X-710, 1962.

Lockw'ood, Vernard E.: Low - Speed Longitudinal Aerodynamic Character - i s t i c s of an 83.3' Delta-Wing Airplane Model Having A u x i l i a r y Variable - Sweep Wing Panels. N A S A T M X-729, 1963.

Re, Richard J., and Simonson, Albert J.: Transonic Longitudinal Aero - dynamic Characteristics of a Variable - Sweep Tactical - Fighter Model With Wing Sweeps of 2 3 O , 65O, 850, and 1 0 6 ~ . NASA TM X-731, 1962.

Ruhlin, Charles L., and Gurley, John R., Jr.: Transonic Flutter Inves - NaSA T M X - 739, t i g a t i o n of Models of a Proposed Variable-%eep Wing.

1962.

Polhamus, Edward C., Alford, W i l l i a m J., Jr., and Foster, Gerald V.: Subsonic and Supersonic Aerodynamic Characteristics of an Airplane N A S A Configuration Utilizing Double - Pivot Variable - Sweep Wings.

m x - 743, 1962.

Runckel, Jack F., Lee, Edwin E., Jr., and Simonson, Albert J.: Sting and J e t Interference Effects on the Afterbody Drag of a Tw-ln-Engine Variable - Sweep Fighter Model at Transonic Speeds. NASA T M X - 755, 1963.

Henderson, W i l l i a m P . : Low - Speed Aerodynamic Characteristics of a Double - Pivot Variable - Wing - Sweep Airplane Configuration Having Con - N A S A T M X - 773, 1963.

ventional and Gutboard Horizontal T a i l Surfaces.

Spearman, M. Leroy, and Harris, Roy V., Jr.: The Longitudinal and Lateral Aerodynamic Characteristics at Mach Numbers of 1.41 and 2.20 of a Variable-%eep Fighter Model With Wing Sweeps Varying From 25' t o 75'.

N A S A T M X - 759, 1963.

Alford, W i l l i a m J., Jr., Hammond, Alexander D., and Henderson, W i l l i a m Low - Speed S t a b i l i t y Characteristics of a Supersonic W i l l i a m P.: Transport Model With a Blended Wing - Body, Variable - Sweep Auxiliary Wing Panels, Outboard T a i l Surfaces, and Simplified High - Lift Devices.

N A S A x - 802, 1963.

Spearman, M. Leroy, Driver, Cornelius, and Robins, A. Warner: Aero - dynamic Characteristics a t Mach Nmibers of 2.30, 2.96, and 3.50 of a ) Supersonic Transport Model With a Blended Wing - Body, Variable - Sweep Auxiliary Wing Panels, and Outboard T a i l Surfaces. N A S A TM x - 803, Spearman, M. Leroy, and Finch, Virginia M.: S t a t i c Longitudinal and

Lateral Aerodynamic Characteristics at Mach Numbers of 2.36 and 2.86

of a Variable - Sweep Airplane configuration Having Vertical - Ramp Inlets. NASA m x-835, 1963.

Variable Sw'eep Aircraft.

Baals, Donald D., and Polhamus, Edward C. : Astronautics and Aerospace Eng., vol. 1, no. 5 , June 1963, pp. 12 - 19.

VARIABLE - SWEEP PIVOT SCHEMES INBOARD OUTBOARD 0 0 (SCAT 1 6 TYPE) / 7 . I /‘ 1 /> AUXILIARY WING VIRTUAL (SCAT 1 5 TYPE) I 7 I‘/ DOUBLE INBOARD M Figure 1 EFFECTS O F CONFIGURATION ON SUPERSONIC TRIM DRAG Mz2.96; MODEL SCALE ;s ..: C ’ B ’ SCAT 1 6 i 0 - ,I - . 2 - .3 -.4

- a Cm

CL

C.G. FORWARD -

Figure 2 * * * e 9 EFFECT OF AEROELASTICITY ON STABILITY AND TRIM M 3.0

- R I G I D

+

I \ cL - - ELASTIC

C m OFF I + C L Figure 3 EFFECTS OF CONFIGURATION ON STABILITY VARIATIONS WITH WING SWEEP ANGLE AND MACH NUMBER .5 A. C.

.

AND ..

C. G. .6 SCAT 1 6 LOCATION, . 4 r /-C.G. POSITION A FRACTION OF F b .70 40 80 0 1.0 20 3.0 ALE,= DEG M

Figure 4

EFFECT OF PIVOT LOCATION ON STABILITY VARlATl ON WITH SWEEP M = O .2 -

. I - '\ , I

dCrn - EXPER I MENT, PIVOT dCL WING BODY LOCATION __Q___ 2 THEORY - - .I WING ALONE

--

- - - - - - -

--- I 1 I I - . 20 20 40 60 80 ALE, DEG Figure 5 LOW - SPEED PITCH CHARACTERISTICS LOW SWEEP n CL Figure 6 LOW - SPEED PITCH CHARACTERISTICS HIGH SWEEP .04 .O 4 C m .08 I I I 1 1 .2 .4 .6 .8 1 . 0 1.2 Figure 7 HIGH - SPEED PITCH CHARACTERISTICS A ~ ~ m 7 5 ' M = 1.2 M = 3.0 .04r r -.04' I 1 0 .2 .4 . 6 0 .2 . 4 . 6 C L C L Figure 8 EFFECTS OF FOREWING GEOMETRY ON THE LOWSPEED . ^ PITCH CHARACTERISTICS c, .04 -.04 .04r

- . I 2 l 2 . I .2 I .6 +

-.IS .4

Figure 9 EFFECTS OF HORIZONTAL - TAIL LOCATION ON LOW - SPEED TAIL PITCH CONTRIBUTION (Acm)T*l L --_- .04- / ,043 / -.04 I -.08 I I I I I - 4 0 4 8 1 2 1 6 20 24 a. DEG Figure 10 EFFECT OF WING DEVICES ON LOW - SPEED PITCH CHARACTERISTICS

o* .. - LEADING FOREWING

b= EDGE DEFLECTION, DEG

cm

- CLEAN

0 I 7- SLAT 0 -.04t \ I

---

-.08/

----- NOTCH

- . I 2 ' I I I 1 0 .4 .8 I . 2 I .6 C L Figure 11 LOW - SPEED STABILITY CHARACTERISTICS OF M WING CON FI GU RAT1 ON

n

- .04

0 .4 .8 1 . 2 TAIL ON CL

- .3

0 20 40 60 80

Au, DEG

Figure 12

GEOMETRIC CHARACTERISTICS OF HIGH-LI FT MODEL 8f =40°* SECTION A - A Figure 13 HIGH - LIFT CHARACTERISTICS 6f = 40° 2.4 AE,DEG I .6 r< RN bf/b, CL,TRIM .8 / : s a f = o 0 8 1 6 24 0 8 1 6 24 a, D E G a, DEG - . 2 - 0 .8 1 . 6 2.4 CL Figure 14 . a b *

E F F E C T -06 H~GW~LI&' D E ~ C E S

ON THE DRAG CHARACTERISTICS . 6 .4 C D .2

I

0 .4 .8 1.2 1.6 2.0 2.4 2.8 / Figure 13 EFFECT OF FOREWING ON HIGH - LIFT CHARACTERISTICS ASPECT RATIO I O FOREWING 1 . 6 - - * 2 4 v $ @ 5 c - .8 .333c HIGH - LIFT SYSTEM I I I I 1 1 0 4 8 1 2 a, DEG Figure 16 EFFECT OF FOREWING ON HIGH - LIFT CHARACTERISTICS - .5 I I I I FOREWING .4 ,3 CD .2 / - CL2

A

. I

% = cD,o+ a

- 2 5 '

L I I I I I I I I c .8 1 . 6 2.4 3.2 C L Figure 17 EFFECT OF CONFIGURATION ON LATERAL CHARACTERISTICS M = 0.2 n I I SCAT 1 6 SCAT 1 5 4 .008 CROSS SECTION cn P.004 C 1~ -.004 \76 0 IO 2 0 3 0 0 IO 20 a , DEG a , DEG Figure 18 ' " - * - E F F E C T O F C O N F I G u RATIC$~O~L~+&HARACTER ISTI cs SUPERSONIC; ALE = 75O

k SCAT 1 5

cnP I C - b . 2 % 6 2.40

- . 0 0 4 0 -

1 0 20 30 0 I O 20 30 a , DEG a ,DEG Figure 1 9 DYNAMIC STABILITY DERIVATIVES OF A VARIABLE - SWEEP TRANSPORT CONFIGURATION a = 0"

/ a , -/-'

- - - A LE , DEG LANGLEY DATA AMES DATA 25 0 0 5 0 8 : Cnr-cngs --4 PER RADIAN -.8 - 1.2

2r

c m q + c m u , O PER RADIAN - .2 I I I I I I I I I

- .4 '

0 . 4 .8 1 . 2 1 . 6 2.0 2.4 2.8 32 3.6 M Figure 20 1 ', NASA WING PLANFORM STUDIES Raymond 3 4 . Hicks, and Edward J. Hopkins @MA, Ames Res,earch Center SUMMARY Supersonic commercial air transport (SCAT) studies have shown the need for increased performance for wings with supersonic leading edges and the requirements for low - speed stability research on planforms with subsonic edges. Summarized are recent planform studies by the National Aeronautics and Space Administration which indicate wing warp to have a Studies at 'low powerful effect on supersonic trimmed lift - drag ratio.

speed have shown that notched planforms tend to reduce the pitch - up tendency .

4 J T

INTRODUCTION Recent SCAT feasibility studies completed by The Boeing Company and Lockheed California Company have concentrated on the SCAT 1 6 variable - sweep - wing configuration and the SCAT 17 fixed - wing configuration.

SCAT 17 studies have shown that, although the planform chosen lends itself to lightweight constructibn, a need to increase the aerodynamic performance throughout the flight Mach number range exists. Other fixed - wing configurations incorporating some of the features of the SCAT 4 and SCAT 15 supersonic configurations are also of interest because of the high aerodynamic efficiency possible at the cruise condition. Low - speed stability problems, however, have been encountered with highly swept arrow wings.

The purpose of the present paper is to present the results of recent planform studies conducted by the NASA and to relate these studies to the problem of supersonic aircraft design. The first part of the paper then will concentrate on planform studies which indicate some possible gains in high - speed performance on the SCAT 17 type and on tailless configurations, and the last part will show the results of some Planform studies aimed at the low - speed stability problems of wings having sub - sonic leading edges at the Mach 3 design point.

SYMBOLS aspect ratio span drag coefficient at zero lift lift coefficient pitching - moment coefficient pitching - moment coefficient at zero lift static margin lift - drag ratio maximum lift - drag ratio trimmed lift - drag ratio Mach number Reynolds number thickness - chord ratio leading - edge sweepback angle, deg D I S C U S SION I SCAT 17 wing is characterized by a planform having a near - delta shape and very thin airfoil sections.

In order to assess changes in planform, several variations on this type of wing have been tested and are shown in figure 1 . In the upper part of the figure are shown a an ogee leading edge, and a trapezoidal wing.

delta wing, a wing with

All these wings had the samp span and had an aspect ratio of 1.55. All

i

the wings were t e s t e d on the same body and each had a circular - arc wing section w i t h thickness - chord r a t i o s of 3 percent a t the fuselage and 2 percent near the t i p .

In addition t o the planar wings shown, wings employing t w i s t and camber and designed f o r maximum l i f t - d r a g r a t i o a t a Mach number of 2.20 were a l s o tested i n a l l these planforms. Wings i n the lower part of the figure have an aspect r a t i o of 2.17 and are of modified trapezoidal, ogee, and d e l t a planform.

The wings had 30 - 70 hexagonal sections and

a l l had a thickness - chord r a t i o of 2 percent.

The right - hand side of figure 2 shows a summary of the aerodynamic data obtained with the wings of aspect r a t i o 2.17; t h i s summary is pre- ) sented as ( L / D ) w and C D , ~ pl o t t e d against Mach number. Compari- sons of the data f o r the wings of aspect r a t i o 2.17 and 1.55 are not v a l i d because of differences i n model fuselage and i n model t e s t condi - tions. The curves i n t h e upper l e f t of the figure show only minor dif - ferences i n m a x i m u m lift - drag r a t i o f o r a l l three planforms throughout the t e s t Mach number range, w i t h a s l i g h t advantage i n favor of the trap - ezoid of about 0.3 t o 0.5 i n ( L / D ) w . For a l l p r a c t i c a l purposes, the differences i n m a x i m u m l i f t - d r a g r a t i o may have only a s l i g h t bearing on the selection of wing planform within the range of wings tested. In the right - hand portion of the f i g u r e note the l a r g e r differences i n with values f o r the modified trapezoidal and d e l t a a t a l l speeds, cD, o wings appearing s l i g h t l y higher. However, the differences i n C D, 0 are offset t o some degree by better drag - due - to - lift f a c t o r s f o r these planforms and the maximum l i f t - d r a g r a t i o appears t o be about the same These data showed a s l i g h t advantage i n favor of f o r a l l planforms.

There appear the ogee planform, again about 0.3 t o 0.5 i n ( L / D ) u .

again r e l a t i v e l y small differences from which t o s e l e c t wing planform on a performance basis as indicated by these data.

Another parameter of i n t e r e s t i s the shift i n s t a t i c margin w i t h Mach number inasmuch as it may be a measure of trim drag throughout the speed range. Figure 3 shows s t a t i c margin &&EL as a function of . Mach number f o r the wings tested. A comparison of a l l wings, s t a r t i n g a t about the same s t a t i c margin, shows that changes from 4 percent t o about 20 percent are evident as Mach number increases a t both the aspect r a t i o s . A s l i g h t l y smaller s h i f t i n s t a t i c margin i s indicated f o r the ogee planform a t the higher speeds. A t M = 2.20 t h i s advantage is about 2 percent f o r both aspect r a t i o s .

The e f f e c t of high s t a t i c mar - gin on trimmed l i f t - d r a g r a t i o may be sizable depending not only on t a i l arm but on pitching moment a t zero l i f t as well.

The warping of plan - forms such as those shown f o r maximum lift - drag r a t i o at some super - sonic speed generally r e s u l t s i n washout along the span. For wings w i t h highly swept leading edges then, a positive pitch i s expected at zero l i f t ; t h i s pitch, when combined with a stable pitch variation, tends t o reduce the drag due t o trimming. Using the s t a b i l i t y variations with

Mach number shown i n figure 3, figure 4 w a s prepared t o show the favor -

able e f f e c t of wing warp on trimmed l i f t - d r a g r a t i o . Shown are trimmed l i f t - d r a g r a t i o s as a function of s t a t i c margin at full - scale Reynolds number and a t Mach numbers of 0.97 and 2.20.

The data shown i n figure 4 are applicable t o a t a i l l e s s configura -

t i o n of aspect r a t i o 1.35. The planform and trailing - edge controls selected as i l l u s t r a t i v e are shown i n the sketch at the top of t h e fig - ure; however, similar curves may be shown f o r the other planforms. The v e r t i c a l l i n e s on the figure show the s t a b i l i t y l e v e l consonant with each Mach number from figure 3 .

A t a Mach number of 0.97 the planar wing shows a decrement of about 3.0 i n due t o trimming at the required s t a b i l i t y level, whereas the L/D warped wing t r i m s with no control deflection a t a s t a t i c margin of 14 per - cent because of the higher positive Cm a t zero lift. A t M = 2.20 neither wing had sufficient Cm,o t o t r i m the potential L/D of t h e planform but the warped wing showed a value of L/D about 1.5 greater a t t r i m with the required s t a b i l i t y . The s e n s i t i v i t y of L/D t o s t a t i c margin i s approximately the same f o r a l l these wings and it can be seen t h a t t h e 2 - percent advantage i n s t a t i c margin mentioned previously f o r L/D at the the ogee wing will only amount t o about 0.20 t o 0.30 i n t r i m s t a t i c margin.

7 by Francis E. McLean and Harry W.

As pointed out i n paper no.

CL may be Carlson, f u r t h e r warping of t h i s wing f o r a higher design expected t o reduce the m a x i m u m value of L/D s l i g h t l y . But, based on the r e s u l t s shown i n figure 4, the higher C associated with t h e % 0 CL may be expected t o more than make up the slight loss higher design i n ( L / D ) w by providing a trimmed L/D which i s much closer t o the the planform could be m a x i m u m value shown. According t o t h i s figure, warped even more i n an attempt t o move the s o l i d curve t o t h e r i g h t while L/D.

minimizing the reduction i n the l e v e l of m a x i m u m I n addition t o t h e data shown, lateral - directional data a r e avail - ) no r a d i c a l d i f - able on a l l these planforms and, as might be expected, ferences appeared between wings as a r e s u l t of the planform changes.

N o w consider wings employing subsonic leading edges at the M = 3.0 cruise design point. As pointed out previously the wings thus far con - sidered have been arrow planforms as on SCAT'S 4 and 15; these wings have been characterized by a longitudinal i n s t a b i l i t y known as pitch - up i n t h e low - speed or landing f l i g h t regime.

Although it has been shown that placement of the horizontal tail below the wing-chord,plane is effective in the elimination of pitch - up, many of the configurations developed for SCAT study preclude, for one reason or another, this tail placement. The problem then is confined to the development of wing planforms having high aerodynamic aspect ratio and low structural spans which are pitch - up free and to the placement of the horizontal tail in a suitable location consistent with overall good design.

The NASA has recently completed a preliminary investigation to study means by which pitch - up may be eliminated or, at least, helped by the judicious selection of planform and the careful matching of wing and horizontal - tail characteristics.

Some of the planforms tested are shown in figure 5.

For the tests of this series of wings, the leading - edge sweepback was held constant at 7 5 O . All the wings were tested in combination with a fuselage and employed flat - plate sections with rounded leading edges. Shown in the upper left - hand corner of the figure is a basic trapezoid, which is pitch - up free but shows limited cruise performance potential. Wings were tested with trailing edges cut away to increase aspect ratio; these wings are shown across the top of the figure. Leading edges were also notched or cut as shown in the left - hand column in the figure. In addi - tion, several combinations were tried of leading - edge and trailing - edge notches as shown by the remaining planforms.

Some data which are typical of these tests are shown in figure 6.

The wing planforms are those delineated by the aspect - ratio numbers with the sketches. Pitching - moment coefficient is plotted as a function of lift coefficient; these coefficients have been reduced by using the mean aerodynamic chord and the area of each wing, respectively. The static stability level at zero lift has been adjusted to 5 percent for all wings in order to judge nonlinearities more easily.

The curves in the upper left - hand plot in the figure are typical of those for the arrow wings and show pitch - up, as the basic trapezoid trailing edge is cut, at all the aspect ratios indicated. These data

are typical of the curves obtained with the SCAT 4 configuration and

I were, to a large extent, responsible for the configuration being tabled at SCAT midterm.

The curves on the upper right - hand plot in the figure show data for variations in the leading - edge notch location from the basic trapezoid Note that of 25 percent, 33 percent, and 40 percent of the semispan.

the planforms with the 25 - percent and 33 - percent leading - edge notch locations become more stable with increasing lift, whereas the planform with the 40 - percent leading - edge notch location tends to less stability.

The lower plot is an attempt to show the sensitivity of the planform with the 33 - percent leading - edge notch location to trailing - edge notch location. The trailing - edge notch is moved spanwise with +50° trailing - edge sweep and a pronounced change in the shape of the pitch curve is evident.

1 4 by A . Warner Robins and Another wing, discussed in paper no.

Richard T . Whitcomb, is similar to the A = 1.48 wing but of higher aspect ratio and pitched down if chord - extensions were used on the outer panel.

Figure 7 shows a comparison of data at subsonic and supersonic

speeds on similar, but not identical, planforms. These data are included to indicate the marked similarity between the subsonic and supersonic pitch curves and to indicate the type of data which are becoming avail - able on arrow - type planforms as a function of speed. Further work of this sort will no doubt be accomplished in the future.

It appears from the test results shown that there are combinations of leading - and trailing - edge notch which provide acceptable longitudinal stability variations, at aspect ratios significantly higher than might be indicated by the Shortal-Maggin stability curve. Further tests are required, however, to validate the expected aerodynamic performance of the notched planforms at transonic and supersonic speeds.

CONCLUDING FGMAR?S It has been shown that wing planform differences on thin sharp - edged wings appear to have small effects on the maximum lift - drag ratio at sub - sonic, transonic, and supersonic speeds. Planform differences were also shown to have a small effect on aerodynamic - center shift with a change in Mach number. Wing warp, on the other hand, was shown to have a power - ful effect on trimed lift - drag ratio of all planforms in the lower aspect - ratio series.

On wings with subsonic leading edges at a Mach number of 3.0, the notched wings appear to show promise of reducing the pitch - up tendency at low speed, although further testing is required to validate the expected aerodynamic performance at transonic and supersonic speeds.

A = 2.17 t/c = 0.02 Figure 1 SUMMARY OF PERFORMANCE CHARACTERISTICS A=2.17; R = 5X106 A = 1.55

0 I_

cD,o "i?, *01 I-3: 1 2 3 0 I 2 3 0 MACH NUMBER Figure 2 e * r l i EFFECT OF MACH NUMBER ON STATIC MARGIN OGEE ---______ DELTA A = 2.17 A = 1.55 TRAPEZOID

~-~~ a C L 72 ~~ - --- -

- . 3 1 . 0 2.0 3.0 0 1 . 0 2.0 3.0 MACH NUMBER Figure 3 EFFECT OF WING WARP ON TRIMMED L/D FULL - SCALE REYNOLDS NUMBER; Az1.55 - - - - PLANE WARPED M = 0,97 M = 2.20 I I I 1 0 -08 - . I 6 -,24 0 -,08 -.I6 -.24 Figure 4 A = 1.03 3 3 + 4 A = 1,14 1.28 1.38 1.48 4 1 A = L 2 6 TYPICAL WING Figure 5 SUBSON IC PITCH CHARACTER1 STI CS M = 0.40

.08 r r

-.08 0 .5 1.0 CL .08 A =L48 -.08 0 .5 I .o CL Figure 6 SUBSONIC AND SUPERSON!G..:PITCH COMPARISONS . .*’- i . = M = 0.40; ALE = 75” M 2.86; ALE = 67O

r

Figure 7 I P I

W

A-Langley Research Center SUMMARY A substantial quantity of research information has been obtained on the stability characteristics of aircraft configurations at supersonic speeds. The investigations include the effects of variable wing sweep, wing modifications, vertical - tail modifications, and body - cross - section changes. Illustrative examples of the results are presented in order to point out some of the factors that affect the lateral - directional stability characteristics.

w. / ( I j T W m -

INTRODUCTION SCAT feasibility studies have indicated some stability problem areas, such as pitch - up, low directional stability, and high effective dihedral. As a result, several wind - tunnel investigations have been undertaken in order to gain insight into these problems and to indicate some factors that may aid in providing solutions. Paper no. 11 by Alford, Lockwood, McKinney, and Greif, and paper no. 12 by Taylor, Henderson, Hicks, w d Hopkins have considered the problem of pitch - up at subsonic speeds. The stability characteristics of delta - wing config - urations are presented in the paper by Anderson (paper no. 10). The present paper will smarize some recent stability investigations con - ducted at supersonic speeds with the primary emphasis on the lateral - directional stability problems. The factors to be discussed include the effects of variable wing sweep, the effects of filling the notch of swept wings, the effects of outboard - vertical - tail shape, and the effects of body cross section.

SYMBOLS CL lift coefficient effective - dihedral parameter C variation of effective - dihedral parameter with lift C2 $CL coefficient C directional - stability parameter tail contribution to directional - stability parameter

(NnP> TAIL

M Mach number U angle of attack, deg ALE sweep angle of wing leading edge, deg DISCUSS I O N Variable - Sweep Aircraft Since several transport configurations utilizing variable - sweep wings have been proposed, an investigation of a generalized variable - sweep model has been undertaken in which the effects of wing sweep have been studied systematically over a large Mach number range. The variable - sweep research model is presented in figure 1 . The model was tested in combined pitch and sideslip for a range of wing sweep angles

from 3 0 ° to 75' over a Mach number range from 1.4 to 3.5 with the ver -

tical tail both on and off. Both a plane wing and a twisted wing were investigated. Three wing - tip controls having different hinge - line skew M = 1.41 angles have also been investigated with various wigg sweeps at and 2.2, but none of these results are included in the present paper.

Wing sweep was found to have little effect on the directional sta - bility characteristics, but, as might be expected, there are significant The effects of effects of sweep on the rolling - moment characteristics.

wing sweep and Mach number on the variation of rolling moment due to sideslip with lift coefficient are presented in figure 2. The ("lPCL) The results coefficients are based on the geometry of the 75O wing.

indicate a general increase in effective dihedral with increasing wing sweep and a general decrease with increasing Mach number.

The rolling moment provided by the wing is a function of the wing - panel lift - curve slope and thus would be expected to change with sweep angle and Mach number. Theoretically, the rolling moment due to side - slip would be negative for subsonic leading edges but would decrease and possibly change sign for supersonic leading edges. The tick marks i .” D I D . u on the curves in figure 2 ihCnc&te %&*~onic-leading-edge points and show the relatively high U R t f y e vakes of rolling moment due to side - slip for wingS swept behin% the MEh l h e s and relatively low values for wings swept ahead of the Mach lines.

In the case of the lowest sweep and the highest Mach number (Am = 45O, M = 3.5), a small positive value was obtained.

It is apparent that wings highly swept to provide subsonic leading edges for good lift - to - drag characteristics will inher - ently have high values of effective dihedral.

Of course, the level of effective dihedral can be regulated to some extent by the contributions of the tail, wing geometric dihedral, and wing height. Another point to be noted is the fact that the variation of effective dihedral with Mach number can be regulated by varying the wing sweep provided that proper consideration is given to the effects of sweep on the other aero - dynamic characteristics.

Fixed - Wing Modifications One of the modifications proposed for highly swept fixed wings has been to reduce the notch ratio for the purpose of improving the pitching - moment linearity and also to improve the structural integrity.

Accordingly, a systematic study has been made of the effects of various wing modifications, including variations in notch ratio. The wing - modification model is presented in figure 3 . The basic model consisted of an ogive - cylinder body with a 62O swept wing. Various modifications to the wing provided for two full - span leading - edge extensions having sweep angles of 64O and 6 7 O and three trailing - edge inserts that resulted in progressively filling the notch until a trapezoidal plan - form was obtained. Each of the trailing - edge inserts have been tested with each of the leading - edge extensions. In addition, two semispan leading - edge gloves having sweep angles of 670 and 71° were tested in conjunction with the basic 62O swept wing. Tests have been made of the model in combined pitch and sideslip with the vertical tail both on and off for Mach numbers of 1.6, 2.2, and 2.9.

The effects of progressively filling in the notch of the 62O swept wing until a trapezoidal wing is obtained are illustrated in figure 4 .

These results are all based on the geometry of the 62O swept wing as CL represents I indicated by the shaded region; hence, a given value of a constant weight condition but varying wing loadings. The sideslip characteristics indicate less deterioration of Cn with increasing

P

and generally lower values of effective dihedral for the filled - in CL wings. The improvement in C results from the fact that for a given n$ lift the angle of attack is lower for the larger filled - in wings and the detrimental effects of the forebody lift on are lessened.

Cn

P

i The reduction in -Czp is partly due to the'reduction in wing - panel aspect ratio and the attendhnt decrease in pdel lift - curve slope.

SCAT 15 Vertical - Tail Study Previous investigations of the SCAT 15 model at supersonic speeds indicated relatively high values of effective dihedral and some deterio - As a ration of directional stability with increasing angle of attack.

result, an investigation has been conducted to determine the effects of decreasing the vertical - tail area above the wing chord plane and increasing the area below the chord plane in the hope that the level of effective dihedral might be reduced while the directional stability was improved. The model was The SCAT 15 model i s represented in figure 5 .

investigated with a series of equal - area vertical tails ranging from a tail having all its area above the wing chord plane to a tail having a11 its area below the wing chord plane. In addition, one test was made with the tails removed.

The model was tested at combined angles of attack and sideslip for Mach numbers of 2.3, 2.6, 3, and 3.5.

The results for M = 2.3 (fig. 6) indicate that relocating the entire vertical - tail area from above the wing chord plane to below the wing chord plane does reduce the effective dihedral throughout the angle - of - attack range and provides an increase in directional stability at high angles of attack. The increment of effective dihedral provided by the vertical tails i s relatively small compared with the basic level for the wing - body configuration, however, and it is doubtful whether the reduction in obtained by placing the entire tail area below the -CzB wing would be worth the complications involved. The small effect of the vertical tail on Cz is characteristic of outboard - tail configu -

B

rations and results from an interference effect such that the pressure field from the tail produces a rolling moment from the wing tip which is opposite to that provided by the tail.

Body - Cross - Section Study j.

Among other general studies made of factors affecting the lateral - directional stability characteristics has been the investigation of the effects of body cross section. The body - cross - section research model is

presented in figure 7 . The body cross sections shown extended over the

entire body length and the wings were located at the midbody height. A

49' swept wing was investigated at Mach numbers of 1.4 and 2.2 and a 7 5 '

wing was investigated at M = 2.5, 3, 4, and 4.7. The model was tested f in combined pitch and sideslip with the vertical tail on and off. I n addition, the model with triangular cross sections was also tested with the wing removed.

Lateral and directional stability characteristics are presented in figure 8 for the 49' swept configuration at M = 2.2 with a circular and a vertical elliptical body cross section. It is apparent that body cross section can have a significant effect on the sideslip derivatives in that not only are the variations of Cn and C2 with angle of

P P

attack considerably different for the wing - body configurations but the contributions of the vertical tail to Cn- are affected as well. These

P

j differences are related to differences in the body cross flow for par - ticular body shapes and, to some extent, might be affected by Reynolds number. However, results obtained for the triangular body shape indi - cated no effects of Reynolds number over a range from 1 X 1 0 6 to 4 X 1 0 per foot. Although these results are qualitative in nature, they serve as an indication that body - cross - section effects on the lateral stability characteristics should not be overlooked.

CONCLUDING REMARKS A substantial quantity of research information has recently been obtained on the stability characteristics of aircraft configurations at supersonic speeds. The investigations include the effects of variable wing sweep, effects of wing modifications, effects of vertical - tail modifications, and effects of body cross section. Each of the investi - gations indicates important effects on the lateral - directional stability characteristics that should be considered in specific configuration design. Continuations of these investigations will include the effects of wing height and wing geometric dihedral for highly swept wings, addi - tional studies of lateral control devices, and additional planform modifications.

VAR I ABLE - SWEEP MODEL

A L E = ~ O " , 45O, 65". AND 75 " ; Mz1.4, 2.3, 3, AND 3.5; VERTICAL TAIL ON AND OFF TI P - CONTROL

/qf HINGE LINES

/ I I I Figure 1 EFFECT OF SWEEP AND MACH NUMBER ON EFFECTIVE DIHEDRAL LEADING EDGE I I I 40 60 80 ALE, D E G Figure 2 ir * *H, .

WING - MODIFICATION MODEL Mz1.6, 2.2, AND 2.9; VERTICAL TAIL ON AND OFF Figure 3 EFFECT OF WING MODIFICATION M= 2.2 0 0 0 .002 C "P

- .002

- .004

C 1 8 -.002

- .OO 4

0 .4 .8 I .2 cL Figure 4 SCAT I5 VERTICAL - TAIL STUDY M = 2.3, 2.6, 3, AND 3.5; VERTICAL TAIL O N AND OFF Figure 5 EFFECT OF TAIL VARIATIONS SCAT 15; M = 2.3

--------- -----------_

C

------ _ _

"B OFF I I I I -.004 O L I I I I

-.004 '

- 4 1 2 0 4 8 a , DEG Figure 6

* . 0 * - - -

* - .* . * . * ' BODY - CROSS - SECTION MODEL ALE = 4S0 A T M = 1.4 AND 2.2 ; ALE = 75' AT M = 2.5, 3, 4, AND 4.7; VERTICALTAIL ON AND OFF; WING ON AND OFF . - Figure 7 EFFECT OF BODY CROSS SECTION ALE=^^'; M z 2 . 2 I / 'Oo2 TAIL OFF -.002 ---/ -.004 L - . 0 0 4 - -.004 0 I O 20 0 1 0 20 a, DEG a, DEG Figure 8 ADDITIONAL CONFIGURATION APPROACHES

By A . Warner Robins and Richard T. Whitcomb Ad

-----L.

Langley

E

- - - - SUMMARY Some of t h e major problems encountered i n the SCAT f e a s i b i l i t y studies are reviewed and changes t o improve the f o u r basic SCAT config - y a t i o n s a r e discussed.

Improvements i n nearly a l l problem areas appear ksible. Advanced versions of each of the four SCAT concepts a r e being investigated.

w

INTRODUCTION The supersonic transport f e a s i b i l i t y studies have replaced NASA's original qualitative assessment of t h e problems of each of the four SCAT concepts with a reasonably complete quantitative evaluation.

With the provision of a b e t t e r understanding of the r e l a t i v e importance of the problems of supersonic transport design, new approaches t o existing concepts a r e suggested.

Concepts represented are f o r SCAT'S 17, 16, 13, and 4 .

The purpose of t h i s paper i s t o review some of these problems and t o indicate t h e directions t h a t t h e e f f o r t s of N A S A a r e now taking toward making improved configurations of a l l four concepts.

S Y M B O L S A aspect r a t i o aspect r a t i o of wing i n minimum sweep position cD,W wave - drag coefficient reference chord cREF L/D l i f t - d r a g r a t i o Mach number M mass - flow r a t i o longitudinal distance measured from wing apex X &.C. aerodynamic - center s h i f t , percent c m sweep angle a t wing leading edge

*m

Subscripts: m a x i m MAX M I N minimum DISCUSSION S C A T 17

Figure 1 shows a sketch of each of t h e f i n a l S C A T 17 configurations

as developed by the t w o contractors. (See paper no. 3 by R. Richard Heppe and J i m Hong and paper no. 4 by Lloyd T. Goodmanson, W i l l i a m T.

Hamilton, and Maynard L. Pennell.) The problems of these configura - t i o n s include undesirably low l i f t - d r a g r a t i o s i n both low - speed and cruising f l i g h t and low - speed handling q u a l i t i e s t h a t are i n f e r i o r t o those of present 3et transports. The two configurations do not share these problems equally.

A configuration indicative of the direction presently being taken by $hose concerned with the SCAT 17 concept i s shown i n figures 2 and 3 .

Changes include the adoption of a wing planform having an increased effective aspect r a t i o and a complex leading edge, the inboard region being subsonic and the outboard region, supersonic. Increased wing volume and improved area progressions r e s u l t . The new configuration reverts t o the arrangement of the original configuration with both canard and a f t tail. Figure 2 shows the a f t t a i l folded down so as t o increase directional s t a b i l i t y and decrease longitudinal s t a b i l i t y as required i n cruising f l i g h t . The outboard wing leading edge i s coni - c a l l y cambered and provides f o r leading - edge droop about a hinge l i n e shown as the second l i n e behind the leading edge. These changes were made with a view t o improving low - speed l i f t - d r a g r a t i o s and speed sta - b i l i t y i n approach, as well a s low - speed handling qualities. I n order t o improve t r i m drag and t o r e t a i n reasonably low drag due t o l&ft.at i

.. 4

9 .

.. . - *

a * .

- C r

the cruise point, several wing warp schemes a r e being studied. A d d i - t i o n a l work along these and, similar l i n e s i s planned.

SCAT 16 The contractors' f i n a l S C A T 16 configurations are sketched i n f i g - Major problems, with magnitude

ure 4 . (See paper nos. 3 and 4.)

varying between the two configurations, were as follows: j e t impinge - ment on empennage and aft - end structure, pitch - up, and undesirably low cruise l i f t - d r a g ratios. Changes which would r e s u l t i n significant improvements i n each of these problem areas are incorporated i n the configuration shown i n figures 3 and 6. These changes include (1) s h i f t ; t o an inboard pivot with a small glove area, (2) notch - located nacelles i n packaged pairs beneath t h e fixed underwing panel, and ( 3 ) increased wing area.

Selection of an inboard pivot arrangement substitutes the problem of large aerodynamic - center shifts f o r the pitch - up problems inherent i'n outboard pivot arrangements w i t h large glove areas having high sweep.

Ordinarily t h i s pitch - up is solved by the i n s t a l l a t i o n of low horizontal t a i l s but, i n SCAT configurations, t h i s solution generally tends t o r e s u l t i n serious jet - impingement problems. Other means f o r solving the pitch - up have not as yet proved very satisfactory.

Figure 7 shows the difference i n aerodynamic - center s h i f t due t o sweep between t y p i c a l outboard and inboard pivot arrangements as w e l l as the s h i f t due t o Mach number for both. (These curves are based on the forward - sweep reference chord and nonelastic wings.) The much larger aerodynamic - center shift with sweep f o r t h e inboard pivot prac - t i c a l l y precludes the use of a low minimum sweep; t r i m drags arising from attempts t o t r i m out t h i s e n t i r e shift would be prohibitive. This problem i s the reason f o r the selection, a t some loss i n high - lift potential, of a forward - sweep r e s t r i c t i o n of 3 6 O i n the modified SCAT 16 (shown i n f i g . 5 ) . Since t h e t r i m penalties are dependent on the increment between the aerodynamic - center curve and the curve of the center - of - gravity location, the variation i n the center - of - gravity loca - t i o n due t o t h e sweeping of the wings i s a l s o shown i n figure 7. The increment remaining a t the cruise Mach number, which includes both the aerodynamic - center s h i f t due t o sweep and t o Mach number, is now seen t o approach t h a t which might be trimmed out without penalty through the use of wing warp as discussed i n paper no. 7 by Francis E. McLean and Harry W. Carlson. Thus, the net r e s u l t of the pivot - location change w i l l be some reduction i n high - lift capability, a reduction of t r i m drag i n the midsweep (transonic) range, and a slight increase i n the t r i m penalty a t cruise; with pitch - up reduced t o a type ( c l a s s i c swept - wing rather than glove - induced) which should permit solution without an accompanying j e t - impingement problem.

I n t h i s area t h e The second change w a s i n engine arrangement.

problem w a s t o provide a short, two - engine pack which could be located so as t o provide f o r reasonably good a i r c r a f t balance as w e l l as the desired buoyancy on t h e receding surfaces of the wings and fuselage it w a s desired t h a t a l l i n l e t s be under the afterbody. Furthermore, wings t o provide b e t t e r pressure recovery and common i n l e t control Figure 8 shows such a nacelle which u t i l i z e s , i n addition t o schemes.

the precompression due t o the wing, a 2.5' precompression from the s p l i t t e r between the i n l e t s . I n paired i n l e t s , such s p l i t t e r s are I n t h i s arrangement, the required so as t o avoid unstart interference.

I n com - s p l i t t e r i s wedged t o provide t h e additional precompression.

parison with two - engine packs without the additional precompression, .

this arrangement provides f o r reductions i n i n l e t area, diffuser length, ) and wetted area. A b e t t e r match of mass flow over the transonic and supersonic speed ranges i s a l s o provided.

The achievement of high super - The t h i r d change was i n wing area.

sonic L/D l e v e l s lies t o a great extent i n obtaining low values of two r a t i o s . The first i s the r a t i o of wetted area t o wing area and t h e second, t h e r a t i o of t h e 2 / 3 power of the t o t a l volume t o the wing area.

E a r l i e r configurations of the SCAT 16 type had s m a l l wing areas and, i n addition t o high wing u n i t weights, had high values of these two r a t i o s .

Increasing the wing area would be beneficial i n both respects and would reduce, as well, t h e high - lift - coefficient requirement a t low speeds.

Estimates of trimmed l i f t - d r a g r a t i o a t the cruise point exceed by 0.6 t h e values f o r t h e o r i g i n a l USA SCAT 16 configurations.

I n summary, the pitch - up and attendant problems of the outboard pivot arrangement have been replaced by the aerodynamic - center shift, problem of the inboard pivot arrangement, since such a trade i s believed, Other considerations of a t present, t o provide a b e t t e r overall r e s u l t .

Design of wind - tunnel models of the present approach appear promising.

this configuration is therefore underway.

S C A T 15 The contractors' f i n a l versions of SCAT 15 are sketched i n f i g - high s t r u c t u r a l w e i g h t of both Major problems were as follows: ure 9.

i t h e fixed and variable - sweep wing panels, the structural - mechanical problem of extensive wing overlap, and aeroelasticity. A modified con - figuration i s shown i n figures 10 and 11. Changes include (1) sub - stantial reductions i n t h e s t r u c t u r a l spans of both the fixed and variable - sweep panels, ( 2 ) increased absolute thickness of both panels, ( 3 ) decreased wing overlap with a full - depth pivot, and (4) higher maxi - mum sweep with an attendant reduction i n thickness r a t i o . The packaged engines provide f o r a s t r u c t u r a l box covering about one - half of t h e fixed - wing semispan, and they are located t o provide positive i n t e r - ference beneath the aft porkion of theAwing.&$ cruise Mach number as i n the original $CAT 13. The effects of these”changes should be t o reduce structural and aeroelastic problems substantially, a t the expense of Additional work on t h i s concept high aerodynamic performance a t cruise.

i s contemplated.

SCAT 4

The f i n a l developments by the contractors of t h e S C A T 4 configura -

Major problems were as follows: tions are sketched i n figure 12.

pitch - up, subsonic roll due t o sideslip, speed s t a b i l i t y i n approach, Figures 13 and 14 show a revised /and j e t interference with the t a i l .

version of S C A T 4 which incorporates an M wing discussed i n paper no. 12 Taylor e t al. Changes include (1) adoption of the M plan - by Robert T.

form having an increased aspect r a t i o , (2) i n s t a l l a t i o n of a ventral The addition of the f i n , and ( 3 ) a substantial reduction i n dihedral.

ventral f i n should provide adequate directional s t a b i l i t y a t supersonic speeds; and a l l three of the changes l i s t e d should improve subsonic roll - due -t o - sides l i p charact eri s t i c s.

Not included i n the changes listed i s the substitution of paired nacelles located lower, more inboard, and more forward than i n the Reduction of j e t interference on the t a i l should original arrangement.

r e s u l t .

The following discussion w i l l deal with the effects of t h e princi - p a l modification - t h a t of planform.

Tests a t low speeds of M and related planforms have already been With t h e exception of the leading - edge discussed i n paper no. E.

extension, t h e planform of the modified S C A T 4 i s nearly i d e n t i c a l t o

Additional data show a progressively improving one shown i n t h a t paper.

pitching - moment curve with increasing leading - edge extension, with the more extreme extension (representing a Kriiger flap, shown i n f i g . 13) exhibiting a gratifying pitch - down tendency. The same sort of improve - ment i n the pitching - moment curve i s seen i n such data as are available f o r M wings i n the transonic speed range.

I Figure 15 compares the wave - drag coefficients of M and arrow wings having the same thickness r a t i o , aspect r a t i o , and leading - edge sweep.

The values presented are theoretical because no comparative experimental M - wing data have been found beyond high transonic speeds since i n t e r e s t i n the M wing has diminished with the advent of low horizontal tails.

A s might be expected from study of the inset area diagram, the M wing shows higher wave drag than the arrow wing a t a Mach number of 1.0 ’but lower values from high transonic speeds upward.

The low Mach number of the crossover point u8Z &&se drag - coefficient curves i s character - i s t i c of t h e comparative experimental data available.

Thus, M wings o f f e r d e f i n i t e promise from t h e wave - drag standpoint.

Figure 16 i l l u s t r a t e s t h e interference f i e l d s of t h e nacelles, whether paired or i n separate pods. With regard t o wave drag, note t h a t the nacelle compressions place buoyancy on t h e receding slopes of the wings as w e l l as on the fuselage afterbody.

Favorable interference, as noted i n the discussion of the SCAT 4 and 15 concepts, should be realized.

With regard t o lift interference f o r t h e M planform with an optj--.m loading, l i n e a r theory c a l l s f o r an i n f i n i t e slope of the chords con.

taining t h e apexes and the trailing - edge notches.

T h i s slope i s a l s o called f o r i n t h e root chord of arrow wings and is, by one means or another, f a i r e d out. Body camber i n the case of arrow wings helps pro - duce t h e interference required by theory and the moderate amounts p r a c t i c a l l y permissible have, i n f a c t , proven advantageous.

For the M wing shown i n figure 16, t h e nacelles should d i r e c t l y provide t h i s l i f t interference so t h a t some hope i s held f o r approaching, more nearly, t h e values of the drag - due - to - lift f a c t o r s indicated by theory.

Tests a t supersonic speeds of a family of f l a t and warped M wings are pres - ently underway.

I n summary, it would appear t h a t the changes noted i n t h e SCAT 4

presentation could provide a configuration with high cruise performance and satisfactory s t a t i c s t a b i l i t y characteristics throughout t h e speed range. Design of wind - tunnel models of a complete configuration based on t h i s approach i s underway.

CONCLUDING FENARK The N A S A is investigating advanced versions of all four basic SCAT concepts (SCAT'S 17, 16, 15, and 4 ) .

Improvements appear possible i n nearly a l l problem areas.

282 .

SCAT 1 7 CONFIGURATIONS MAJOR PROBLEMS : 0 L I D AT LOW SPEED AND CRUISE 0 LOW - SPEED HANDLING QUALITIES Figure 1 MODIFIED SCAT 1 7 CONCEPT r - ; '3 CHANGES : 0 PLANFORM 0 HORIZONTAL CONTROLS 0 WING LEADING EDGE Figure 2 x Figure 3

L - 63 - 7734

SCAT 1 6 CONFIGURATIONS *- 7 BOEING ;'i LOCKHEED , ,,--I , MAJOR PROBLEMS: 0 JET INTERFERENCE WITH TAILS 0 PITCH - U P 0 C R U l S E L I D

Figure 4

* e. +?-?-m MODIFIED SCAT 16 C O N C E P T ,F-7 (A)h = 7.0 LE, M I N ,' ; ALE = 36' TO 7 5 ' ,I i CHANGES: e INBOARD PIVOT 0 NACELLE ARRANGEMENT e WING AREA INCREASED Figure 5 Figure 6 L-63-7735 LONGITUDINAL STABILITY CONSIDERATION§ OF SCAT 1 6 TYPE CONFIGURATIONS - LOW SPEED MAXIMUM A L E 6 C

- r

Aac., o / o C~~~ - 20 TWO - ENGINE NACELLE ARRANGEMENT @ SMALL INLET AREA i @ SHORT DIFFUSER * SMALL WETTED AREA @ GOOD VARIATION OF m/ma WITH M Figure 8 SCAT 1 5 CONFIGURATIONS MAJOR PROBLEMS: 0 HIGH STRUCTURAL WEIGHT OF WINGS e WING OVERLAP 0 AEROELASTICITY MODIFIED SCAT 1 5 CONCEPT ALE, M,N = 25" ; ALE, MAX 78O; A= 6.36 A CHANGES: 0 REDUCED STRUCTURAL SPANS INCREASED WING THICKNESS 0 REDUCED OVERLAP - FULL - DEPTH PIVOT Figure 10 Figure 11 L-63-7737 SCAT 4 CONFIGURATIONS MAJOR PROBLEMS : 0 PITCH - UP 0 ROLL DUE TO SIDESLIP 0 SPEED STABILITY IN APPROACH 0 JET INTERFERENCE WITH TAIL Figure 12 i I c.

nn CHANGES : 0 PLANFORM 0 VENTRAL FIN * DIHEDRAL Figure 13 Figure 14 L-63-7736 I Y WAVE DRAG M AND ARROW WINGS .U 0- ARROW - / r M %, W I I I 1 OIlO 1 1 2 1.4 I .6 1.8 2 . 0 M Figure D R A G AND LIFT INTERFERENCE

NACELLES -WING - BODY

r - I ,

-

. . , ; . .

" a .

. * _ i - - f Figure 16 15. PROPULSION ASPECTS O F THE SUPERSOKCC TMSPORT

By James F. Dugan, Jr.C

&SAe Lewis Research C e n t e a / -

-

Arvid L. Keith, Jr., and f i a n u e l Boxer cd-7Q- SUMMARY The effectiveness of various advances i n propulsion technology has been evaluated by using t h e ground rules of the NASA contracted SCAT f e a s i b i l i t y studies. Increases i n t h e operating turbine tempera - t u r e beyond present - day values w i l l permit substantial reductions i n airplane gross weight. Unless significant advances beyond those antic - ipated are m a d e i n turbine cooling technology and i n turbine materials, it is unlikely that additional gains w i l l come from s t i l l higher t u r - bine temperatures. Because efficiency of engine components i s already appears t h a t t h e needed improvements w i l l have t o come from high, it advances i n lightweight engine construction and i n variable - geometry technology.

Although t h e advanced turbojet engine and the advanced turbofan engine of t h e studies have shown substantial reductions i n airplane gross weight compared with those of present state - of - the - art engines, t h e use of these engines r e s u l t s i n a marginally acceptable supersonic It is hoped t h a t t h i s p a r t i c u l a r r e s u l t of the SCAT studies transport.

w i l l stimulate those engaged i n t h e propulsion e f f o r t t o intensify t h e i r research and develop a very much improved propulsion system.

w- J T Y d @

INTRODUCTION The importance of propulsion t o t h e success of a supersonic com - mercial transport can be appreciated when it i s recognized t h a t w e l l * over one - half of t h e take - off gross weight is comprised of t h e propul - sion system and t h e f u e l it w i l l consume. I n t h i s paper consideration will be given t o such problems as how good must the propulsion system be t o result i n an a t t r a c t i v e take - off gross weight, and t o what extent w i l l improvements i n t h e propulsion system decrease airplane gross weight. Basically, t h e means available t o t h e propulsion engineer f o r developing a good system are few; he can t r y t o improve component pe,r- formance, integrate t h e components i n t o a better package, and reduce) weight.

SYMBOLS AND ABBRIIVIATIONS ( L/D m a x i m l i f t - d r a g r a t i o M Mach number Tmax maximum augmentation temperature turbine i n l e t temperature TIT gross weight wg W reference gross weight g, change i n nozzle velocity coefficient ACV PR pressure r a t i o BPR bypass r a t i o SFC specific f u e l consumption T J turbojet DISCUSSION I n order t o evaluate the significance of improvements i n t h e pro - pulsion system, it is necessary t o consider both t h e airplane and t h e

mission. Two airplane configurations were selected - one had a fixed

wing and t h e other had a variable - sweep wing. Assumed aerodynamic characteristics of these configurations are shown i n figure 1 as varia- t i o n of (L/D),, with f l i g h t Mach number. A n y postulated advances i n propulsion w i l l be evaluated by determining t h e e f f e c t on gross w e i g h t of one or t h e other of these two airplanes.

The selected mission, shown i n t h e bottom half of t h e figure, i s characterized by Mach 3 cruise, a 3,200 nautical - mile range, and reserve requirements as s t a t e d i n t h e SCAT f e a s i b i l i t y studies work statement.

Some of t h e propulsion - system requirements and problem areas .are pointed out i n figure 2. When a specific engine i s selected and properly sized t o f l y t h e airplanes on the selected mission, engine thrust and airplane drag vary w i t h flight Mach number as shown. The difference bekween the thrust and drag curves i s the net force avail - able f o r acceleration and climb t o t h e i n i t i a l Mach 3 cruise a l t i t u d e .

I n sizing the engines, it i s necessary t o provide adequate accelera - t i o n margin i n the transonic range where a l i m i t on sonic - boom over - pressure s e t s a lower l i m i t on a l t i t u d e . The engines.must a l s o be large enough t o meet the r e s t r i c t i o n s on take - off distance, a i r p o r t and nearby community noise, and second - segment climb.

When Mach 3 cruise is attained and engine thrust equals airplane drag, it i s seen from figure 2 t h a t the required thrust i s very much ~ less than the available thrust. This is also t r u e f o r cruise t o an ' a l t e r n a t e a i r p o r t and holding p r i o r t o landing. These three steady - state conditions are very important, since t h e engine - airframe match at these conditions determines the amount of f u e l consumed during cruise and t h e amount of f u e l that i s carried onboard as reserves.

Thus, t h e r e i s a basic requirement f o r high engine efficiency under part - power operating conditions.

The general character of the engine - airframe match at these three c r i t i c a l steady - state conditions i s shown i n figure 3 f o r a variable - sweep - wing airplane powered by e i t h e r advanced turbojet or advanced turbofan engines.

For Mach 3 cruise, the required thrust i s such that near - minimum specific f u e l consumption is attained by both t h e turbofan and the turbojet engines. The turbojet m i n i m s p e c i f i c f u e l consumption i s achieved w i t h no augmentation, whereas the turbofan minimum specific f u e l consumption is achieved w i t h moderate augmentation.

I n comparing the turbojet engine w i t h the turbofan engine, it i s generally t r u e that the turbojet has lower specific f u e l consumption duping supersonic cruise, whereas t h e turbofan has lower specific f u e l consumption during subsonic cruise. The r e l a t i v e merit of the two engine ty - pes is very dependent on the mixture of supersonic and sub - sonic portions of the cruise.

The e f f e c t s of the major turbojet engine parameters on the gross / It can be seen i n w e i g h t of a fixed - wing airplane w i l l be examined.

figure 4 how airplane gross weight changes as the major parameters t h a t characterize turbo j e t engines a r e varied: namely, turbine i n l e t temperature, compressor pressure r a t i o , engine weight, and the use of augmentation. For the range of turbine temperature considered, design compressor pressure r a t i o of the dry turbojet should be about 10.

The reduction i n airplane gross weight that w i l l r e s u l t from being able t o obtain l i g h t e r dry turbojets is shown i n t h e right - hand p l o t i n figure 4. A r e l a t i v e gross w e i g h t of 1.0 corresponds t o a payload f r a c t i o n of about 6 percent of t h e gross weight. If it i s desired t o r a i s e t h e payload f r a c t i o n t o 8 percent of t h e gross weight, engines are needed t h a t w i l l result i n a reduction of airplane gross weight of about 20 percent. For t h e dry turbo j e t , engine weight reduc - t i o n s of t h e order of 40 percent and turbine i n l e t temperatures of the order of 2,800~ R a r e needed. Such an engine i s not e a s i l y attainable, e s p e c i a l l g i n view of t h e safety and r e l i a b i l i t y requirements and the desired goal of 2,000 hours time - between - overhaul. Since mid - 1962, the engine companies have been engaged i n a propulsion program i n support of the supersonic transport. It is hoped t h a t t h i s program will answer , two important questions: What l e v e l of turbine i n l e t temperature and / how much reduction i n engine weight can be achieved f o r t h e first U.S. supersonic transport?

I n the upper left - hand p l o t of figure 4 i s shown t h e reduction i n

airplane gross weight t h a t r e s u l t s from using augmented turbojets rather than dry turbojets. If a higher climb path is required t o meet t h e sonic - boom limitation, it w i l l be even more advantageous t o use augmentation. However, if a lower climb path proves t o be acceptable, it i s possible that a very lightweight, high - turbine - temperature dry turbojet w i l l meet t h e requirements.

Before other means can be considered by which improvements i n t h e propulsion system can lead t o an improved supersonic transport, perhaps it would be worthwhile t o consider some form of ideal engine.

By determining t h e reduction i n airplane gross weight that would result from t h e successful development of t h i s i d e a l engine, t h e designer will have a yardstick against which he can measure t h e effectiveness of various p r a c t i c a l means f o r improving t h e propulsion system.

To help define t h e i d e a l engine, o v e r a l l propulsion efficiency is shown as a function of Mach number i n figure 5.

The lower two curves are f o r turbofan engines B and D of t h e SCAT f e a s i b i l i t y studies as they operate during t h e acceleration and climb - to - cruise phase of flight; t h e engines are operating with f u l l augmentation. Engine e f f i c i e n c i e s read from t h e curve f o r the i d e a l engine are t h e highest i values attainable by engine D at each Mach number.

The optimum pressure - ratio curve represents t h e r e s u l t s of a parametric study at a constant turbine inlet temperature of 2,700' R i n which the overall and f a n pressure r a t i o s were optimized at each flight Mach number t o f u r t h e r increase engine efficiency; t h e optimum curve thus represents t h e envelope of many d i f f e r e n t engines. The open symbols are the supersonic cruise, subsonic cruise, and hold - at - landing efficiencies of engines B and D. The main conclusion t o be drawn from t h i s figure i s that there i s considerable room f o r improvement i n the propulsion efficiency during climb and acceleration.

The ideal engine chosen f o r consideration is one having variable mass flow through t h e use of variable geometry t o allow it t o operate during acceleHation and climb at the propulsion efficiencies, shown by with t h e same thrust schedule and i n t e r n a l com - the long - dash curve, ponent effici$ncies as those of engine D. To r e a l i z e the f u l l p o t e n t i a l of t h i s ideal engine, its i n l e t spillage and bypass drag are assumed t o be zero.

The r e l a t i v e gross w e i g h t of the variable - sweep - wing airplane is shown i n figure 6. The higher temperature turbofan, engine D, permits a gross - weight improvement of about 10 percent as compared w i t h ) engine B. An additional reduction i n gross weight of 18 percent i s achieved by the ideal engine; t h i s value is only 2 percent less than that f o r t h e optimum pressure - ratio engine. I n absolute terms, t h i s i s a reduction of about 70,000 pounds f o r a 400,000 - pound airplane.

The improvement offered by the advanced technology engine D as compared with t h e present - day technology engine B represents a reduction i n take - off gross w e i g h t of approximately one - third of that offered by the i d e a l turbofan. Similar r e s u l t s have been obtained f o r idealized and optimized turbojet engines. It should be emphasized that t h e tech - niques needed t o build t h e i d e a l engine today a r e unknown, but perhaps some day when considerable advances i n variable - geometry technology are realized the performance gains indicated may be approached.

It i s natural t o expect that the added complexity of variable geometry w i l l increase t h e propulsion package weight. For t h e condi - t i o n s j u s t discussed the i n s t a l l e d thrust - weight r a t i o of t h e i d e a l engine w a s assumed t o be equal t o that of engine D. The results of investigating the e f f e c t s of increased power - plant weight on the rela - t i v e take - off gross weight are shown i n t h e right - hand p l o t i n fig - ure 6. It is noted that sizable reductions i n gross weight are pos - s i b l e even with power - plant w e i g h t increases of as much as 50 percent.

The next problem t o be considered i s that of p r a c t i c a l means f o r improving payload. It may be recalled t h a t one feature of the ideal j turbofan w a s that it had zero i n l e t spillage and bypass drag. I n a 1 p r a c t i c a l engine, these drags e x i s t over most of the f l i g h t Mach num - ber range because it is not yet possible t o build an engine that i s capable of swallowing a full stream tube of air under a l l operating conditions. For t h e reference turbofan engine with an axisymmetric mixed compression inlet, the s p i l l a g e drag i s quite high during opera - t i o n i n t h e transonic speed range. One way of reducing t h i s spillage drag i s t o build a turbofan w i t h a different engine airflow schedule.

This can be done by replacing the constant exhaust nozzle area of the main gas stream w i t h a variable - area nozzle. With such a nozzle, t h e engine airflow can be reduced at high flight Mach numbers. The revised engine airflow schedule r e s u l t s i n l e s s air being s p i l l e d during tran -

sonic operation. There are other e f f e c t s - some good and some bad - so

t h a t it is necessary t o evaluate the concept, which has come t o be known as low flowing, by way of a mission study. I n figure 7, t h e advantages of low flowing t h e reference turbofan engine are considered.

The engine drag coefficients of t h e reference and low - flow turbofans are shown i n t h e left - hand p l o t of t h e figure.

Below a Mach number of 2.5, t h e engine airflow schedule of the low - flow turbofan is t h e same as t h a t f o r t h e standard - flow turbofan.

Above Mach 2.5, engine airflow of t h e low - flow turbofan decreases u n t i l at Mach 3 it i s 20 percent l e s s than t h a t f o r t h e standard - flow engine.

Since t h e i n l e t i s sized by t h e Mach 3 cruise conditions, t h e low - flow turbofan has a smaller and l i g h t e r i n l e t but a l s o has more nacelle wave drag. Over most of t h e flight Mach number range, however, t h e low - flow turbofan has a lower engine drag coefficient than the standard - flow turbofan. This condition i s due t o the f a c t that the reduction i n s p i l l a g e drag more than compensates f o r the increase i n wave drag. The net e f f e c t i s shown i n t h e bar graph on t h e right - hand side of t h e f i g - ure. By low flowing, the gross weight of t h e variable - sweep - wing air - plane i s reduced about 2 percent.

A s w a s pointed out i n figure 3, a shortcoming of t h e turbojet engine is i t s r e l a t i v e l y poor SFC during subsonic flight. A w q t o improve t h i s shortcoming is t o equip t h e turbine with variable turbine s t a t o r s . For subsonic f l i g h t at low l e v e l s of thrust, the f u e l control and turbine s t a t o r s can be adjusted simultaneously so as t o r e s u l t i n lower SFC. I n figure 8, the concept of using variable turbine s t a t o r s t o decrease t h e gross weight of t h e fixed - wing a i r c r a f t i s considered.

I n the left - hand p l o t fixed and variable turbine - stator performances at a Mach number of 0.27 and an a l t i t u d e of 1,500 f e e t a r e presented.

A t t h e low l e v e l of required thrust, t h e advantage i s that t h e SFC f o r t h e variable turbine stators is reduced 8.5 percent. I n t h e bar graph on t h e right - hand side of t h e figure, it can be seen t h a t t h e use of variable turbine s t a t o r s reduces t h e airplane gross weight by about 2 percent. This reduction results from t h e lower SFC attained i n meeting t h e two reserve requirements f o r subsonic f l i g h t : cruise t o j an a l t e r n a t e a i r p o r t and holding p r i o r t o landing.

I n calculating engine performance, it i s necessary t o use some schedule of inlet pressure recovery with flight Mach number. The schedule u t i l i z e d w a s t h a t given i n t h e work statement of t h e SCAT f e a s i b i l i t y studies and i s shown i n the upper right - hand p l o t i n f i g - ure 9. Since it is quite possible t h a t i n l e t s with higher performance can be b u i l t , t h e improved performance shown by t h e upper curve w a s postulated. If engine B is mated with t h i s higher performance inlet, t h e gross weight of t h e variable - sweep - wing airplane is reduced about 4 percent.

Comparable r e s u l t s are obtained if t h e nozzle velocity coeffi - With cient can be increased over the e n t i r e flight path by 1 percent.

t h i s b e t t e r nozzle performance, a reduction i n gross weight of about 5 percent w i l l result. It i s realized t h a t reaching out f o r such gains may r e s u l t i n complicated and weighty mechanisms. The importance of keeping component weight t o a minimum can be seen i n t h e lower right - hand p l o t of f i g u r e 9. If t h e gains i n i n l e t and nozzle performance require overall power - plant weight increases of 4- 1 and 5 - 1 percent, 2 2 respectively, no gross - weight advantage can be realized. It should be pointed out t h a t , if the i n l e t recovery schedule and nozzle performance assumed f o r t h e SCAT f e a s i b i l i t y study a r e not realized, proportionate penalties i n gross weight w i l l accrue.

CONCLUDING R E M A R K S The ground rules f o r the N A S A contractors' f e a s i b i l i t y studies have been used t o evaluate t h e effectiveness of various advances i n This evaluation propulsion technology f o r t h e supersonic transport.

indicated that increases i n operating turbine temperatures beyond t h e temperatures used i n present - day transports w i l l permit substantial However, without significant reductions i n airplane gross weight.

improvements i n turbine cooling technology and i n turbine materials, f u r t h e r gains from s t i l l higher turbine temperatures appear unlikely.

Because efficiency of engine components is already high, it appears t h a t t h e needed improvements w i l l have t o come from advances i n l i g h t - weight engine construct ion and i n variable - geometry technology.

Although t h e advanced turbojet engine and t h e advanced turbofan engine of the studies have shown substantial reductions i n airplane gross weight compared with those of present state - of - the - art engines, the use of these engines r e s u l t s i n a marginally acceptable supersonic ' transport. Additional research t o develop a much improved propulsion system i s therefore necessary.

ASSUMED AERODYNAMIC PERFORMANCE AND STANDARD MISSION PROFILE VARIABLE - SWEEP WING 1 4 FIXED WING , 0 I 2 3 MACH NUMBER RESERVES M=3,10°,6 80,000 TRIP TIME 60,000 250 N.MI.

ALTITUDE, BEST SUBSONIC M = 0.3 30 - MIN. HOLD 20,000 0 500 IO00 1 5 0 0 2000 2500 3000 3500 DISTANCE, N. MI.

Figure 1 PROPULSION RESTRICTIONS AND REQUIREMENTS r 1 . 0 ~ THRUST RELATIVE .6 FORCE U I 2 3 MACH NUMBER Figure 2 VARIATION OF SPECIFIC FUEL CONSUMPTION WITH THRUST TYPICAL SCAT MISSION STRATOSPHERE STRATOSPHERE 1,500 FEET M = 3.0 M=O.8 M = 0.3 1.4r I

- ADVANCED TURBOJET

1.2 ADVANCED TURBOFAN RELATIVE .8 SFC -----_----- 0 loo 200 0 1 0 0 200 0 100 200 300 THRUST, 7 ' REQUIRED THRUST Figure 3 EFFECT OF TURBOJET - ENGINE - CYCLE VARIABLES ON GROSS WEIGHT REL. ENG. WT . 7 L 2300 2600 2900 2300 2600 2900 TURBINE INLET TEMPERATURE, O R Figure 4 TURBOFAN ENGINE EFFICIENCY TURBINE INLET TEMP, ENGINE ACCEL. CRUISE B - 0 D O

2700'R IDEAL -- --

0 . 4 .8 1 . 2 1 . 6 2 . 0 2 . 4 2 . 8 3 . 2 FLIGHT MACH NUMBER Figure 5

TURBOFAN - ENGl N E -AIRPLANE OPT1 M IZATION

E N G I N E S - 1.1 - 1.0- - I DEAL wg .9- wg, REF - OPT1 M U M PRESS. RATIO - .8 - - .7 .4 .8 I .2 INCREASE IN POWER - PLANT WEIGHT Figure 6 , ADVANTAGE OF LOW FLOWING A TURBOFAN ENGINES

'"r

I - - . I 6 ENGINE DRAG

COEFF. .12 -

- .08 - .04 1 , - I I 1 LOWER SFC FROM VARIABLE TURBINE {STATORS AUG. TJ PR = I O TI, = 2 7 6 0 " R Tu&= 3 5 6 0 " R STATORS - r VARIABLE 1.2 REL.

SFC,

'g6t FIXED

GROSS LB/HR/LB WT.

- 1.0 IREQ. THRUST I I .88 . g 2 ~ .8 THRUST, ' Y o MIL. THRUST Figure 8 r EFFECT OF INLET AND NOZZLE IMPROVEMENT TURBOFAN ENGINE ; TURBINE INLET TEMPERATURE, 2,40 O R

P

TIMPROVED

a

J

-1 7

,941 I I I I I I I 0 .02 .04 .06 INCREASE IN POWER - PLANT WE1 G HT Figure 9 STUDY O F AN AXIsyMNFilTRIC SUPERSONIC-!I!JUNSPORT INLET c - - -

h 5 b a -

By Norman E. Sorensen NASAdmes Research Center \r SUMMARY

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I n support of the supersonic commercial air transport program, a 20 - inch capture diameter axisyrmnetric mixed - compression i n l e t model has been designed and t e s t e d at the Ames Research Center. The main conclu - sions t o be drawn from these preliminary t e s t s are t h a t the supersonic i n per - portion of the i n l e t performed as predicted, and the main l o s s formance l i e s i n the t h r o a t and subsonic diffuser. Since recoveries as high as 97 percent can be attained i n the throat, it seems certain t h a t with further development i n the throat and subsonic diffuser and with more e f f o r t i n determining b e t t e r bleed configurations, an axisymmetric i n l e t with performance quite adequate f o r t h e supersonic transport w i l l be developed.

9.

INTRODUCTION I n support of the supersonic commercial air transport (SCAT) program, a 20 - inch capture diameter axisymmetric mixed - compression i n l e t model has A photograph of the been designed and tested a t Ames Research Center.

- - model mounted i n the wind tunnel i s shown i n figure 1 . The i n l e t i s suitable f o r application t o a vehicle designed f o r a Mach number of 3.0 and i s capable of performing a t off - design Mach numbers by translating the centerbody. The length - diameter r a t i o of the inlet - engine - thrust reverser package i s consistent with those t e s t e d on the S C A T configura - tions. The model, which i s as large as practical, w a s tested i n the wind tunnel a t a Reynolds number per foot of about 2 x 10 6 corresponding t o the Reynolds number per foot at an a l t i t u d e of 65,000 feet and a Mach number of 3.0. The tests were conducted primarily t o determine the internal performance and the bleed requirements of the i n l e t from Mach numbers of 0.6 t o 3.0. C r i t i c a l t o the transonic performance of the propulsion system i s the spillage drag. Since t h i s drag does not yield well t o theoretical analysis and l i t t l e data a r e available, the s p i l l - age drag w a s experimentally determined a t Mach numbers of 0.6 t o 1.1.

A number of bleed configurations were tested a t a Mach number of 3.0 so t h a t a suitable configuration could be chosen f o r t e s t s a t off - design Mach numbers. Limited data were obtained a t flow angles up t o 8 O .

i SYMBOLS A , capture area free - stream tuhe area A0 local duct area pk additive drag coefficient

m a

h local height H local duct height bleed mass flow % ' D engine mass flow % G m, free - stream mass flow M, free - stream Mach number total pressure recovery at engine face PtENG total pressure recovery in throat pth free - stream total pressure P t C f J R inlet capture radius incremental distance along center line Ax 0 centerbody conical half - angle DESIGN The aerodynamic design of an inlet naturally is a compromise of a number of requirements. One of the most important is the requirement of sufficiently low transonic spillage drag so that the thrust margin of the vehicle is large enough to accelerate through this critical Mach number range. Spillage drag is composed of cowl drag and additive drag created by the spillage of excess airflow over the cowl lip at Mach numbers below the design. Both of these quantities can be kept low by keeping the cowl angle and the centerbody angle sufficiently low.

Fig - ure 2 shows such a design. An internal cowl angle of Oo was selected because a feasible internal design could be created with a 1 2 . 5 ' conical half - angle centerbody which produces reasonably low additive drag. The external cowl angle can then be as low as 2.5O. The internal contours were estimated for a free - stream Mach number of 3 . 0 beginning with this selected combination of angles and were then " tested " by a computer program employing the method of characteristics to describe the flow field. To determine a satisfactory set of contours, a goal must be set for the throat conditions and a limitation set on the pressure ratio across the shock - wave impingements to prevent boundary - layer separation.

Uniform flow in the throat at a Mach number of 1.3 and a throat - pressure \t recovery above 95 percent were the goals; the pressure ratio across the first shock - wave impingement on the centerbody was less than 3 and across the impingement on the cowl was about 2 . A s final requirements, the design had to provide at least 40 percent of the capture mass flow at a Mach number of 1 . 0 to match the off - design flow demand of a typical turbofan engine, and the centerbody translation requirements had to be on the order of one - half the inlet diameter to avoid large weight pen - alties associated with long translation distances. The design shown is about as short as is thought practical largely because of the limit set on the pressure ratios across the shock - wave impingements. Once the supersonic design is determined, a satisfactory throat and: subsonic dif - fuser must be designed to fit the supersonic portion of the inlet. A throat which had boundary - layer bleed for a length of two throat heights

was selected, and a loo equivalent conical angle subsonic diffuser was

selected as being reasonably short from the trade - off of weight with performance. Boundary - layer compensation was included in the throat and between the throat and the forward shock - wave impingements.

No high - performance inlet has as yet functioned properly without judiciously bleeding a portion of the boundary layer. It has not always been clearly understood where to bleed and how much to bleed. In this light the porous bleed configuration shown by the dotted areas was not necessarily the best but represented a fairly good compromise within present understanding.

SLTPERSONIC RESULTS Preliminary performance of the inlet at Mach numbers of 1.55 to The results shown and those to follow 3.0 is presented in figure 3.

were obtained without the effects of the wing or fuselage. The bleed configuration was the one shown in figure 2 . Eighty - six - percent pres - sure recovery was attained at the engine face with a rather high bleed The recovery which mass flow of 13 percent a t a Mach number of 3.0.

increased at t h e lower Mach numbers and t h e bleed which decreased were r e s u l t s which w e r e obtained with t h e same bleed exit s e t t i n g s as at a The recovery assumed i n t h e SCAT f e a s i b i l i t y Mach number of 3.0.

Even though t h e recovery indicated by the feasi - studies i s also shown.

b i l i t y schedule i s somewhat higher than t h e experimental recovery a t off - design Mach nunbers, t h i s condition does not necessarily mean that It i s usually t h e transport mission w i l l suffer an appreciable penalty.

more important t o maintain as high a recovery as possible at t h e design Mach number and t o s a c r i f i c e off - design recovery, if necessary, because Although the cruise time i s s o much longer than the acceleration t i m e .

the performance of t h e i n l e t w a s reasonably good, it i s f e l t t h a t b e t t e r performance can be attained at a Mach number of 3.0 especially with The supersonic portion of regard t o lowering the bleed requirements.

The shock - t h e i n l e t performed as predicted by the computer program.

wave impingements and the associated pressure rises correlate well with theory and the flow p r o f i l e s i n t h e throat indicate recovery as high as 97 percent at the end of the supersonic diffuser. A t y p i c a l history of pitot - pressure p r o f i l e s at a Mach number of 3.0 from the throat t o the

engine face i s shown i n figure 4 . The p i t o t pressures are plotted as a

function of the r a t i o of l o c a l height t o l o c a l duct height. Comparison of the t h r o a t recovery, which averages about 95 percent, with the

engine - face recovery, which averages 86 percent , indicates t h a t the

major losses occur i n the region of the terminal shock wave and i n the subsonic diffuser. The p r o f i l e at t h e d i f f u s e r rake j u s t downstream of t h e terminal shock t r a i n shows t h a t a r e l a t i v e l y large l o s s i n recovery e x i s t s on the centerbody side of t h e duct. Turbulent mixing i n the dif - fuser r e s u l t s i n some increase i n recovery i n t h e region adjacent t o the However, as i s centerbody by the time the flow reached the engine face.

usual, t h i s increase i s accompanied by a decrease i n t h e recovery l e v e l Although t h i s condition i s no doubt con - i n the balance of the duct.

t r o l l e d t o a considerable degree by the boundary - layer bleed, it i s believed that the rather rapid divergence of the throat aggravates t h i s flow distortion.

A s part of t h e design requirements previously mentioned, t h e i n l e t must match the mass - flow demands of a t y p i c a l turbofan engine. Figure 5 shows that the i n l e t does supply t h e m a s s flow when the i n l e t i s sized Some of the t o include the bleed mass flow at a Mach nuniber of 3.0.

i n l e t flow w i l l have t o be bypassed a t most of t h e off - design Mach num - bers as shown by the difference between t h e i n l e t flow and the engine The only point where t h e i n l e t may not (including bleed) requirements.

supply enough air i s near a Mach number of 1.6 where some reduction i n bleed rate should be allowable without seriously harming t h e performance.

The rapid deterioration of i n l e t performance with increasing flow Fig - angle i s a matter of c r i t i c a l concern with axisymmetric i n l e t s .

ure 6 shows how t h e present i n l e t performance deteriorates above 2 ' .

Some results for Mach numbers of 2 . 5 and 3.0 are shown. Although only a single result is shown for a Mach number of 3.0, the trends for both Mach numbers are expected to be the same. A 12 - percent loss of recovery and a 7 - percent loss of mass flow at 5 O are considered large penalties to pay for any sustained period of flight. It is, therefore, desirable that the inlet location on the vehicle be as insensitive to flow - angle changes as possible. Such a location could be in the flow field beneath the wing. Shown in the lower part of the figure is the variation with flow angle of centerbody translation distance required to keep the inlet flow " started. " Rather large translation of the centerbody is required to keep the inlet " started " ; hence, severe requirements are imposed on the inlet control system. In the event the control system was unable to produce the necessary translation, the internal shock - wave system would be expelled, large drag increases and severe flow oscillations thereby being produced. This condition means that not only should the inlets be located in flow fields relatively insensitive to vehicle attitude for performance reasons but also to provide some measure of protection from the effects of sudden external - flow disturbances.

From the foregoing discussion it seems reasonable to assume that modification of the throat and subsonic diffuser and careful attention to the bleed configuration should improve the supersonic results. How - ever, before an improved model configuration is proposed, the transonic results should be examined.

TRANSONIC RESULTS One of the main objectives of the transonic tests was to measure additive drag since it is so important to the vehicle acceleration as previously mentioned. The additive drag was determined by using momen - The additive drag is shown tum methods and is presented in figure 7 .

up to a Mach number of 1 . 1 for maximum pressure recovery, minimum addi - tive drag, and an optimum computed for maximum thrust minus drag. The optimized results come about fromthe trade - off between aaditive drag and pressure recovery; that is, low additive drag is associated with a loss in thrust because of the accompanying reduction in pressure recov - ery. All of the results were obtained with the inlet at 0 ' to the free stream. The cowl drag is not included in the optimization since it is relatively small. The additive drag appears to peak at a value of 0 . 1 2 and a Mach number of 1.2 if the minimum transonic additive drag results are faired into the supersonic theoretical computed curve. The peak additive drag is somewhat lower than predicted but is still about 10 percent of the engine thrust. This more favorable result appears to be caused by the reduced pressure induced by the curvature of the exposed portion of the centerbody which can give a net additive thrust as indicated below a Mach number of 1.0. Other transonic performance f a c t o r s presented i n figure 7 show t h a t dqpending on the position of the centerbody, t h e pressure recovery can be changed radically. The more withdrawn positions of the centerbody are more favorabTe from a recovery and d i s t o r t i o n standpoint but a r e very unfavorable t o the addi - t i v e drag. Total - pressure p r o f i l e s at the engine face indicate that If separation off t h e centerbody is the main cause f o r low recovery.

sepa$ation could be alleviated, a lower optimum drag would, of course, it appears t h a t about a 3 - percent be indicated. A t a Mach number of 1.0, improvement i n engine thrust can be attained if t h e recovery l e v e l can be raised t o the maximum shown i n t h e figure. It should be noted t h a t the p e r f o m n e e i s sensitive t o r a t h e r small centerbody t r a n s l a t i o n changes, the wide range of performance shown being covered by a t r a n s l a t i o n of about LO percent of t h e i n l e t capture diameter. I n addition, t h e i n l e t mass flow does not change appreciably over t h i s range because the i n l e t throat area does not change with t r a n s l a t i o n and the throat i s choked or nearly choked f o r the data presented.

MODIFIED MODEL CONFIGURATION The foregoing discussion suggests contour and bleed configuration changes that should result i n improved performance i n forthcoming tests.

Since t h e supersonic portion of the i n l e t appears t o perform as pre - dicted, it i s f e l t that the throat and subsonic diffuser performance can be improved with modified throat and subsonic diffuser contours Figure 8 shows the t e s t e d diffuser which a r e i n i t i a l l y l e s s divergent.

contours compared with the modified contours. Modifications a r e made only on t h e centerbody and r e s u l t i n a throat area r a t i o p r o f i l e which approaches a constant area throat and i s considerably l e s s divergent than t h e one tested. Since the modified subsonic diffuser i s t h e same it i s necessary t o curve the aft portion of length as the one tested, the centerbody rapidly and, as a r e s u l t , incur some loss i n pressure recovery because of t h e sudden expansion. It i s estimated, however, t h a t the l o s s i s only about 1 percent. Transonically, the new contours should a l l e v i a t e the flow separation off the centerbody and improve t h e performance i n t h i s area.

Some need f o r bleed configuration changes i s a l s o suggested. The forward bleeds are effective and it appears that a scoop on the center - body would be most effective because the low pressure i n t h i s area pre - vents porous bleed from removing more than about one - half percent of the i n l e t mass flow. O n the cowl side, t h e forward bleed area does not need t o be very wide t o be effective. I n the throat a concentrated bleed at the beginning of the throat w i l l require l e s s bleed than the broad area previously shown since the s t a t i c pressure ahead of the terminal shock i s much lower than t h a t behind. With t h e modifications suggested, t r a i n it i s estimated t h a t not u n r e a l i s t i c performance goals at a Mach number > > % of 3.0 should be 86 - t o 90 - percent pressure recovery with 6 - t o 10 - percent bleed mss flow.

CONCLUDING RESIARKS The main conclusions t o be drawn from these preliminary tests are t h a t the supersonic portion of t h e i n l e t performed as predicted and the main l o s s i n performance l i e s i n t h e throat and subsonic diffuser.

Since recoveries as high as 97 percent can be attained i n t h e throat, it seems certain that with f u r t h e r development i n the throat and sub - sonic diffuser and with more e f f o r t i n determining b e t t e r bleed config - ~ urations, an axisymmetric i n l e t with performance quite adequate f o r t h e supersonic transport w i l l be developed.

The axisymmetric type of i n l e t i s not the only type considered suitable f o r application t o the supersonic commercial air transport.

The two - dimensional type of i n l e t has demonstrated high performance capability and i s therefore a t t r a c t i v e f o r transport application.

Design considerations similar t o those discussed f o r t h e axisymmetric i n l e t have been applied t o a two - dimensional design, and a large scale model i s now being readied f o r t e s t s .

AXISYMMETRIC INLET MODEL A - 30906 Figure 1 , A X ISY MMET R I C I NL ET 0" INTERNAL SLOPE ENGINE THROAT FACE

\ ,,' &

'~''"2.5O EXTERNAL SLOPE Figure 2 .

SUPERSONIC PERFORMANCE 1.0 - FEASIBILITY STUDY SCHEDULE PRESSURE RECOVERY, 1 PtENG .9 - o \ p t O D I I I ' O D .8 1 . 2 - BLEED m~~~

0 - -

-.I 3 MASS FLOW, , ma i .I - 0 0 0 m ~ ~ l ~ 0 0 0

-

m a 0 ' t I I 1.5 2 .o 2.5 3 . O FREE - STREAM MACH NUMBER, M a Figure 3 PITOT - PRESSURE PROFILES

-

PI TOT - PRESSURE RECOVERY, qh/Pta Figure 4 INLET - ENGINE MATCHING 1 . 0 - P . EXPERIMENT\, /I .8 - ENGlNE+BLEED / I , ’ REQUIREMENT / MASS - FLOW -6 - ’/ RATIO, Q/* / A 0 A/

. 4 - / * , / ’ /%ENGINE

A C A- REQUIREMENT .2 - I I I 01 I 1 . 0 I .5 2.0 2.5 3.0 FREE - STREAM MACH NUMBER, M a Figure 5 EFFECT OF INLET FLOW ANGLE PRESSURE

INCREMENT, RECOVERY -: ~

APtE,,/ pt,

-. 2

INLET FREE - STREAM MASS - FLOW 0 MACH NUMBER INCREMENT, 0 3 .O AmENG/mQ, - .I 0 2.5

TRANSLATION e4 r 0

i REQUIRED INLET RAD1 US AX/R I I I 0 2 4 6 0 FLOW ANGLE, deg Figure 6 TRANSONIC PERFORMANCE PRESSURE MAX ptENG/pt, RECOVERY, MIN C D ~ P~ENG

-

0 OPTIMIZED (THRUST - ADDITIVE DRAG) pt, AD DI TI VE DRAG, CD a - . I L I I I I I I .6 .8 1.0 1.2 1.4 1.6 FREE - STREAM MACH NUMBER, M , Figure 7 MODIFIED MODEL

.5 -

PRESENT TEST CONTOUR \ - - - MODIFIED CONTOUR ~

DUCT AREA, 5 M , = 3.0

CAPTURE A, AREA

. 3 -

Figure 8 W R n r I E W O F N A S A E X H A U S T NOZZLE RESEARCH

By Jack F. Runckel d w

/ ---- NASA.Langley Research Center S l l M M A R Y Exhaust n o z z l e s proposed f o r t h e supersonic - transport propulsion system a r e examined i n t e r n s of performance, complexity, and special requirements imposed by the supersonic commerical air transport (SCAT) Both nozzle i n t e r n a l performance and thrust - minus - drag per - mission.

formance a r e considered. Results from N A S A investigations of t h e tran - sonic performance of several representative types of exhaust nozzles applicable t o the supersonic commercial a i r transport are presented.

ec.y;-6 - 4 4 T KdZ

INTRODUCTION The j e t exhaust nozzle of the supersonic transport must meet more severe operating requirements than nozzles used i n the exhaust systems of current supersonic a i r c r a f t . It i s the purpose of t h i s paper t o point out some of the problems of the SCAT j e t - e x i t nozzle and t h e special features t h a t must be incorporated i n the exhaust system. The subject i s very complex i f dealt with i n d e t a i l , and therefore only a simplified approach i s presented here. R e s u l t s from recent N A S A investigations, primarily obtained i n the transonic speed range, a r e presented t o indicate the l e v e l of performance expected from various types of nozzles applicable t o the supersonic transport.

SYMBOLS Ae nozzle e x i t area i nacelle cross - sectional area AN A t nozzle throat area Nozzle b o a t t a i l drag incremental drag coefficient o f nozzles, ACD C I S Nozzle drag D t o t a l - noz zle - drag rat io, Ideal primary thrust RzLm drag secondary - air ram - drag rat io, DR I d e a l primary t h r u s t Boattail drag b o a t t a i l drag r a t i o , DP Ideal primary t h r u s t nacelle diameter a.3 Measured t h r u s t gross t h r u s t r a t i o , F Ideal primary t h r u s t 2 nozzle external length M Mach number q dynamic pressure S wing area, 3,500 sq f t primary nozzle weight flow wP WS secondary - air weight f l o w corrected secondary - weight - flow r a t i o

(%) eorr

P b o a t t a i l angle

v nozzle efficiency

F - D thrust - minus - drag rat i o

F - DR nozzle i n t e r n a l t h r u s t efficiency

thrust - minus - boattail - drag rat i o

F - Dp

F - D R - Dp nozzle propulsive performance NOZZLE REQUIREMENTS The nozzle must have the ability to provide high performance at This includes not only the inter - all speeds up to a Mach number of 3 .

nal performance, or how efficiently the nozzle utilizes the energy in the exhaust gas to produce thrust, but also the effect of the drag on m e nozzle must be efficient not only the nozzle external surfaces.

for cruise at a Mach number of 3 but also during take - off, climb, and The importance of nozzle efficiency is .

at subsonic cruise and loiter.

shown in paper no. 15 by Dugan, Keith, and Boxer where it is pointed out that a 1 - percent change in nozzle gross thrust coefficient over the speed range resulted in a 5 - percent change in airplane gross weight.

For many of the engine cycles under study, the nozzle system must have the capability of varying the throat sizes of either the engine or duct, or both, during augmented engine operation and for some engines The expansion surfaces may during the cruise portion of the flight.

also have to be variable. The shape of the nozzle might have to change as shown in the top sketches of figure 1 .

The choice of unmixed or mixed flows from the engine and duct (fig. 1) must be dealt with in nozzles for turbofan engines. For the unmixed - flow engine, consideration has to be given to the matching of static pressures between the engine and duct exits to prevent the flow from one throat from suppressing the gas flow from the other throat.

The nozzles must include provision for thrust reversal during both The lower landing and in - flight operation to provide rapid descent.

half of the sketch in figure 1 of the mixed - flow turbofan indicates one possible type of thrust - reverser scheme utilizing cascade vanes to reverse the flow.

Cooling air must be provided during augmented engine operation to protect the nozzle surfaces from the very hot exhaust gases. Such sec - ondary air may a l s o be used to provide thrust increases in ejector noz - (See refs. 1 and 2 . ) The zles and to improve off - design performance.

use of inlet bypass air in an ejector nozzle as a trade - off for spillage (See ref. 3 . ) drag at the lower speeds may also be considered.

i The mechanical actuation systems for accomplishing the necessary geometric variations of the nozzle components should be minimized to reduce the complexity. The aerodynamically adjustable types of nozzles

meet this requirement. (See ref. 4.)

It would be desirable to design the nozzle to provide some noise attenuation during take - off and landing if this could be done without 31-7 decreasing the performance. However, t h e most successful noise - suppressor nozzle configurations have very complex geometric shapes o r Some indication of perform - involve substantial increases i n weight.

ance losses of t y p i c a l noise - suppressor configurations are presented i n reference 5.

The weight penalties f o r t h e aforementioned features should not be An increase of 1 pound i n nozzle w e i g h t can result i n a >-pound severe.

increase i n the gross take - off weight of the airplane.

Consideration should be given t o the location of the nozzle on the airplane i n regard t o the e f f e c t of the l o c a l flow f i e l d on the t.hrust and drag of the exhaust system (ref. 6) and t h e effect of t h e hot high -

The effect of a

pressure j e t on adjacent airplane surfaces (ref. 7).

such as wings, pylons, fuselage surfaces, tails, l o c a l interferences, and adjacent engines, on nozzle performance and t h e j e t interference e f f e c t s on these components need evaluation i n t h e f i n a l SCAT designs.

SCAT NOZZLE DESIGN STUDY I n t e r n a l Nozzle Geometry I n order that the design problems of supersonic - transport engine e x i t s be i l l u s t r a t e d , the basic nozzle requirements f o r a t y p i c a l SCAT The duct - burning turbofan engine w a s f l i g h t p r o f i l e a r e presented.

The v a r i a t i o n of pressure r a t i o and throat selected f o r t h i s study.

area with Mach number f o r both t h e primary - gas generator and the annular fan duct w a s used t o expand isentropically the exhaust gas f o r each f l o w t o ambient pressure at t h e f l i g h t altitvdes, and the required e x i t areas were determined. Figure 2 shows t h e nozzle i n t e r n a l expansion r a t i o s which are required t o produce m a x i m u m nozzle t h r u s t a t each Mach Ae/At number. These nozzle area r a t i o s are necessary t o obtain optimum gross (See pp. 143 t o 160 of t h r u s t regardless of t h e type of nozzle used.

ref. 8.)

The v a r i a t i o n of expansion r a t i o shown f o r t h e duct ( f i g . 2) is t h a t required f o r m a x i m u m augmentation w i t h f u e l burned i n the fan duct.

The expansion r a t i o f o r the duct stream with no augmentation would be somewhat less than t h a t shown, and the fan annular throat would be fur - The flows a r e unmixed a t the t h r o a t s of the engine and t h e r closed.

ducts, where, f o r example, t h e r a t i o s of gas - stream t o t a l pressure t o free - stream s t a t i c pressure at a Mach number of 3 were 24 and 45, respectively. Because of t h i s difference i n t h e pressure levels, con - sideration has t o be given t o matchingthe s t a t i c pressures between t h e engine and duct t h r o a t s t o prevent flow from one throat from suppressing - *. - the gas flow from t h e other %&oat. e., i-: Nozzle External Shape If an engine - nacelle diameter of approximately 82 inches i s assumed and the t o t a l nozzle - exit areas of both exhaust stream flows are related t o the nacelle cross - sectional area, the v a r i a t i o n shown i n figure 3 results. The curves of t h e r a t i o of e x i t area t o nacelle area indicate how t h e outside shape of the nozzle may have t o vary with Mach number f o r a range of engine - power conditions. The nozzle system must 4ave the capability of closing the outside shape on the nacelle e x i t t o about 0.2 i of t h e nacelle area and t o open the e x i t area t o 1.6 times the engine - nacelle s i z e if a mechanically variable nozzle i s used.

To determine the conical b o a t t a i l angles on the outside nozzle sur - faces corresponding t o the area variations of figure 3, a nozzle length must be selected. For optimum performance, the length of an isentropic nozzle would be about 2 nacelle diameters. A s a result, the nozzle would be about 14 feet long and would have low b o a t t a i l angles through - out the Mach number range ( m a x i m u m These low angles are p = 8 O ) .

desirable, but t h e nozzle i s too long and very heavy. If the nozzle i s made half as long - that is,

1 nacelle diameter i n length - the weight

i s reduced by about one - half at the expense of having higher b o a t t a i l angles and drag and somewhat lower efficiency.

The variation of boat - t a i l angle with Mach number f o r the short conical nozzle i s shown i n figure 4 .

The no - augmentation power - setting curve represents the upper l i m i t i n boattail - angle variation.

A t take - off, w i t h p a r t i a l augmenta - tion, the b o a t t a i l angle might be about 1 4 ' .

During the acceleration period, t h e variation i n b o a t t a i l angle would follow t h e maximum power l i n e up t o a Mach number of 3 where a 6 ' flare would occur. A t cruise some augmentation would be used, but the nozzle afterbody would be closed t o a cylinder t o minimize drag and t h e nozzle would then be somewhat underexpanded.

The v a r i a t i o n of incremental airplane drag coefficient with Mach number due t o drag of t h e long and short nozzles i s presented i n fig - ) ure 5. These drag coefficients are based on the nozzle surface drag of f o u r nacelles and on a wing area of 3,500 square f e e t . The data were obtained by using supersonic wave - drag theory (refs. 9 and 10) and interpolated experimental data (refs. 1 1 and 12) i n the subsonic and transonic ranges. A t a Mach number of 1.2 during acceleration (indicated by t h e c i r c l e i n f i g . 5 ) , t h e drag of the short nozzles may be as much as 1 1 percent of t h e t o t a l airplane drag.

This nozzle b o a t t a i l drag repre - sents 7 percent of t h e net thrust.

The need f o r minimizing external drag i s obvious, p a r t i c u l a r l y i n t h e transonic region, where the margin small. (See ref. 13.) It should be noted of airplane t h r u s t t o drag i s that drag forces e x i s t on a l l types of nozzles i n some form or another when a large variation of nozzle expansion r a t i o i s required throughout

the f l i g h t range. From t h e preceding discussion it can (See ref. 14.)

be seen t h a t the nozzle design may have t o be a compromise involving i n t e r n a l performance, external drag, and nozzle weight.

RECENT NOZZLF: INVESTIGATIONS BY NASA Convergent - Divergent Nozzle Some experimental data obtained with t h e convergent - divergent type of nozzle are shown i n figure 6. This configuration represents a variable - geometry ejector, designed for a high - pressure - ratio turbo j e t engine, w i t h the nozzle w a l l s s e t f o r transonic operation. The e j e c t o r corresponds t o the short - nozzle concept previously presented, except that it had contoured inner surfaces and a lower b o a t t a i l angle i n com - bination w i t h a s m a l l base. The nozzle u t i l i z e d secondary airflow f o r cooling and increasing the t h r u s t through entrainment i n the ejector.

(See r e f . 1.) The amount of corrected secondary airflow

( wsIwp)corr

w a s 7 percent of t h e flow passing through t h e primary nozzle. (See ref. 15 for d e f i n i t i o n of corrected secondary - weight - flow r a t i o . ) The performance o r efficiency l e v e l s as fractions of the ideal q t h r u s t of the primary nozzle are shown on the ordinate of figure 6 f o r the t e s t Mach numbers. The curve labeled F represents the ejector - t h r u s t r a t i o . This r a t i o i s the measured t h r u s t corresponding t o the t o t a l momentum of both flows a t t h e e x i t of the nozzle divided by the ideal t h r u s t of t h e primary nozzle. To account f o r the secondary - air momentum, f u l l r a m drag of t h i s a i r (divided by i d e a l primary t h r u s t ) These has been subtracted out f o r t h e dashed curve labeled F - DR.

data, then, represent t h e i n t e r n a l thrust efficiency of t h i s nozzle.

The lowest curve ( f i g . 6) i s t h e propulsive performance of the noz - z l e and includes, i n addition, the drag of the nozzle external surfaces

as a f r a c t i o n of i d e a l primary thrust. This parameter F - DR - DP

accounts f o r everything except i n s t a l l a t i o n losses o r benefits and forms t h e b a s i s f o r comparison of d i f f e r e n t types of nozzles.

The data points shown at a Mach number of 3 (note broken scale) are f o r t h e e j e c t o r operating undereqanded with cylindrical outer sur - faces and a r e based on data from reference 1. The results indicate t h a t a shortened, contoured, e j e c t o r nozzle should have good performance throughout the Mach number range.

Blow - In - Do& Ejec$;br $@zzle The previous discussion has been concerned with t h e variable convergent - divergent types of j e t - e x i t nozzles. Other types of nozzles which are not as complex mechanically and with which a reduction i n weight i s possible because of elimination of t h e divergent - shroud actu - a t i n g mechanism w i l l next be considered.

A type of nozzle that.may be considered t o be p a r t i a l l y aerodynam - i c a l l y variable is known as the blow - in - door e j e c t o r nozzle. (See ref. 4.) The e j e c t o r consists of a fixed shroud with hinged t r a i l i n g - edge f l a p s and hinged doors around t h e primary nozzle, as shown i n the sketch of figure 7. The hinged portions of t h e nozzle open and close according t o t h e pressure d i f f e r e n t i a l across t h e m . A t supersonic speeds t h e high engine pressures force t h e doors t o close and the f l a p s t o diverge, thereby approximating the convergent - divergent nozzle. A t transonic and subsonic speeds, the jet i s separated from the ejector w a l l s and induces air through t h e blow - in - doors t o a l l e v i a t e overexpan - sion, and the trailing - edge f l a p s close inwardly.

Results from an investigation of a blow - in - door e j e c t o r nozzle are shown a l s o i n figure 7. For these tests t h e doors were fixed open and the trailing - edge f l a p s closed, as shown i n the sketch. The net force of t h e blow - in - door air i s included i n the coefficients. Data a r e pre - sented f o r a schedule of jet - pressure r a t i o with Mach number where the

pressure r a t i o w a s 4 at a Mach number of 1.2. The ejector - thrust r a t i o

F i s t h e i n t e r n a l performance as previously presented. The thrust -

minus - drag parameter (F - D) included only the' shroud and b o a t t a i l drag

f o r the region indicated i n the sketch. Subsonically the thrust-minus- drag performance w a s actually higher than t h e i n t e r n a l performance because external forces i n t h e t h r u s t direction were obtained on the b o a t t a i l and shroud i n t h e indicated drag region. This occurred because of pressure recovery at the r e a r of the boattailed region. Some N A S A investigations of blow - in - door e j e c t o r s are reported i n references '15 and 16.

Plug Nozzles The u t i l i z a t i o n of t h e aerodynamically adjustable nozzles can pro - vide reductions i n nozzle weight. The best known of t h i s type i s t h e isentropic plug nozzle f o r which t h e outer boundary of the exhaust stream continually adjusts t o ambient pressure. There are, however, several problems associated with t h i s type of nozzle f o r the supersonic transport. The excellent s t a t i c performance of t h i s nozzle suffers from detrimental e f f e c t s of an external airstream. (See r e f s . 4, 1 4 , and 17.) Some of t h e design problems found i n applying t h e plug nozzle t o the SCAT mission are i l l u s t r a t e d i n figure 8. A mixed - flow exhaust i s assumed and the nozzle i s t o be designed f o r a cruise Mach number of 3. which The isentropic plug requires a steep t u r n at the throat, r e s u l t s i n a nozzle diameter greater than the basic nacelle diameter and i n a steep b o a t t a i l , both of which can have a sizable e f f e c t on drag (top l e f t sketch).

Several methods of reducing t h e drag are i l l u s t r a t e d by t h e other sketches. The nozzle shown i n t h e top right - hand sketch i s designed f o r a lower Mach number and has lower plug and b o a t t a i l slopes. For the pressure r a t i o s at a cruise Mach number of 3 , the jet continues t o expand beyond the nozzle - lip diameter (as indicated by the broken l i n e s ) , and some penalty i n performance occurs. The off - design per - formance a t lower speeds should be b e t t e r than the plug nozzle designed ) Y f o r a Mach number of 3. (See ref. 17.)

Another a l t e r n a t i v e i s t o provide some i n t e r n a l expansion as shown i n the lower sketches. The use of conical plug surfaces has a l s o been shown t o provide good performance w i t h a divergent shroud (lower l e f t - hand sketch of f i g . 8).

The nozzle shown i n the lower right - hand sketch takes advantage of t h e underexpanded nozzle through the use of terminal fairings. These bodies, distributed around the exit, can be designed t o take some t h r u s t on t h e i r inner surfaces, and the f a i r i n g s reduce the effective base area o r b o a t t a i l angle. With open spaces between t h e fairings, t h e over - expansion losses a t lower speeds may a l s o be alleviated.

(See refs. 18 and 1 9 . ) There i s an additional problem of varying the throat area f o r a l l these plug nozzles; and the proposed methods, such as flaps, col - lapsing and t r a n s l a t i n g plugs, and t r a n s l a t i n g shrouds, w i l l require evaluation.

Plug Nozzle With Terminal Fairings Some performance results obtained throughout the SCAT Mach number range on a plug nozzle w i t h and without terminal f a i r i n g s are presented i n figure 9. (See a l s o refs. 19 and 20. ) The basic plug nozzle w a s designed f o r a pressure r a t i o of 10 corresponding t o a Mach number i of 1.8. The thrust - minus - drag performance of the basic plug i s indi - cated by the solid line. The performance peaks at the design Mach num - ber and drops off a t the higher speeds (underexpanded nozzle) and at the lower speeds, around a Mach number of 1 (overexpanded nozzle); and, as t y p i c a l f o r plug nozzles, the performance improves again at subsonic speeds. (See ref. 21.) Addition of the terminal f a i r i n g s provided increased performance above the design Mach number and at transonic and subsonic speeds.

Concave - Base Plug Nozzle Several types of plug nozzles have been designed and tested with various degrees of truncation of t h e plug. A new con - (See ref. 4. ) cept of t h i s type of nozzle has been - under development at the Langley Research Center. This configuration i s known as the concave - base plug nozzle, which eliminates the plug extension. (See ref. 22.) A sketch i l l u s t r a t i n g the principle of operation of the concave - base plug nozzle i s shown i n the upper portion of figure LO. The nozzle can be consid - ered as having an aerodynamically variable plug since it has aerodynam - i c a l l y adjustable inner and outer jet boundaries and therefore elimi - nates a large portion of the plug weight. I n operation, t h e base - recirculation principle develops a ring vortex which pressurizes the base. The throat area i s varied w i t h hinges on the plug surfaces; these surfaces can a l s o be used t o block t h e exhaust flow f o r cascade t h r u s t reversal. The performance is highly dependent on l i p geometry and j e t pressure r a t i o .

Results are shown f o r the schedule of the jet pressure r a t i o pre - viously used f o r most of the other nozzles. Even though the perform - ance may be somewhat lower at transonic speeds than f o r the other noz - zles, t h e reduction i n weight makes it competitive. A t higher pressure r a t i o s and Mach numbers t h e performance i s expected t o improve.

SUMMARY OF NOZZLE PERFORMA.NCE state of the art of nozzle performance A n attempt t o summarize t h e f o r the supersonic transport is presented i n f i g u r e 11. Nozzle per - formance q i s t h e t h r u s t efficiency, including the secondary momentum drag of e j e c t o r s and t h e drag of t h e nozzle external surfaces. The predicted performance is indicated by a band which includes the various types of nozzles proposed f o r the supersonic transport, such as t h e variable convergent - divergent ejector, the blow - in - door ejector, the shrouded plug nozzle, and terminal - fairing configurations. The lowest nozzle efficiencies occur i n the transonic speed range where nozzle overexpansion conditions o r high drags on t h e nozzle external surfaces \ contribute t o t h e losses. The spread i n nozzle performance at a Mach ’ number of 3, i n addition t o nozzle configuration differences, is a result of changing engine power from full augmentation t o the cruise operating condition.

”.

CONCLUDING REMARKS The selection of an exhaust nozzle of the supersonic commercial air transport engine is still an area in which much further research is Several promising types needed to determine the optimum configuration.

of nozzles appear to be competitive when compared on a thrust - minus - drag performance basis. These nozzles are the short variable convergent - divergent ejector, the blow - in - door ejector, the conical plug nozzle, and the plug nozzle with terminal fairings. In addition, other config - urations offer interesting possibilities.

For the final nozzle evaluation, model investigations mf complete

.l

supersonic - transport configurations will have to be conducted with a simulated propulsion - system exhaust. Performance of the installed exhaust nozzle must be obtained with the simulation of the external The flow field that is expected to be encountered with the airplane.

model should include proper shaping of the aircraft and nacelles ahead of the nozzles, boundary layer and Reynolds numbers consistent with scaling practices, and local interferences, such as wings, pylons, fuselage surfaces, tails, and engine - cluster arrangements which can affect both the external and internal performance of the nozzle. The jet interference effects on drag, loads, dynamics, and temperatures of the airplane surfaces will also require evaluation.

1 . Stofan, Andrew J., and Mihaloew, James R . : Performance of a Variable Divergent - Shroud Ejector Nozzle - Designed for Flight Mach Numbers up to 3.0. NASA TM X-255, 1961.

2 . Beheim, Milton A . : Off - Design Performance of Divergent Ejectors.

NACA RM E58G10a, 1958.

3 . Hearth, Donald P . , and Connors, James F.: A Performance Analysis of NACA Methods for Handling Excess Inlet Flow at Supersonic Speeds.

TN 4270, 1958.

Thrust Nozzles for Supersonic 4 . Migdal, David, and Horgan, John J.: Transport Aircraft. Paper No. ~ ~ - A H G T - T ~ , ASME, Mar. 1963.

5. Schmeer, James W., Salters, Leland B . , Jr., and Cassetti, Marlowe D.: Transonic Performance Characteristics of Several Jet Noise Sup - pressors. NASA TN D-388, 1 9 6 0 .

6. Swihart, John M., Mercer, Charles E . , and Norton, Harry T . , Jr.: Effect of Afterbody - Ejector Configurations on the Performance at Transonic Speeds of a Pylon - Supported Nacelle Model Having a Hot - Jet Exhaust. NASA TN D-1399, 1962. (Supersedes NASA MEMO 1-4-591;, 1959.)

7 . Swihart, John M . , and Crabill, Norman L . : Steady Loads Due to Jet

Interference on Wings, Tails, and Fuselages at Transonic Speeds.

NACA RM L571)24b, 1957.

8. Talbot, J. E.: Aerodynamics of Power Plant Installations. Super - sonic Engineering, J. T . Henshaw, ed., John Wiley & Sons, Inc., c .1962, pp . 143- 160.

9. Fraenkel, L. E . : The Theoretical Wave Drag of Some Bodies of Revo - lution. R . & M. No. 2842, British A . R . C . , 1955.

1 0 . Jack, John R . : Theoretical Pressure Distributions and Wave Drags for Conical Boattails. NACA TN 2972, 1953.

1 1 . Silhan, Frank V., and Cubbage, James M . , Jr.: Drag of Conical and Circular - Arc Boattail Afterbodies at Mach Numbers From 0 . 6 to 1.3.

NACA RM ~56~22, 1957.

12. Cubbage, James M . , Jr.: Jet Effects on the Drag of Conical After - bodies for Mach Numbers of 0 . 6 to 1.28.

NACA RM L57B21, 1957.

13. Nichols, Mark R. : Supersonic Transport Propulsion Requirements.

Aerospace Eng., vol. 20, no. 9, Sept. 1961, pp. 12-15, 58 - 63.

1 4 . Keenan, J. G. : Some Problems Associated With High Mach Number A i r Breathing Powerplants. Interayia, vol. 18, no. 3, Mar. 1963, PP- 345-347.

15. Schmeer, James W., Mercer, Charles E., and Kirkham, Frank S.: Effect of Bypass A i r on Performance of a Blow-inADoor Ejector Nozzle at Transonic Speeds. NASA T M x - 896, 1963.

16. K i r k h a m , Frank S., and Schmeer, James W.: Performance Character - i s t i c s a t Mach Numbers up t o 1.29 of a Blow - in - Door Ejector'Nozzle \ With Doors Fixed i n N l - O p e n Position. NASA T M x - 830, 1963.

J 17. Valerino, Alfred S . , Zappa, Robert F., and Abdalla, Kaleel L.: Effects of External Stream on the Performance of Isentropic Plug- Type Nozzles at Mach Numbers of 2.0, 1.8, and 1.5.

NASA MENO 2-17-59E, 1959.

18. W i l l i s , Conrad M., and Mercer, Charles E.: Effect of Afterbody Ter - minal Fairings on the Performance of a Pylon - Mounted Turbojet- Nacelle Model.

NASA T M X-215, 1960.

19. W i l l i s , Conrad M., and Norton, Harry T., Jr.: Effect of Afterbody Terminal Fairings on the Performance of Plug - Type Exhaust Nozzles a t Transonic Speeds. NASA T M x - 762, 1963.

20. Norton, Harry T., Jr., and Swihart, John M.: Effect of Terminal Fairings on the Performance of a Plug - Type Exhaust Nozzle i n Qui - escent A i r and a t Mach Nmbers of 1.62, 1.93, 2.55, and 3.05.

N A S A TM X-463, 1961.

21. Salmi, R. J., and Cortright, E. M., Jr.: Effects of External Stream Flow and Afterbody Variations on the Performance of a Plug Nozzle at High Subsonic Speeds.

NACA RM ~ 5 6 ~ 1 1 a , 1956.

22. Mercer, Charles E., and Salters, Leland B., Jr.: Performance of a Plug Nozzle Having a Concave Central Base With and Without Termi - d n a l Fairings at Transonic Speeds. N A S A TN D-1804, 1963.

SCAT EXHAUST - NOZZLE PROBLEMS TAKE - SUBS. ACCELER - SUPERS.

OFF CRUISE ATION CRUISE

-- -77- v

-- a -

UNMIXED MIXED \

* = - ! DUCT \e - < -

REVERSE C ?

Figure 1 NOZZLE EXPANSION RATIOS FOR DUCT - BURN I NG TURBOFAN / / 4- A e P t 2 - I I I I I I I 0 I 2 3 M Figure 2 1 8 RATIO OF EXIT AREA TO NACELLE AREA FOR DUCT - BURN1NG TURBOFAN

r

+=!!I

MAXIMUM AUGMENTATION '

I .2 i-

I , NO AUGMENTATION Figure 3 BOATTAIL ANGLES FOR CONICAL CONVERGENT - DIVERGENT NOZZLES TAKE - OFF I

- p , P

NO AUGMENTATION (ACCELERATION) ' .

- - 4 - 8 I I I I I I 0 I 2 3 M

Figure 4

INCREMENTAL DRAG COEFFICIENTS OF FOUR NACELLE BOATTAI LS WING AREA=3,500 FT2 LONG NOZZLE, Z = 2 d N SHORT NOZZLE, Z = d N .OI 6

-

NO AUGMENTATION

- - -- - MAXIMUM AUGMENTATION

M Figure 5 PERFORMANCE OF SHORTENED CONVERGENT - DIVERGENT NOZZLE

(") = 0.07

wP CORR F- DR - --- - - - - - - - - -___----- - ---, ps

--

----

F - D - D R @ L l I I I I A I 0 .9 I .o 1. I 1.2 1 . 3 " 3.0 M Figure 6

BLOW - IN - DOOR NOZZLE

DRAG REGION 1

F - D 7 I . O r .9 I I I I I I I Ll .8 I ,o I .2 I . 4 0 .6 M Figure 7 PLUG NOZZLES

-to----

M = 3

t

L dN 0

Figure 8 P L U G NOZZLE WITH AND WITHOUT TERMINAL FAIRINGS M = 1 . 8 DESIGN PLUG BASIC PLUG NOZZLE F - D

. 7 "t

T I I I I I I 0 I 2 3 M Figure 9 CONCAVE PLUG NOZZLE T * 8 1 , , , I , I , . 7 L 0 . 7 .9 1 . 1 I .3 M Figure 10 - > -, PREDICTED PERFORMANCE FOR SCAT EXHAUST NOZZLES I . O r I # 2 2 . M Figure 11 '(I&) STRUCTURAL CONCEPTS AND MATERIALS S W T I O N

m TEE FEASIBILITY STUDIES

- . - I

By Richard A. Pride d4

NASAeLangley Research Center 2- SUMMARY //?LO Feasibility studies of the four SCAT configurations by two con - ) tractors have shown that 8~1-1340-i~ duplex - annealed titanium alloy is the most efficient material for use in the primary structure. Skin- stringer type of construction was utilized in both wing and fuselage designs. Calculated thermal stresses were shown to result in an over - all increase in structural weight of less than 2 percent, even though different methods were used to combine thermal stress with load stress for ultimate load conditions. Structural weights were 25 to 29 percent of the gross weight on all except a variable - sweep dual - wing configura - tion, which had a structural weight that was 37 percent of the gross weight.

kQ LJ-i-kf-Od INTRODUCTION \ Paper no. 3 by R . Richard Heppe and Jim Hong and paper no. 4 by Lloyd T . Goodmanson, William T . Hamilton, and Maynard L . Pennell present the SCAT feasibility studies of the Lockheed - California Co. and The Boeing Co., respectively. These summary papers touch on the structural designs and materials used in the four Mach 3 configurations as well as in the two aluminum configurations that were added to the study.

The present paper discusses in greater detail the material selection, some of the structural concepts and design allowables, the thermal stress problem, and the structural weight.

The discussion primarily covers the Mach 3 configurations with some comparisons with the aluminum designs.

SYMBOLS Young's modulus of elasticity, psi acceleration due to gravTty, ft/sec2 T temperature, 9 maximum temperature difference between specified locations in Nw structure, OE structural weight of airplane, lb WS take - off g r o s s weight of airplane, lb ‘ T O E strain

d stress , ksi

Subscripts : Y yield Tu tensile ultimate LL limit load UL ultimate load TH thermal m,L thermal at limit load TH,u thermal at ultimate load MATERIAL SICWETION The Mach 3 cruise speed requirement in the initial studies resulted in structural temperatures high enough to remove aluminum alloys from consideration in the material selection process. Both stainless steels and titanium alloys were initially considered; however, the final choice

of the 8 percent aluminum - 1 percent molybdenum - 1 percent vanadium

(8~1-IMO-1v) titanium alloy was made by both contractors.

i Figures 1 and 2 show two of the many material - selection parameters used: tensile strength and stiffness. In figure 1 the tensile ultimate stress divided by the density is shown as a function of temperature for several materials. The top curve is the 8~1-IMO-1v titanium alloy, the next curve is one of the better stainless steels (PH 14 - 8 Mo, SRH 1050 condition) and the bottom curve, shown as a band, includes several of the aluminum alloys which might be used at slower speeds and lower temperatures. Included are x2020-~6, 2024-T81 and 2618461 aluminum alloys. The aluminum data are based on 10,000 hours of exposure to test temperature, the longest exposure data av'ailable, because aluminum - alloy strengths are reduced by exposure time. No such effect has been observed for the titanium or stainless steel, and the results shown in figure 1 are for short - time exposure.

Two things should be noted. The higher strength of the titanium indicates that a tension member built from titanium would be about 1 0 percent lighter than one built from stainless steel over the entire temperature range shown. Also, the titanium tension member would be about 20 percent lighter than an aluminum tension member in its useful temperature range.

A similar comparison for stiffness is shown in figure 2 based on the modulus - density ratio. In this figure, there is not such a clear cut advantage in any one material since they all are grouped rather closely. Thus, for parts of the structure which must be designed for stiffness rather than strength, any of these materials would result in about the same weight. For elastic compression buckling of fuselage or wings, this modulus - density parameter becomes either the square root of the modulus or: the cube root of the modulus divided by the density, and the order of weight saving would be aluminum fjrst, titanium next, and steel last.

Additional material - selection factors were considered in the design studies. Such factors as fatigue life, compression yield strength, and weld strength modify the degree of weight saving experienced with titan - ium, but, in general, short - time - exposure material - selection parameters indicate that the 8~1-NO-1. titanium alloy is the best material for the SCAT configurations. Paper no. 19 by Herbert F. Hardrath and George J.

Heimerl discusses NASA research programs concerning the effect of long - time exposure to temperature.

When a new material is recommended for commercial use, one of the questius raised is, " How reproducible is it? " Figure 3 shows the vari - This ation in tensile ultimate stress for the Ti-8Al-lMo-1' alloy.

alloy is always used in the annealed condition. However, at least three different annealing processes are available and some of the data reported in various papers in this compilation were obtained from mate - rial in each of the three anneals. The simplest form is the single - ' anneal or mill - anneal process. In order to improve the crack propaga - tion characteristics of the mill - annealed material, the triplex - anneal process was developed which involves three exposures in the heat - treating furnace. Subsequently, a duplex - anneal process was developed which gives essentially the same material characteristics as the triplex - anneal process but involves only'two exposures in the furnace.

The duplex - annealed material is that which is recommended today.

i In the duplex - annealed condition, results of a compilation of tests at The Boeing C o . , Lockheed - California Co., Langley Research Center, a d TitaniumMetals Corp. of America show about 15 percent scatter for eight different heats and about 6 percent scatter for different sheets within one heat. Strength of 1 - inch - thick plate indicates the type of strength reduction associated with thicker material. Similar scatter can be seen for the mill - annealed and triplex - annealed materials. Note that in fig - ure 3 the stress scale is expanded, by not starting at zero, to emphasize

the effects. Comparative data for aluminum alloys show about 5 percent

scatter between different sheets of one run and about 12 percent scatter Establishing sound minimum between typical and m i n i m properties.

design properties is one of the problems which needs continuing effort for the 8 u - 1 ~ 0 - 1 ~ duplex - annealed titanium alloy.

Fabrication of structural test specimens from 8 ~ ~ ~ 4 0 - 1 v duplex - annealed titanium alloy under this SCAT contract and other related super - sonic transport contracts indicates that there is no difficulty in Additional experience is needed to establish working with the material.

optimum shop practices and to build a general level of confidence. Rea - sonable rates of machining integral skin - stringer panels from titanium plate have been accomplished. Stringers have also been brake - formed and joined to cover skins by resistance and fusion welding without serious degradation of the welded - area material strength.

STRUCTURAL CONCEPTS A structural design approach has been followed for all SCAT con - figurations based on previous experience with subsonic commercial trans - port, modified locally where necessary to suit peculiar requirements such as wing pivots.

The basic structural concepts are shown in figure 4 . Use of tita -

nium favors skin - stringer construction in both wing and fuselage sec - tions. If stainless steel were used, much more complex construction would be required to provide the local stiffening necessary to develop the high strength in compression and torsion. The structural weight of the design usingsteel would be at least 1 0 percent greater than that using titanium.

Fuel tank designs used include both integral and isolated internal tanks in both wing and fuselage sections. Fuselages are designed with sufficient area in the rings and crack - stopper straps where needed to prevent catastrophic crack growth. Minimum fuselage skin thickness is set at 0.028 to 0.032 inch for cabin pressure differentials of 10.5 to 11.9 psi. Both pressures meet specified requirements (ref. 1 ) but reflect slightly different attitudes on passenger comfort.

The structural designer for SCAT is faced with at least three problems which are of larger magnitude or are different from those encountered in^ previously designed transpor'lP ai *.These problems include long - lifetime exposure to loads and ele mperatures (discussed in paper no. IS), aeroelastic problems associated with the thin wings and long fuselages (discussed in paper no. 20 by Robert W .

Boswinkle, Jr, , and E . Carson Yates, J r . ) , and thermal stress problems

associated with the temperature gradients in the wing and fuselage panels (discussed in the following section).

TBERMAL STRESS Structural temperatures on the lower surface of the wing in the region of an integral fuel tank are shown in figure 5. The lower sur - face is a double - skin design with the inner skin forming the bottom of the fuel tank. The upper surface is single - skin Z - stringer construction with a radiation barrier attached to minimize heat transfer to the fuel.

The temperature of the outer skin on the lower surface climbs very rapidly as the airplane attains supersonic speeds and quickly reaches equilibrium at about 450' F. The temperature of the inner skin rises very slowly since it is in contact with the fuel throughout the flight.

The difference in temperature between these two skins results in thermal stresses because the skins are joined together through the Z - stringers.

Temperatures shown are typical for this type of fuel tank. Large vari - ations would be possible depending on the structural details.

A typical comparison of flight load stress and thermal stress for the lower surface of a wing in the region of an integral fuel tank is The short - dashed line represents the l g load stress shown in figure 6 .

The solid line represents the combined load and history for both skins.

thermal stress history for the inner skin, and the long - dashed line represents the combined load and thermal stress history for the outer skin. Designing the inner skin thicker than the outer skin minimizes the additive effect of thermal stress in the inner skin, and results in compressive thermal stresses in the outer skin which are large enough to cancel out the tensile load stress near the end of the supersonic flight.

z A similar situation exists in the fuselage temperature distribu -

tion, as shown in figure 7 . The cutaway section is a portion of fuse -

lage skin, supported by a longitudinal stringer and a circumferential ring. Calculated temperature histories are shown for four points. The solid lines designated A and B represent the temperature history of the skin and rings. The maximum difference in temperature between the skin and rings for this case is 340° I ? . This difference produces compres - sion thermal stress circumferentially in the fuselage skin, which is i beneficial when combined with the pressurization stress. However, this temperature difference also produces tensile thermal stress in the rings, which may be detrimental to their role as crack stoppers. The dashed lines designated C and D represent the temperature history of the skin and inside flange of the stringer. The maximum temperature difference of 240° F produces longitudinal compression in the skin which may prpduce local buckling when combined with the fuselage bending - moment stresses.

Thermal stresses in the supersonic transport are of a magnitude that they can be combined with lg load stresses by elastic superposition. But there is no uniformity of opinion as to how to combine thermal stress and load stress at limit or ultimate load. Two methods used by the contrac - tors to assess the effect of thermal stress on strbptural weight are dis - 'I The first method is shown in figure 8; the details are worked OL 2 cussed.

for a tension member, but similar relations would apply in compression.

A typical stress - strain curve for the material is shown in figure 8 with the yield stress and tensile ultimate stress designated by circles.

The standard definition - that ultimate load must be 1.5 times the limit

load - is maintained. Whether or not thermal stresses are present, the

limit load stress must not exceed 67 percent of the ultilriate load stress.

Also, at limit load, the combination of thermal stress and limit load stress must not exceed the yield stress; that is, the combination shall produce no permanent set which would require repair, adversely affect the mechanical operation of components, or adversely affect the aerodynamic characteristics of the vehicle. The thermal stress is determined by com - bining a thermal strain with the limit load strain where the thermal strain is calculated from the temperature distribution in the structure.

At ultimate load, the thermal stress plus 1 . 5 times the limit load stress is equal to the tensile ultimate stress of the material. Again, the thermal stress is based on combining the thermal strain with the ultimate load strain, which results in an alleviated thermal stress due to plas - ticity. The increment of thermal strain at limit load produces a large thermal stress; the same increment of thermal strain at ultimate load produces a much smaller thermal stress.

The thermal strain at ultimate load is based on the same temperature distribution as was present at limit load. The method may be deficient, therefore, because no " safety factor " is applied to the temperature distribution. i A more conservative method is shown in figure 9. The same stress -

strain curve is shown as in figure 8 and the same definition - that ther -

mal stress plus limit load stress must not exceed the yield stress - is

applicable. However, the criterion at limit load is that limit load

stress plus thermal stress must not exceed 67 percent of the tensile

ultimate stress. Thermal stress is calculated elastically from the tem - perature distribution present at limit load conditions. At Lltimate load, both the l i m i t load s t r e s s and the thermal stress a r e multiplied by t h e f a c t o r 1.5 before summing and equating t o t h e t e n s i l e ultimate s t r e s s .

The F A A t e n t a t i v e standard f o r supersonic transport design (ref. 2) adopts a method similar t o t h i s second method (fig. 9) except t h a t the thermal s t r e s s i s multiplied by a f a c t o r of 1.25 instead of 1.5 before summing f o r the ultimate load.

A comparison between figures 8 and 9 shows a large difference i n load stress available t o carry the ultimate load on the portion of t h e structure considered. However, since thermal s t r e s s e s are self - equilibrating within the structure, there w i l l be other m e m b e r s i n which t h e thermal stress w i l l subtract from t h e load stress rather than add, )with the r e s u l t a n t e f f e c t that more load stress can be carried i n those members. Thus, a complete s t r u c t u r a l analysis i s required t o determine the impact of d i f f e r e n t thermal stress design c r i t e r i a on the s t r u c t u r a l weight.

H o w much of a weight penalty associated with thermal s t r e s s e s , then, did the contractors' studies show?

Figure 10 shows the stiffened panel weights f o r one p a r t i c u l a r case of an 8~1-1-140-iv titanium dry - bay wing lower surface. Wing panel weight is plotted as a function of panel end load. The curve f o r the room - temperature panel weight i s shown f o r reference.

Ascent thermal conditions appear t o produce a 40 - percent increase i n panel weight.

However, t h i s i s not the proper way t o use these curves. A t room temperature, any p a r t i c u l a r location i n the wing w i l l have a panel weight determined by the end load corresponding t o the gross weight of the airplane a& take - off and t o subsonic aerodynamic loading. By the t i m e the peak thermal condition i s obtained, near the start of supersonic cruise, a considerable amount of f u e l w i l l have been consumed and the w e i g h t of the airplane reduced. Supersonic aerodynamic loading w i l l apply and t h e panel end load w i l l be reduced so t h a t t h e panel weight w i l l increase only s l i g h t l y , if a t all.

I n addition, many p a r t s of t h e structure are not designed by thermal stress conditions; consequently, the studies showed less than a 2 - percent increase i n over - a l l s t r u c t u r a l weight a t t r i b u t e d t o thermal stress.

STIiLTcTuRAL WEIGHT A comparison of s t r u c t u r a l weight breakdowns i n t e r m s of f r a c t i o n of gross weight at take - off f o r the various SCAT configurations is shown i n figure l l . Grouped together a r e the four titanium airplanes, the two aluminum airplanes, and t h e tvo subsonic transports. The numbers a t the top of t h e bars are the SCAT numbers.

Although the contractors did not always agree on gross weight poundage, their structural fractions are very close and average values are therefore used in this presentation.

It must be recognized that SCAT 4 and SCAT 15 did not receive as many refinement cycles in the feasibility study and therefore their weights are probably a little less well - defined. Also, since the aluminum SCAT airplanes were only brought into the study recently, they do not reflect as much refinement.

An interesting item is the weight of the SCAT 15 wing, about two This heaviness is due to the double - times heavier than any other wing.

wing feature for variable sweep and is the main reason for the gross weight being unacceptable (having a weight fraction of 0.37). However, the weight fractions of the other components in the SCAT 15 structure are not very different from the same components in the other SCAT air -

i

planes. The titanium SCAT airplanes, with the exception of SCAT 15,

have structural weight fractions ranging from 0.2'3 to 0.29 - values

which are not too different from those for the subsonic Boeing 707-320B This small o r the Electra airplane with weight fractions of about 0.24.

difference, however, amounts to about 50 percent of the payload weight for the SCAT airplanes.

Structural weight fractions for the aluminum airplanes also are in

the same range - about 0.28 - probably by coincidence since both alumi -

num SCAT airplanes had less efficient strength - weight materials and less efficient propulsion systems. Actually, one version of the alumi - num SCAT did not make the range requirement and the other did not make the sonic - boom overpressure requirement.

CONCLUDING ENARKS Both contractors have used skin - stringer type of construction in their Structural concepts for the feasibility studies of the four SCAT configurations. Both selected 8~1-1~0-1v duplex - annealed titanium alloy as the primary structural material for use in the Mach 3 configu - rations. Differences in design techniques for thermal stress may require closer specification by a contracting agency, but these differ - ences do not have serious consequences on ultimate loads f o r the super - i sonic transport.

Thermal stresses were shown to inflict a structural weight penalty of less than 2 percent on wing and fuselage panels.

Structural weights were 25 to 29 percent of the take - off gross weight for all but SCAT 13 which had a structural weight that was 37 percent of the g r o s s weight. This weight fraction of 0.37 was generally unacceptable. Structural weight fractions of aluminum SCAT airplanes were about 0.28 but the aluminum designs did not meet the range o r sonic - boom requirements for SCAT.

The titanium airframe appears to be I .

the most efficient WGX to c a m & +hg design payload supersonically over the design range witkinrthd sonic - boom overpressure restrictions.

REFERENCES 1 . Anon.: Airplane Airworthiness; Transport Categories. Civil Air Regulations Part 4b, Rules Service Co. (Washington, D . C . ) , Jan. 7 , 1963.

Tentative Airworthiness Objectives and Standards for Super - 2 . Anon.: Flight Standards Service, FAA, sonic Transport Design Proposals.

Aug. 15, 1963.

MATERIAL SELECTION T E N S I L E STRENGTH

rx lo

- .8 - P H 1 4 - 8 MO STEEL

STRESS ;6

DENSITY I N * .4- - .2

1 I + I I I 1

0 100 200 300 400 500 600 TEMPERATURE, O F Figure 1 MATERIAL SELECTION STIFFNESS PH 1 4 - 8 MO STEEL 80 SRH 1050 E t ALUMINUM DENSITY IN,

40 t

I I I I 0 200 400 600 TEMPERATURE, O F F i g u r e 2 VARIATION OF TENSILE STRENGTH T i - 8 A t - 1 M o - 1 V (ANNEALED) 1 8 ' ~ SINGLE DUPLEX TRIPLEX w 8 HEATS STRESS, KSI I40

120 L

Figure 3 B A S I C S T R U C T U R A L CONCEPTS TWO - SPAR WING - - - - ~ L I ~ ~ X I - % - L I T -L SKIN STRINGER

\ FUSELAGE

Figure 4 STRUCTURAL TEMPERATURE WING LOWER SURFACE; INTEGRAL FUEL TANK 500- RADIATION BARRIE ----------> T, O F 0 20 40 60 80 1 0 0 120 140 TIME, MIN Figure 5 FLIGHT STRESS PROFILE WING LOWER SURFACE 40 r r I N N E R SKIN - 201 I I I 1 I I I 0 20 40 60 80 100 120 140 TIME, MIN Figure 6 STRUCTURAL TEMPERATURE FUSELAGE CABIN AREA 300 CALCULATED - TEMPERATURE O F LOCATI ONS I O 0 0 20 40 60 80 100 120 140 TIME, MIN Figure 7 COMBINED THERMAL AND LOAD STRESS TENSION MEMBER, METHOD I STRESS, KSI Figure 8 COMBINED THERMAL AND LOAD STRESS TENSION MEMBER, METHOD 2 uTU\ u y \ Io5TH, L STRESS, KS I 5 0

t

aUL= 1.5 uLL I

w STRAIN

€TH LIMIT : VTH,L + QLL S 0 . 6 7 ~ ~ ~ < a ~ ~ , ~ + u~~ = ULTIMATE: I.~QTH,L + ISULL = QTU Figure 9 PANEL WE I GHTS WING LOWER SURFACE DRY BAY 10- .

r / 0 - SINGLE SKIN, ASCENT THERMAL CONDITION WING 6 - PANEL WEIGHT, L B / F T ~ 4 - 2- 0 I O 20 30 END LOAD, KIPSIIN.

Figure 10 STRUCTURAL WEIGHT FRACTION WING BODY 1 5 UillUlJ TAIL E 3 LANDING GEAR EZl ENGINE SUPPORTS, INLETS, ANDNACELLE .3 - 707 - 3208 1 6

y ELECTRA

WS

-

wTO .2 .I Ti SCAT A1 SCAT SUBSONIC Figure 11 @NASA RESEARCH ON MATERIALS APPLICABLE TO SUPERSONIC TRANSPORTS By Herbert F. Hardrath and George J. Heimerl

J = 4 _ ? 1 . ___

arch Center S U M M A R Y \ The National Aeronautics and Space Administration has conducted ' tests to determine the mechanical properties of materials suitable for use in a supersonic transport over the temperature range of interest On the basis of studies and after prolonged exposure to temperature.

on fatigue behavior, rate of fatigue crack propagation, residual static strength, and resistance to deterioration of properties due to prolong( exposure to temperature, no serious materials problems are anticipated However, stress corrosion is found to be a potentially serious problem in titanium alloys in a hot salt environment and in stainless steels i i The Ti-8AZ-lMo-lV alloy is found to be gener - an ambient environment.

ally superior to other contending materials in all respects studied except for salt stress corrosion in which respect it was poorer than all others. Further research is recommended on stress corrosion; on the combined effects of fatigue, thermal exposure, and creep; and on the development of structural configurations and fabrication procedure; suitable for the supersonic transport.

A d 7 f Y f l k Z

INTRODUCTION T % design of a successful supersonic - transport structure is in large part contingent upon the proper selection of material and in the use of proper allowable stress levels. Current airplanes may utilize established materials at stress levels that experience has proven will produce efficient and reliable structures. Such background information does not exist for materials of interest to the designers of a super - sonic transport.

This paper presents some of the results of several studies into the behavior of various materials that were candidates for use in the structure of a supersonic transport at the time the studies began. The results of earlier studies for screening purposes are not included because these data and the conclusions drawn from them have been pub - lished in reference 1 . Most of the work described herein was done at P the Langley Research Center but some information from contracted studies and from the Lewis Research Center are also included.

The e a r l i e r screening studies t r e a t e d a wide variety of sheet mate - r i a l s subjected t o quick t e s t s a t temperatures higher than those t h a t w i l l be encountered i n t h e vehicle t o exaggerate t h e temperature e f f e c t .

The present s e r i e s deals with a few selected materials t e s t e d i n more nearly r e a l i s t i c ways a t r e a l i s t i c temperatures and includes the e f f e c t s of long - time exposures. It i s believed t h a t these r e s u l t s w i l l be help - f u l i n making f i n a l selections of materials and i n choosing the allow - able stress levels t o be used i n design. This discussion i s limited t o of traveling materials suitable f o r use i n t h e skin of a vehicle capable a t a Mach number of 3 .

h S Y M B O L S AND ABBRENIATIONS s t r e s s concentration f a c t o r KT N number of cycles t o f a i l u r e number of cycles t o f a i l u r e of unexposed specimens NO R r a t i o of minimum s t r e s s t o maximum s t r e s s S stress mean s t r e s s s, t e n s i l e ultimate strength STU stress f o r 1 g loading a t take - off gross weight Slg aged A CR cold r o l l e d CRT cold r o l l e d and tempered double aged DA elongation i n 2 - inch gage length E l GAG ground - air - ground H T heat t r e a t e d MA mill annealed N S notch strength TA t r i p l e x annealed STATEMENT O F P R O B L W In order t o focus attention on the specific questions t o be dis - 'cussed, it i s of i n t e r e s t t o review the operating environment f o r major portions of the structure. Figure 1 represents a time history of s t r e s s f o r t y p i c a l f l i g h t a t a given s t a t i o n on the wing of the airplane.

Obviously, t h i s time history w i l l change significantly depending upon the specific airplane configuration, t h e s t a t i o n within t h e airplane, the p a r t i c u l a r f l i g h t , dynamic response characteristics, and so forth.

However, the general features w i l l usually be s i m i l a r t o those shown.

The s t r e s s values a r e normalized with respect t o the nominal s t r e s s present f o r l e v e l undisturbed f l i g h t at take - off gross weight and the time axis i s not t o scale.

The features of i n t e r e s t t o the s t r u c t u r a l designer a r e a s follows: 1. Most of the significant s t r e s s cycles occur during the climb portion of the f l i g h t while the airplane i s a t ambient temperature.

2. Few dynamic s t r e s s cycles a r e encountered during cruise because gust a c t i v i t y i s reduced a t high a l t i t u d e s .

3 . The exposure t o elevated temperatures up t o about 550' F may be as long as 90 minutes per f l i g h t .

4. The change i n temperature can produce thermal stresses of sig - nificant magnitude depending on d e t a i l designs. These s t r e s s e s are not shown i n figure 1.

5 , Dynamic loads encountered during descent a r e l e s s severe because t h e airplane has l o s t substantial weight because of the burning of the fuel.

6. A t stations outboard of the landing gear, a rather significant ground - air - ground cycle i s introduced.

3 51 i With t h e exception of the length of exposure t o elevated tem - 7.

peratures, the fatigue damage accrued during a given f l i g h t i s t o a f i r s t approximation independent of f l i g h t length.

Some of t h e engineering questions t h a t must be answered t o provide a structure capable of surviving the loadings shown f o r useful l i v e s of F i r s t of a l l , the basic mate - the order of 36,000 hours a r e as follows.

r i a l properties must be established over t h e temperature range of i n t e r - e s t . I n the N A S A studies the notch and t e n s i l e strengths, basic fatigue behavior, r a t e s of fatigue crack propagation, and residual s t a t i c strength of sheets containing cracks were obtained. O f the various studies made of creep properties, it appears t h a t overall creep w i l l not be a significant problem if the structure i s b u i l t of any of the leading titanium alloys, s t a i n l e s s steels, or superalloys. Having 1 established these properties over the temperature range of i n t e r e s t , one must know how some of these properties might be affected by t h e prolonged exposure t o elevated temperatures. The Langley Research Center has studied notch and t e n s i l e strengths, fatigue behavior, the rather perplexing salt stress corrosion, and t o some extent s t r a i n - r a t e s e n s i t i v i t y . The l a s t of these properties w a s studied t o only a l i m - i t e d extent with no significant e f f e c t s found, so it i s not discussed further. Some work has been s t a r t e d on methods of welding the mate - rials of i n t e r e s t and on studying the e f f e c t s of fabrication procedures on the strength of s t r u c t u r a l components.

RESULTS OF FZSEARCH IN INDIVIDUAL PR0BLF;M AREAS I n the following sections, each of these problem areas i s discussed i n t u r n with sample data presented t o indicate the conclusions reached i n each investigation. Generally, no insurmountable problems have been discovered and the Ti-8AZ-lMo-lV a l l o y appears t o be one of the best available on t h e basis of t h e r e s u l t s given here. S a l t s t r e s s cor - rosion has been found t o be one of the areas of concern and one on which additional study i s recommended.

Materials i Most of t h e various investigations were carried out on each of t h e following sheet materials: Condition Titanium alloys :

Ti-kAZ-3Mo-lV . . . . . . . . . . . . . . . . . . . . Aged

T i -6Az - 4V . . . . . . . . . . . . . .

. . . . . . Annealed

a+ ._ *., *?- * -

* - Condition

Ti-8AZ-lMo-lV . . . . . . . . . . . . . . . . Mill annealed

Ti-8A2-1Mo-lV . . . . . . . . . . . . . . Triplex annealed

Stainless s t e e l s :

AM 350 . . . . . . . . . . . . . . . . . . . . Double aged

AM 350 . . . . . . . . . . Cold r o l l e d (20%) and tempered

~ ~ 1 5 - ~ M O . . . . . . . . . . . . . . . . . . . . . m 1050

PH14-8Mo . . . . . . . . . . . . . . . . . . . . .SRH:950

. . . . . . . . . . . . . . . . Cold r o l l e d (50%)

AIS1 301

AM 367 . . . . . . . . . . Cold rolled (20%) and tempered

Superalloy : Ren641 . . . . . . . . . . . . . . . . . . . . . . . . . B Only a few t e s t s have been conducted on the superalloy, Re& 41, which may be useful i n the construction of leading edges and nacelle structures.

Notch and Tensile Strengths Most information on these properties was obtained i n connection with studies of the e f f e c t s of long exposure t o elevated temperatures.

Since no s t a r t l i n g r e s u l t s were obtained, these properties a r e discussed i n a subsequent section.

Fatigue Behavior Most of our study of fatigue behavior of materials of i n t e r e s t i n the design of a supersonic transport w a s devoted t o Ti-8AZ-LMo-lV and The titanium alloy was studied i n the m i l l - t o AM 350 s t a i n l e s s s t e e l .

annealed and i n the triplex - annealed conditions and the AM 350, i n the cold - rolled - and - tempered condition. A major p a r t of t h i s research w a s conducted under NASA contract a t B a t t e l l e Memorial I n s t i t u t e and a t lance Vought Corporation.

A summary of some of t h e data on t h e t i t a n i u m a l l o y i s shown i n The bar graphs indicate t h e fatigue strength a t a l i f e of figure 2.

about a million cycles f o r each of several configurations t e s t e d a t three temperatures (room (80° F), 5500 F, and 8000 F), as indicated by the shading. A l l tests were conducted on sheet specimens subjected t o a x i a l loads producing a mean s t r e s s of 25 ksi.

The fatigue strength of unnotched specimens was not adversely affected by the elevated temperatures; i n f a c t , a s l i g h t increase i n strength was noted. The mill - annealed material behaved as w e l l as t h e triplex - annealed material i n unnotched specimens. Fusion welding of t h e mill-annealed material producedpnly a moderate reduction i n strength. A specimen having an unloaded square of material spot - welded t o one face had a fatigue strength about t h e same as t h a t f o r a speci - men with an open hole with a stress concentration factor of 2.5. A spot - welded double butt - strap j o i n t had fatigue properties not too dif - ferent from those of edge - notched specimens with a t h e o r e t i c a l stress

concentration f a c t o r of 4. This l a t t e r configuration i s i d e n t i c a l t o

the configuration Langley has come t o regard as representative of the overall quality of construction achieved i n current aluminum - alloy i structures. The e f f e c t of temperature i s minor i n a l l cases where it was studied a s a parameter.

The same general comments apply t o s i m i l a r results of t e s t s on AM 350 s t a i n l e s s s t e e l which a r e presented i n figure 3 . The mean s t r e s s f o r t h i s material was 4 0 ksi.

These limited r e s u l t s do not, of course, provide a l l the informa - t i o n needed t o design a structure from a fatigue point of view. How - ever, they do indicate t h a t notch s e n s i t i v i t i e s a r e reasonable and t h a t welds and spot - welds produce strength reductions not too different from those of s t r e s s r a i s e r s representative f o r current structures. Obvi - ously, considerable t e s t i n g w i l l be required t o prove the adequacy of specific design d e t a i l s .

s t i l l needed on the possible interaction of creep Information i s and fatigue, especially around fasteners and other points of s t r e s s concentration. A current s e r i e s of t e s t s are being conducted a t the Lockheed Aircraft Corporation under an FAA - NASA - DOD sponsored contract and other tests a r e being conducted a t Langley.

Fatigue Crack Propagation The r a t e of fatigue crack propagation i s a most significant con - sideration i n assessing the f a i l - s a f e characteristics of a structure.

Figure 4 presents a sample of data obtained i n axial - load fatigue tests

of 8 - inch - wide - sheet specimens t e s t e d a t R = 0 (zero t o tension loading). The r a t e of crack propagation i s a reasonably straight - line function of crack length when both are p l o t t e d on a log scale.

The r e s u l t s presented a r e f o r room - temperature t e s t s with a mean s t r e s s of 25 ksi f o r titanium alloys and 40 k s i f o r s t e e l s . The most desirable material would be the one having the lowest rate of crack propagation.

O f the t i t a n i u m alloys, t h e Ti-8AZ-lMo-lV a l l o y i n the triplex - annealed condition i s superior. Two s t a i n l e s s s t e e l s , AM 350 CRT and A I S 1 301, - 0 .

appear superior t o the Ti-8AZ-lMo-1V alloy, but t h e rate of propagation i n both materials increases by a large f a c t o r a t elevated temperatures.

For reference, t h e r a t e s of propagation f o r two aluminum alloys, 2024-T3 and 7075-T6, are a l s o p l o t t e d i n figure 4 . The data shown are A t these stress levels, t h e aluminum - alloy f o r a mean stress of 15 ksi.

curves are somewhat steeper, indicating a f a s t e r increase i n rate of crack propagation as a function of crack length. The 7075-T6 material a l s o experiences crack growth rates higher than those f o r the other materials shown.

The e f f e c t of temperature w a s studied a t -llOo F and a t 550' F,

- with only moderate e f f e c t s on the r e s u l t s , except t h a t AISI 301 and

I ' AM 350 CRT experienced a marked increase i n r a t e of crack propagation This preliminary study indicates t h a t r a t e s of fatigue a t 550° F.

crack propagation at t h e s t r e s s l e v e l s considered a r e not too d i f f e r - ent from the behavior of current aluminum alloys a t a corresponding s t r e s s level. The e f f e c t s of temperature are moderate except f o r The Ti-8AZ-lMo-lV a l l o y stressed a t 25 k s i AISI 301 and AM 350 CRT.

appears t o be t h e best choice from among the materials tested.

Residual S t a t i c Strength Another important consideration with regard t o f a i l - s a f e character - i s t i c s i s the residual s t a t i c strength of a p a r t containing a fatigue crack o r other damage. Langley conducted s t a t i c t e s t s on the same s e t of specimens used t o study r a t e s of fatigue crack propagation. A sample of t h e results i s shown i n figure 5 . The s o l i d bar graphs show the residual strength of specimens containing 1 - inch cracks a s compared with the o r i g i n a l s t a t i c t e n s i l e strength of t h e material. The strengths a r e The data were divided by density t o compare s t r u c t u r a l efficiencies.

obtained f o r nine materials a t each of three temperatures.

I n general, the strength of these materials exhibited moderate s e n s i t i v i t y t o cracks. Exceptions a r e AM 330 (DA) and PH15-7Mo, both of which had low strength and f a i l e d by shattering a t t h e lowest tem - perature and AISI 301 which deteriorated seriously a t elevated temper - The titanium alloys are generally superior on a strength - density ) ature.

basis, Ti-8AZ-lMo-lV i n the triplex - annealed condition being t h e overall leader. The triplex - anneal process improved t h e residual s t a t i c strength of t h i s l a t t e r a l l o y a t room and subzero temperature. I n sev - eral materials t e s t e d a t elevated temperature, cracks produced l i t t l e o r no reduction i n residual s t a t i c strength beyond t h a t of reducing the cross - sectional area. The superalloy Re& 4 1 does not compare favorably here. It is, of course, affected very l i t t l e by temperature.

I *r The data presented i n figure 5 are f o r 1 - inch cracks only, but t h e general ranking of materials would be quite similar i f strengths f o r other crack lengths had been compared. I n most cases t h e reduction i n s t a t i c strength caused by cracks i s moderate and, i n general, i s con - siderably l e s s than t h a t for current. aluminum alloys. Several mate - r i a l s experience no l o s s of s t a t i c strength except t h a t produced by l o s s of area. Again, triplex - annealed Ti-8AZ-lMo-lV i s the best over - a l l choice of the materials evaluated.

Notch and Tensile Strength After Exposure t o Temperature The Langley Research Center i s conducting a study of t h e e f f e c t s of prolonged exposure t o elevated temperatures on s t a t i c t e n s i l e and

i

notch strength properties. Specimens are exposed without s t r e s s a t 550° F f o r various periods of t i m e and then t e s t e d a t room temperature The notched specimens were 1 inch wide and contained and a t - 1 1 0 ' F.

60° V notches with root radii <0.001 inch (ASTM ty - pe). Table I pre - Exposure t i m e s up t o sents a summary of t h e r e s u l t s obtained t o date.

14,000 hours have shown no significant e f f e c t on s t a t i c t e n s i l e or on In a few cases, elongations a t f a i l u r e have shown notch strengths.

rather large percentage changes, but most of the materials i n question

have r a t h e r low elongations (4 t o 8 percent) i n i t i a l l y , s o the absolute

changes a r e not much greater than the accuracy of measurement.

Some of the studies a t the Lewis Research Center indicated s t a b i l i t y problems with the two precipitation hardening s t a i n l e s s Steels, PH13-8Mo and PH14-8140, and with AM 367. Special heat treatments of AM 367 have improved the s t a b i l i t y characteristics but a t the expense of a serious decrease i n corrosion resistance.

Unstressed exposure t o elevated temperatures f o r as long as 14,000 hours produced no significant e f f e c t on s t a t i c strength proper - t i e s at' 80° F and -llOo F. Consequently, t h e materials studied a t Langley appear t o be reasonably s t a b l e f o r the exposures obtained a t t h e present t i m e .

Fatigue Behavior After Exposure t o Temperature In another study, fatigue t e s t specimens a r e being exposed t o 550° F f o r prolonged times without stress. Figure 6 presents the r e s u l t s of t e s t s on Ti-8AZ-lMo-1V i n the mill - annealed condition.

Four configurations of specimens were tested: unnotched, edge - notched (% = 4)' fusion welded, and a double butt - strap spot - welded j o i n t .

Complete S - N curves were derived from axial - load t e s t s of unexposed specimens, but exposed specimens were t e s t e d a t a single s t r e s s l e v e l chosen t o produce failure i n approximately lo5 cycles.

The mean stress i n a l l cases w a s 25 k s i and a l l t e s t s were a t room temperature.

The data presented are normalized with respect t o t h e l i f e of unexposed specimens and are p l o t t e d against exposure t i m e i n hours.

Each point i s the geometric mean of f i v e t e s t r e s u l t s .

These data f o r Ti-8AZ-l.Mo-lV display almost the extreme variations noted i n t h i s series of tests which included seven other materials. A factor of about two o r less on l i f e i s noted a t each data point except the one f o r unnotched specimens exposed for 9,000 hours which had a l i f e 10 t i m e s as long as t h e original. There i s some doubt t h a t t h e highest point i s valid because two of t h e tests produced l i v e s much more nearly equal t o t h e o r i g i n a l l i f e ; thus, t h i s point may be associated with t h e small s t a t i s t i c a l sample used.

The tests w i l l continue u n t i l exposure t i m e s reach about 30,000 hours. Unfortunately, t h i s test s e r i e s cannot be accelerated; therefore, the final data are not available at the present t i m e .

The data available a t t h i s date show no cause f o r alarm over dete - r i o r a t i o n of fatigue properties with exposure.

This general conclusion i s supported by supplementary information obtained a t BatteTle where notched specimens of AM 350 were t e s t e d i n fatigue a f t e r exposure t o 5500 F under 40 k s i s t r e s s f o r 3,000 hours with no significant e f f e c t of t h i s exposure on fatigue behavior.

S a l t Stress Corrosion The f i n a l investigation t h a t i s t o be discussed i n detail i s a study of s a l t stress corrosion a t elevated temperature, a recognized problem area f o r titanium alloys.

Figure 7 presents a summary of r e s u l t s of salt - stress - corrosion studies conducted a t Langley.

The specimens t e s t e d a r e indicated by the small sketch.

S t r i p s of sheet about 1/4 inch wide

and 4 inches long were bent through a specified angle near each end so

t h a t when two such s t r i p s were spot - welded a t t h e ends t h e center por - t i o n s were subjected t o a constant bending moment.

The bending stresses involved were e i t h e r 50 ksi o r 100 ksi.

The specimens were dipped i n concentrated s a l t solutions t o form a substantial crust over t h e e n t i r e 'specimen. The coated specimens were stored i n ovens a t 550° F f o r v a r i - ous times. Upon removal from the oven, the specimens were cleaned and subjected t o a x i a l compression loads as indicated i n the sketch.

The reduction or shortening i n the overall length of t h e specimen p r i o r t o f a i l u r e of one of the legs w a s noted.

This value i s normalized r e l a t i v e t o the value obtained i n unexposed specimens and plotted against t h e exposure time i n hours.

If severe stress - corrosion cracking occurs, the shortening a t fracture i s reduced and t h e r e l a t i v e shortening may approach zero.

Metallographic examinations have shown t h a t the depth f and number of stress - corrosion cracks i n t h e specimens correlates w e l l with t h e r e l a t i v e shortening parameter p l o t t e d i n figure 7.

The alloy Ti-kAZ-3Mo-lV i s v i r t u a l l y unaffected by salt stress cor - rosion but Ti-8AZ-lMo-lV (MA) i s affected very severely i n only N o specimens have 2,000 hours of exposure and Ti-6AZ-h, i n 4,000 hours.

f a i l e d i n t h e furnace i n as many as 7,000 hours.

The same t e s t s have been conducted on the s t a i n l e s s s t e e l s with no e f f e c t noted when the compression t e s t w a s conducted immediately upon removal from the oven. However, several specimens of AM 350 (DA) have f a i l e d i n about 2 months a f t e r removal from the oven. A l s o one speci - men of AM 350 (DA) f a i l e d a f t e r several months exposure t o the Langley outdoor atmosphere without salt and without a r t i f i c i a l heat.

It i s d i f f i c u l t t o assess t h e importance of the s a l t - s t r e s s - corrosion problem. Certainly the t e s t condition i s t o o severe but nat so severe a s t o suggest t h a t no problem e x i s t s . I f moisture, and thus soluble contaminants, can reach the inside of the structure o r lodge i n crevices on the outside, salts w i l l tend t o collect over a period of t i m e i n service. Tests elsewhere i n fast - moving a i r have indicated t h a t the attack i s l e s s severe under these circumstances. However, the insides of structures w i l l not have such ventilation. It was disap - pointing t o f i n d Ti-8AZ-lMo-lV most susceptible t o t h i s attack when t h i s material appears t o be one of the most suitable of the materials based on other considerations. The problem deserves more study with emphasis on a more r e a l i s t i c representation of conditions t o be encoun - tered i n service. Some information on the combined e f f e c t s of corrosion and fatigue i s desirable. Langley i s planning t e s t s t o provide some data and other data w i l l be supplied under FAA - NASA - DOD sponsored contracts.

The problem of s t r e s s corrosion may be quite serious f o r s t e e l s i n view of the f a c t t h a t f a i l u r e s occurred without salt or heat.

The f i n a l solution t o the stress - corrosion problem may require some form of corrosion inhibitor and probably very careful inspection.

CONCLUSIONS AND R E C O M M E N D A T I O N S FOR FUTURE RESEARCH d Recent investigations a t Langley on t h e properties of titanium and s t a i n l e s s - s t e e l sheet materials suitable f o r a supersonic transport indicate t h a t : The fatigue problem appears not t o be a formidable b a r r i e r a t 1.

t h e s t r e s s levels considered.

~* J I i r 2. Fail - safe considerations involving "rates of k t i g u e crack prop - agation and residual s t a t i c strength appear more promising f o r the b e t t e r titanium and stainless-steel materials t e s t e d than i s t r u e f o r current aluminum alloys.

3 . Prolonged exposure t o 550° F - has not been found t o degrade s t a t i c and fatigue properties significantly a t least i n the first 10,000 t o 14,000 hours.

4. Stress corrosion may be a problem with titanium alloys i n dry salt a t elevated temperatures and i n some of the s t a i n l e s s s t e e l s a t normal temperatures.

5. O f the contending materials, Ti-8AZ-lMo-lV a l l o y appears supe - I rior from a l l viewpoints consiaered except s t r e s s corrosion i n s a l t .

6. Further study i s recommended t o evaluate the combined e f f e c t s of dynamic loadings, thermal exposure, and creep.

7. Further study i s a l s o needed t o assess the severity of the hot salt - stress - corrosion problem f o r supersonic - transport applications.

8. Studies are needed t o develop p r a c t i c a l , e f f i c i e n t , and r e l i a b l e structures f o r supersonic transports b u i l t of titanium alloy.

REFEKENCE 1. Raring, Richard H., Freeman, J. W., Sehultz, J. W., andvoorhees, H. R.: Progress Report of t h e N A S A Special Committee on Materials Research f o r Supersonic Transports. N A S A TN D-1798, 1963.

TABU I EFFECT OF EXPOSURE AT 550° F Exposure Percent change Material Titanium alloys Ti -6AZ - 4V A 14 x 103 - 110 0 0

---

Ti-8AZ-lMo-lV A 10 - 110 80 0 - 10

--

-110 Ti -4AZ -3Mo-lV HT 7 1 0 80 0

I Stainless steels

- 110

PH15 -WO HT 7 x 103

CRT 7 - 110 AM 350

A I S I 301 CR 4 - 110

SCHEDULE OF LOADS AND TEMPERATURES+OR SUPERSONIC TRANSPORT FLIGHT - i CRUISE CLIMB DESCENT G A G 55OOF 8OoF 8OoF 8OoF 90 MIN

1.5r

I .o

&

’lg .5

- . 5

T I M E Figure 1 FATIGUE STRENGTH AT N =IO6 CYCLES Ti-8Al-!Mo-IV; S m = 2 5 KSI TA UNNOTCHED MA I FUSION WELD MA OF I 80 UNLOADED TA SPOT KT ~ 2 . 5 TA SPOTJOINT MA MA K T = 4 TA 0 40 80 FATIGUE STRENGTH,KSI Figure 2 FATIGUE STRENGTH A T N =IO6 CYCLES AM 350 CRT; S , = 40 KSI UNNOTCHED FUSION WELD

I

UNLOADED O F SPOT

- 80

KT 2.5 800 SPOT JOINT I - ; FATIGUE STRENGTH, KSI Figure 3 FATIGUE CRACK PROPAGATION IN CENTRALLY NOTCHED TITANIUM, STEEL, AND ALUMINUM SHEET SPECIMENS 8 - INCH WIDTH; R=O; 80 " F I I I I I I I I I I ,I .5 1 . 0 CRACK LENGTH, IN.

Figure 4 ' 3 vl RESIDUAL AND TENSILE STRENGTHS I - INCH CRACK IN 8 - INCH SPECIMEN -RESIDUAL (GROSS) Ti-8Al-lMo-lV TA Ti-8AI- 1Mo-IV MA Ti-4Al- 3MO- 1V Ti-6A2-4V AM350 AM 3 5 0 C AIS1 301 PH15-7'0 H REN€ 41 STRESS/DENSITY, IO6 IN.

Figure 5 FATIGUE LIFE AFTER EXPOSURE TO 550° F N . - No I 0 5,000 10,000 EXPOSURE T I M E , HR Figure 6 SALT STRESS CORROSION AT 550 " F

Ti - 4AZ- 3Mo-lV

I .o RE LATl V E SHORTENING *5

- - - - - -

0 I O ~ Figure 7 ROELASTIC CONSIDERATIONS

By Robert W. Boswinkle, Jr., and E . Carson Yates, Jr. A-2

SUMMARY

// 762

Problem areas which involve combinations of aerodynamic, elastic, and inertial forces are considered in relation to the design of the supe Some problems of ground loads, buffet, flutter, panel sonic transport.

.flutter, gust loads, and a consideration of aeroelastic effects on sta- bility and control are discussed. Aeroelastic - dynamic considerations will be very important in the design and flight - test phases of the super sonic transport; however, in general, the techniques for handling the - problems which will arise appear to be reasonably adequate.

4 d T f l d L

w-

h INTRODUCTION

Many problem areas in the design of the supersonic transport will involve combinations of aerod&amic, eJ-a_stic and inertial forces. A

--\’-

few of these problem areas, which are illustrated in figure 1, are dis - cussed herein. The parameters are altitude and Mach number. The flight Flutte region would be above the design dive flight condition boundary.

of lifting surfaces, gust response, panel flutter, and buffet will prob - The effects of aeroelasticity ably be critical near transonic speeds.

on stability and control will probably be greatest at the high supersonj Mach numbers where the dynamic pressure is greatest; although, problems may be encountered elsewhere.

the design is to have the instabilities, the excessive loads, and other undesirable conditions represented by the problem areas The area fall a sufficient distance below or above the flight region.

of ground loads is shown at the low - speed end of the Mach number scale.

Two major factors which complicate the solutions to most of these aeroelastic problems are: (1) The low stiffnesses which result from thin lifting surfaces and a slender fuselage (2) The wide range of operating conditions involving variations in Mach number, altitude, aerodynamic heating, weight, and pos - sibly aircraft configuration.

I a length of skin panel i n stream direction b width of skin panel perpendicular t o stream direction E Young's modulus M Mach number 9 dynamic pressure t thickness of skin panel leading - edge sweep angle ALE wind - off buckling s t r e s s bc r compressive midplane s t r e s s i n direction of flow bX AEROELASTIC PROBLEM AREAS Following b r i e f comment on two of the areas shown i n figure 1, ground loads and buffet, the other areas a r e discussed i n more d e t a i l .

Ground Loads In regard t o ground loads, the problems associated with taxiing, take - off, and landing f o r the supersonic transport are not expected t o be greatly d i f f e r e n t from those of t h e current, large, high - performance a i r c r a f t . However, current design practices may be affected by the anticipated long fuselages, new gear configurations, and high nose - up

landing a t t i t u d e s (ref. 1) .. I n addition, as i n current a i r c r a f t , an

important aspect of the airframe design w i l l be the loads generated by These may even be the c r i t i c a l loads f o r runway and taxiway roughness, In gen - some designs (paper no. 3 by R. Richard Heppe and J i m Hong).

e r a l , the a n a l y t i c a l techniques f o r handling ground loads a r e believed t o be reasonably adequate.

Buffet may be defined a s s t r u c t u r a l response t o randomly fluctu - a t i n g aerodynamic forces caused by separation of the flow ( r e f . 2).

Since flow separation must be minimized t o maintain low drag, buffet would not be expected t o present major d i f f i c u l t i e s over much of t h e f l i g h t region. However, one p a r t of t h e f l i g h t region where buffet may be troublesome i s a t transonic speeds. Flow separation becomes a greater problem a t transonic speeds because of shock boundary - layer interactions and possibly because r e l a t i v e l y high angles of attack may be required a s a result of high - altitude f l i g h t a t transonic speeds t o reduce sonic - boom overpressures. Few buffet problems are amenable t o analytic study; however, buffet indications often have been obtained . !

I n addi - a s a byproduct of wind - tunnel force - test studies on models.

tion, some buffet studies have been made by using e l a s t i c a l l y scaled models.

F l u t t e r F l u t t e r i s a self - induced o s c i l l a t i o n of a structure which can A f l u t t e r boundary f o r a variable - sweep lead t o catastrophic f a i l u r e s .

wing i s shown i n terms of a l t i t u d e and Mach number i n f i g u r e 2. The boundaries, shown by t h e three s o l i d curves f o r sweep angles of 20°, 4 5 O , and 6 5 O , were based on t h e scaled model f l u t t e r studies of refer - The dashed extension f o r t h e 6 5 O sweep ence 3 . (See a l s o r e f . 4 . ) curve i s extrapolated t o indicate the t y p i c a l trend. An example design - dive - speed boundary f o r t h e supersonic transport i s also shown.

This figure may be used t o make several points: (1) I n order t o insure freedom from f l u t t e r , the c i v i l a i r regula - tions require t h a t t h e f l u t t e r boundary be a t an equivalent speed a t l e a s t 20 percent higher than t h e design dive speed ( r e f . 5 ) .

(2) Sweepback as indicated i n figure 2 increases t h e f l u t t e r mar - gin, and early sweepback might be used f o r variable - sweep wings t o pro - vide t h e required f l u t t e r margin a t transonic Mach numbers.

( 3 ) For most supersonic a i r c r a f t , the minimum f l u t t e r margin occurs I a t transonic speeds with increasing f l u t t e r margins occurring i n the low supersonic range. However, the characteristic r i s e i n the f l u t t e r boundary a t the higher Mach numbers may cause some designs t o become c r i t i c a l again at Mach numbers as high a s 3 .

(4) Over t h i s wide operating range, d i f f e r e n t components of the and t h e a i r c r a f t may become f l u t t e r - c r i t i c a l a t d i f f e r e n t Mach numbers, flutter boundary for the complete aircraxt must be considered rather than that of only the wing component as illustrated here.

The following general procedure for treating flutter has been evolved over the years and appears applicable to the supersonic transport: (1) Inputs, based on experience and existing data, during initial design to avoid conditions which appear unfavorable from the standpoint of flutter (2) Use of analyses and model experiments to study flutter of com - ponents, with the initial simplified approaches being super - seded by more complex methods as the design progresses (3) Use of analyses and model experiments for the complete airplane

(4) Verification of full - size aircraft structural properties

as components become available followed by ground vibration tests of the complete aircraft (5) Flight flutter tests The flutter analyses, referred to in items 2 and 3, may be dis - cussed with the aid of figure 3 . Many analytical methods are available for representing the various forces involved in flutter. Usually the most difficult of these forces to represent is the aerodynamic force.

Some of the methods used at different Mach numbers for representing the aerodynamic force are illustrated; however, the actual range of appli - cability depends on the configuration. Shown as applicable are the

kernel - function method (refs. 6 and 7) at subsonic speeds, various

aerodynamic - influence - coefficient methods (refs. 8 to 10) at supersonic speeds, and the relatively simple but nonlinear piston theory (refs. 11 to 13) at the higher supersonic speeds.

It should be noted that none of these theories may be applicable over a relatively wide band at transonic speeds which is often the critical region. This difficulty has led to the development of approx - imate methods. Many of the approximate methods have as their basis the steady - state flow, which may be obtained either from theory or experi - ment. One such scheme is the so - called modified - strip method (refs.

For a 45' swept wing at subsonic speeds, the modified - strip - to 1 7 ) .

analysis results, shown with a solid line, yielded excellent agreement with the theoretical kernel - function results shown by the dashed line.

Results of the modified - strip analysis are in good agreement with exper - imental data throughout the transonic speed range.

Approximate aerodynamic theories are often used throughout the - Mach number range in the place of more rigorous theory to save com puting time and to provide a more flexible analysis; however, the use of approximate theory requires discretion and caution.

Figure 4 illustrates a system for mounting complete flutter models (This system has application to item 3 in the pre - in wind tunnels.

One of the problems in experimental transonic studies ceding listing.)

of complete models has been deficiencies in the various mounting systems proposed. The system illustrated, which was developed recently at the NASA Langley Research Center (ref. 1 8 ) , appears to avoid most of the deficiencies. It consists of a cable in the vertical plane and one in the horizontal plane; both cables are attached to the tunnel walls and 'are passed around pulleys in the model. The mount permits the model to be tested with minimum restraint on its free - flight motions. Three different aircraft configurations have been investigated at transonic speeds with this mount system.

It may be remarked that flutter studies of complete models at supersonic speeds have not often been necessary because of the large flutter margins which usually exist in the Low supersonic range; how - ever, if, as previously discussed, a second critical flutter region exists at the maximum Mach number, a supersonic study of a complete model may be required.

A few comments that may be made in regard to the flutter character - istics of specific configurations are as follows: (1) The solution of delta - wing flutter problems requires careful adjustments of local mass and stiffness in order to avoid unnecessary weight. Few other generalities for delta wings can be made.

(2.) For a variable - sweep wing, the pivot does not present any par - ticular difficulty in making a flutter analysis, and on the basis of several design studies, it - appears that a pivot designed to have enough strength to accommodate the loads may be considered almost rigid from the standpoint of flutter.

(3) For wings with engines attached, the flutter speed may vary considerably with engine location (refs. 19 and 2 0 ) ; since engine location affects other structural quantities such as sonic fatigue, it may be necessary to involve the flutter characteristics in trade - off studies .

Panel Flutter Panel flutter is a self - excited oscillation of external skin panels in which the deformation of the skin surfaces generates aero - dynamic forces which sustain the oscillations. Thus, the response of a panel to acoustic or mechanical forces is excluded from consideration in this discussion of panel flutter.

The following is a list of some factors which affect panel flutter: Panel Flow

-

Dynamic pressure Aspect ratio Edge restraint Thickness Elastic modulus Mach number Boundary layer Curvature Cavity size Orientation Construction Cavity pressure Midplane stress Since the number of factors is large, panel flutter is a difficult prob - lem to handle. The phenomenon may manifest itself as either fatigue or as sudden disintegration of the skin. It usually occurs only at super - sonic speeds on panels with low design loads which have minimum skin gage and large spacings between stiffeners.

In the last decade, panel flutter has caused many problems which required considerable time and expense to correct. Examples of air - planes involved are the A3J, X-15, and the F - 101. The difficulties with these airplanes generated considerable experimental research, which generally correlated poorly with theory. The experiments also gave rise to panel - flutter criteria which were not reliable. Theoretical - and experimental - prediction methods are still somewhat deficient and larger design margins of safety are recommended than for lifting - surface flut - ter; however, the situation appears to be improving. The improvements are coming about through a better understanding of the importance of adequate representation in analyses of such items as the edge conditions and through an awareness of the extreme care with which any experiments must be conducted.

A n example of recent research (ref. 21; see also, refs. 22 and 23) i is shown in figure 5 where the frequently used panel - flutter parameter is plotted against a parameter which is proportional to the compressive midplane stress in the flow direction. One of the major effects of temperature on panel flutter occurs through the production of midplane stress. The effect of increasing compressive midplane stress is to reduce the dynamic pressure at which flutter occurs for unbuckled panels. For higher values of midplane stress, the panel is buckled and the dynamic pressure for flutter increases with increases in midplane stress. The minimum value of ‘dfia;nid ‘prekwe is a critical value for panel design and is the subject of current analytical and experimental studies.

The recommended design procedure. for avoiding panel flutter is as follows : (1) Identify, on the basis of experience, which panels are least safe.

(2) Estimate, from avai2able experimental and theoretical data, the flutter margins of the identified panels.

I ( 3 ) For panels indicated to be critical, either make specific experimental and theoretical studies to verify the indications or change the design.

(4) Verify freedom from flutter during flight tests.

Gust Loads The gust response of an aircraft may be computed by combining two The turbulence input from the atmosphere and the transfer quantities: function of the aircraft.

Figure 6 shows a sample power spectrum of atmospheric turbulence in a thunderstorm at 40,000 feet (ref. 2 4 ) .

Curves for other atmospheric conditions tend to be parallel to this one.

This curve shows relative power as a function of the wave length of the turbulence with the wave length increasing to the left. The solid line represents measurements made recently to a wave length of 14,000 feet.

The shape of the dashed line is theoretical.

Shown above the abcissa are the wave lengths which correspond to the short - period stability mode for present subsonic jet transports and Figure 6 may be used to illustrate the for the supersonic transport.

following comments: (1) The wave lengths of atmospheric turbulence which are most important to the response tend to be longer for the supersonic trans - port than for the present subsonic jets.

(2) “he available atmospheric measurements, of which this is a sample, indicate that the relative power of the turbulence input is higher for the longer wave lengths so that the airplane response would be expected to be greater for the supersonic transport than for the present subsonic jets for flight in the same atmospheric conditions.

Whether or not the supersonic transport will actually experience a more I I ~ V * ?

> i b z # ) > a severe response history will depend -also an such factors as the fre - quency of encountering turbulence in the flight regime of interest and the success of avoiding turbulent regions.

( 3 ) Various deficiencies exist in the description of the atmos - pheric turbulence; the most importkt are power - spectral descriptions for the longer wave lengths (to confirm the degree to which the curve bends over), power - spectral descriptions for clear - air turbulence at altitudes above 5,000 feet, and measurements of the frequency of encoun - ter for various vertical velocities at the higher altitudes.

As previously noted, the supersonic transport will probably have the greatest gust problem at transonic speeds; however, there has not yet been adequate experimental confirmation that the transfer func - tion at transonic speeds for proposed supersonic - transport configura - tions can be calculated satisfactorily. The problem is being studied at Langley.

Figure 7 illustrates the importance of including structural flexi - bility in the gust analyses (ref. 25). The configuration is a large, flexible, swept - wing airplane. The vertical acceleration of the air - plane center of gravity is shown on the left, and wing bending strain is shown on the right as functions of the frequency of the sinusoidal gust input. The data points shown by the circles were obtained from flight tests. The solid lines are theoretical predictions for the flexible'airplane. The dashed lines are predictions which assume that the airplane is rigid. The conclusions drawn from figure 7 are: (1) Although the inclusion of airplane flexibility did not greatly affect the maximum acceleration at the center of gravity, large effects of flexibility were shown for the wing bending strain.

(2) The theoretical methods yielded adequate results.

Aeroelastic Effects on Stability and Control A wide spectrum of problems are classified as aeroelastic effects on stability and control (see refs. 2 6 to 2 8 ) . Aeroelasticity affects: i (1) The static - stability quantities such as lift - curve slope, center of pressure, and aerodynamic center; some of these factors have been studied in connection with supersonic trans - port configurations in paper no. 11 by William. J. Alford, Jr., Vernard E. Lockwood, Linwood W. McKinney, and Richard K .

Greif, and in the midterm SCAT review.

2 % > i $ 3 (2) The dynamic - stability quantities sue frequency and damping of aircraft oscillations (see ref. 29) (3) The effectiveness and reversal characteristics of various controls (e.g., refs. 2 8 .and 29) as influenced by the flexibility of the wings, fuselage, stabilizer, and fin (4) Various other items such as the performance, particularly at off - design conditions, due to changes in trim drag and the transient response due to engine failure.

One of the aeroelastic effects on the dynamic stability of a delta - wing canard configuration (ref. 29) is illustrated in figure 8. The relative damping for the first three elastic modes of the fuselage and / the relative damping in the short - period stability mode are shown. The abscissa is the square root of the relative structural stiffness of the fuselage. Figure 8 shows that as the fuselage stiffness decreases, the aerodynamic damping in the short - period stability mode decreases, and, actually, for low stiffnesses, is shown to become unstable. A n insta - bility is also indicated at low fuselage stiffness for one of the elastic modes.

The methods of analysis used to study aeroelastic stability and control effects are generally well known. Steady - state aerodynamic loadings can be calculated from proven lifting - surface thepries. The usual procedure has been to compute static aeroelastic corrections to wind - tunnel data obtained with rigid models. Recently, however, for at least one airplane, static stability wind - tunnel data are being obtained with an elastically scaled model.

When oscillatory aerodynamic loadings are required, they are usually obtained from aerodynamic methods used in flutter analyses.

The stability augmentation system should, of course, be considered in the dynbic analysis (ref. 30) In fact, the blending of dynamic sta - bility into the flutter area has led in some cases ta the use of a single analysis for both problem areas.

CONCLUDING REMARKS Aeroelastic - dynamic considerations will be very important in the design and flight - test phases of the supersonic transport; however, in general, the techniques for handling the problems which will arise appear to be reasonably adequate.

1. Anon.: Tentative Airworthiness Objectives and Standards for Super - Flight Standards Service, FAA, sonic Transport Design Proposals.

Aug. 15, 1963.

2. Garrick, I. E., and Rainey, A Gerald: Remarks on the State - of - the - A r t of Buffet - Loads Prediction. Presented t o Structures and Materials Panel of AGARD (Paris, France), July 3 - 6, 1962.

3. Ruhlin, Charles I, and Gurley, John R., Jr.: Transonic F l u t t e r Investigation of Models of a Proposed Variable - Sweep W i n g . N A S A TM X - 739, 1962.

4. Stonesifer, John C., and Goetz, Robert C.: Transonic and Super - sonic F l u t t e r Trend Investigation of a Variable - Sweep W i n g . NASA T M X-598, 1961.

5 . Anon.: Airplane Airworthiness; Transport Categories. Civil A i r Regulations Part 4b, R u l e s Service Co. (Washington, D . C . ) , Jan. 7, 1963.

6. Watkins, Charles E., Runyan, Harry L., and Woolston, Donald S.: O n t h e Kernel Function of t h e Integral Equation Relating the L i f t and Downwash Distributions of Oscillating F i n i t e Wings i n Subsonic Flow. NACA Rep. 1234, 1955. (Supersedes NACA TN 3131.)

7. Watkins, Charles E., Woolston, Donald S., and Cunningham, Herbert J.: A Systematic Kernel Function Procedure f o r Determining Aerodynamic Forces on Oscillating o r Steady F i n i t e Wings a t Subsonic Speeds.

NASA TR R - 48, 1959.

8. Pines, Samuel, Dugundji, John, and Neuringer, Joseph: Aerodynamic F l u t t e r Derivatives f o r a Flexible Wing With Supersonic and Sub - sonic Edges. Jour. Aero. Sci., vol. 22, no. 10, Oct. 1955, PP* 693-7000 9. L i , Ta: Aerodynamic Influence Coefficients f o r an Oscillating F i n i t e Thin Wing i n Supersonic Flow. Jour. Aero. Sci., vol. 23, I no. 7, July 1956, pp. 613 - 622.

3 74

10. Zartarian, Garabed, and Hsu, Pao - Tan: Theoretical Studies on t h e Prediction of Unsteady Supersonic Airloads on E l a s t i c Wings.

WADC Tech. Rep. 56 - 97, Pts. I and 11, U.S. A i r Force.

Part I. Investigations on t h e U s e o f Oscillatory Supersonic Aero - dynamic Influence Coefficients. ASTIA Doc. No. AD 110591, Dee. 1955.

Part 2 . R u l e s f o r Application of Oscillatory Supersonic Aero - dynamic Influence Coefficients. ASTIA Doc. N o . AD 110592, Feb. 1956.

11. Lighthill, M. J.: Oscillating Airfoils a t High Mach Number. Jour.

Aero. Sci., vol. 20, no. 6, June 1953, pp. 402 - 406.

i 12. Ashley, Holt, and Zartarian, Garabed: Piston Theory - A New Aerodynamic Tool f o r t h e Aeroelastician. Jour. Aero. Sei., vol. 2 3 , no. 12, Dee. 1956, pp. 1109-1118.

13. Chawla, Jagannath P.: Aeroelastic I n s t a b i l i t y a t High Mach Number.

Jour. Aero. Sci., vol. 25, no. 4, Apr. 1958, pp. 246-258.

14. Yates, E. Carson, Jr.: Calculation of F l u t t e r Characteristics f o r Finite - Span Swept o r Unswept Wings a t Subsonic and Supersonic Speeds by a Modified S t r i p Analysis. NACA RM L57L10, 1958.

15. Yates, E. Carson, Jr.: Use of Experimental Steady - Flow Aerodynamic Parameters i n t h e Calculation of F l u t t e r Characteristics f o r Finite - Span Swept o r Unswept Wings at Subsonic, Transonic, and Supersonic Speeds. NASA. TM x - 183, 1960.

1 6 . Yates, E. Carson, Jr.: Subsonic and Supersonic F l u t t e r Analysis of a Highly Tapered Swept - Wing Planform, Including Effects of Density Variation and F i n i t e W i n g Thickness, and Comparison With Experiments. NASA TM X - 764, 1963.

17. Yates, E. Carson, Jr., and Bennett, Robert M.: U s e of Aerodynamic Parameters From Nonlinear Theory i n Modified - Strip - Analysis F l u t t e r Calculations f o r Finite - Span W i n g s a t Supersonic Speeds.

NASA TN D - 1824, 1963.

18. Reed, Wilmer H., 111, and Abbott, Frank T., Jr.: A New " Free - Flight " Mount System f o r High - speed Wind - Tunnel F l u t t e r Models.

N A S A paper presented t o t h e Symposium on Aeroelastic and Dynamic Modeling Technology (Dayton, Ohio), Sept. 23 - 25, 1963.

19. Walberg, Gerald D.: Transonic F l u t t e r T e s t s of a Highly Swept Arrow Wing With and Without Simulated Trailing - Edge - Mounted Engine Masses. NASA TN D-1023, 1962.

.E 20. Anon.: Commercial Supersonic Transport F l u t t e r Studies. NAA Rep. No. 63-646 (Contract No. AF 3 3 (657) -8514), North American Aviation, Inc . , J u l y 30, 1963.

21. Guy, Lawrence D., and Dixon, Sidney C.: A C r i t i c a l Review of Experiment and Theory f o r F l u t t e r of Aerodynamically Heated Panels.

Symposium on Dynamics of Manned Lifting Planetary Entry.

S. M. Scala, A. C. Harrison, and M. Rogers, eds., John Wiley & Sons, Inc., c.1963, pp. 568-595.

22. Stocker, James E.: A Comprehensive Review of Theoretical and Experimental Panel F l u t t e r Investigations. Rep. No. NA6lH-444, North American Aviation, Inc., Sept. 15, 1961.

2 3 . Anon.: Commercial Supersonic Transport Panel F l u t t e r Studies.

NAA Rep. No. 63-952 (Contract No. AF 3 3 ( 6 5 7 ) - 8 5 1 5 ) , North American Aviation, Inc., Aug. 2 3 , 1963.

2 4 . Rhyne, Richard H., and Steiner, Roy: Turbulence and Precipitation Problems Associated With Operation of Supersonic Transports.

NASA paper presented t o Fourth Conference on Applied Meteorology (Hampton, V a . ) , Sept. 10- 14, 1962.

25. Houbolt, John C., Steiner, Roy, and P r a t t , K e r m i t G.: Flight Data and Considerations of t h e Dynamic Response of Airplanes t o Atmospheric Turbulence. Presented t o Structures and Materials Panel and t o Flight Mechanics Panel of AGARD (Paris, France), July 3 - 13, 1962.

2 6 . Taylor, A. S . : The Present Status of Aircraft S t a b i l i t y Problems i n t h e Aeroelastic Domain. Tech Note N o . Aero.2538, B r i t i s h R.A.E., Dec. 1957.

27. J. B . Rea Co., Inc.: Aeroelasticity i n S t a b i l i t y and Control.

WADC Tech. Rep. 55-173, U.S. A i r Force, Mar. 1957.

2 8 . Anon. : B-70 Technology Applicable t o t h e Supersonic Transport.

Rep. NA - 62 - 601, North American Aviation, Inc., June 1, 1962.

2 9 . Ruiiyan, H. L . , P r a t t , K. G., and Bennett, F. V.: Effects of Aeroelasticity on t h e S t a b i l i t y and Control Characteristics of Airplanes.

Presented t o AGARD Specialists Meeting on S t a b i l i t y and Control (Brussels, Belgium), Apr. 10- 14, 1961.

30. McRuer, D. T., Benun, D., and Click, G. E.: The Influence of Servomechanisms on the F l u t t e r of Servocontrolled Aircraft.

AF Tech. Rep. No. 6287, W r i g h t A i r Dev. Center, U.S. A i r Force, Mar. 1954.

J AEROELASTIC PROBLEM AREAS ALTITUDE, FT 80 xi03

r

60 c

FLUTTER BOUNDARY 80 X I O ~

r

DESIGN DIVE

/-"- FLIGHT CONDITION

60t

ALTITUDE,

c FT 40

0 I 2 3 MACH NUMBER Figure 2 COMPARISON OF FLUTTER ANALYSIS AND EXPERIMENT FOR SWEPT WING ,- MODI IF^^ S T R I P , ,

t N l i E O R Y 4

AERODY NAMlC INFLUENCE COEFFICIENTS ALTITUDE

‘I

0 EXPERIMENT - MODIFIED - STRIP ANALYSIS --- KERNEL FUNCTION 0 I 2 3 M Figure 3 TWO - CABLE MOUNT i

Figure 4

L b PANEL FLUTTER AT M = 3 . 0 a/b = I O

FLOW - 0 4 1

- - I J b D I RECTION

f

FLUTT'ER 0 BUCKLED PANEL FLAT NO FLUTTER

1 PANEL

T I 1 I I I I 0 2 4 6 ux/qcr Figure 5 SAMPLE SPECTRUM OF TURBULENCE FOR THUNDERSTORM

- '\

105 lo4 POWER SPECTRAL DENSITY, .

lo3 ( F T I S E C I ~ RADlAN/FT '

IO2 - I +M=0.8b

Figure 6 SUBSONIC RESPONSE OF LARGE SWEPT - WING AIRPLANE o EXPERIMEN T THEORY

-

FLEXIBLE RIGID WING C. G.

BENDING ACCELERATION STRAIN 0 I .o 2 -0 C 1.0 2 .o FREQUENCY, CPS Figure 7 - * ,’<.T

. I

EFFECT OF FUSELAGE STIFFNESS ON DAMPING FOR DELTA - WING CANARD MODAL DAMPING RATIO -.04 I + JRELATIVE STIFFNESS Figure 8 a - CONFIGURATION EFFECTS ON SONIC BOOM By Harry W. C a r l s o n & & NASArLangley Research Center SUMMARY / m 3 An outline of current theoretical methods for sonic - boom estimation is presented and the correlation of this theory with experimental data is illustrated. Sonic - boom lower bound concepts are used in examining the dependence of sonic boom on configuration variables and in pointing out design consideraLions tending to minimize the problem.

d r U 6 L

% -

INTRODUCTION One of the most important problems to be faced in the development of a supersonic transport is that posed by the sonic boom. Feasibility studies have shown that sonic - boom considerations alone may dictate allowable minimum altitudes along most of the flight path and have indi - cated that in many cases the airframe sizing and engine selection depend directly on the sonic - boom characteristics.

It is the purpose of this paper to outline current theoretical illustrate the correlation of this theory with estimation methods, t o experimental data, and to examine the dependence of sonic - boom over - pressures on configuration variables pointing out, in general, design considerations tending to minimize the problem.

SYMBOLS A cross - sectional area of airplane or model determined by \ supersonic area - rule concepts nondimensionalized cross - sectional area A / Z ~ at nondimen- sionalized station t = X effective nondimensionalized cross - sectional area due to a A E ( t) ,AE combination of volume and lift effects,

A ( t ) + B ( t )

. < . ' I )1 a k a a J < a * w 0 " a * a cross - sectional area at base of airplane or model *b B equivalent cross - sectional area due to lift at airplane or model station given by the equation nondimensionalized equivalent cross - sectional area due to lift B / Z ~ at nondimensionalized station t = drag coefficient at zero lift ' D , o lift coefficient CL lifting force per unit length along airplane or modellongi- tudinal axis effective area distribution function given by the equation h airplane flight altitude o r lateral distance from model to measuring probe reflection factor length of airplane or model equivalent body Mach number integers reference pressure incremental pressure due to flow field of airplane or model dynamic pressure wing planform area J 'I 3 P k > 3 i > i s 9 3 , 1 9 l ' i i v nondimensionalized distance measured along longitudinal axis t from airplane or model nose, x/2 W airplane w e i g h t Y r a t i o of specific heats f o r a i r T dummy variable of integration measured i n same direction and using same u n i t s a s t z To value of T giving the l a r g e s t positive value of t h e T i n t e g r a l , F(T)d-r X distance measured along longitudinal axis from airplane or model nose

Mach angle, sin-1 1

I-L M Subs c r i p t : MAX maximum

A prime ' i s used t o indicate a f i r s t derivative and a double

prime ", a second derivative with respect t o distance.

DISCUSSION The nature of the airplane shock f i e l d which i s responsible f o r the sonic - boom phenomena i s i l l u s t r a t e d i n the schematic sketch of f i g - ure 1. A t supersonic speeds t h e airplane - generated flow f i e l d i s con - centrated within a bow shock front and a t a i l shock front fanning out from the airplane t o reach the ground and there t o be reflected upward.

Near the airplane the pressure signature i s quite complex.containing pulses from the airplane nose, wing - fuselage juncture, engines, t a i l surfaces, and other airplane components. A s the distance from the air - plane increases, t h e separate pulses merge and only a nose and t a i l shock remain. The resultant N - wave s i g n i f i e s the attainment of t h e so - ground l e v e l t h e incident and reflected c a l l e d f a r - f i e l d conditions. A t signatures a r e coincident and an amplification of the pressure rise occurs. O n a hard l e v e l surface a doubling of t h e pressures takes place.

For other surfaces t h i s r e f l e c t i o n f a c t o r may be somewhat less than 2.0. It i s t h i s p a i r of pressure jwrrps traveling with the air - plane and passing over the ground t h a t i s sensed by the observer.

An outline of t h e t h e o r e t i c a l method used i n estimating t h e inten - s i t y of the f a r - f i e l d bow - shock pressure r i s e d i r e c t l y below an air - plane i n l e v e l supersonic f l i g h t i s shown i n figure 2.

The method i l l u s t r a t e d i n t h e figure, i n a form suitable f o r a numerical solution using electronic computing machines, has been derived from the work of references 1 and 2, t h e main differences stemming from changes i n ter - minology and i n the expression of l i f t e f f e c t s i n terms of equivalent cross - sectional area. The following equation r e l a t e s overpressure t o t h e geometry of t h e airplane and t h e f l i g h t conditions: The right - hand side of equation (1) depends only on the geometry of the airplane and i s evaluated by taking i n t o consideration a distribution of nondimensionalized airplane area A( t ) formed by supersonic - area - r u l e cutting planes and a distribution of nondimensionalized equivalent area due t o l i f t B ( t ) evaluated through an integration of the l i f t i n g force per u n i t length along the airplane longitudinal axis. Since the pressure f i e l d d i r e c t l y below the airplane i s of primary concern, only one s e t of cutting planes having an angle with respect t o the hori - p zontal i s used. The many s e t s necessary i n drag evaluation are not pertinent.

Improved accuracy r e s u l t s when t h e area distributions include the increases i n cross - sectional area due t o boundary layer and engine exhaust e f f e c t s . Configuration sonic - boom characteristics a r e d i r e c t l y dependent on an effective area distribution curve A E ( t ) formed by a d i r e c t addition of actual area and equivalent area due t o l i f t . The A E ( t ) curve may be approximated by a s e r i e s of parabolic a r c s having a first derivative composed of connected straight - line seg - ments and a second derivative composed of a step or pulse function.

The i n t e g r a l involved i n the function can be evaluated quite F(T) e a s i l y when A i ( t ) i s a constant; by superposition, a complete F curve may be b u i l t up corresponding t o the pulse distribution.

A $ ( t ) A n integration of the F(T) function t o t h e point To (cross - hatched area i n f i g . 2) i s then used i n evaluating the right - hand side of , I equation (1). The degree of approximation of the AE(t) curve can be improved by increasing the number of pulses used. A machine computing procedure using this technique is described in the appendix.

When the computational procedure is carried out for a series of values of lift coefficient, the results may be conveniently plotted in the parametric form used in figure 3 to illustrate the correlation of theoretical estimates with measured wind - tunnel overpressures for a An overpressure parameter derived from theory bomber model (ref. 3 ) .

has been plotted as a function of a lift (or airplane weight) parameter.

The area distribution used in the theory includes the estimated cross - sectional area under the displacement thickness of the laminar boundary layer assumed to exist on the small model. In figure 3 , the variation of the measured pressure - rise parameter with lift coefficient closely follows the theoretical estimate. As a matter of interest, theoretical curves are shown for volume effects alone or lift effects alone. A notable feature of the data shown here is that through a favorable com - bination of area and lift distributions the resultant overpressure becomes less than that associated with lift alone. The uniform atmos - phere results shown here may be applied in making estimates of ground overpressures in the real atmosphere provided that proper account is taken of the pressure and temperature variation between the airplane and the ground. This may be accomplished by replacing the ambient pres - sure p by an appropriate reference pressure. A rigorous treatment of atmospheric effects has been made in reference 4. A reference pressure

based on the method of reference 4 has been used in the correlation of

estimated ground overpressure with flight - test results for the bomber (ref. 5) shown in figure 4 .

A reflection factor of 2 . 0 corresponding to that measured for the dry lake bed over which flight tests were conducted has been used in the estimates. Estimates based on tunnel data are made with differ - ences in area distributions due to estimated differences in the model and airplane boundary layer being considered. Both the theory and the tunnel estimates are shown as a band to account for variations in air - plane weight and Mach number at a given altitude. The flight data

points represent an average pressure for the 6 or 7 ground track micro -

phones in use during a single overpass. The agreement of measured and ,estimated overpressures is reasonably good, there being a slight tend - ency toward underestimation.

The theory of references 1 and 2 as further developed in refer -

ences 6 and 7 has made possible the definition of a lower bound of

Figure 5 illustrates some of these lower sonic - boom overpressure.

bound concepts. As shown previously, sonic - boom strength depends on an effective area distribution combining both volume and lift components.

A n example of an effective - area - distribution curve for an arrow - wing Y i - 8 , a 1 , ?) B 9’3 3 * transport configuration i s shown a t the r i g h t of t e. The sonic - boom overpressure i s primarily dependent on the value of the effective cross - sectional area a t the base of the airplane which i n t u r n i s d i r e c t l y dependent on the airplane base area (including boundary - layer The over - and engine exhaust e f f e c t s ) and t h e f l i g h t l i f t coefficient.

pressure also depends on the shape of t h e effective - area curve. A t h e o r e t i c a l study ( r e f . 6) has shown t h a t the shape of t h e area curve yielding a minimum sonic boom i s represented by a function i n which the i n t h e area i s proportional t o t h e square root of the distance except immediate neighborhood of t h e airplane nose, as i l l u s t r a t e d by the dashed l i n e . Thus, it i s possible t o define a relationship f o r a lower bound of attainable sonic - boom overpressure t h a t may be shown on the p l o t of overpressure parameter as a function of l i f t parameter a t t h e l e f t of figure 5 . The area below t h e curve represents unattainable \ values of sonic - boom overpressures. This lower l i m i t of attainable i overpressures depends only on the airplane length, weight, and base area (including boundary - layer and engine exhaust e f f e c t s ) and on t h e f l i g h t conditions.

This lower bound concept not only s e t s limits on attainable over- pressures but a l s o suggests design methods of approaching these l i m - Theoretically f o r a selected f l i g h t condition (a design i t i n g values.

point), it should be possible t o redesign a configuration t o approach the sonic - boom minimization requirements. I n the example shown, the modification consisted of a t a i l o r i n g of the fuselage area distribution a s shown by the dashed l i n e . I n t h i s example, the design point repre - sents the transonic acceleration portion of the f l i g h t f o r a 300,000 - pound airplane (M = 1.4, The overpressure h = 35,000 f t ) .

characteristics of t h e o r i g i n a l and modified configuration a r e shown on the l e f t of figure 5. The maximum t h e o r e t i c a l reduction i n boom strength (about 25 percent) occurs a t the design point and benefits f a l l off rapidly on e i t h e r side of t h a t point. Whether any substantial portion of these benefits could be achieved i n practice and whether the compromises with other airplane design considerations would be p r o f i t - able i s yet t o be demonstrated.

Some experimental data believed t o be applicable i n connection with these concepts a r e shown i n figure 6. Measured and t h e o r e t i c a l overpressures i n parametric form have been p l o t t e d against a l i f t parameter f o r two wing-boQ models. The model with t h e wing i n the r e a r location t h e o r e t i c a l l y approaches the lower bound even though it was not designed s t r i c t l y i n accordance with the concepts previously discussed. The experimental data however show only a p a r t of the t h e o r e t i c a l gains. Some of t h i s discrepancy may be due t o unknown boundary - layer and separated f l o w e f f e c t s on t h e small models.

O f fundamental importance i n any evaluation of configuration - boom reduction i s the resultant change i n air - changes aimed a t sonic plane drag. Figure 7 shows t h e o r e t i c a l sonic - boom characteristics and corresponding values of z e r o - l i f t wave drag f o r an arrow - wing transport . The o r i g i n a l configuration has configuration and two modifications.

reasonably good sonic - boom characteristics and large reductions i n over - pressures should not be expected. Drag calculations were made by using the machine computing procedure discussed i n paper no. 6 by Roy V .

Harris, Jr. The modifications consisted only of changes i n fuselage area distribution. The second configuration which was modified t o approach t h e sonic - boom lower bound f o r a design point of a t M = 3 an a l t i t u d e of 60,000 feet f o r a 300,000-p0~d airplane required a greatly enlarged forward fuselage with a resultant t o t a l airplane vol - i ume increase of 60 percent. This increase i n volume resulted i n an extremely large z e r o - l i f t drag penalty and a l s o showed up a s an increase i n overpressure f o r zero l i f t . The large drag penalty associated with the attainment of near lower bound sonic - boom values i s not the contra - diction it might outwardly appear t o be. The explanation l i e s i n the f a c t t h a t the drag i s dependent on t h e shock strength of the f l o w f i e l d t o a l l sides of the airplane, not j u s t below, a s i n the case of the boom Although the lower bound configuration r e s u l t s i n reduced ground over - pressures, it probably r e s u l t s i n an increased shock strength i n the remainder of the airplane flow f i e l d p a r t i c u l a r l y above the airplane.

The t h i r d configuration represents a compromise design having no vol - ume change, i n which an attempt w a s made t o produce a smooth effective - area - distribution curve (similar t o that of t h e area distribution f o r a minimum - wave - drag body of revolution) a t the design point of M = 3 a t an a l t i t u d e of 60,000 feet. The decrease i n overpressure which was not as pronounced as that f o r the other modification extended over the whole range of l i f t coefficients and only a small drag penalty was shown.

There are, of course, s o many other factors involved (including the important matter of airplane balance) which have not been studied t h a t it i s not possible a t t h i s time t o d r a w any conclusions regarding the f e a s i b i l i t y of these sonic - boom minimization concepts.

I n order t o provide an i l l u s t r a t i o n of configuration e f f e c t s on sonic - boom characteristics, an i n t e r e s t i n g comparison f o r two transport configuration models i s shown i n figure 8. Both t h e o r e t i c a l and experi - mental wind - tunnel data a r e shown and a r e compared with a lower bound i curve f o r which the base area i s a minimum which includes only t h e A b area of t h e model s t i n g support.

Cross - sectional areas used i n the theory include the area within the estimated displacement thickness of a laminar boundary layer. The lower overpressures f o r the arrow - wing design may be a t t r i b u t e d t o t h e reduced base area and t o smooth area and l i f t distribution curves. I n order t o assess these configuration e f f e c t s i n terms of chmges i n Ap, estimated ground overpressures for these configurations sized t o accommodate 125 passengers (take - off weight of 400,000 pounds) a r e shown i n figure 9 and a r e compared with a lower bound f o r which Ab included only t h e estimated engine exhaust A reflection f a c t o r of 1 . 9 stream tube area less i n l e t capture area.

The reference pressure w a s taken as the standard atmosphere w a s assumed.

pressure a t mid - altitude which f o r a standard atmosphere i s a reasonable

approximation t o t h e r e s u l t s of reference 4 . (Mid - altitude i s t h e a l t i -

tude halfway between the ground and the airplane.) A weight of 360,000 pounds a t a Mach number of 1.4 w a s chosen t o represent the c r i t i c a l climb portion of the flight, whereas a weight of 3OO,OOO pounds a t a Mach number of 3 was chosen t o represent the cruise portion. For M = 1.4, t h e an assumed overpressure l i m i t of 2.0 lb/sq f t i n climb at arrow - wing configuration could f l y at an a l t i t u d e of about 48,000 f e e t whereas t h e canard configuration would be r e s t r i c t e d t o a l t i t u d e s above 37,000 f e e t . For a representative cruise a l t i t u d e of 70,000 feet the canard design would have an overpressure of about 1.6 lb/sq f t as com - pared with 1.4 lb/sq f t f o r t h e arrow - wing design. I n view of t h e many compromises involved it i s believed t h a t a p r a c t i c a l sonic - boom opti - mized airplane would not be able t o achieve t h e corresponding lower bound value of 1.15 lb/sq f t and t h a t attainable minimum overpres - sures a r e l i k e l y t o be nearer t o t h e 1.4 lb/sq f t level. Any change i n airplane weight would, of course, c r i t i c a l l y affect t h i s comparison, of t h e t h e overpressure being roughly proportional t o t h e square root weight.

CONCLUDING REMARKS Experimental data from wind - tunnel and f l i g h t tests have indicated t h a t existing theory can provide reasonably accurate estimates of sonic - boom overpressure f o r the steady - state condition. Additional develop - ments of the theory have made possible t h e definition of a lower bound of sonic - boom overpressure which depends only on airplane length, weight, and base area (including boundary - layer and engine exhaust e f f e c t s ) and on the f l i g h t conditions.

The lower bound concept a l s o suggests airplane design methods t h a t t h e o r e t i c a l l y allow lower bound overpressure values t o be achieved. However, it has been indicated t h a t compromises with other design considerations w i l l prevent anything more than a limited approach t o lower bound overpressures.

I /

APPENDIX A NUMERICAL EVALUATION O F SONIC - BOOM THEORY FOR USE ON HIGH-SPEZD ELECTRONIC COMPUTING MACHINES The effective nondimensionalized cross - sectional area discussed previously and i l l u s t r a t e d i n figure 2 may be expressed as: i' L J If it is assumed t h a t l o c a l l i f t i n g pressures a r e d i r e c t l y proportional t o the t o t a l l i f t coefficient, the quantity i n brackets i n the above equation i s a constant f o r any airplane s t a t i o n . The inputs t o the program a r e thus a tabulation of A ( t ) and t h e r a t i o B ( t ) / B ( l ) as a Equa - function of equally spaced nondimensionalized airplane stations.

t i o n ( A l ) then allows an evaluation and tabulation of the effective area distribution A E ( t ) f o r selected values of t h e l i f t parameter

P S

A series of parabolas f i t t e d t o these points s o t h a t the - CL -.

2 .2 I resultant curve has no discontinuities i n slope may be expressed as: (0 < t < A t ) (At < t < 24t) t 2

(mt < t < Xt)

((n - 1 ) A t < t < nAt)

The second derivative then i s (0 < t < A t > (At < t < 2At) (at < t < Xt) r=n-1 1t 2

((n - 1 ) A t < t < dt))

= - AE,n-

I) -2>-Ai ,r

At2 r=l (A21 The F(T) function then becomes

~ ( 7 ) = - 1 A ~ , ~ F

( 0 < T < A t ) 3t i

((n - 1 ) A t < T < nAt)

l and the i n t e g r a l of the function may be written as F(T) Because of the deviations introduced by the curve - fitting technique used, it has been found t h a t improved accuracy results when t h e i n t e - gral expression is averaged and is rewritten as The overpressure i s then found by selecting the maximum value of equation (A3) and substituting i n t h e following equation: J The computational program may be summarized as follows. The inputs a r e a tabulation of nondimensionalized area and t h e inte - A ( t ) grated l i f t d i s t r i b u t i o n r a t i o B ( t ) / 3 ( 1 ) as a function of equally spaced nondimensionalized airplane stations. Equation (Al) then a l l o w s t h e determination of a table of effective cross - sectional areas f o r preselected values of t h e lift parameter p - CL 12. The second deriva- t i v e s t e p function i s then generated by using equation (A2). Tabulated values of A " ( t ) a r e used i n evaluating t h e i n t e g r a l of equation ( A 3 ) , E t h e maximum value of t h i s i n t e g r a l being selected and used i n equa - t i o n (Ab) i n evaluating the configuration sonic - boom characteristics.

For airplane configurations employing camber t h e loading d i s t r i b u t i o n at zero l i f t may be taken i n t o account by using a modified area dis-

t r i b u t i o n . I n t h i s case - A i s replaced by A +

12 12 2q22 where is t h e l i f t i n g force per unit length at zero l i f t .

Fi i REFERENCES 1. Whitham, G. B.: The Behaviour of Supersonic Flow Past a Body of Revolution, Far From the Axis. , Proc. Roy. Soc. (London), ser. A, vol. 201, no. 1064, Mar. 7, 1950, pp. 89-1-09.

2. Walkden, F.: The Shock Pattern of a Wing - Body Combination, Far From the Flight Path. Aero. Quarterly, vol. I X , p t . 2, May 1958, pp. 164-194.

3. Carlson, Harry W.: Wind - Tunnel Measurements of the Sonic - Boom Characteristics of a Supersonic Bomber Model and a Correlation With Flight - Test Ground Measurements. NASA T M X-700, 1962.

4. Friedman, Manfred P., Kane, Edward J., and Sigalla, Armand: Effects of Atmosphere and Aircraft Motion on the Location and Intensity of a Sonic Boom. A I A A Jour., vol. 1, no. 6, June 1963,

PP - 1327-1335

Maglieri, Domenic J., Huckel, Vera, and Hilton, 5. Hubbard, Harvey H., Measurements of Sonic - Boom Pressures f o r the David A. : Ground Altitude Range of 10,000 t o 75,000 Feet. NASA T M X - 633, 1962.

6. Jones, L. B.: Lower Bounds f o r Sonic Bangs. Jour. R.A.S. (Tech.

Notes), vol. 65, no. 606, June 1961, pp. 433-436.

7. Carlson, Harry W.: The Lower Bound of Attainable Sonic - Boom Over - pressure and Design Methods of Approaching This L i m i t . NASA TN D - 1494, 1962.

a AIRPLANE SHOCK FIELD Figure 1 THEORETICAL ESTIMATION METHOD d t i Figure 2 SON IC - BOOM CHARACTERISTICS BOMBER AIRPLANE; M = 2 EXPERIMENT

(9) (.Y

0 7 ' 5 0 h MAX KrIP4 THEORY I /' ----- LIFT ONLY COMBlNATlON

- - VOLUME ONLY

0 .01 .02 .03 Figure 3 CORRELATION OF TUNNEL DATA, FLIGHT DATA, AND THEORY

BOMBERAIRPLANE; M = 1.5 - 2.0

%/////A TUNNEL 0 FLIGHT T 'ESTS NASA USAF

! I

I I I I L-d 0 30 40 50 60 70 ALTITUDE, FT Figure 4 , i > I 8 ) > LOWER BOUND CONCEPT ORIGINAL CONFIGURATION

---- MODIFIED CONFIGURATION

7 -

.I2 . O 8

($E) (+TI4

t M A X Kr .04 I I I 0 .o I .02 .03 Figure 5 EXPERIMENTAL CONFIGURATION STUDY M= 2 E X P - e - - +

( . ) ( + -

M A X Kr B 'I4 i 0 .004 .008 ,012 .016 .020 Figure 6 A I > * * c e?'

OVERPRESSURE-~RA% RELATIONSHIP

BOOM OPTIMIZED CONFIGURATIONS

*-===a*

r'f3; h-60,000 FT

.I2

- - - _ _ _ - -

.08 .O 4 0 ' 1 . 0 1.5 2.0 2 . 5 3.0 M Figure 7 SONIC - BOOM CHARACTERISTICS TRANSPORT CONFIGURATIONS; M=1.4AND2 THEORY EXP a & .16r

-e

.. . . _ _ I , . . . .

. . . . . . .

I 1 0 .O I .02 .O 3 Figure 8 ESTIMATED GROUND OVERPRESSURES TRANSPORT CONFIGURATIONS CLIMB CRUISE M=l.4; W=36O,OOOLB M=3; W=300,000 LB

4 - L

0 3 0 . ~ 5 40 50 60X103 0 do $0 $0 8(OX103 1 .

ALTITUDE, FT Figure 9 . I '.r - 'i i 22. FACTORS AFFEZTING COMMUNITY ACCEPTANCE O F THE SONIC BOOM By Harvey H. Hubbard and Domenic J. Maglieri

y m fi

- -

ASAILangley Research Center SUMMARY //96Y Discussions are given of ground overpressure exposures f o r current m i l i t a r y operations and how these exposures are affected by the atmos - phere and by a i r c r a f t maneuvers.

B r i e f remarks are included about var - ious operations f o r which some response information has been obtained.

?

Indications are t h a t building response i s a very important f a c t o r i n influencing public reaction. Based on these data, some estimates a r e given regarding t h e frequency and nature of damage incidents resulting from sonic - boom exposures.

P UTW3BL

% -

INTRODUCTION One of the important questions r e l a t i n g t o the operation of the supersonic transport i s whether o r not i t s sonic booms w i l l be accept - able over inhabited areas. This question has l e d t o the j o i n t collection of pertinent data by the National Aeronautics and Space Administration, the United States Air Force, and the Federal Aviation Agency during several supersonic - flight research programs. Although these studies have not produced the f i n a l answers, they have markedly broadened knowl - edge o f t h e subject both with regard t o the overpressure patterns t o which communities a r e exposed and with regard t o the complex manner i n which the communities react t o these e&sures.

It i s the purpose of t h i s paper t o summarize available information on both aspects of the problem.

Figure 1 illustrates schematically a cross - country f l i g h t .

Altitude i s indicated on t h e v e r t i c a l scale and distance, on the horizontal scale; several other items of i n t e r e s t a r e a l s o noted.

The first discussion r e l a t e s t o the sonic - boom overpressure values that have recently been experienced and how they may have been affected by atmospheric phenomena and by the operation of the a i r c r a f t , including maneuvers.

The second topic, which constitutes the main portion of the paper, r e l a t e s t o com - munity reaction induced by sonic boom.

GROUND OVERPRESSURE EXPOSURE3 Atmospheric Effects One of the effects of the atmosphere on the propagation of sonic - boom waveforms can be illustrated with figure 2 . Shown schematically in this figure is a theoretical curve of overpressure Ap as a function of airplane altitude; the shaded region extending above and below the theo - retical curve indicates the normal spread of measured data. Also shown in the figure are three sample sonic - boom pressure waveforms that are representative of many that have been measured. (See refs. 1 and 2.)

It can be seen that these waveforms differ widely with regard to detailed shape, and thus the associated overpressure values as defined in the fig - ure will also vary. Such differences as these, which follow an orderly statistical pattern, are believed to result from the effects of sound- speed gradients associated with atmospheric turbulence in the lower layers of the atmosphere.

There is a similar effect due to large - scale temperature and wind gradients which takes the form of overpressure buildups or changes in the ground pressure distributions. The basic nature of such an effect, namely, that of a sound - speed inversion, is demonstrated in figure 3 .

This figure illustrates schematically the shapes of the shock fronts for propagation in two different atmospheric conditions. The standard atmospheric condition for zero wind is illustrated in the left - hand sketch. Waves propagating to the ground in the presence of a standard sound - speed gradient have a slight cwyature due to the increasing value of sound speed in the lower levels of the atmosphere. The case of a sound - $peed inversion due either to temperature or wind effects results in a different wave - pattern diagram, as indicated in the sketch on the right. In this case, the extremities of the wave propagate faster than the middle portion with the result that the ray paths tend to be squeezed together, thus causing a local increase in the shock - wave pressure. This phenomenon is presently under study by Dr. Manfred P. Friedman at the Massachusetts Institute of Technology, according to the method of refer - ence 3 . His initial results suggest that a sound - speed inversion at a relatively low altitude can have a marked detrimental effect on the pres - I sure patterns. If a similar sound - speed inversion exists at relatively high altitudes, the effect is negligible since the disturbances at high altitudes are compensated f o r by the wave as it propagates. At the lower altitudes, the mechanism for such an adjustment does not have time to act effectively. One of the implications here is that weather fore - casting for sonic - boom purposes will more than likely be required only at the lower altitudes.

The Superboom Phenomena Certain maneuvers of an aircraft in which longitudinal, lateral, or normal accelerations occur can result in so - called " superbooms. " (See

This latter phenomenon can be illustrated with the aid of fig - ref. 4 . )

ure 4 which shows the shock - wave grou&-intersection patterns for two

flight conditions of an aircraft. For simplicity, a homogeneous atmos - phere is assumed, and only the bow shock wave is considered. In the left diagram is indicated the ground pattern for an aircraft in steady flight.

It can be seen that the ray paths, represented by the straight lines, are generally parallel to each other, and the ground pattern, represented by the heavy curved line, is essentially hyperbolic in shape. On the other hand, the pattern on the right is for an aircraft experiencing a lateral It can be seen that the ray paths are no longer parallel. I acceleration.

In some regions they tend to converge and in others, to diverge. Like - wise, the shock - wave ground - intersection pattern is no longer symmetrical and may contain some irregularities and cusp formations in which the pressures are higher than those for the steady - flight condition. Such pressure buildups are referred to as superbooms. The case for lateral acceleration is shown only for convenience. Similar results would be obtained for flight conditions involving normal or longitudinal accel - One important consideration is the shape of these superboom erations.

These areas are shown in figure 5 for some common areas on the ground.

flight maneuvers.

In the upper part of the figure is illustrated a level, or linear, acceleration. Linear acceleration is a common maneuver which must be As indicated in the executed during every flight to supersonic speeds.

sketch by the stippled areas, superbooms occur over relatively small expanses of the ground. Dimensions are such that total superboom area is approximately 1 square mile. The pressure buildups in these stippled areas are believed to be a function of the rate of acceleration of the aircraft but, for a practical operating range, are approximately 2 times the corresponding steady - flight values. Also of possible concern in the operation of the supersonic transport are such maneuvers as horizontal turns and push - overs, as might occur during changes in course and air - plane attitude. In the latter instances the ground patterns of pressure 5, and, because buildups are different in shape, as indicated in figure of the higher accelerations involved, the buildup factors may tend to be higher and the areas smaller than f o r the case of linear acceleration.

Summary of Recent Sonic - Boom Overpressure Experience Before some of the effects induced by sonic boom are discussed, it will be helpful to become familiar with current sonic - boom exposures as indicated in figure 6. The ground overpressures due to routine training operations involving two types of service a i r c r a f t are shown as a func - t i o n of a i r c r a f t a l t i t u d e . It should be noted t h a t such (See ref. 2.)

operations are limited t o a l t i t u d e s above 30,000 f e e t but have been carried on over many metropolitan areas i n the United States. The stip - pled region r e l a t e s t o f i g h t e r a i r c r a f t f o r steady - flight operations, and it can be seen that, depending on t h e a l t i t u d e of t h e operation, the associated overpressure range varies from less than 0.3 lb/sq f t t o about 1.8 lb/sq f t . The crosshatched region relates ta B- 58 operations.

It can be seen t h a t the overpressure range associated with these opera - t i o n s varies from about 1 . 3 lb/sq f t t o nearly 3 lb/sq f t , again depending on t h e a l t i t u d e . The hatched region i s not w e l l defined but has been estimated, based on fragmentary experimental data, t o encompass t h e superboom overpressure range associated with training maneuvers of both f i g h t e r and bomber a i r c r a f t .

? J Estimated values of overpressure f o r t h e supersonic transport a r e within the range of current experience f o r f i g h t e r and bomber a i r c r a f t .

Three differences are, however, anticipated between t h e service ty-pe of operations and the supersonic - transport operations. One of these dif - ferences i s t h a t the transport operations w i l l be of much longer dura - tion; t h a t is, they w i l l sweep over more area, t h e frequency of exposure w i l l eventually be greater, and, furthermore, t h e associated wavelength w i l l be about 2.0 times as long as f o r the 13 - 58 and about 4.0 times as long as for the f i g h t e r - a i r c r a f t wavelengths t h a t are indicated i n the figure .

A recording of the pressure time history of thunder i s i l l u s t r a t e d i n figure 6 for comparison. The peak pressures involved a r e approxi - mately 0.5 lb/sq f t f o r lightning discharges located a t distances of about 1 mile from the observer. It i s believed, of course, that higher overpressures would be associated with closer s t r i k e s , Although sonic booms a r e often judged t o sound l i k e thunder, it i s obvious t h a t the pressure t i m e history of thunder consists of many peaks within a rather short time i n t e r v a l and, hence, i s markedly different from t h e pressure time h i s t o r i e s of the sonic booms.

Wave shape and wavelength may be significant influences i n response, although they have not as yet been evaluated.

C O K ~ T I T Y RESPONSE TO SONIC BOOMS Results of Judgment Studies It i s very d i f f i c u l t t o devise an experiment of a subjective nature t h a t w i l l give a l l the desired answers.

Some of these answers w i l l only come as a result of simulated o r actual supersonic - transport operations. During t h e past few years, however, several attempts have fl - y * * c s been made i n conjunction with &ystcal" n&sur;ments t o collect sub jec- t i v e data. The following t a b l e i s a list of the available studies during which some subjective observations were obtained.

(See refs. 2 and 5 t o 9 . ) overpressure, It can be seen t h a t observations have been made f o r r e l a t i v e l y large numbers of f l i g h t s and f o r observers ranging from small groups t o e n t i r e a wide range of overpressure values populations of large c i t i e s ; also, vas involved i n these studies. It i s i n t e r e s t i n g t o note that f o r study A there w a s very l i t t l e , if any, recorded reaction t o the por - t i o n represented i n the table. There were numerous complaints, how - ever, associated with another portion of the mission f o r which the pressures w e r e considerably higher. It i s a matter of record t h a t some scheduled supersonic f l i g h t s prompted complaints and claims of damage even though the f l i g h t s were never accomplished.

Where s m a l l groups were involved, judgments were s o l i c i t e d regarding acceptability. There were, of course, inherent differences of opinion of the observers, and the test conditions differed i n many respects. The so - called acceptable range of overpressures varied from This range i s significant ?out 0.9 lb/sq f t t o about 1.9 lb/sq f t .

d n c e it includes t h e overpressure range estimated f o r the operation of the supersonic transport.

The f a c t that s c a t t e r e x i s t s i n these and other judgment data i s not surprising since it i s believed t h a t several d i f f e r e n t f a c t o r s can a f f e c t these judgments. Some of these factors relate t o the physical c h a r a c t e r i s t i c s of t h e pressure wave, such as i t s overpressure magni-.

tude, its wavelength, and i t s detailed Shape. Other f a c t o r s which a r e believed t o be significant, based on sociological considerations, are the time of day of the observed *boom, the type of operation causing it, the type of community i n which the observer lives, the observer's per - sonal environment, and the frequency of supersonic flights.

Summary of Reported Effects In one instance there w a s actually an opportunitysto interview representative segments of the population of a large metropolitan area i n order t o get a d i r e c t sampling of reaction t o sonic booms. The r e s u l t s of these interview studies are summarized i n figure 7. People representing about 1,100 homes were interviewed t o determine t h e i r reactions t o 64 known sonic booms over a period of several months. O f t h e 1,100 interviewed, about 90 percent indicated an awareness of the sonic booms t o the extent that they discussed various effects.

For instance, the most commonly mentioned e f f e c t w a s the vibration of building structures. About 40 percent of those interviewed indicated some concern over possible detrimental e f f e c t s of such vibrations.

Those who said they actually went t o the trouble t o f i l e a formal com - p l a i n t were 5 i n number and thus constituted roughly 0.5 percent of the t o t a l . The types of incidents reported as a basis f o r such formal con - p l a i n t s from over 3,000 cases i n A i r Force f i l e s have been broken dom into several categories and a r e plotted i n figure 8 i n the form of bar graphs t o indicate frequency of occurrence.

For instance, p l a s t e r cracks were reported most frequently and, i n fact, w e r e mentioned i n 43 percent of the complaints registered.

Other damage, such as cracks i n window glass, w a l l s , t i l e , and so forth, w a s reported t o have occurred l e s s frequently. It is believed significant that a large percentage of the complaint reports mentioned some type of damage, whereas only about 7 percent mentioned miscellaneous effects, including annoyance only. The implication i s that the a b i l i t y of the sonic boom t o cause buildings t o vibrate i s very significant since it is readily observable; it i s suggestive that damage can occur, and such damage is, of course, of concern t o the general public.

Because of the obvious significance of building vibration, a test w a s made i n which the sonic - boom - induced stresses i n several components of a flat - roofed frame building were measured.

In figure 9, as an I example of the type of data obtained, r a f t e r stresses r~ are plotted as a function of sonic - boom overpressure Ap f o r a range from about 0.3 t o 3 lb/sq f t . The measured stresses did not exceed 20 lb/sq in. and were noted t o be of t h e same order of magnitude as those associated w i t h such everyday occurrences as door closing. Door slamming would result i n values corresponding t o the highest measured values of the figure.

The sonic - boom - induced s t r e s s values were also noted t o be r e l a t i v e l y low compared w i t h t h e design stresses of the building. The implication i s t h a t i f a building was w e l l designed and i n good repair, no damage would be anticipated because of t h e r e l a t i v e l y small loads induced by sonic booms.

P r o b a b i l i w of Barnag% *

I , . -

O n t h e b a s i s of t h e results of a r c h i t e c t u r a l and engineering surveys i n areas where sonic - boom damage w a s reported and of the knowledge of t h e exposure patterns and the population d e n s i t i e s of the areas involved, there is a strong suggestion t h a t the probability of damage i s related t o population density. The data of figure 10 have been determined on t h e assumption t h a t t h e average occupancy rate i s f i v e persons per building. Plotted on the v e r t i c a l scale i s the accumulative probability of damage occurrences per million buildings shown as a function of the overpressure on the horizontal scale.

The data are presented i n the form of shaded bars. The four lower bars represent routine training type of operations, whereas the two upper bars represent isolated accidental - damage incidents. The calculated probability curve i s adjusted t o a best f i t of available data. Several features of t h i s curve are worthy of mention. The number of damage incidents, of course, decreases as the overpressure value decreases, but even at very low pressures there still seems t o be t h e p o s s i b i l i t y of damage occurrences.

One explanation f o r t h i s i s {hat t h e investigators were lenient i n t h e i r judgments against the complainants. I n other words, some of the damage incidents which were judged v a l i d may i n f a c t not have been valid.

There i s always the p o s s i b i l i t y of other contributing factors i n any damage occurrence. These are, f o r instance, weathering effects, vibra - t i o n due t o road t r a f f i c , degradation of building materials, s e t t l i n g of t h e building, poor workmanship, and so forth. All these factors can contribute t o the same type of damage a t t r i b u t e d t o sonic booms, and hence t h e job of honestly evaluating claims i s a d i f f i c u l t one. Con - sequently, it i s believed t h a t the numbers quoted i n figure 10 are con - servative; t h a t is, they a r e no doubt higher than the actual number of damage incidents. It is obvious t h a t additional studies a r e needed f o r t h e purpose of b e t t e r defining the values of the lower portion of the curve and p a r t i c u l a r l y f o r repeated exposures over a period of time.

Even though there i s reason t o suspect the absolute value of t h e / - numbers quoted i n figure 10, the trends indicated may be very signifi i s e s s e n t i a l l y pro - cant. For instance, t h e number of damage incidents portional t o the overpressure i n t h e overpressure range of greatest i n t e r e s t . This f a c t would suggest t h a t even a sizable increase i n overpressure should not result i n an unusually large increase i n the damage occurrences.

There a r e s t i l l some differences of opinion regarding acceptable sonic - boom exposure. Ekperience with military a i r c r a f t has been i n the range of overpressures of i n t e r e s t but has not otherwise been d e f i n i t i v e enough f o r making a quantitative evaluation of the problem. Recent experience has been very useful q u a l i t a t i v e l y and indicates that a major f a c t o r i n shaping a t t i t u d e s toward sonic booms i s the matter of building vibrations. Only occasionally does building vibration due t o sonic booms result i n damage, and ensuing damage is, i n turn, limited i n extent. There does not, however, seem t o be any assurance of completely eliminating damage, and hence t h i s damage problem w i l l be a continuing one.

FUZFEFUZNCES 1 . Maglieri, Domenic J., and Parrott, Tony L . : Atmospheric Effects on Sonic - Boom Pressure Signatures.. Sound, vol. 2, no. 4, July - Aug.

1963, pp. 11 - 14.

2 . Hubbard, Harvey H., Maglieri, Domenic J., Huckel, Vera, and Hilton, David A.: Ground Measurements of Sonic - Boom Pressures for the Altitude Range of 10,000 to 75,000 Feet. NASA T M X - 633, 1962.

Friedman, Manfred P . , Kane, Edward J., and Sigalla, Armand: Effects 3 .

of Atmosphere and Aircraft Motion on the Location and Intensity of \ i a Sonic Boom. AIAA Jour., vol. 1 , no. 6, June 1963, pp. 1327-1335.

4 . Maglieri, Domenic J., and Lansing, Donald L . : Sonic Booms From

Aircraft in Maneuvers. Sound, vol. 2, no. 2, Mar. - Apr. 1963, PP. 39 - 42.

5. Maglieri, Domenic J., and Carlson, Harry W . : The Shock - Wave Noise Problem of Supersonic Aircraft in Steady Flight. NASA M f 3 M o 3-4-59b 1959- 6. Maglieri, Domenic J., Hubbard, Harvey H . , and Lansing, Donald L . : Ground Measurements of the Shock - Wave Noise From Airplanes in Level Flight at Mach Numbers to 1.4 and at Altitudes to 45,000 Feet.

NASA TN D- 48, 1959.

7 . Lina, Lindsay J., and Maglieri, Domenic J . : Ground Measurements of

Airplane Shock - Wave Noise at Mach Numbers to 2 . 0 and at Altitudes to 60,000 Feet. NASA TN D-235, 1960.

8. Power, J. K. : Sonic Boom & Community Relations.

k p r i n g 683~, SOC. Automotive Eng., Apr. 1963.

9. Maglieri, Domenic J., Huckel, Vera, and Parrott, Tony L . : Ground Measurements of Shock - Wave Pressure for Fighter Airplanes Flying at Very Low Altitudes and Comments on Associated Response Phe - , nomena. NASA '151 x - 611, 1961.

OPERATING CONSIDERATIONS SONIC BOOM MANEUVERS ALTITUDE - = WEATHER

e

A D I S T A N C E Figure 1 METEOROLOGICAL EFFECTS EXAMPLES OF MEASURED WAVEFORMS I r a I AIRPLANE ALTITUDE Figure 2 EFFECT OF SOUND - SPEED GRADIENT STANDARD INVERSION \ ALTITUDE SOUND SPEED SOUND SPEED Figure 3 GROUND PRESSURE PATTERNS

*--

+% - , STEADY FLIGHT ACCELERATED FLIGHT Figure 4 Figure 5 SONIC - BOOM EXPOSURE LEVELS ROUTINE OPERATIONS 6 - TRAl NlNG b -

/- MANEUVERS

h.

4- - *P, LB/FT 2 - - L - 0 20 40 60 x IO3 ALTITUDE, F T Figure 6 RESULTS O F INTERVIEW STUDIES 64 FLIGHTS 1,100 INTERVIEWED EXPRESSED CONCERN FILED FORMAL COMPLAlNTS

1 I I I 1

0 25 50 75 100 PERCENT OF TOTAL INTERVIEWED Figure 7 BREAKDOWN OF COMPLAINTS RECORDED BROKEN WINDOWS BROKEN BRIC - A - BRAC DAMAGED APPLIANCES MISCELLANEOUS 0 IO 20 30 40 50 PERCENT OF TOTAL COMPLAI NTS Figure 8 a RAFTER STRESSES

30r

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l.1 .3 I 3 I O GROUND OVERPRESSURE, Ap, L6/FT2 Figure 9 DAMAGE INCIDENTS PER MILLION BUILDINGS / - - - I: NUMBER IO3 OF - : INCIDENTS102 . - I EXPERIMENT - - - CALCULATED I .o .I+ I I I I I I 0 5 1 0 15 20 25 30 OVERPRESSURE, A,.,, LB/FT~ Figure 10 ENGINE AND BOUNDARY - LAYER NOISE CONSIDERATIONS

By Jimmy M. Cawthorn and Domenic J. Maglieri & '

esearch Center S m Y The main findings of the contractor studies are summarized with regard to noise problems during the take - off, initial climb, and landing approach.

Data are presented in the form of perceived noise levels, and ; comparisons are made with a current subsonic transport of comparable range. Airport noise levels are estimated to be higher than those for the current aircraft, whereas the climbout and landing - approach noise levels can be of the same order of magnitude or lower, depending on the aircraft configuration and the operating procedures used.

Brief mention is also made of the noise - induced structural - response problem during take - off and cruise.

w. 4 d7kfcJafl

INTRODUCTION Since the supersonic transport will be integrated into existing air traffic systems, it is required that its acoustic characteristics be compatible with those of current aircraft. This is particularly desir - able for take - off and landing operations during which power - plant noise is an important consideration. Noise in the community due to the power plants is, of course, a function of the type of power plant used and the manner in which it is operated and also, as will be shown in the paper, the configuration of the aircraft in which it is installed.

The objectives of this paper can be discussed with the aid of fig - ure 1 .

Noise considerations during the take - off run, the initial climb - out from the airport, and the landing approach will be discussed.

Some mention will also be made of the structural response problems of the airframe due to noise excitation by the power plants and by the aero - . ' dynamic flows. An attempt is made to include the main findings of the contractors during their feasibility studies.

TYPES OF ENGINE CYCLES First let us consider the significance of the type of engine cycle Noise level data for advanced turbojet (engine C) and turbofan chosen.

engines (engine D) are given i n t a b l e 1 f o r both augmented and unaug - mented t h r u s t conditions. The basic engine noise data presented a r e taken from the Boeing studies and have been normalized t o a t h r u s t of Also shown f o r comparison are similar data f o r a current 45,000 pounds.

turbofan a t l7,OOO-pound thrust. Listed i n t h e second column are the It m a x i m u m sound pressure l e v e l s estimated a t a distance of 1,000 f e e t .

can be seen t h a t the lower overall noise l e v e l s are associated with the turbofan type of power plant. It can also be noted, however, t h a t even the advanced turbofan engine cycles have associated with them higher overall noise levels than do the current turbofan engines. These higher levels f o r both proposed engines a r e due t o the f a c t t h a t they w i l l be larger and have higher exhaust v e l o c i t i e s because of t h e i r greater t h r u s t requirements.

Estimated exhaust noise spectra f o r the engines of table 1 a r e shown i n figure 2. The range of sound pressure levels estimated f o r the advanced turbojet and turbofan engines i s indicated by the two shaded regions. The noise levels associated with the lower extremity of each whereas shaded region correspond t o the unaugmented operating condition, those associated with the upper extremity of each band correspond t o f u l l augmentation. It can be seen that these levels a r e higher a t a l l frequencies than those f o r the current turbofan engine, as i l l u s t r a t e d by t h e dashed curve.

1) i s applied t o the When the perceived noise l e v e l procedure ( r e f .

spectra of figure 2, the perceived noise l e v e l values l i s t e d i n the last Since the shape of the spectra i s not column of table 1 a r e obtained.

markedly different, the perceived noise levels exhibit the same rank order as t h e sound pressure levels. It i s therefore obvious from the data of table 1 and figure 2 t h a t on the basis of minimum noise gen - erated, it would be desirable t o use a turbofan engine cycle rather than a turbojet.

TAKE - OFF AND INITIAL CLIMBOUT The manner i n which the a i r p o r t noise s i t u a t i o n i s affected can be i l l u s t r a t e d by the plan view sketch of figure 3, i n which a r e plotted constant perceived noise l e v e l ___ contours _____ of 110 decibels f o r a current f an - powered subsonic, 2, ?OO-mile-range, intercontinental j e t and f o r a proposed supersonic transport. The origin represents the start of take - off roll. The horizontal scale is, then, the distance p a r a l l e l t o the runway, and the v e r t i c a l scale i s the distance perpendicular t o the run - way. The s o l i d l i n e represents t h e perceived noise l e v e l contour of The 110 decibels f o r the subsonic airplane without power cutback.

unbounded shaded area represents t h e range of estimated perceived noise l e v e l contours of 110 decibels f o r SCAT 16 with no power cutback, which came from the two contractor studies. The width of the shading is due to differences in the aircraft configuration involved as well as to the different assumptions which go into making these estimates. It can be seen that the estimated perceived noise level contour of 110 decibels for the supersonic airplane extends laterally to a greater distance than that for the current airplane and would extend further in the direction of the runway unless a power cutback were initiated. Current procedures employ a power cutback which shortens the pattern in the direction of the runway, as indicated by the dashed line. This sane power cutback procedure is judged feasible for the supersonic transport and, as sug - gested by the bounded shaded region, would further shorten the contour in the direction of the runway.

The effect of such a power cutback in the initial climb is further !

illustrated in f i @ e 4 , which contains a summary of the data produced by the two contractors for SCAT 16 and SCAT 17 and a comparison with similar data for a current ftm-powered ‘2,500 - mile - range aircraft. Shown in the top sketch are the altitude - distance profiles, and in the lower sketch are shown the variations of perceived noise levels as a function of distance from brake release. It can be seen that both of the super - sonic aircraft take off in a shorter distance and climb initially at a greater angle, and, as a result, they attain a higher altitude over the community than the subsonic aircraft ddes. It can be seen that SCAT 1 6 attains a more rapid initial climb angle and a resulting higher altitude in the initial climbout than does SCAT 17 because of its inherently better low - speed aerodynamic characteristics. The differences in the performance of the two aircraft also show up in the resulting noise exposures as illustrated in the lower sketch of the figure. The solid line again represents the noise levels of a current subsonic aircraft - and the hatched areas represent the ranges of estimates by the two con tractors for SCAT 1 6 and 17, hot - day operation being assumed. It can be seen that the SCAT 1 6 noise levels are generally lower than those of SCAT 17. The data of the figure seem to indicate, however, that either configuration might have noise levels comparable to or lower than those of the current airplane when operated as in the top sketch of the figure.

In summary of these two plots, it can be seen that the noise levels for the supersonic aircraft will be higher within the confines of the airport and the immediate surrounding areas.

At greater distances there is an indication, however, that either configuration might have noise levels comparable to or lower than those of the current airplane, depending on the operational procedures used.

The state of development of the configurations studied has not pro - gressed to the point where the optimum take - off and climbout procedures have been determined. It should be noted, however, that the supersonic transport aircraft has inherent operational flexibility that would permit a power cutback, as illustrated, or some other altitude - distance profile that might be more attractive.

LANDING APPROACH The noise during landing approach can be discussed.with the aid of figure 5 and involves the geometry of the engine installation and the aircraft operating characteristics. Shown in the figure are perceived noise levels as a function of distance from touchdown for SCAT 1 6 , SCAT 17 (based on the contractor studies), and a comparable current fan - powered subsonic aircraft.

A 3O glide slope has been assumed in all cases. The solid line represents the perceived noise levels on landing - approach for the current subsonic airplane. The upper bounds of each of the shaded areas represent the landing - approach noise levels for SCAT 1 6 and SCAT 17, hot - day operations with no inlet noise suppression being assumed. It can be seen that the SCAT 1 6 noise levels are lower than those of SCAT 1 7 as a result of its better low - speed aerodynamics which permit it to approach at lower engine power settings. However, the SCAT 16 noise levels at best only approach those that are currently experienced on the subsonic aircraft.

The extent of the shading which is shown for the two SCAT configu - rations indicates the variation in perceived noise levels that have been estimated for varying amounts of inlet noise suppression.

These esti - mates of the perceived noise levels for inlet noise suppression based on contractors' past experience varied from 6 to 12 decibels.

The point to be made is that some inlet noise suppression will be required to bring the landing noise levels of the SCAT configurations below those of the current subsonic aircraft.

" he significance of geometry is suggested in figure 6, in which are shown schematic diagrams of a subsonic and a supersonic engine installa - tion. The shading represents the extent of the inlet surface area.

For current subsonic airplanes it is generally agreed that the inlet noise during landing approach is more objectionable than the exhaust noise, and thus accounts for most of the registered complaints.

E As suggested by the shading of the upper sketch, there is only a limited area for application of noise reduction procedures. In the case of the supersonic transport, however, the situation may be considerably differ - ent since there is a greater working area.

In addition, there is the possibility of deliberately choking the flow, and in this regard a con - certed effort is being made in NASA research studies to evaluate inlet noise reductions by means of variable inlet geometry with particular application to the landing - approach condition and associated operating problems.

It is obvious from the data o ? * & & e . - % u o Gigures that the landing- P I * approach noise problem should be considered a% a major design goal. It is believed possible to reduce the landing - approach noise levels through further improvements in the low - speed aerodynamics and through suitable inlet noise - suppression procedures.

STRUCTURAL RESPONSE TO NOISE The noise - induced structural - response problems of the supersonic transport which are important from the standpoint of maintaining accept - able cabin noise levels and minimizing sonic fatigue can be discussed I

with the aid of figure 7 . The shaded areas of the airplane plan view

sketch at the top of the figure represent regions where noise loads may be a design consideration. At the bottom of the figure are sample flow noise and engine noise input spectra estimated for the proposed oper - ating conditions of the airplane. The flow noise loading will exist for nearly the entire duration of the flight. The curve of the left diagram is based on recent NASA free - flight measurements at the appro - For some local priate Mach numbers and Reynolds numbers. (See ref. 2. ) flow separation, surface roughness, o r shock - wave - interaction condi - tions, the levels would be higher as indicated by the shading. For these latter conditions the estimated spectral content (shaded area of fig. 7) which is based on results of recent contract wind - tunnel studies by the Douglas Aircraft Company is similar to that given by the solid line.

The noise from the engines is believed t o be significant for only a short period during each mission, and only the structure in the vicin - ity and to the rear of the engines, as indicated by the cross - hatching in the sketch, will be affected. It can be seen that the estimated spectra peak at lower frequencies and reach higher sound pressure levels than the flow noise spectra. (See ref. 3 . ) Some brief comments with regard to structural design are suggested Although the acoustic loads are more severe by the data of the figure.

than those for current aircraft, the design of structures to withstand these loads is not a brand new problem since similar environments have been encountered in other current operational vehicles. It is believed that engine noise structural response experience to date is directly applicable; however, the flow noise problem has not been satisfactorily defined, particularly for long-term exposures at elevated temperatures.

Both the flow noise inputs and associated structural responses are cur - rently being studied in NASA research programs.

c v CONCLUDING REMARKS The results presented in the paper are based on proposed engine and airplane configurations, and it should be emphasized that no noise data were furnished to the contractors. Because of their present state of development, it is only possible to indicate broad ranges of noise pre - dictions. The higher thrust engine proposed for the supersonic trans - port will result in relatively higher noise levels in the airport con - fines and in the areas immediately adjacent to the airport. Because of its predicted operational flexibility, however, the climbout and landing - approach community noise levels may be lower than those of current air - craft of similar range. The most urgent research problems relate to optimizing low - speed aerodynamic characteristics of the aircraft, mini - mizing inlet noise, and defining the flow noise - induced structural - response problem.

REFERENCES 1. Bolt Beranek and Newman, Inc.: Studies of Noise Characteristics of the Boeing 707-120 Jet.Airliner and of Large Conventional Propeller - Driven Airliners. Prepared for The Port of New York Authority, Oct. 1958.

2. Hilton, David A . , Bracalente, EImedio M . , and Hubbard, Harvey H. : In - Flight Aerodynamic Noise Measurements on a Scout Launch Vehicle.

NASA TN D - 1818, 1963.

3 . Trapp, W. J., and Forney, D. M., Jr., eds.: WADC - University of

Minnesota Conference on Acoustical Fatigue. WADC Tech.

Rep. 59-676, U.S. Air Force, Mar. 1961.

TA€iLE I NOISE FROM VARIOUS ENGINES DISTANCE = 1,000 F T SOUND PERCEIVED ENGINE THRUST, PRESSURE NOISE L B LEVEL, LEVEL, d b db PROPOSED

I TURBOJET,AUGMENTED I 45,000 I 1 2 9 I 1 3 6 I

TURBOJET I 45,000 I 125 I 133 I

TURBOFAN, AUGMENTED 45,000 124 I30 TU R B 0 FAN 45,000 1 1 4 123 c TURBOFAN 17,000 104 I I5 . . .

NOISE PROBLEM AREA$'DUE TO POWER PLANT A N D AIRFLOW

NOISE, STRUCTURAL RESPONSE INITIAL ; CUMB ,' # Figure 1 ENGINE - NOISE SPECTRA I20 I10 SOUND PRESSURE

-Y'' ADVANCED T U R B O F A N - ' '"///////\

LEVEL, 1 0 0 db LPRESENT TURBOFAN

%

70 I I I I I I I J 3 7 . 5 75 150 300 600 1,200 2 , 4 0 0 4,800 TO TO TO TO TO TO TO TO 9 . 6 0 0 75 150 300 600 1 . 2 0 0 2.400 4.800 FREQUENCY B ANDS , CPS Figure 2 NOISE - LEVEL CONTOURS FOR TAKE - OFF PERCEIVED NOISE LEVEL, I IO d b -SCAT 1 6 (POWER REDUCED) -SCAT 1 6 - - - SUBSONIC (POWER REDUCED) SUBSONIC

2r

I I 0 I 2 3 4 5 6 DISTANCE FROM BRAKE RELEASE, MILES Figure 3 NOISE DURING CLlMBOUT ,-POWER REDUCTION 140r M O W E R REDUCTION PERCEIVED 120 NOISE LEVEL, db 1 0 0 LSCAT 16 Figure 4 ,

LANDING - APPROACH NOISE

PERCEIVED NOISE LEVEL, d b 1 1 0 T I I I 0 I 2 3 DISTANCE FROM TOUCHDOWN, MILES Figure 5 INLET GEOMETRY

+ - -t--

SUBSONIC SUPERSONIC Figure 6 L ' I ACOUSTIC INPUTS TO STRUCTURE FLOW NO I SE 170r aFLOW NO1 SE ENGINE NOISE r /--- J B SOUND150 PRESSURE LEVEL, db 1 3 0 ~~ 110, , , , I , I ,

-

37.5 150 600 2,400 37.5 150 600 2,400 TO TO TO TO TO TO TO TO 75 300 1,200 4,800 75 300 1,200 4,800 FREQUENCY BANDS, cps Figure 7

4 ?

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+

-

2 ?

OPERATIONAL FACTORS I N RELATION TO T 6Y ERSONIC-TRANSPORT FUEL REQUIREMENTS

By Joseph W. Wetmore q d Joseph W. Stickle $ z E &-

SUMMARY An analysis of t h e interaction of operational environment and air - ' c r a f t characteristics of the supersonic transport (SST) i n t h e area of f u e l requirements has been made.

The increment i n design range o r fuel - carrying capability above t h a t of t h e basic t r i p distance i s con - sidered i n relation t o t h e e f f e c t s of wind, temperature, anticiDated

I

& f l i g h t - l e v e l assignment, and payload load factor.

Reserve fuel requirements are discussed and a s t a t i s t i c a l approach i s used i n determining en route and holding reserves.

It i s indicated t h a t protection equivalent t o t h a t provided f o r subsonic j e t s by cur - rent reserve requirements i s obtained f o r t h e supersonic transport with less than the present C i v i l A i r Regulation (CAR) en route and holding requirements.

T

INTRODUCTION A complete understanding of a l l the f a c t o r s which w i l l a f f e c t the operations of t h e SST can only be gained from service experience with t h e actual airplane i n i t s own operational environment. Nevertheless, it i s necessary at an early stage t o anticipate and account f o r these operational factors t o t h e f u l l e s t extent possible where they may have an important bearing on the design and ultimate capabilities of the airplane. The purpose of t h i s paper is t o examine the interaction of operational environment and a i r c r a f t characteristics of t h e SST i n the very important area of f u e l requirements.

i , With the use of s t a t i s t i c a l approaches, two aspects of f u e l requirements,are analyzed.

F i r s t , t h e design range o r destination f u e l w i l l be consjdered i n r e l a t i o n t o t h e e f f e c t s of wind, temperature, anticipated f l i g h t - l e v e l assignments, and payload load factors. The other area of consideration i s f u e l reserves i n which en route reserves, accounting f o r various uncertainty factors, as well as f u e l f o r holding and diversion t o a l t e r n a t e a i r p o r t s are discussed.

Fuel allowances f o r en route engine f a i l u r e are related t o t h e reserve needs f o r t h e more normal operational variations.

DESIGN RANGE The design fuel - carrying capability of t h e supersonic transport should provide a rakional margin to.allow f o r l i k e l y variations of those factors which a dispatcher normally accounts f o r i n planning destination fuel. Destination f u e l is considered t o be t h a t f u e l required t o f l y t o and land at t h e destination. I n t h e following dis - cussion, t h e s t a t i s t i c a l e f f e c t s of these f a c t o r s on destination f u e l requirements are presented i n terms of variations i n standard atmos - phere and still - air range r e l a t i v e t o the basic t r i p distance of t h e a i r c r a f t , here taken t o be from New York t o P a r i s or 3,160 nautical m i l e s .

Atmospheric Effects One of t h e primary factors t o be accounted f o r i n determining destination f u e l i s t h e expected atmospheric conditions en route, t h a t is, winds and temperatures. The s t a t i s t i c a l characteristics of the east - west effective wind component and effective temperature difference from international standards ( r e f s . 1 and 2) a r e shown f o r t h e winter season, December through February, and t h e summer season, June through August, i n t a b l e I. I n t h e case of wind, t h e effective means and stand - ard deviations a r e shown separately f o r climb and descent and f o r cruise. The temperature variations from standard are averages f o r the c r i t i c a l climb - and - accelerate range of a l t i t u d e s and a r e assumed t o apply t o the whole f l i g h t since t h e e f f e c t s of temperature are most important i n t h e climb - and - accelerate stage. The mean winds and vari - a b i l i t y a r e considerably greater i n the winter, particularly at cruise altitudes, than i n t h e summer. The summer mean temperatures are higher, as would be expected, but t h e standard deviation i s somewhat l e s s than i n t h e winter.

The increment of still - air range required t o account for the e f f e c t . of wind is, of course, the product of t h e average wind during the f l i g h t and the flight time. Primary consideration is given t o west - bound f l i g h t s t o account f o r t h e adverse e f f e c t s of prevailing head winds.

J

The e f f e c t s of temperature on incremental range required were calculated specifically f o r t h e variable - sweep supersonic - commercial - air - %ransport SCAT 16 configuration with turbofan engine B of the f e a s i b i l i t y studies, sized t o y i e l d representative climb - and - accelerate performance. However, similar e f f e c t s were indicated by limited cal - culations f o r the fixed - delta SCAT 17 configuration powered f o r equiv - alent basic performance. I n general, each degree of temperature f variation above standard required between 6 and 7 nautical m i l e s of additional range.

Both t h e wind and temperature variations have approximately Gaussian, o r normal, distributions. Since t h e effects of wind and temperature on incremental range required are nearly l i n e a r throughout t h e range of conditions of interest, t h e corresponding distributions of incremental range required are also normal. Therefore, t h e combined e f f e c t s of temperature and wind were incorporated i n the probabilities of requiring given range increments by t h e usual methods applicable t o normal frequency distributions.

A comparison i s given i n figure 1 of winter and summer conditions f o r t h e combined e f f e c t s of wind and temperature on t h e probability t h a t t h e increment i n standard-day no - wind range required t o achieve the westbound t r i p with f u l l payload w i l l equal o r exceed given values.

These r e s u l t s indicate t h a t a greater range increment i s required f o r t o compensate f o r pos - winter than f o r summer conditions. For example, s i b l e adverse wind and temperature conditions f o r a l l but 1 percent of winter f l i g h t s , t h e airplane should have a design - range capability about 244 nautical miles greater than the basic t r i p distance. For summer conditions the range increment f o r t h e same probability l e v e l would be 200 nautical miles. The winter condition i s therefore assumed i n the following discussion of t h e e f f e c t s of other factors. The 1 - percent probability l e v e l w i l l a l s o be used i n evaluating these effects, but only f o r purposes of i l l u s t r a t i o n . It i s not within t h e purpose of t h i s analysis t o suggest the probability l e v e l t h a t should ultimately be selected f o r design.

Effect of Flight - Level Assignment Thus far it has been assumed t h a t the airplane would always cruise It is quite likely, however, t h a t t h e supersonic at optimum a l t i t u d e .

transports, as f o r present trans - Atlantic jets, at t i m e s w i l l be assigned f l i g h t l e v e l s other than optimum a l t i t u d e t o provide adequate separation from e i t h e r high - flying military a i r c r a f t o r from other supersonic transports. Figure 2 i l l u s t r a t e s t h e e f f e c t on t h e incre - ,mental range required of flying at a l t i t u d e s other than optimum f o r a " t y p i c a l supersonic transport. Here again, incremental range i s a representation of t h e added f u e l required i n terms of t h e increased range it would provide under standard - temperature still - air optimum - It i s shown t h a t f o r s m a l l deviations from opti - a l t i t u d e conditions.

mum, t h e e f f e c t s of off - design a l t i t u d e are about t h e same, whether the f l i g h t a l t i t u d e is above o r below optimum level. However, f o r flying above optimum a l t i t u d e e n t a i l s a somewhat l a r g e r deviations, greater penalty than flying below optimum a l t i t u d e . A preliminary I estimate has indicated t h a t all t h e one - direction supersonic - transport t r a f f i c of t h e same type could be accommodated at three levels, sepa - rated, as at present, by 4,000 - foot increments. Thus, all flights i n one direction could cruise within 4,000 f e e t of optimum a l t i t u d e which, as indicated i n t h e figure, would give a penalty of about 65 and 80 nautical miles i n range increment required, f o r below - and above - optimum a l t i t u d e conditions, respectively.

It w a s assumed t h a t f o r all westbound f l i g h t s one l e v e l would be required f o r one - third of t h e t i m e , two levels, f o r one - third of the t i m e , and three levels, f o r t h e remaining t h i r d of t h e t i m e . I n other words, one - half of a l l f l i g h t s would be made at optimum a l t i t u d e and t h e remaining f l i g h t s , 4,000 f e e t above o r below optimum.

A probability curve of incremental range required, incorporating t h e e f f e c t s of t h i s d i s t r i b u t i o n of discrete - altitude assignments, was computed and i s together with winter wind and temperature effects, shown i n figure 3 . A corresponding curve f o r all f l i g h t s at optimum a l t i t u d e i s given f o r comparison. Again, with t h e 1 - percent probability level, t h e r e s u l t s indicate t h a t if three cruise f l i g h t l e v e l s are required f o r westbound supersonic - transport t r a f f i c , t h e increment i n design range over t h e basic t r i p distance should be about 288 nautical

miles o r about 44 nautical miles greater than if a l l f l i g h t s could

cruise at optimum a l t i t u d e .

Effect of Payload Load Factor I n t h e preceding discussion of the f a c t o r s of weather and f l i g h t - l e v e l assignments, t h e objective of carrying f u l l payload f o r a large percentage of f l i g h t s w a s assumed. However, a 100 - percent load factor, f o r passenger payload at l e a s t , i s not a p r a c t i c a l goal because of the v a r i a b i l i t y of demand f o r accommodations. Therefore, unless t h e assump - t i o n can be made t h a t v a r i a b i l i t y i n passenger payload can be compen - sated by providing excess cargo capacity i n t h e airplane and standby cargo, t h e load f a c t o r w i l l be substantially less than 100 percent.

Allowing for a reduced average load f a c t o r and a suitable variation about t h i s average would tend t o reduce t h e design - range increment required t o l i m i t t o a reasonable frequency t h e need for denying requested accommodations. Average passenger load f a c t o r f o r t h e super - sonic transport has been variously estimated between 55 and 70 percent of capacity.

The e f f e c t of load f a c t o r on t h e range increment required i s i l l u s t r a t e d by assuming a normal d i s t r i b u t i o n of load f a c t o r about a mean value of 60 percent with a standard deviation of 1 2 percent (20 - percent standard deviation of demand). It is f u r t h e r assumed t h a t t h e design o r capacity payload is 7 percent of design take - off gross weight.

The e f f e c t s on range capability of t h e variations of take - off weight resulting from t h i s varying payload w e r e computed and combined s t a t i s t i c a l l y with t h e e f f e c t s of winter wind and temperature condi - tions and variation i n f l i g h t - l e v e l assignments discussed previously.

The overall e f f e c t of these f a c t o r s on t h e probability of requiring incremental range capability i s shown i n figure 4. For comparison, t h e probability curve from figure 3 f o r t h e corresponding 100 - percent - mean - load - factor case i s given. The r e s u l t s indicate t h a t allowing f o r the assumed v a r i a b i l i t y i n load f a c t o r considerably reduces t h e incremental range required t o provide a given probability of meeting t h e demand f o r seats. For the 60 - percent kean load factor, a design range incre - ment of approximately 160 nautical miles should be s u f f i c i e n t f o r all but 1 percent of t h e winter westbound f l i g h t s , whereas, i f t h e goal were t o be the 100 - percent mean load f a c t o r of mixed passengers and cargo payload, range increments of 288 nautical m i l e s would be needed at t h e same probability level.

FUEL RESERVE REQUIREMENTS En Route Factors There are a number of factors which introduce some uncertainty i n the determination of en route o r destination fuel. These factors, together with estimated standard deviations from t h e i r predicted values and t h e corresponding v a r i a b i l i t i e s calculated f o r en route f u e l are l i s t e d i n t a b l e 11. The standard deviation of incre - requirements, mental f u e l requirement due t o each f a c t o r is shown as a percent of t h e predicted destination f u e l requirement. It is indicated t h a t t h e e f f e c t of variations between forecast and actual wind velocity i s r e l a t i v e l y s m a l l f o r t h e supersonic transport, whereas, f o r the subsonic jet, uncertainty i n accounting f o r wind e f f e c t s has been indicated t o be t h e primary f a c t o r i n en route reserve requirements. The primary fac- t o r , i n t h e case of t h e supersonic transport, i s shown t o be v a r i a b i l i t y i n specific f u e l consumption.

The estimate of 4 - percent standard deviation i n specific f u e l con - i i sumption for each engine is a t h i r d greater than values that have been used f o r subsonic j e t engines; t h i s estimate allows f o r t h e p o s s i b i l i t y of more v a r i a b i l i t y with the more complex supersonic engines. With f o u r engines, t h e overall standard deviation i n specific f u e l consumption, hence, i n f u e l consumed, reduces t o 2 percent. The next most important f a c t o r i s indicated t o be drag variation. This variation is assumed t o have a standard deviation of 1.5 percent among different airplanes of t h e same design and is based on observations of speed variations The e f f e c t s of each of the other factors a r e f o r subsonic transports.

0.5 percent, or less, i n standard deviation of f u e l required. The con- bined e f f e c t s of t h e l i s t e d factors give a standard deviation i n des - t i n a t i o n f u e l required of 2.55 percent. When these factors are applied t o a subsonic transport f o r comparison, t h e standard deviation i n des - t i n a t i o n f u e l is 4.05 percent.

Combined En Rou%e and Holding Requirement referred It is apparent t h a t protection i n the fqrm of reserve fuel, t o as en route reserve, w i l l be required t o &low f o r those occasions i n which the foregoing v a r i a b i l i t y of f u e l consumption results i n burning ''3.

O n other flights, however, more f u e l than t h e planned destination fuel.

t h i s added f u e l would not be used and would be available f o r holding at the destination. It would appear logical, therefore, t o consider these t w o reserve requirements together, with t h e ultimate objective of l i m - i t i n g t o a reasonable frequency t h e need f o r diversion t o an a l t e r n a t e a i r p o r t because of f u e l l i m i t e d holding t i m e . To t h i s end, holding-timet, s t a t i s t i c s f o r subsonic jet transports have been obtained from an a i r l i n e source for t h e N e w York International Airport. These data indicated that 23 percent of t h e arriving flights during 1962 were delayed f o r clearance by A i r Traffic Control f o r an average of 10 minutes. Inas - much as t h e frequency d i s t r i b u t i o n of these delays w a s not available, it w a s assumed t h a t they would be described by t h e positive half of a Gaussian, o r normal, d i s t r i b u t i o n w h i c h would give the 10 - minute mean delay. Accordingly, t h e frequency distributions of f u e l consumption a delta and corresponding t o these holding times were calculated f o r a variable - sweep supersonic transport and combined with t h e en route 5 .

reserve s t a t i s t i c s t o give t h e probability curves shown i n figure These curves indicate the p r o b a b i l i t i e s of exhausting given amounts of combined reserve fuel. The reserve f u e l AWF i s shown divided by

the destination f u e l WDnF . Further delay would then require e i t h e r pro -

ceeding t o the alternate a i r p o r t o r consuming a l t e r n a t e reserve fuel. I n order t o obtain a r e a l i s t i c probability level, hence, the reserve f u e l which should be provided f o r t h e supersonic airplanes, a similar analysis t o that described w a s made f o r a subsonic j e t transport and t h e probabil - i t y curve f o r this case i s a l s o shown. The CAR en route plus holding f u e i w a s then determined f o r the subsonic transport and is indicated by the v e r t i c a l l i n e . The intersection of t h i s l i n e w i t h the probability curve f o r t h e subsonic transport indicates the probability l e v e l of equaling o r exceeding the CAR reserve which is assumed t o be acceptable.

The reserve fuel required t o give t h e same probability of exhausting en route and holding reserves i s shown t o be about 10 percent f o r t h e variable - sweep supersonic transport and about 1 1 percent f o r t h e i The combined en route reserve (10 per - delta - wing supersonic transport.

cent of t r i p t i m e ) and holding reserve (30 minutes at 1,500 feet), as specified f o r subsonic turbine transports i n CAR 41,* w e r e computed f o r both a i r c r a f t and faund t o be 13 percent and 14 percent of destination f u e l f o r t h e variable - sweep and d e l t a w i n g s , respectively, o r about 3 percent of t h e destination f u e l greater than the values determined by t h e statistical. procedure u t i l i z e d i n t h i s paper.

Two - Engine Failure The p o s s i b i l i t y of two - engine failure, even though remote, requires t h a t t h e airplane carry sufficient f u e l on board t o proceed, with two engines stopped, from any point en route t o a suitable airport. If the f u e l required i n t h i s case were t o exceed t h e planned destination f u e l by more than t h e amount provided by t h e en route, holding, and diver - sion reserves, the reserve f u e l allowance would be, i n effect, deter - mined by t h i s contingency.

It w a s found t h a t the two - engine range potential of both the variable - sweep and d e l t a configurations i s b e t t e r a t subsonic speeds than at supersonic speeds. The range remaining a t various points along the route after two - engine f a i l u r e w a s computed f o r both configurations f o r a Mach number of 0.80 and t h e corresponding optimum altitudes. It w a s assumed t h a t f u e l f o r diversion t o an a l t e r n a t e airport 250 nautical miles distant would be available i n addition t o en route and holding reserve f u e l determined by the procedure o f t h e preceding section e n t i t l e d " Combined En Route and Holding Requirement. " The range remaining a f t e r two - engine f a i l u r e i s plotted against distances from t h e origin a i r p o r t a t which f a i l u r e occurred i n figure 6 f o r the two configurations. It i s assumed t h a t at landing, complete. The two 45' f u e l burnout of destination and reserve f u e l i s l i n e s (dashed) represent the distance f o r e i t h e r returning t o the origin o r proceeding t o t h e destination. The lower boundary indicates the distance t o t h e nearest adequate alternate a i r p o r t along the route from New York t o Paris. Inasmuch as neither t h e variable - sweep nor the delta - wing transport falls below t h e alternate boundary, the reserve fuel allowance provided i s sufficient t o meet t h i s requirement. It except f o r t h i s s m a l l portion near t h e mid - should a l s o be noted that, point of t h e f l i g h t , t h e variable - sweep transport could e i t h e r return t o t h e origin or proceed t o t h e destination.

~~

*

Anon.: Certification and Operations Rules f o r Certificated Route A i r Carriers Engaging i n Overseas and Foreign A i r Transportation and A i r Transportation Within H a w a i i and Alaska. Civil A i r Regulations Part 41, Rules Service Co. (Washington, D.C. ), M a y 6, 1963. .

CONCLUSIONS I n conclusion, t h i s analysis of operational factors i n r e l a t i o n t o f u e l requirements f o r t h e supersonic transport has indicated t h e prob - a b i l i t i e s of requiring additional design - range capability above t h e basic t r i p distance t o allow f o r t h e e f f e c t s of atmospheric variations, It i s shown, f o r f l i g h t - l e v e l assignments, and payload load factors.

example, t h a t with a l l t h e factors considered and the assumptions used, an addition of about 160 nautical miles t o t h e basic t r i p distance. from N e w York t o Paris would be s u f f i c i e n t t o permit acceptance of a l l pay - load offered except f o r 1 percent of the f l i g h t s .

I n t h e area of f u e l reserves, it w a s found t h a t 10 percent of the destination f u e l f o r t h e variable - sweep supersonic transport and 1 1 per - cent f o r t h e delta - wing supersonic transport would provide equivalent protection t o t h a t given subsonic j e t transport operations by the com - bined c i v i l air regulations (CAR) en route and holding requirements.

These reserves f o r t h e supersonic transport represent a savings equiv - alent t o about 3 percent of t h e destination f u e l over those calculated by using CAR en route and holding requirements.

It was also found t h a t these reserves, i n combination with a planned a l t e r n a t e airport 250 nautical miles distant, would provide ample f u e l t o proceed at subsonic speeds t o an adequate a i r p o r t i n the event of a two - engine f a i l u r e anywhere en route from N e w York t o Paris.

1. Crutcher, Harold L.: Climatology of t h e Upper A i r as Related t o t h e Design and Operation of Supersonic Aircraft. Weather Bur., U.S.

Dept. Commerce, Mar. 1963.

2. Tolefson, H. B.: Atmospheric Temperature Observations t o 100,000 Feet f o r Several Climatological Regions of t h e Northern i Hemisphere. NACA TN 4169, 1957.

TABLE I e - VARIATION O F ATMOSPHERIC CONDITIONS F R O M S T A N D A R D I Mean Standard Mean wind, Standard Flight temperature, deviation, knots deviation, Season phase O F OF knots (a) (b)

Climb and 37- w 522

des cent Winter - 4.1 510. 2

Cruise 33 w 518

Climb and 26 w 3 9

descent Summer +5*9 k7.6 Cruise 2 w k8 TABLE 11.- EN ROUTE RESEIiVE FUEL FACTORS Standard deviation i n Standard Factor f u e l required, deviation percent of planned destination f u e l Wind 10 knots 0.46 Temperature 4 ' F Drag 1.5 percent 1-33 Specific f u e l 4 percent (each engine) 2.00 consumption Fuel - gage error 0.5 percent Navigation errors 58 nautical miles (lateral) .06 Weather avoidance 58 nautical miles ( l a t e r a l ) .06 A i r t r a f f i c control 58 nautical miles (lateral) .06 Altimetry e r r o r 100 f t a t 65,000 f t .10 Combined standard deviation of destination f u e l = 2.55 percent EFFECT OF WIND AND TEMPERATURE ON INCREMENTAL DESIGN - RANGE REQUIREMENTS 1.0 r_ WINTER - PROBABILITY \

'\\\ \

\ \ \ \ \ I I \ Figure 1 EFFECT OF OFF - DESIGN ALTITUDE ASSIGNMENT ON DESIGN - RANGE INCREMENT REQUIRED BASIC RANGE 3,160 N. MI.

400r I RANGE INCREMENT, N. MI.

- 8,-000 0 8,000 DIFFERENCE BETWEEN ASSIGNED AND OPTIMUM ALTITUDE, FT Figure 2 EFFECT OF FLIGHT - LEVEL ASSIGNMENT O N DESIGN - RANGE REQUIREMENTS WINTER CONDITIONS I.Or - . I AT - PROBABILITY - . O I - \

,001 I I I I I

0 1 0 0 200 300 400 RANGE INCREMENT, N. MI.

Figure 3 EFFECT OF PASSENGER LOAD FACTOR ON DESIGN - RANGE REQUIREMENTS 3 FLIGHT LEVELS; WINTER CONDITIONS I Or MEAN LOAD FACTOR, 100%

-h

'i

\ '\

.OOl I I I I I

0 100 200 300 400 RANGE INCREMENT, N. MI.

Figure 4 PROBABILITY OF RESERVE FUEL REQUIREMENT EN ROUTE AND HOLDING FUEL - 1 . 0 TRANSPORT CIVIL AIR REGULATION - .OOl - 0 -0 5 .IO -15 .20 A w ~ / w ~ . ~ .

Figure 5 SUBSONIC RANGE REMAINING AFTER TWO - ENGINE LOSS VARIABLE - SWEEP N. MI.

SUPERSONIC TRANSPORT 1,000 / / i ~~ DISTANCE EPiREsT, TO ALTERNATE 0 1,000 2,000 3,000 4,000 DISTANCE FROM TAKE - OFF WHEN FAILURE OCCURS, N. MI.

Figure 6 h R CONSIDEFUTION OF SOME AIRWORTHINESS REQUIRENf3NTS FOR SUPERSONIC TRANSPORTS

BY momas L . Coleman Jf

'.

NASAJLangley Research Center SUMMARY Some airworthiness requirements for supersonic transports in the ;areas of structural design criteria, performance, and handling qualities are discussed. Several areas are indicated wherein present Civil Air Regulations are not directly applicable to the supersonic transport.

For some areas, it appears that the requirements can be adapted to supersonic - transport configurations by only minor modifications or extensions. For other areas, it appears that basic concepts on which the requirements are based will require reexamination.

4 d t Y d <

eQs-6-

INTRODUCTION That supersonic transports be sufficiently airworthy to maintain, or improve, current levels of safety is a matter of concern to airline operators, manufacturers, Government regulatory agencies, and the NASA.

Although NASA is not charged with the responsibility of specifying air - worthiness requirements, it does have a role in anticipating airworthi - ness problems, conducting research to find solutions, and in dissemi - nating research results to the aviation community.

Airworthiness is interwoven with the aerodynamic and propulsion characteristics, the operational environment, and the manner that the Research in the airworthiness area is R continuing airplane is flown.

effort involving the accumulation and extrapolation of fundamental knowl - edge and operational experience to new configurations and anticipated operating conditions. In some cases, this research has stayed ahead of %he problems but, in other cases, has fallen behind.

NASA research, the supersonic - tsansport feasibility studies, and discussions with various aviation grotps have given an insight into some airworthiness aspects of supersonic transports. In this paper, the information from these sources are discbssed to indicate the state of the art in several airworthiness areas. The paper will highlight a few airworthiness items which, if the thesis of maintaining current levels of safety is accepted, could have a bearing on the merits of the config - urations and the operating procedures.

The discussion will touch on three broad areas of airworthiness: (1) structural design criteria, (2). performance, and ( 3) handling qual - Each of the three areas will be more specifically covered in ities.

subsequent sections.

SYMBOLS VA maneuvering speed VB rough - air speed VC cruise speed dive speed VD M Mach number lift coefficient CL maximum lift coefficient

( VMAx

operational lift coefficient

( cL) OPERAT

a angle of attack useable lift coefficient (CL)USEAJ3LB AV airspeed error actual airspeed VACTUAL target airspeed VTARGET F thrust W weight cycles to damp to one - half amplitude c1/2 DISCUSSION Structural Design Criteria Structural design speeds. - The strength requirements and operational airspeed limits for transport airplanes are tied directly to a set of structural design speeds. Figure 1 illustrates a set of these speeds in terms of the variation of Mach number with altitude for a supersonic transport. The curves labeled maximum lift, buffet, engine blowout, maxi - mum temperature, airframe and engine strength, and flutter are boundaries imposed on the airplane by aerodynamic, propulsion, and structural ,characteristics.

The flight region is bounded on the low - speed side by the maneuvering speed VA and the rough - air speed VB. These are the speeds at which the airplane can develop full load factor or withstand the maximum gust velocity without stalling or overstressing the structure. The high - speed side of the flight region is bounded by the design cruise speed Vc, which is the maximum speed for normal flight operations. The speed spread between Vc and the design dive speed VD is a margin provided to insure structural integrity in the event of overspeeds caused by such factors as: airplane upset in turbulence, traverses of wind and tempera - ture gradients, and inadvertent speed exceedances of by the pilot.

Vc Several problem areas appear to exist in adapting current Civil Air Regulations and concepts to specifying the structural design speeds to supersonic transports and these areas are: (1) The procedure of basing the maneuvering speed on a maximum VA lift coefficient may not be feasible for low - aspect - ratio configurations since, as will be discussed subsequently, the maximum lift occurs at such high angles of attack as to be operationally impractical. A possible approach is to base the speed on an attitude limit of the airplane.

(2) The concept of the rough - air speed VB may not be applicable to the supersonic transport because: (a) At subsonic speeds, slowing down from VC to VB, trav - ersing rough air, and then regaining speed will be unattractive from the fuel - requirement standpoint.

(b) At supersonic speeds, the airplane response characteris - tics are such that no appreciable reduction in gust loads can be obtained unless speed is reduced to subsonic values. In addition, it is doubtful whether the turbulence can be detected sufficiently in advance to permit slowing down prior to traversing the turbulence.

Consequently, it is thought that the rough - air speed and slowdown con - cepts may not be applicable to supersonic transports. This condition would imply that the supersonic transport may have to be designed for the maximum gust velocity at a speed higher than VB, possibly at the design cruise speed Vc.

( 3 ) In regard to the design cruise speed

V , it is thought that the

current practice of permitting the manufacturer Fo select the speed w i l l be as applicable to the supersonic transport as it is for present transports.

(4) The present requirement for determining the minimum speed margins between Vc and VD may not be adequate for the supersonic transport.

- This apparent inadequacy is due primarily to the much higher thrust weight ratio anticipated for the supersonic transports than for present transports. The higher thrust - weight ratio increases the possibility of inadvertent overspeeds following transition from climbing flight to level flight, or from level flight to descending flight. Overspeeding is an operational problem on present transports (ref. 1) and is likely to be one with supersonic transports unless attention is given to this poten - tial problem area.

Limit loads. - It is likely that the supersonic transport will be designed, as are present transports, for limit loads resulting from maxi - mum expected in - flight and ground loads. In this section, three sources

of loads - vertical velocity at landing impact, gust loads, and maneu -

vers - will be discussed.

Vertical velocity: Present Civil Air Regulations (ref. 2) requOre that transports be designed for a limit vertical velocity, or sinking speed, of 10 feet per second at touchdown. There is some question as A s to the applicability of this value to the supersonic transports.

shown in figure 2, this question stems from the evidence that the verti - cal velocities of current turbojet transports are significantly higher than those experienced by piston transports. (See ref. 1 . ) For exaple, 1 landing in 1 0 0 exceeds a vertical velocity of about 2.5 feet per second for piston transports and about 5 feet per second for turbojet As a point of interest, an extrapolation of the turbojet transports.

\ curve indicates that a vertical velocity in excess of about 6 feet per second would be expected in the estimated 2,000 landings made each day by the U.S. turbojet fleet.

Simulator studies aimed at determining the causes of the increased vertical velocities have been conducted by NASA (ref. 3 ) and, also, are being conducted under a contract sponsored by the Federal Aviation Agency.

So far, however, the reasons for the increased vertical velocities of the turbojets over the piston transports have not been determined. I t due in large part to the turbojets having a is thought, though, to be more sluggish response to pilot control inputs during the flare maneuver.

(See ref. 1 . ) Studies have indicated that the flare - response character - istics of supersonic - transport configurations may be decreased from that for present turbojets. In this event, it may be that the supersonic transports will experience higher vertical velocities than present If this is true, the question arises as to whether the transports.

10 - feet - per - second design value is adequate for the supersonic transport.

This question cannot be simply answered but, rather, it must be consid - In ered from the viewpoint of maintaining a minimum level of safety.

this regard, the answer will depend, to a large extent, on the basis (number of flights, flight hours, flight miles, or a fleet concept) used to define the level of safety.

P Gust loads: In view of the new configurations, increased operating speeds, and altitudes of the supersonic transport, it is well to con - sider the implications of these changes relative to gust loads. The plot on the left of figure 3 summarizes information on the average per - cent time that rough air has been experienced by subsonic airplanes at the various altitudes (ref. 4 ) . The portion of this curve below 4 0 , 0 0 0 feet is based, to a large extent, on data collected during trans - port operations. Above 40,000 feet, the curve is based on military oper - ations and special flight investigations. As shown, the amount of rough air decreases rapidly with increasing altitude, but is still present about 2 percent of the time at altitudes between 60,000 and 75,000 feet.

of the turbulence also decreases Although not shown here, the intensity with increasing altitude. Additional data on the turbulence environment are being obtained by the NASA and by the U.S. Air Force from measure - ments made during routine operational flights and also from special investigations.

Studies considering the turbulence environment together with the supersonic - transport flight profiles and response characteristics have 3 ) is likely to result indicated that the transonic flight region (fig.

in the most critical gust loadings. Also, the studies indicate that the low - altitude turbulence environment associated with subsonic flight is likely to be a more prolific source of repeated gust loads than the turbulence environment at high altitude associated with supersonic flight.

The high supersonic flight region is not expected to be critical from the 'loads standpoint. Because of the high speeds and the likelihood of decreased damping of airplane motions at the cruise altitudes, the small amount of turbulence at high altitude is thought likely to be more of a problem as regards airplane control and stability than as regards loads.

Civil Air Reg - The plot on the right in figure 3 shows the present cruise speed ulations design gust velocities applicable to the design are presently V c and the design rough - air speed VB. Gust velocities specified up to an altitude of 50,000 feet. Data are available, however, on which to base the selection of design gust velocities up to altitudes of 75,000 feet.

A s previously mentioned, there is some question regarding the appli - cability of the slowdown and rough-gir speed concept to the supersonic transport.

Thus, although available data do not indicate that the design gust velocities need to be altered for the supersonic transport, there is a need to re - assess the speeds at which they are to be applied.

Fac - tors which should be considered in this re - assessment are: (1) feasi - bility of accomplishing effective and practical slowdown, (2) effect of weather radar and flight planning on turbulence avoidance, and ( 3 ) maneu - verability of airplane for avoiding detected turbulence.

Gust loads are currently calculated by using two complementary methods: (1) discrete gust and (2) continuous turbulence. The discrete - gust method has been used to set the level of loading and the continuous method, to detect unusual response characteristics of the airplane for use in modifying the load level derived from the discrete method.

Although progress has been made in developing the continuous turbulence method (ref. 5 ) , it is doubtful that it will be sufficiently developed in time to permit its employment on an absolute basis to the supersonic transport. Consequently, it appears that the present practice of calcu - lating gust loads by using the two methods on a complementary basis will have to be applied to the supersonic transport design.

Maneuver loads: A s is the case with present transports, the super - sonic transport will experience maneuver loads during operational passenger - carrying flights and during pilot and airplane training and check flights. For use in fatigue analyses, it is thought that, at present, the best estimate of the maneuver loads can be obtained by assuming that the supersonic - transport maneuver experience will be simi - lar to that for current turbojet transports. Information on turbojet maneuvers experienced is being obtained by NASA and some data have been published in references 1 and 6. Maneuvers performed during pilot and airplane training and check flights constitute a major source of the overall maneuver experience for turbojet transports.

(See ref. 6. ) It is anticipated that training and check flights also will be a major source of maneuver loads for supersonic transports. Based on a crude assessment of the many factors which may influence the amount of check flying required, it is estimated that on the order of 4 percent of the total flight time for the supersonic transport will be spent in pilot training or airplane check flights. Further, it is estimkLed that about one - fourth of the training and check flying will be at supersonic speeds.

Inasmuch as limit maneuver load factors in combination with thermal stresses at supersonic speeds are apt to be a critical design condition, careful attention to the limit - maneuver - load - factor requirements at supersonic speeds is warranted. In the absence of operational data - applicable to supersonic transports, the specification of the supersonic maneuver load factors will have to be done primarily on the basis of engineering judgment based on extrapolating experience gained from pres - ent transports to the new configurations and operational requirements.

Available data from turbojet transports (ref. 1 , for example) indicate that these airplanes are occasionally subjected to maneuvers as large or larger than the limit values of -1g and -2.5g. Further, these maneuvers appear to be induced by a variety of causes such as recovery from loss of control, improper operation of autopilot, and collision avoidance.

Consideration of available maneuver data in conjunction with anticipated supersonic - transport operational requirement has not indicated that the present design maneuver load factors of -1g and +2.5g should be changed for either subsonic or supersonic flight.

Ultimate loads. - Present Civil Air Regulations (ref. 2) require that

the airplane structure be designed to support ultimate loads equal to 1.5 times the limit loads. The specification of ultimate load design requirements for the supersonic transport is complicated by the neces - sity of considering thermal effects on the structural material. There are three major effects which require consideration: (1) short - term degradation of the material properties, (2) long - term effects of the thermal history on the material, and ( 3 ) thermal stresses which are highly dependent on the flight profile. The first two effects can be accounted for by using degraded material properties and do not appear to be a major problem. In the case of thermal stresses, however, it is not clear which of several possible approaches is the most rational. There are two major aspects of the thermal stresses which need to be resolved.

These are: (1) what factor of safety should be used? and (2) how should the thermal stresses for the high - speed portion of the V - n envelope be treated?

As was discussed in paper no. 1 8 of this compilation by Richard A .

Pride, two different approaches vere used in the feasibility studies to account for the thermal stresses in computing ultimate loads. In one conibined thermal approach, a factor of safety of 1 . 5 was applied to the and mechanical stresses. In the other approach, ultimate loads were determined without directly considering a factor of safety for the ther - ! m a l stresses. It would appear that this latter approach, in effect, .

assumes that either the thermal stresses are precisely known or that the safety factor applied to the mechanical stresses a l s o c a , n adequately cover any unknowns associated with the thermal stresses. At this time, it is not clear which thermal factor of safety would be the most rational.

It would appear, however, that the thermal stresses may be known to a higher degree t h n are the mechanical stresses and, consequently, the fac - tor of safety could be less than 1 . 5 . On the other hand, there w i l l be 4 4 5 some error in determining the thermal stresses and, consequently, a factor greater than 1.0 would appear to be required.

As the normal flight profile will not encompass speeds higher than the design cruise speed Vc, it appears that thermal stresses for the high - speed portion of the V - n envelope based on transient overspeed or emergency conditions must be considered. This approach was used in the feasibility studies.

Fatigue loads. - The major load sources to be considered in deriving fatigue loading spectra for the supersonic transport are: (1) Ground - air - ground cycles

(2) Thermal stress cycles 1

( 3 ) Maneuver loads (4) Gust loads ( 5 ) Pressurization cycles (6) Landing impact (7) Taxi load (8) Sonic - induced loads Except for the addition of the thermal stress cycles, the spectra involve the load sources which have been considered for present trans - ports. Because of the increased concentration of the airplane weight inboard on the supersonic transports, the ground - air - ground cycle may be of more relative importance in their fatigue life than on present trans - ports. Calculations have indicated that the thermal stress cycles which the structure will experience on each flight will be an important source of fatigue damage. Because of the strong dependence of the thermal stress cycles on the detailed mission profile, possibly more consideration will need to be given to deriving the fatigue loading spectra based on mission analyses than has been required in the past.

Performance From the overall viewpoint, the performance requirements for super - sonic transports would not appear to be greatly different from those for present transports. In applying current requirements and insuring that *’ minimum performance margins are attained in practice, however, there are several areas which will require further consideration. Two of these areas are discussed in the following sections.

Performance reference. - Many of the current low - speed performance margins and operating speeds are tied directly to a stall speed. Because a clearly defined stall speed may not exist for some supersonic - transport configurations, the present concept of relating performance to a stall speed may require modification.

To illustrate this point, figure 4 shows two plots of lift coeffi - cient CL against angle of attack a for two configurations. The curve on the left is representative of a subsonic turbojet transport having an

aspect ratio of about 7 . The other curve is representative of a

supersonic - transport configuration having an aspect ratio of about 2 .

For the high - aspect - ratio configur&ion, the lift - coefficient curve reaches a maximum value A at a particular angle of attack.

fairly definite stall speed also is associated with this maximum lift coefficient. Present low-speed performance requirements use this stall speed as a reference for specifying operational speeds so that the opera -

tional lift coefficient will be sufficiently below ( C L ) ~ to provide

adequate maneuverability.

In the case of the low - aspect - ratio configuration, the lift - ( C L ) ~ nor does a definite coefficient curve does not have a distinct stall speed exist within practical angles of attack. Consequently, the concept of basing performance requirements on a stall speed is not appli - cable to the low - aspect - ratio configurations.

One procedure used in the feasibility studies to assess the perfor - mance of the low - aspect-ratio configurations involved: (1) Selection of the lift coefficient at which it was desired that the airplane operate.

(2) Examination of the aerodynamic and control characteristics of the airplane to insure that the airplane was capable of producing a useable lift coefficient sufficiently high to provide maneuverability comparable to that inherent in present requirements. The angle of attack corresponding to the useable lift coefficient would, in effect, be a m a x i m u m operational limit. Above this limit, the flying qualities would likely deteriorate.

Another, and possibly more fundamental, approach would be to divorce In this completely the performance specification from a reference speed.

approach, the manufacturer would be permitted to select the operational speeds, but would be asked to meet minimum maneuverability requirements associated with the particular flight condition. At this time, however, it is doubtful whether sufficient data are available on which to make a rational determination of the minimum maneuverability requirements.

Nevertheless, the possible merits of the approach are thought to warrant further consideration.

i i Climb gradient. - Present Civil Air Regulations specify minimum climb gradients for various segments of flight and for engine - out con - ditions. These gradients have evolved from past operating experience with transports. In view of the novel configurations being considered for supersonic transports, it appears that the implications of the new configurations on the climb gradients may need to be examined.

Figure 5 illustrates one aspect of the effect of configuration on the climb gradient for the second - segment flight condition. The upper plot shows two drag or thrust - required curves as a function of airspeed.

The lower curve is for an aspect - ratio - 7 subsonic turbojet transport, and the upper curve is for an aspect - ratio - 2 supersonic transport config - uration.

As only the shapes of the curves are pertinent to this discus- the curves have been sion, shifted closer together along both the drag and airspeed axis for convenience of presentation.

During the climb - out, the pilot will be attempting to fly at a par - ticular speed, referred to as ''VWGm.'' At this speed, the airplanes are required to have sufficient thrust with one engine out to provide a minimum climb gradient - 3 percent for this case.

In practice, the actual airspeed achieved by the pilot will miss the target speed by some amount AV with the result that the actual airspeeds achieved form a distribution about the target speed. The available climb gradient is (for each configuration) proportional to the difference between the engine - out thrust and drag curves and decreases as the actual speed falls below the target speed.

The lower plot in figure 5 shows the remaining climb gradients for both configurations as a function of the airspeed error AV. The results indicate that a given negative airspeed error causes a significantly larger reduction in climb gradient for the low - aspect - ratio configura - tion than for the high - aspect - ratio configuratign. Other factors such as deviations of weight and thrust from nominal values also have a more deleterious effect on the climb gradient for the low - aspect - ratio config - uration than for the high - aspect - ratio configuration. From this compari - son, it appears that the configuration may have a significant effect on For other flight the actual level of performance achieved in service.

segments, the order of the results may be reversed from those shown in the figure. Nevertheless, it would appear that the configuration effects : may be sufficiently large to warrant consideration in the specification A more detailed discussion of the effect of of perfomnee margins.

configuration on climb performance is given in reference 7 .

4 4 8 Handling Qualities The handling qualities which appear to be desirable for the super - sonic transport are not thought to be very different from those desired for present transports. The major problems appear to be in obtaining configurations with characteristics which are inherently conducive to good handling qualities. Consequently, it appears that considerable stability augmentation will be required to provide satisfactory handling qualities as is discussed in several other papers in this compilation (papers no. 3, 4, 26, and 27). Three aspects of handling qualities which will require special consideration are subsequently discussed.

Speed stability.- Although the subject of speed stability during approach is discussed in several papers (papers no. 3, 4, 26, and 2 7 ) , of airworthiness.

it is briefly discussed here from the standpoint Figure 6 shows thrust required curves for four supersonic - transport configurations during the approach condition6 The portion of the curves to the left of the vertical lines is the familiar "back-side" of the power curve and that to the right is the region of speed stability.

A s shown by the symbols, the approach speeds for three of the configura - tions are on the back side of the curves. Although some military air - planes operate on the back side, current transports operate in the region of speed stability during approach. The question, then, is whether oper - ations on the back side of the power curve during approach can be toler - ated in transport operations. If not, shifting the drag curves to a more favorable position by use of drag brakes may be one method of improving the approach characteristics. (See ref. 8 . ) Some pilots are of the opinion that operations on the back side would be acceptable for transport operations, provided the rate of change of thrust required with airspeed is not too large. It is felt, however, that additional simulator studies and flight tests will be required before a final answer is obtained.

Ground handling.- '%he handling qualities of the supersonic transport during the take - off roll and during execution of the rotation maneuver may be adversely affected by the airplane response characteristics to runway roughness. Figure 7 shows the variation of the root - mean - square acceleration of the pilots' compartment with speed for a turbojet and I supersonic transport configuration. These results are taken from refer - ence 9 and are based on a simplified analysis considering rigid - body pitch and vertical translation and linearized landing - gear characteristics.

The level of runway roughness was comparable to that of a good commercial runway.

The results in figure 7 indicate that over most of the speed range, the response of the supersonic transport is much higher than that for the turbojet transport. Turbojet pilots have complained that present response levels sometimes cause discomfort, affect control of the airplane during the take - off run, and interfere with the precise execution of the rota - tion maneuver. Consequently, it would appear that the higher indicated response level for the supersonic transport may not be acceptable. Thus, the response characteristics of the supersonic transport to runway rough - ness may be more of a problem than it is for present transports. Detailed study of the response of configurations to specific runway roughness pro - files will be required to assess this problem area further.

Stability augmentation.- It is anticipated that the supersonic trans - port will require increased use of stability - augmentation devices to pro - vide acceptable handling qualities during low - speed operations and during cruise flight. At this time, the effects which the increased use of stability - augmentation systems may have on airworthiness requirements are not fully known. Unless a very high degree of reliability of the augaentation systems can be assured, however, it would appear that the requirements will need to cover both the failed and unfailed systems.

In this regard, the amount and reliability of the augmentation systems could have a bearing on the selection of speed placards and on some load requirements.

Vertical - tail loads appear to be one of the areas where special attention must be given to the case of a failed lateral - stability aug - mentation system. In this connection, figure 8 indicates the effect of period and damping on the vertical tail due to turbulence. The ordinate values are the ratio of the root - mean - square sideslip angle at the ver - tical tail to the root - mean - square gust input angle. The abscissa is the reciprocal of the number of cycles to damp the lateral motion to half amplitude. The results, except for the supersonic transport, were taken from reference 1 0 and were calculated on the assumption that the predominant parameters are the period and the damping of the airplane.

The cross - hatched area in the figure represents the range of results obtained for several supersonic transports by using estimates of the period and damping of the lateral mode for the unaugmented airplane.

The results indicate that, for the maugmented case, the vertical - tail loads for the supersonic transport may be amplified to a greater extent than for present turbojet airplanes. St may be expected, therefore, that the case of failed lateral - stability augmentation may constitute one of the critical design conditions for the vertical tail.

CONCLUDING REMARKS A review of several aspects of airworthiness for supersonic trans - ports has indicated some areas wherein present Civil Air Regulations are not directly applicable. In some cases, it appears that the requirements can be adapted to supersonic - transport configurations by only minor znodif ications or extensions. 5th- :cBses, basic concepts w i l l require - re - examination. It should be recognized that the Federal Aviation Agency is cognizant of the airworthiness problems discussed herein and has circulated tentative supersonic transport airworthiness objectives.

(See ref. 1 1 . ) a 1. Staff of Langley Airworthiness Branch: Operational Experiences of NASA TN D-1392, Turbine - Powered Commercial Transport Airplanes.

1962.

2. Anon.: Airplane Airworthiness; !Transport Categories. Civil Air Regulations Part 4b, Rules Service Co. (Washington, D. C. ), Jan. 7 , 1963.

3. Bray, Richard S.: Piloted Simulator Studies Pertaining to the Low - Speed Longitudinal Handling Qualities of a Supersonic Transport ) Airplane. Paper presented at AIAA Meeting (Colwabus, Ohio), Aug. 26 - 28, 1963.

4 . Coleman, Thomas L., and Stickle, Joseph W.: Turbulence Environment for Supersonic Transports. Proceedings Symposium on Supersonic Transports. SOC. Exp. Test Pilots, Sept. 1961, pp. 144-160.

5 . Houbolt, John C., Steiner, Roy, and Pratt, Kermit 6.: Flight Data and Considerations of the Dynamic Response of Airplanes to Atmos - pheric Turbulence. Presented to the Structures and Materials Panel and to Flight Mechanics Panel of AGARD (Paris, France), July 3 - 13, 1962.

6. Hunter, Paul A., and Coleman, Thomas L. : Operational Experiences of Jet Transports. SAE Trans., vol. 71, 1963, pp. 34 - 40, 61.

7 . Chaplin, J. C.: Safety and Airworthiness Aspects. Supersonic Engineering, J. T. Henshaw, ed., John Wiley & Sons, Inc., c.1962, pp. 18 - 38.

8. Greene, L. P . , and Bonner, E.: Performance Requirements for Super - sonic Transports.

[Beprinq 674B, SOC. Automotive Eng., Apr. 1963.

9. Silsby, Norman S.: An Analytical Study of Effects of Some Airplane and Landing - Gear Factors on the Response to Runway Roughness With i Application to Supersonic Transports. NASA TN D-1492, 1 9 6 2 .

10. Funk, Jack, and Cooney, T . V . : Some Effects of Yaw Damping on Air - plane Motions and Vertical - Tail Loads in Turbulent Air. NASA MENO 2-17-59L, 1959.

11. Anon.: Tentative Airworthiness Objectives and Standards for Super - sonic Transport Design Proposals. Flight Standards Service, FAA, Aug. 15, 1963.

STRUCTURAL DESIGN SPEEDS VA MANEUVERING SPEED VB ROUGH - AIR SPEED Vc CRUISE SPEED 80 xi03 VD DIVE SPEED

r

ALTITU DE, FT IRFRAME AND ENGINE STRENGTH VERSPEEO MARGIN 0 I 2 3 4 M Figure 1 DESIGN SINKING SPEED I - I I I I I I - D A Y OPERATION I NO. OF LAND1 NGS Figure 2 GUST LOADS

r

suP€usoN/c * O [ ~ I o 3 60

\ FL/GHT

t

A L T Z D E bo 1 ROUGH -

SPEED <<<<e<: CRU ISE 20 SU8SON~C SPEED FLIGHT 1 0 20 25 50 75 0 70 ROUGH AIR DERIVED GUST VELOCITY, FPS Figure 3 PERFORMANCE REFERENCE HIGH ASPECT RATIO LOW ASPECT RATIO I

I I 1 I 1 I

0 0 \ ANGLE OF ATTACK, a Figure 4 , CLIMB GRADIENT SECOND SEGMENT THRUST: I ENGINE OUT

I U

DRAG O R THRUST LOW VIGH 4 1 AIRSPEED -03

. \ J H I G H ASPECT RATIO

CLIMB -02 GRADIENT .01 '.

0 - 10 - 2 0 - 30 AIRSPEED ERROR, AV, KNOTS Figure 3 SPEED STABILITY DURING APPROACH - .4 - .3 THRUST REQUIRED, .2 F - w - . I ( I I I I I I Figure 6 PILOT COMPARTMENT RESPONSE TO RUNWAY ROUGHNESS

- SUPERSONIC TRANSPORT

--- TURBOJET

ACCEL., g UNITS - 1 0 50 100 150 200 SPEED, KNOTS EFFECT OF DAMPING ON VERTICAL - TAIL LOADS SUPERSONIC TRANSPORT CONFIGURATIONS 0 JET AIRPLANES WITH Y A W DAMPER 0 JET AIRPLANES WITHOUT YAW DAMPER OPISTON - ENGINE AIRPLANES

I

0 I 2 3 4 I DAMPING PARAMETER, - 5 / 2 Figure 8 I ~ M E N T O F SATISFACTORY RANDUNG QUALITIES FOR A SOPERSONIC-TRANSPORT CONFIGURATION ‘r THROUGH SIMULATOR STUDIES By Maurice D. White, Alan E. Fay$, Jr., .” - and George E. Cooper Jw , 4 / 5 b . ’ SA4mes Research Center

-

-c- Piloted simulator studies have been conducted of an early supersonic - transport configuration t o provide advanced indications of possible handling - qualities problems. These studies have shown t h a t the configurations being considered f o r t h e supersonic transport r e s u l t i n some handling - qualities problems t h a t are novel f o r a transport class Careful t a i l o r i n g of t h e design may be required t o achieve of airplane.

satisfactory handling q u a l i t i e s .

c-6, m / 2

INTRODUCTION A number of d i f f e r e n t types of configurations are being considered f o r t h e supersonic transport. I n order t o obtain prompt indications of t h e possible handling - qualities problems, simulator studies were made of early versions of t h e various concepts. I n t h i s paper r e s u l t s will be reported f o r a simulator study of a supersonic - transport configura - t i o n t h a t had a variable - sweep wing. The program w a s conducted by using t h e aerodynamic and i n e r t i a l characteristics of a specific design; how - ever, t h e basic s i m i l a r i t y of many of t h e aerodynamic and i n e r t i a l characteristics of t h e t e s t configuration t o those of other supersonic - transport concepts suggests t h a t some of t h e problems discussed may be of concern f o r t h e other supersonic - transport concepts. Two views of Although the general. I t h e t e s t configuration a r e shown i n figure 1.

proportions of t h e t e s t configuration are similar t o those of variable - sweep designs t h a t have evolved more recently, t h e aerodynamic charac - t e r i s t i c s are generally sensitive t o design d e t a i l s . Thus, results f o r the t e s t configuration are not necessarily d i r e c t l y applicable t o t h e newer versions; they a r e of more value f o r demonstrating t h e classes of problems and t h e design trade - offs t h a t must be considered i n t a i l o r i n g t h e design f o r good handling qualities.

I C mean aerodynamic chord referred t o swept - wing configuration pitch damping derivative .

yaw damping derivative longitudinal s t a t i c margin static - directional - stability derivative a t zero angle of attack Mach number 1 M effective value of yaw - damping derivative, body axes

N r '

effective value of s t a t i c - d i r e c t i o n a l - s t a b i l i t y derivative,

NP '

body axes damping r a t i o of short - period longitudinal o s c i l l a t i o n damping r a t i o of Dutch r o l l lateral - directional o s c i l l a t i o n frequency of Dutch r o l l lateral - directional o s c i l l a t i o n wd r a t i o of bank angle t o s i d e s l i p angle i n control - fixed lateral - directional o s c i l l a t i o n s parameter used t o define aileron - yaw effect ( r e f . 1) TESTS The f l i g h t p r o f i l e and weight variation t h a t were used f o r the s i m - u l a t o r study a r e shown i n figure 2. The f l i g h t p r o f i l e w a s an idealized one, but there w e r e no features i n it t h a t would compromise the conclu - The Mach number scale i s divided i n t o three sions drawn from t h e study.

ranges: 0 t o 3.0, representing t h e climb and acceleration t o Mach 3.0 at an a l t i t u d e of 60,000 feet, followed by a f u r t h e r climb a t constant Mach number t o an a l t i t u d e of 70,000 feet; cruise at a constant a l t i t u d e of 70,000 f e e t and Mach number of 3.0 during which the f u e l expenditure (and corresponding decrease i n weight) r e s u l t s i n some s t a b i l i t y param - e t e r changes; and, finally, a deceleration at constant a l t i t u d e , which would be followed by a descent a t constant dynamic pressure. Flight con - d i t i o n s a t Mach numbers below 2.3 i n the descent were not studied a f t e r ? ' J it w a s determined t h a t handling q u a l i t i e s consistently improved a s t h e descent progressed. Also, t h e landing - approach characteristics were not considered i n t h i s program; however, t h e landing - approach handling qual - i t i e s of a more advanced variable - sweep configuration, t h e SCAT 16, a r e discussed i n paper no. 27 by Walter E. McNeill and Robert C. Innis. The wing - sweep schedule provided f o r conkinuous variation of sweep angle with Mach number t o provide e s s e n t i a l l y m a x i m u m l i f t - d r a g r a t i o at each Mach number .

The simulation w a s conducted under instrument f l i g h t conditions with six - degree - of - freedom equations of motion on a simulator t h a t had motion i n roll, pitch, yaw, and side t r a n s l a t i o n ( f i g . 3 ) ; these motions were adjusted t o represent those of a p i l o t compartment located about , 90 f e e t ahead of t h e center of gravity.

I n t h e tests, t h e handling q u a l i t i e s were examined at each of a number of discrete points along t h e mission, i d e n t i f i e d by t h e test A t each point t h e airplane w a s balanced and t h e points i n figure 2.

p i l o t performed maneuvers appropriate t o transport operation, such as turns, precise control of altitude, and so f o r t h .

The p i l o t rating scale shown i n t a b l e I w a s used i n rating the con- figurations. A r a t i n g of 5 would define t h e boundary between satisfac- t o r y conditions and conditions t h a t are only acceptable f o r emergency conditions, and a r a t i n g of 6 - would mark t h e lower l i m i t of acceptabil- i t y i n an emergency, t h e emergency i n the present instance being assumed t o be a stability - augmenter f a i l u r e . The philosophy underlying t h e study w a s t h a t t h e aerodynamic characteristics of the airplane should never r e s u l t i n ratings poorer than $, and t h a t a r t i f i c i a l s t a b i l i t y , augmentation could be r e l i e d on t o improve t h e rating t o b e t t e r than 5.

RESULTS Before entering i n t o a detailed discussion of t h e r e s u l t s it would be desirable t o review t h e background f o r one assumption t h a t had a very powerful e f f e c t on t h e results obtained. Preliminary evaluation ' of t h e rigid - airplane aerodynamics f o r t h e test configuration indicated t h a t if a constant center - of - gravity position were assumed t h a t would provide longitudinal s t a b i l i t y at low speeds, t h e high - speed trim - drag penalties would be prohibitive. This i s t h e well - known e f f e c t of aerodynamic - center shift between subsonic and supersonic speeds which has been a matter of concern f o r aJ-1 supersonic - transport configurations.

Other papers i n t h i s compilation indicate t h e large l e v e l of e f f o r t t h a t is being applied t o t h e development of d e t a i l modifications t o reduce t h e aerodynamic - center s h i f t s . These efforts, i n combination with beneficial. aeroelastic effects, should provide solutions t h a t do not involve in - flight adjustment of center - of - gravity position. How - ever, i n t h e absence of a well - defined aerodynamic solution, it w a s decided at the time of t h e simulation tests t h a t center - of - gravity adjustment by f u e l t r a n s f e r could be assumed.

I n figure .$(a), t h e center - of - gravity movement t h a t w a s used t o provide a constant s t a t i c margin of 5 percent is shown as a percentage of the swept - wing mean aerodynamic chord. The maximum physical movement

associated with t h i s curve is 1 7 . 7 feet, a very large distance which

would, of course, severely tax any fuel - transfer - system design t h a t might be considered f o r achieving t h i s center - of - gravity movement. I n fact, it i s unlikely t h a t a final. design would be accepted t h a t incorporated fuel - transfer requirements of t h i s magnitude. However, center - of - gravity adjustments of smaller magnitude could conceivably be incorpo - rated i n a f i n a l design, and the present results, even though they a r e of exaggerated magnitude, remain of i n t e r e s t i n t h a t they define problem areas and trade - offs t h a t need consideration.

The most important e f f e c t s of t h e assumption regarding center - of - gravity movement a r e shown i n figure 4(b). Here i s shown a comparison of t h e s t a t i c - d i r e c t i o n a l - s t a b i l i t y derivative Cnpo as it would vary with Mach number i f t h e center of gravity were held fixed and as it varied with t h e assumed center - of - gravity t r a v e l . It i s apparent t h a t a considerable l o s s i n directional s t a b i l i t y occurred at high Mach num - bers as a consequence of t h e assumed center - of - gravity travel, and t h i s l o s s had a significant e f f e c t on t h e r e s u l t s obtained. The implica - t i o n s of t h i s r e s u l t w i l l be discussed further.

Figure 5 shows some of t h e r e s u l t s t h a t were obtained i n t h e t e s t s .

Plotted against Mach number a r e t h e p i l o t ratings and several s t a b i l i t y parameters t h a t indicate t h e nature of t h e handling - qualities problems -___ t h a t were encountered.

Plotted against a Mach number scale t h a t cor - responds t o t h a t of figure 2 are: (1) P i l o t r a t i n g (2) The Dutch roll damping r a t i o (d, which i s a measure of t h e /I control - fixed damping of t h e lateral - directional oscillations. (Desir - able values f o r t h i s parameter a r e equal t o o r greater than the range 0.4 t o 0.7.)

( 3 ) The pitch damping r a t i o (, which i s a measure of the damping of the control - fixed short - period longitudinal oscillation. (Desirable values f o r t h i s parameter are a l s o equal t o o r greater than t h e range 0.4 t o 0.7.)

i 3 ' i d > ' > V Y I I

u, which is a measure of control - fixed

( 4 ) The roll - yaw r a t i o

IPI

(Minimum cross - coupling e f f e c t s i n lateral - directional oscillations.

values f o r t h i s quantity a r e desired.)

( 3 ) The aileron - yaw - effect parameter (zr, which describes the

(Desir - cross - coupling e f f e c t introduced i n aileron - control maneuvers.

able values f o r t h i s parameter a r e i n the region of 1.0.)

Consider first t h e v a r i a t i o n of p i l o t r a t i n g f o r t h e basic airplane; the spread i n t h e data covers t h e s c a t t e r of d a t a from f i v e participating t e s t p i l o t s . A s shown i n figure 5, t h e airplane w a s s l i g h t l y unsatis - factory at t h e lowest test airspeeds, and as t h e Mach number increased through t h e climb and acceleration, t h e airplane handling q u a l i t i e s became progressively worse u n t i l at a Mach number of 3.0 and an alti - tude of 'j'O,OOO feet t h e airplane w a s v i r t u a l l y unflyable. With decr6asing weight (and correspondingly decreasing angle of attack) through t h e cruise, t h e characteristics improved somewhat, and there w a s a very s l i g h t f u r t h e r improvement during t h e deceleration at con - s t a n t a l t i t u d e p r i o r t o descent. The p i l o t s ' comments i d e n t i f i e d t h e i n i t i a l degradation i n p i l o t rating with Mach number at subsonic speeds with t h e decrease i n Dutch roll damping. The reduced l e v e l of Dutch roll damping attained at sonic speeds persisted at all supersonic speeds and should, without f u r t h e r repetition, be inferred t o be a f a c t o r i n t h e problems encountered at higher speeds. A s Mach number increased through t h e transonic range, t h e emphasis of t h e p i l o t complaints s h i f t e d t o l a t e r a l - d i r e c t i o n a l cross - coupling effects, both i n control - fixed

o s c i l l a t i o n s - '" and aileron control maneuvers (Zr. Further

I P i

increase i n Mach number t o 1.6 indicated more serious complaints about lateral - directional characteristics; i n addition, t h e reduced longitu - d i n a l damping contributed t o d i f f i c u l t i e s i n holding a l t i t u d e precisely i n turns or i n leveling off after climbs. Further continuation along t h e climb t o cruising speeds and a l t i t u d e s resulted i n marked deteriora - t i o n of lateral - directional coupling parameters and of p i t c h damping, which reflected t h e poorer ratings assigned by t h e p i l o t s .

The handling - qualities problems i d e n t i f i e d i n these operations I have been t h e subject of studies which have l e d t o t h e development of Boundaries derived numerical c r i t e r i a f o r satisfactory conditions.

fromthese studies f o r two of t h e f a c t o r s are shown i n f i g u r e 5 . The boundary i n figure 5(d) w a s derived from reference 2, and t h e boundary i n figure 5(e) w a s derived from reference 1. These boundaries represent

a &, p i l o t rating; they w e r e , however, developed from fighter - airplane

studies and primarily with consideration of t h e landing - approach condi - tion. The agreement shown, over of these boundaries and t h e area where t h e p i l o t r a t i n g crossed t h e 6- 1 l e v e l provides encouraging indication t h a t these c r i t e r i a may be applied more generally.

A variety of s t a b i l i t y augmentation arrangements w a s examined i n an e f f o r t t o achieve a Rracticable solution. A s shown i n figure 6, a f a i r l y reasonable arrang ment t h a t provided p i l o t ratings a t o r below

1 e

6 - throughout t h e mission w a s t o increase t h e s i z e of t h e vertical. t a i l The directional by 50 percent and increase p i t c h damping by 4 t i m e s .

s t a b i l i t y recovered by t h i s increase i n t a i l s i z e w a s about t h e same as A change i n horizontal- t h a t l o s t through center - of - gravity movement.

control gearing t o provide a 75-percent reduction i n s e n s i t i v i t y o r power at high Mach numbers w a s a l s o required, but t h i s change should probably not be considered as s t a b i l i t y augmentation.

t o With an additional increase i n pitch - damping derivative 1 1 times the basic value and t h e yaw - damping derivative Cnr t o

10 times the basic values, p i l o t ratings of t h e order of 9 were

M = 3.0. Although t h i s s e t of achieved at t h e c r i t i c a l condition of values w a s not evaluated throughout the Mach number range, it seems reasonable t o deduce from t h e trends of t h e data t h a t t h e improvements provided by t h i s augmentation would hold throughout t h e mission. Sub - s t a n t i a t i n g a point made i n t h e " Introduction, " it w i l l be noted t h a t a delta - wing - canard model previously studied on the simulator i n (See r e f . 3 . ) Mach 3 flight required similar amounts of augmentation.

It will be noted t h a t dihedral effect, which i s usually considered a primary f a c t o r i n cross - coupling problems, w a s not reduced as one of t h e augmentation changes. I n fact, an attempt t o reduce dihedral e f f e c t The explanation a t supersonic speeds resulted i n a more adverse rating.

f o r t h i s effect l i e s i n t h e f a c t that, f o r t h e marginal directional stability t h a t existed at high Mach numbers, most of t h e apparent l a t e r a l - d i r e c t i o n a l s t a b i l i t y w a s contribuked by the dihedral effect.

Reduction of t h i s contribution w i t h lowered dihedral. e f f e c t affected t h e lateral - directional s t a t i c s t a b i l i t y more adversely than it improved It i s pertinent i n connection with these results t o roll - yaw coupling.

i r e c a l l t h e assumption regarding center - of - gravity movement discussed previously, which resulted i n greatly reduced directional s t a b i l i t y at The character and magnitude of t h e contributions of high Mach numbers.

dihedral e f f e c t s 'that were j u s t indicated would be g r e a t l y affected by t h a t assumption.

Another point of i n t e r e s t developed by t h e studies i s indicated i n Contrary t o t h e usual ekpectation, t h e Dutch roll damping figure 7.

ratio, which i s an important measure of lateral - directional damping, w a s actually reduced with t h e 50 - percent increase i n v e r t i c a l - t a i l area.

of course, due t o t h e r e l a t i v e contributions of the This r e s u l t is, increased t a i l s i z e t o t h e directional damping and t o t h e s t a t i c direc - t i o n a l s t a b i l i t y , which are indicated t o a first order i n the lower p a r t of t h e figure. A s noted, the percentage change i n Np' (an important s t a t i c - s t a b i l i t y 'or frequency term) i s much greater than the percentage change i n Nr' (an important damping term);.hence, t h e reduction i n damping r a t i o previously noted.

These data demonstrate why it may prove d i f f i c u l t t o provide required l e v e l s of yaw damping by simple increase i n t a i l s i z e , Simul- taneous changes i n s t a t i c directional s t a b i l i t y would increase t h e f r e - quency of directional o s c i l l a t i o n s t o undesirably high levels; i n t h e current operations, p i l o t s noted t h a t the period of 5 1 seconds t h a t occurred with t h e 50 - percent increase i n t a i l s i z e did not seem a rea - sonable value f o r t h e long, r e l a t i v e l y limber fuselage t h a t would prob - ably characterize t h e supersonic transport. This r e s u l t would indicate a possible requirement f o r achieving damping by a r t i f i c i a l means, even i f t h e use of increased t a i l s i z e could have been tolerated otherwise.

Here again the basic assumption regarding t h e effect of center - of - gravity position on directional s t a b i l i t y would have a significant e f f e c t on some rather important conclusions.

CONCLUDING REMARKS I n conclusion, it appears from simulation studies of an early supersonic - transport design t h a t coping with t h e problems of Longitu - dinal aerodynamic - center s h i f t with Mach number can present d i f f i c u l t i e s o t h e r than the well - known trim - drag effects. The s e n s i t i v i t y of the directional s t a b i l i t y t o longitudinal center - of - gravity position i s one of a number of f a c t o r s t h a t r e s u l t i n r e l a t i v e l y novel l a t e r a l - directional handling - qualities problems which may require careful tai - loring of t h e design. I n cruising f l i g h t , augmentation of aerodynamic damping appears t o be a requirement f o r both t h e variable - sweep and t h e delta - wing - canard configurations.

?

REFERENCES 1. Vomaske, Richard F., Sadoff, Melvin, and Drinkwater, Fred J., 111: The Effect of Lateral - Directional Control Coupling on P i l o t Control of an Airplane as Determined i n Plight and i n a Fixed - Base Flight Simulator. NASA TN D-1141, 1961.

2. Anon.: Flying Qualities of Piloted Airplanes. Military Specifica - t i o n MIL-F-8785(ASG), Sept . 1, 1954; Amendment - 2, O c t . 17, 1955.

3. White, Maurice D., Vomaske, Richard F., McNeill, Walter E., and A Preliminary Study of Handling - Qualities Cooper, George E.: Requirements of Supersonic Transports i n High - speed Cruising Flight Using Piloted Simulators. NASA TN D - 1888, 1963.

PILOT OPINION RATING SYSTEM I I I PRIMARY

1 ADJECTIVE NUMERICAL

CAN BE

RATING I RATING I DESCRIPTION

MISSION LANDED ACCOMPLISHED I I I A 0 z I EXCELLENT, INCLUDES OPTIMUM GOOD, PLEASANT TO FLY

1 SATISFACTORY I '3 1

SATISFACTORY, BUT WITH SOME MILDLY z n 0 UNPLEASANT CHARACTERISTICS YES YES ACCEPTABLE, BUT WITH UNPLEASANT CHARACTERISTICS UNSATISFACTORY UNACCEPTABLE FOR NORMAL OPERATION ACCEPTABLE FOR EMERGENCY CONDlTlOh W

ONLY '

Z I I 7 I UNACCEPTABLE EVEN FOR EMERGENCY v CONDITION' DOUBTFUL

5 I UNACCEPTABLE I I

W UNACCEPTABLE - DANGEROUS I 9 I UNACCEPTABLE - UNCONTROLLABLE 'FAILURE OF A STABILITY AUGMENTER T E S T CONFIGURATION c

i

Figure 1 SIMULATOR TEST CONDITIONS ALTITUDE, f t GROSS WEIGHT, Ib

'."'ii 0 1 2 3 3 2 I

+CLIMB--(CRUISE*DESCENT+ MACH NUMBER Figure 2 \ i i MOTION SIMULATOR A - 29190 Figure 3 CENTER - OF - GRAVITY EFFECTS C.G. MOVEMENT REQUIRED - - T O PRODUCE CONSTANT VALUE OF aC,/dCL

PERCENT c

A F T /FIXED C.G. POSITION

cn& : : I b ) - /CONSTANT C.G. ADJUSTED a c , FOR lacL

0 I 2 3 M Figure 4 3 ' i VARIATIONS OF SELECTED HANDLl NG QUALITIES PARAMET PILOT 4 RATING 0 I 2 3 3 2 MACH NUMBER .4 r C U ME CRUISE DESCENT

' d a2 t~~~ DUTCH ROLL DAMPING I

I AILERON - YAW EFFECT r-, 0 1 2 3 3 2 -bo I 2 3 3 2 MACH NUMBER MACH NUMBER CLIMB CRUISE DESCENT CLIMB CRUISE DESCENT Figure 5 STABILITY AUGMENTATION REQUIREMENTS AUGMENTATION 1. INCREASE VERTICAL TAIL SIZE 50% 2. INCREASE PITCH DAMPING 4 TIMES 3. a)lNCREASE PITCH DAMPING I1 TIMES b) INCREASE YAW DAMPING IO TIMES AUGMENT I AUGMENT 1+2 \ / U ACCEPTABLE PILOT SATISFACTORY RATING SATISFACTORY I I I I 0 ' I 2 3 3 2 MACH NUMBER Figure 6 .

EFFECT OF T A I L SIZE ON DAMPING RATIO .2r BASIC D U T C H R O L L DAMPING R A T I O i S T A T I C D I R E C T I O N A L S T A B I L I T Y

-

C L I M B C R U I S E DESCENT 0 1 2 3 3 2 M A C H N U M B E R Figure 7 -

$6

0 q-JP 27. A SIMULATOR STUDY O F THE LATERAL DIRECTIONAL HANDLING QUALITIES OF TWO SCAT CONFIGURATIONS I N THE LANDING A P P R O A C H P- N A S A - -

By Walter E. McNeill and Robert C. Innis &

NASA&Ames Research Center A piloted simulator study of the lateral directional handling qual - i t i e s of one SCAT 16 and two SCAT 17 configurations showed t h a t the SCAT 16 and one SCAT 17 would have acceptable handling q u a l i t i e s i n an emergency condition without s t a b i l i t y augmentation.

The SCAT 16 exhib - i t e d almost satisfactory behavior without augmentation.

It w a s found that these SCAT 16 and 17 configurations could be impoved t o a satis - factory level, as far as l a t e r a l directional s t a b i l i t y i s concerned, by .

ordinary means of s t a b i l i t y augmentation.

v>

INTRODUCTION A s i n the case of many large, slender airplane configurations, the present supersonic transport designs, o r SCAT configurations, are expected t o have undesirable handling q u a l i t i e s which w i l l require some form of a r t i f i c i a l s t a b i l i t y augmentation i n the lateral directional modes of motion. It i s the purpose of t h i s paper t o present and discuss some preliminary r e s u l t s of a piloted simulator study which indicate the types and amounts of lateral. directional augmentation desirable f o r two

I

SCAT configurations, the SCAT 16 and the SCAT 1 7 .

I n t h i s study, only the landing - approach condition w a s investigated.

SmBOLS b wing span, f t

-

reciprocal of cycles required f o r lateral o s c i l l a t i o n t o damp 6112 t o half amplitude I period, sec r o l l i n g velocity, radians/sec dynamic pressure, lb/sq f t yawing velocity, radians/sec wing reference area, sq f t t r u e airspeed, f t / s e c angle of attack, deg s i d e s l i p angle, deg bank angle, deg r a t i o of bank - angle amplitude t o sideslip - angle amplitude i n t h e lateral o s c i l l a t o r y mode aileron deflection, radians rudder deflection, radians Rolling moment

cz =

SSb Yawing moment SSb ac, - v

c 2 P - ap

TESTS Configurations Inve st igated The plan views and major size parameters of the S C A T configurations T h r e e designs f o r studied on the simulator are presented i n t a b l e I.

which pertinent information w a s available a r e considered as they existed It should be noted that there are at t h e t i m e of the midterm review.

some differences between t h e basic configurations shown here and the The results of t h i s study, f i n a l Yersions evolved by the contractors.

therefore, should be interpreted i n a general fashion as indications of probable augmentat ion requirements .

The SCAT 16 i s shown i n table I with t h e w i n g s i n the forward posi - t i o n a t 30° of sweep; the landing weight w a s 242,500 pounds; the approach t airspeed w a s 138 knots; t h e trim angle of a t t a c k f o r the approach w a s 4.2O.

The infor - The other configuration shown i n table I i s t h e SCAT 17.

mation presented pertains t o two versions of the S C A T 17 which differed rather significantly and thus were f e l t t o deserve separate attention i n The SCAT 17-A con - These are shown labeled 17-A and 17-B.

t h e study.

figuration had a canard surface f o r primary p i t c h control and an aft horizontal t a i l f o r t r i m ; the SCAT 17-B had a canard f o r t r i m and elevons These two versions had landing f o r p i t c h control, but no horizontal tail.

weights of 253,300 pounds and 210,000 pounds and approach speeds of 132 knots and 145 knots. The approach angles of a t t a c k of 12O and 7 O f o r the SCAT 17-A and the SCAT l7-B, respectively, should be noted; the difference w a s due primarily t o the difference i n approach speeds.

Piloted Simulator The simulator used i n t h i s study consisted of a fixed, transport - type cockpit containing the e s s e n t i a l p i l o t controls, instruments, and external v i s u a l display f o r performing I L S landing approaches t o touch - down with minimum v i s i b i l i t y . A photograph of the cockpit i s presented The standard f l i g h t and engine instruments are shown, as i n figure 1.

well as the view of t h e runway and approach lights projected on a screer 12 f e e t from t h e p i l o t . Throttles f o r control of engine thrust a l s o were available. The visual motion cues were provided by a landing simu - l a t o r which employs closed - circuit television f o r transmission of thk v i s u a l image. The motions and t h e instrument readings were generated from six - degree - of - freedom equations of motion by using a general - purpose analog computer. Applications of the simulator and more detailed descrip - t i o n s thereof a r e available i n references 1, 2, and 3 .

Evaluation Tasks " he evaluating p i l o t s were two NASA research p i l o t s and t w o company t e s t p i l o t s experienced i n f l y i n g large commercial j e t a i r c r a f t .

The p i l o t s f i r s t familiarized themselves with the general handling q u a l i t i e s of a p a r t i c u l a r basic SCAT design at the given approach speed and assessed the longitudinal dynamics. I n a l l cases, the longitudinal behavior proved t o be satisfactory. The lateral directional character - i s t i c s of t h a t configuration were then rated according t o the widely used Cooper r a t i n g scale (ref. 4) a f t e r a number of specified tasks were performed. These tasks, which are often required i n air transport oper - ation during the instrument approach and the subsequent landing, a r e as follows F i r s t were l a t e r a l or Dutch roll o s c i l l a t i o n s with controls fixed; second were ILS approaches and landings i n s t i l l air, mild t o moderate 'J rough air, and i n steady crosswinds requiring removal of d r i f t or crab angle immediately p r i o r t o touchdown; and t h i r d were o f f s e t s which called f o r corrective sidestep t u r n s t o l i n e up with t h e runway immedi - The remaining a t e l y following t r a n s i t i o n from ILS t o v i s u a l f l i g h t .

t a s k w a s t h a t of correcting, with f l i g h t controls only, f o r f a i l u r e of an outboard engine during a go- around from an aborted landing.

> $ 3 > > . # J ' * . P r* *.

S t a b i l i t y Augment at ion The p i l o t ratings and comments obtained during the previously men - tioned maneuvers were used as c r i t e r i a f o r adding s t a b i l i t y augmentation.

I n general, the procedure w a s t o vary, one by one, the s t a b i l i t y deriv - a t i v e s which the p i l o t believed offered the best m o v e m e n t i n lateral directionalbehavior, and then t o arrive at a combinatZm of augmented derivatives which should produce ztn overall r a t i n g of satisfactory. The r e s u l t i n g a u m n t e d configuration w a s then assessed on the bases of t h e tasks j u s t described t o v e r i f y t h e satisfactory behavior.

I n order to include p r a c t i c a l e f f e c t s of limited augmenter authority and aerodynamic cross coupling, variation of the derivatives w a s i n most cases accomplished through the roll- and yaw - control surfaces. : RESULTS AND DISCUSSION Basic Configurations All the SCAT configurations were found t o require some form of sta - b i l i t y augmentation t o provide satisfactory handling q u a l i t i e s f o r normal operation i n t h e landing approach. Before consideration i s given t o t h e augmentation, however, it would be w e l l t o look at the principal f a c t o r s wh$ch the evaluating p i l o t s f e l t t o be objectionable i n t h e basic config - urAtions. These are summarized f o r each design i n table 11. The reader i s reminded that the configurations studied represent interim designs that existed a t t h e t i m e of the midterm review.

F i r s t , the SCAT 16 w a s characterized by large adverse s i d e s l i p during t u r n e n t r i e s and a d i f f i c u l t y i n controlling heading accurately.

Directional s t a b i l i t y Cn and yaw due t o r o l l i n g C were immediate P nP candidates f o r augmentation. Low Cn would account for the poor P (which had heading cantrol and would a l s o accentuate any e f f e c t of

cnp

a negative sign i n t h i s case) i n producing adverse sideslip. While these c h a r a c t e r i s t i c s were objectionable i n the midterm configuration, they are not necessarily ty - pical of the variable - sweep concept. I n s p i t e of these ,!

q u a l i t i e s , the SCAT 16 w a s considered marginally satisfactory.

The SCAT 17-A exhibited l o w dmping of the lateral dirtectional or Dutch roll oscillation, which w a s e a s i l y excited i n t u r n e n t r i e s and i n rough air. Other objectional q u a l i t i e s were high dihedral e f f e c t and low roll damping, which contributed t o the unsatisfactory Dutch roll chasacteristics and which proved bothersome i n removing the crab angle during a crosswind landing and i n rough air. Adverse s i d e s l i p i n turn L C (I e n t r i e s , which undoubtedly w a s accentuated by t h e highly inclined roll axis at 12O angle of attack, a l s o w a s noted. T h i s configuration w a s rated generally unsatisfactory t o unacceptable f o r normal operation.

The S C A T 17-B w a s troubled by a .high r a t i o of bank t o s i d e s l i p i n the Dutch roll o s c i l l a t i o n (though the damping w a s not objectionable) ; however, the principal problem which made this configuration unacceptable even for an emergency condition without augmentation w a s an e a s i l y excited second lateral o s c i l l a t o r y mode. During such ordinary tasks as an I L S approach, t h i s mode apparently contributed t o a continual r o l l o s c i l l a - t i o n which the p i l o t s had d i f f i c u l t y distinguishing from unacceptable Dutch roll behavior. Pulse - type aileron control appeared t o subdue the oscillation, but extremely close attention t o bank a t t i t u d e w a s required.

U s e of rudder control resulted only i n making the problem more d i f f i c u l t .

the p i l o t s complained of excessive dihedral e f f e c t and I n addition, excessive favorable, o r " proverse " , aileron yaw, both of which probably contributed t o the roll - control problem discussed previously.

Desired Augmentation The degrees of augmentation found desirable for the various deriva - t i v e s of the SCAT configurations studied are shown i n figure 2. The mounts of augmentation of the s t a t i c derivatives which determine dihe -

dral e f f e c t C z , directional s t a b i l i t y Cn , and aileron yaw C are

P P nga presented. The variations indicated are i n terms of the basic or unaug - mented value. For each case, both a nominal and an actual. variation from t h e basic r a t i o of 1.0 are given. The nominal. variations are those which would be obtained with pure derivative changes o r with no cross coupling; the a c t u a l variations represent the net eff e c t s (through aero - dynamic cross coupling) on rolling - moment derivatives due t o rudder and on yawing - moment augmentation (changing CnP can a l s o change c 2 P ) derivatives due t o aileron augmentation. A l l variations indicated i n figure 2 except f o r C were obtained through the control surfaces; nga C w a s varied as a pure derivative change t o allow some freedom i n nga by design of the roll - control surfaces. The augmenter t a i l o r i n g C I nga authority i n both roll and yaw was one - third of the t o t a l control t r a v e l .

T h i s authority appeared t o be more than adequate f o r a l l normal maneuvers.

Inspection of figure 2 reveals some differences and some similar - i t i e s i n the desired augmentation which one might expect from what w a s learned about the basic configurations.

Both the S C A T 17-A and 17-B

required a sizable decrease i n dihedral e f f e c t - t o about 30 t o

c z P 50 percent of t h e basic value. A l l the p i l o t s indicated t h a t t h e S C A T 16, because of the heading - control problems mentioned previously, required about twice the directional s t a b i l i t y C of the basic configuration.

nS To decrease adverse s i d e s l i p i n t u r n e n t r i e s and t o aid i n controlling and hence roll, i n the event of asymmetric engine f a i l u r e , the sideslip,

. The SCAT’ l7-B, how -

S C A T 17-A a l s o required about twice the basic CnP ever, w a s made satisfactory without changing C n , probably because of, P first, t h e overshadowing improvement i n the overall lateral o s c i l l a t o r y behavior brought about by t h e other augmentation and, second, the f a c t that no adverse yaw o r s i d e s l i p problem existed i n the basic airplane.

Indeed, the S C A T 17-B required about a 73 - percent decrease i n proverse aileron yaw C t o be considered satisfactory.

nga Figure 3 shows t h e corresponding augmentation desired f o r t h e rotary o r damping derivatives. A l l variations indicated were accomplished through actuation of t h e control surfaces, except for CZ,. Consider As one might expect, the r e l a t i v e l y high first the roll damping C2 P* The aspect r a t i o SCAT 16 required no augmentation of roll damping.

S C A T 17, however, w a s found t o benefit from t h e addition of a roll damper which provided an additional 100 t o 150 percent of basic roll damping.

T h i s increased damping was especially advantageous i n rough air and i n a l l r o l l i n g maneuvers, so long as sufficient aileron - control power also w a s available.

Both the SCAT 16 and t h e S C A T 17-A benefited from Cn augmentation sufficient t o r e s u l t i n a value of the derivative of equal magnitude but of opposite sign ( t h a t is, changed from negative t o positive). T h i s w a s found t o a i d significantly i n the apparent adverse yaw problems of the basic configurations. Since the value of. C n f o r t h e S C A T 17-B w a s P already positive i n sign, no change w a s made i n that derivative.

The Dutch r o l l behavior of t h e S C A T 16 and the S C A T 17-A were improved by yaw d q e r s which increased t o three times the basic Cnr value. As will be shown later, this amount of y a w damping i n the S C A T 16 I n j did not bring about a large iqrovement i n t h e Dutch roll damping.

’ f a c t , the p i l o t s complained of having t o hold excessive rudder control into steady turns with the yaw damping indicated. O f course, s t a b i l i t y augmentation designers are w e l l aware of t h i s problem and suitable wash - out networks are usually the answer. A s indicated, the SCAT 17-B w a s improved s u f f i c i e n t l y without a p w damper.

A s shown i n the last graph of figure 3 , a nominal variation of roll t o a value of zero w a s found t o be helpful t o the due t o yawing ‘2, i J . 2 - i 3 " r o l l following asymmetric engine failure. S C A T 17-A in suppression of change i n Cnr, however, t h e net decrease Because of the simultaneous i n C z , w a s very small. Variation of C z r w a s not investigated fur - t h e r i n the other two airplanes.

O n e i t e m of a funaamental, though interesting, nature i s shown i n figures 2 and 3j that is, the precise variation desired in a given derivative i s not always obtained when augmenting through the control surfaces, For exaaple, if one i s attempting t o augment at the same time m d C aerodynamic cross coupling can come into play and modify

% a p '

The required augmenter gains the desired changes i n both derivatives.

can, of course, be readily computed, and proper compensation introduced, C but r e l i a b l e results depend on accurate bowledge of the values of n% Lateral Oscillatory Characteristics The improvements i n the l a t e r a l o s c i l l a t o r y or Dutch roll charac - t e r i s t i c s of the augmented SCAT configurations over those of t h e unaug - mente4 configurations are shown i n figure 4 . The characteristics are

indicated i n terms of the damping parameter - and the bank-to-

c1/2 side s l i p rat i o h i ! ! .

The scale change on the abscissa should be noted.

I p I

The basic characteristics are indicated by the black symbols and the corresponding behavior w i t h augmentation sufficient t o produce an over - al+. p i l o t r a t i n g of satisfactory i s indicated by the light symbols. The numbers t o the l e f t of the symbols are the periods i n seconds. The boundary indicates the degree of damping found by p i l o t s i n a flight study w i t h a v a r i a b l e - s t a b i l i t y aimlane (ref. 5 ) t o be satisfactory f o r normal operation i n the landing approach. To indicate current behavior,

t h e Boeing 707-320 without a yaw damper has a value of - of 0.54

c1 19

\ ; and a value of of 1.57 w i t h a period of 8.2 seconds.

I PI

Figure 4 indicates t h a t various degrees of change were achieved in

A s mentioned e a r l i e r , the improvement f o r the the Dutch r o l l behavior.

Without any augmentat ion at all, the SCAT 16 configuration w a s slight.

Dutch roll characteristics of - the S C A T 16 were rated satisfactory by t w o p i l o t s .

' s a t * P > > I r i r .

I ) For the SCAT l7-A, a dramatic improvement i n damping and about a 50 - percent reduction i n bank - to - sideslip rat i o were brought about by the combined e f f e c t s of the several augmented derivatives.

Not all the aug - mentation w a s aimed d i r e c t l y toward the Dutch roll behavior, but the large increase i n damping shown did result. The increase i n period should a l s o be noted; no complaints were made about tbis.

The two lateral o s c i l l a t o r y modes of the SCAT 17-8 i n the unaug - mented condition are shown i n figure 4.

They are not greatly different i n period and have comparable bank - to - sideslip ratios. I n the augmented case, only one o s c i l l a t o r y mode remains and the other splits i n t o the usual aperiodic spiral and r o l l i n g modes. The Dutch roll damping of the augmented S C A T l7-B, while not as high as f o r either of the unaugmented i modes, i s s t i l l respectable and the reduction i n bank - to - sideslip r a t i o and elimination of the coupling problems associated w i t h the two oscil - l a t o r y modes resulted i n satisfactory lateral directional handling q u a l i t i e s .

Some remaining mildly objectionable characteristics were noted: For the SCAT 16, the additional y a w damping required excessive rudder i n t o steady turns and some d i f f i c u l t y holding a heading w a s s t i l l reported; f o r the S C A T LT-A, rudder control required i n turns w a s s t i l l considered excessive and even f u r t h e r reduction i n dihedral e f f e c t w a s considered desirable.

CONCLUDING R E M A R K S It has been shown through use of a piloted simulator that the SCAIT 16 and one version of the S C A T 17 should have acceptable handling q u a l i t i e s i n an emergency condition without s t a b i l i t y augmentation and that the SCAT 16 studied actually should border on satisfactory behavior without augmentation. It has also been shown that all the SCAT con - figurations investigated on t h e simulator can be improved t o a satis - factory level, as far as lateral directional handling q u a l i t i e s are con - cerned, by ordinary means of s t a b i l i t y augmentation. Unquestionably, some of the improvement i n handling q u a l t i t i e s effected i n this study could be achieved by careful design of the airplane (and estimation of the c r i t i c a l s t a b i l i t y derivatives, especially those which are sensitive t o high - lift configuration), thus reducing the requirements placed on the augmentation equipment.

REFERENCES 1. White, M. D . , Sadoff, M., Bray, R. S., and Cooper, G. E.: Assessment of C r i t i c a l Problem Areas of the Supersonic Transport by Means of Piloted Simulators. Paper N o . . 62 - 20, I n s t . Aerospace Sci., Ja.

1962.

2 . Bray, Richard S . : Piloted Simulator Studies Pertaining t o t h e Low - Speed Longitudinal Handling Qualities of a Supersonic Transport Airplane. NASA paper presented at Ai% Meeting (Columbus, Ohio) , Aug. 26 - 28, 1963.

3 . McNeill, Walter E.: A Piloted Simulator Study of the Loss of Altitude g by a Jet Transport i n a Go - Around From an Instrument - Landing Approach. N A S A TN D - 2060, 1963.

Understanding and Interpreting P i l o t Opinion.

4. Cooper, George E. : Aero. Eng. Rev., vol. 16, no. 3 , Mar. 1957, pp. 47 - 51, 56.

5. McNeill, Walter E., and Vomaske, Richard F. : A Flight Investigation To Determine the Lateral Oscillatory Damping Acceptable for an Air - N A S A MEMO 12-10-58~, 1959.

plane i n the Landing Approach.

TABU I.- SCAT CONFIGURATIONS STTJDIXD . . . . . . . . . . . . . . . . . . . . . .

S p a n , f t . . 154 . . . . . . . . . . . . . . . . . . .

W i n g a r e a , s q f t . 2,540

Landing weight, lb . . . . . . . . . . . . . . . . . . . 242,500

. . . . . . . . . . . . . . . .

Approach velocity, knots 138

a , d e g . . . . . . . . . . . . . . . . . . . . . . . . . 4.2

S C A T 17 + + +

-~ ~ SCAT 1 7 - ~ . . . . . . . . . . . . . . . . . . . . . . . .

S p a n , f t 105

Wing area, s q f t . . . . . . . . . . . . . . . . . . . . 5,100

Landing weight, lb . . . . . . . . . . . . . . . . . . . 253,300

Approach velocity, knots . . . . . . . . . . . . . . . .

a , d e g . . . . . . . . . . . . . . . . . . . . . . . . . 12

SCAT 17-~ S p a n , f t . . . . . . . . . . . . . . . . . . . . . . . .

Wing area, sq f t . . . . . . . . . . . . . . . . . . . . ?,a0

Landing weight, lb . . . . . . . . . . . . . . . . . . . 210,000

Approach velocity, knots . . . . . . . . . . . . . . . .

. . . . . . . . . . . . . . . . . . . . . . . . .

a , d e g 7

TABLE 11. - W O R OBJECTIONAB~ CHARACTERISTICS

SCAT 16

-+

, Large adverse s i d e s l i p i n t u r n entries Difficult t o maintain heading SCAT 17 SCAT 1 7 - ~ Low Dutch roll damping High dihedral e f f e c t Low roll damping Adverse s i d e s l i p i n turn e n t r i e s SCAT 1 7 - ~ High Dutch roll bank - to - sideslip r a t i o Second l a t e r a l o s c i l l a t o r y mode High dihedral e f f e c t High proverse aileron yaw SIMULATOR INSTRUMENTS AND VISUAL DISPLAY A - 31061 Figure 1 DESIRED SCAT AUGMENTATION STAT I C DE R I VAT1 V E S NOMINAL C C -(UNCOUPLED) nda /‘ndagmlC ACTUAL I 2 1 2 3 0 I ‘(COUPLED)

-m rn

SCAT 1 6 SCAT 17 - A

+

SCAT 17-8 Figure 2 DESIRED SCAT AUGMENTATION ROTARY DERIVATIVES SCAT 1 6

I

SCAT 17 - A

h

SCAT 17 - 8

i

EFFECTS OF AUGMENTATION ON LATERAL OSCILLATORY BEHAVIOR 8 - 11.5

/* 0 SCAT 1 6

7 - Ps9.0 // / / 6 - / 0 SCAT 17 - 8 \ / \ / DAMPING, 5 - I

-

4 - c1/2 3 - 2 - 8.

NORMAL OPERATION 9.8 6.2 NASA MEMO 12 - 10 - 58A

' t

. _ I I I A 1 1 I 0 I 2 v 4 6 8

ROLL - SIDESLIP PARAMETER, - 144

1/31 Figure 4 28. PILOTING PR0BI;EMS ASSOCIATED WITH THE SUPERSONIC - TRANSPORT MISSION PROFILF: By Glen W. Stinnett, Alan E. Fay6, Jr., a n d m e t t D. Fry d - & , & S A - - - - - .

SUMMARY A simulator study w a s conducted of t h e piloting problems asso- ( ked with t h e supersonic - transport mission profile. Several problem areas not common t o present transport operations were examined and s o h - t i o n s f o r these problems are suggested.

3- -

INTRODUCTION During t h e simulator study described by White, Fay&, and Cooper i n paper no. 26, it w a s convenient t o examine the handling qualities of t h e supersonic transport (SST) at discrete f l i g h t conditions along the mis - sion profile, where a l t i t u d e and Mach number were held constant f o r each f l i g h t condition. I n actual operation, however, t h e p i l o t i s confronted with t h e task of continuous f l i g h t along the p r o f i l e where a l t i t u d e and/ o r Mach number are continuously changing. The purpose of t h i s paper i s t o discuss, from a p i l o t ' s point of view, potential problem areas arising from t h e many adjustments which must be made t o follow such a profile.

Eight p i l o t s participated i n t h i s program which u t i l i z e d t h e same moving - base simulator and airplane configuration described i n paper no. 26. The piloting problems were studied under instrument flight con - ditions, i n smooth air, with ,the basic airplane characteristics aug - mented t o provide satisfactory handling qualities.

The f l i g h t p r o f i l e (fig. 1) used i n t h i s study should be examined This i p order t o become familiar with the potential problem areas.

f ire shows a l t i t u d e plotted against t i m e and depicts the climb, Across the top, Mach num - c h i s e , and descent portions of t h e mission.

ber i s shown as it varies through t h e climb phase, being held constant a t 0.9 f o r a short t i m e and at 3.0 f o r t h e last portion of climb and the establishment of cruise. The t i m e scale i s broken and t h e normal - and emergency - descent t i m e h i s t o r i e s begin from a new zero reference t i m e .

This p r o f i l e w a s chosen, not as an o-ptimum, but as a representative type f o r the purpose of studying p i l o t i n g problems associated with SST operations. I ' Tracing a t y p i c a l run, t h e first task i s t o establish a climb schedule, avoiding sonic speeds at low a l t i t u d e s . Then, above t h e a l t i t u d e requiring subsonic speed, t h e task i s t o accelerate as rapidly as possible t o cruise Mach number while climbing s u f f i c i e n t l y t o keep t h e sonic - boom ground overpressure below 2 pounds per square foot. The next task is t o climb t o cruise a l t i t u d e and l e v e l off while maintaining cruise speed. Duriag t h e cruise portion of t h e mission, t h e t a s k i s During t h e normal descent, a simply t o maintain speed and a l t i t u d e .

constant equivalent airspeed of 310 knots is maintained. The task i n an emergency descent i s t o reach, as quickly as possible, a safe alti - tude where 100 percent oxygen w i l l sustain life.

I n t h e simulator flying of this p r o f i l e a number of p i l o t i n g prob -

lems were disclosed. These are j

(1) Conventional display deficiencies ( 2 ) Overshoot at top of climb ( 3 ) Speed control during cruise ( 4 ) Emergency descent Proceeding sequentially along t h e profile, t h e i n i t i a l establish - ment of t h e subsonic climb presented no p a r t i c u l a r problems although t h e acceleration and climb performance of t h e SST i n t h i s area is very high and appeared t o be quite similar t o fighter - airplane performance.

if duct burning proves t o be uneconomical i n t h i s region, o r I n fact, i f noise - abatement considerations preclude f u l l - t h r o t t l e operation, p a r t - t h r o t t l e performance i s more than adequate f o r the subsonic acceleration.

CONVENTIONAL DISPLAY DEFICIENCIES conven - The first problem encountered concerns t h e deficiencies of t i o n a l instrument displays. Specifically, two kinds of information a r e

involved - precise pitch a t t i t u d e and Mach number error. Figure 2

,i

shows t h e area on t h e climb p r o f i l e where t h i s problem is most serious.

I n t h i s figure, a l t i t u d e is shown as a function of Mach number M. The s o l i d l i n e i s the reference profile, and t h e dashed l i n e i s a l i n e of constant - longitudinal - acceleration Ax capability. Note here t h e por - I n t h i s area longitudinal t i o n of t h e climb around Mach number 1.4.

acceleration is at a minimum, rate of climb i s quite low, and any departure from t h e reference p r o f i l e t o the high side results i n a t , r > ) > $ i i > J > J further reduction of acceleration. Departure to the low side, of course, results in increased ground o v ~ p m ~ w q . , - ~ f incidental interest in this connection was the value for net"lohgifXdTna1 acceleration of 0.4 ft/sec2, which was found to be the minimum necessary to provide for drag increase in maneuvering during profile tracking.

Figure 3 shows profile tracking performance when conventional instruments were used. The oscillations apparent here seem to be the result of two factors. First, pitch - attitude changes, required for flight - path corrections, are so small (on the order of 1/20> that they are barely perceptible on the attitude indicator. Second, no single parameter remains constant during this portion of the climb where the ground - overpressure boundary determines the relationship between Mach ,numberand altitude. In other words, the pilot has no good instrument ,)to use as a primary reference for pitch - attitude control.

Incidentally, this lack of adequate pitch - attitude information +s Practice SST climb profiles are not peculiar to ground - base simulation.

being flown from the NASA Flight Research Center in a North American Vigilante to an altitude of 50,000 feet and a Mach number of 1.7.

There, similar difficulties are being reported, even when operating in visual flight conditions.

One solution that was found to be acceptable consisted of displaying Mach number error directly to the pilot by use of an indicator located within and at the left edge of the attitude indicator as shown in fig - ure 4. This pointer, which moves vertically, was used to display the Mach number error. Full - scale deflection of the pointer equaled a Mach number error of 0.1 or, in other words, each dot represented a Mach num - ber error of 0.05. This signal was generated by using as inputs the actual Mach number, the actual altitude, and the profile Mach number appropriate to the actual altitude. Correction for Mach number error was made using an instantaneous rate - of - climb indicator for slight pitch - attitude adjustments. This provided better scaling as a pitch reference than did the attitude indicator, and the use of these two

indicators - the Mach number error and the instantaneous rate of climb -

resulted in improved prozile tracking as can be seen in figure 5 . This figure shows a comparison of the tracking seen previously in figure 3 with a conventional display and the tracking with the improved display where Mach number error was held to much lower values. Whether these nstruments could be implemented for actual aircraft use as nicely as for simulator use is another question. It would seem, however, that some such improvement in pitch attitude and climb-schedule-error dis - plays is in order for the SST.

OVERSHOOT AT TOP OF CLIMB The next problem area is that of altitude and speed overshoot at the top of the climb. In figure 6, which shows longitudinal accelera - tion along the climb profile used, it can be seen that, following the rather sluggish performance i n t h e middle portion of t h e climb where power' i s marginal, acceleration increases "as Mach number increases, and the SST r e a l l y g e t s a " second wind " as it approaches a Mach number of 3 .

If t h i s power is converted t o v e r t i c a l momentum at a constant Mach num - ber of 3, r a t e of climb becomes q u i t e high. Rate of climb with full t h r o t t l e would increase from 4,000 feet per minute a t 59,000 feet t o 20,000 feet per minute at 7O,OOO feet. This results i n a need f o r timely t h r o t t l e adjustment. An example of t h i s is i l l u s t r a t e d i n fig - u r e 7 w h i c h i s a t i m e history showing the last portion of t h e climb.

The upper portion of t h i s figure shows a l t i t u d e and the middle portion shows both thrust - weight r a t i o T/W at f u l l t h r o t t l e and t h a t being used. These parameters a r e plotted against t i m e . The i n s e r t a t t h e bottom of t h e figure shows tracking perfomance r e l a t i v e t o the refer - ence p r o f i l e with a l t i t u d e plotted against Mach number. Here, t h r o t t l e ) w a s reduced from 100 - percent t h r u s t t o cruise power (or 42 - percent thrust) as t h e vehicle passed 62,000 f e e t . The v e r t i c a l momentum a t t h i s point w a s s u f f i c i e n t t o achieve the last 8,000 f e e t of a l t i t u d e .

This i s an area, then, which requires close p i l o t attention, and even a minor o r routine d i s t r a c t i o n could cause a considerable overshoot, as i l l u s t r a t e d by figure 8 which shows an overshoot resulting from inade - quate t h r o t t l e adjustment. It can be seen here t h a t t h r o t t l e reduction w a s begun at the proper point, but t h e adjustment w a s made too slowly, simply because of a scan - pattern interruption. It i s interesting t o note a l s o that, although t h e overshoot i n a l t i t u d e was about 7,000 feet, t h e overshoot i n Mach number w a s only about 0.1. N o conclusions can be drawn concerning t h e r e l a t i v e likelihood of overspeeding t o overshooting i n altitude, however, because on t h i s simulation speed - control cues were considerably b e t t e r than nompl-acceleration cues. Also, the p i l o t s were briefed t o be p a r t i c u l a r l y careful not t o exceed a Mach num - ber of 3 . 2 . Although not pictured here, normal acceleration varied between zero and 2g during t h i s run. What can be seen from t h i s i s t h a t timely t h r o t t l e adjustment at the top of t h e climb i s quite c r i t i c a l .

SPEED CONTROL DURING CRUISE Regarding t h e third problem area shown, speed control during I cruise, t h e same thrust - speed i n s t a b i l i t y responsible f o r the high lon - gitudinal acceleration during t h e last portion of the climb is present i n t h e . c r u i s e condition, resulting i n a need f o r continuous t h r o t t l e adjustment t o maintain a constant Mach number. Figure 9 is a short t i m e history of p i l o t performance i n t h i s area. It can be seen that Mach number control w a s f a i r l y good, but frequent t h r o t t l e adjustments were necessary. This i s not a d i f f i c u l t task, but it does demand con - s t a n t attention and undoubtedly wastes some f u e l . These last two i problem areas both suggest a need f o r some type of speed - sensitive fuel - cont r o l unit.

€DIEEGENCY DESCEXCS Normal descents were dade using a constant equivalent velocity of 310 knots and no p a r t i c u l a r problems were noted. Ehergency descents were b r i e f l y examined where aWmpts were made t o reduce a l t i t u d e as quickly as possible, constrained m y by an airspeed l i m i t of 500 knots equivalent, and v e r t i c a l acceleratiol'r values of 2g and -1g. The goal w a s t o descend i n less than 2 minutes from a cruise condition a t an a l t i t u d e of 70,000 feet and a Mach number of 3 t o an a l t i t u d e of ) 40,000 feet, roughly t h e maximum safe a l t i t u d e f o r sustained breathing with 100 percent oxygen. This goal w a s achieved through t h e use of a drag - producing device, such as speed brakes, a drag parachute, o r t h r u s t equivalent on t h e vehicle simulated t o an incremental drag reversers, coefficient of 0.023. Oscillations were encountered as a r e s u l t of the rapid i n i t i a l a t t i t u d e change made t o commence t h e descent. These o s c i l l a t i o n s persisted through 40,000 f e e t but did not r e s u l t i n exceeding t h e airspeed l i m i t o r t h e vertical - acceleration values specified.

SUMMARY O F RESULTS

/ I qfa

Referring again t o t h e l i s t of problems discussed: conventional instrument displays appear t o present inade - F i r s t , quate pitch - attitude information a t t h e higher speeds. The addition of velocity - error information and an instantaneous rate - of - climb indi - cator resulted i n improved performance.

overshoot problems a t t h e top of t h e climb require very Second, timely t h r o t t l e adjustment because of the large excess of t h r u s t avail - able a t f u l l t h r o t t l e .

Third, an unstable thrust - speed relationship at cruise caused a need for continuous t h r o t t l e adjustment and suggests a need f o r some type of speed - sensitive fuel - control unit.

Fourth, i n order t o achieve an emergency - type descent from cruise conditions at an a l t i t u d e of 70,000 f e e t t o 40,000 f e e t i n 2 minutes, an incremental drag coefficient of 0.025 w a s found necessary.

~ u t ~ o ~ TYPICAL SIMULATED SST MISSION PROFILE '\ I \\ # .

' v ~ o R M A L I \ W \ DESCENT n \ \ \

2 - 30 \

\ \ i 3 - 1 EMERGENCY \\ a I \ DESCENT I \.

IO - / \ \ / I l l I I n 1 I I I I I I \I 0 8 16 24 0 8 1 6 24 T I ME, minutes Figure 1 SST CLIMB PROFILE LONGITUDINAL ACCELERATION BOUNDARY 7 0 ~ 1 0 ~ ALTITUDE, ft M Figure 2 S S T C L I M B P R O F I L E PILOT TRACKING PERFORMANCE CONVENTIONAL DISPLAY I 2 3 M Figure 3 ATTITUDE INDICATOR W I T H SPEED ERROR INDEX A-31302 Figure 4 SST CLIMB PROFILE PILOT TRACKING PERFORMANCE CONVENTIONAL DISPLAY IMPROVED DISPLAY

rt

CONST ACCEL a I 2 3 M M Figure 5 LONGITUDINAL ACCELERATION ALONG CLIMB PROFILE UTILIZED FOR SIMULATOR PROGRAM M Figure 6 ALTITUDE, ft

----- ~~

T/W .3 ------

FULL THROTTLE - ACTUAL . I 1280 1320 1360 1400 TIME FROM START OF CLIMB, sec i TRACKING PERFORMANCE Figure 7 LEVEL - OFF AT TOP OF CLIMB WITH OVERSHOOT ALTITUDE, f t 6 0 5 0 FU LL THROTTLE A - I I I I I 1300 1340 1380 1420 1460 1500 1540 1580 TIME, sec w P M Figure 8 J TIME HISTORY OF CRUISE T H R O T T L E ADJUSTMENTS 1 SOL a - , - E 3.06 2 ! $ 3.00 Z 2.94 0 20 40 60 80 100 120 140 160 180 TIME, SEC Figure 9 29. G ENERAL S UMMAR Y AND ASSESSMENT By Mark R. Nichols NASA Langley Research Center The present conference has reviewed in - house and contract work being conducted by the National Aeronautics and Space Administration in support of the national supersonic transport program. The purpose of this paper is to provide a condensed summary of the results presented together with an assessment of supersonic transport capabilities and of further research needs.

, RECENT NASA R E S W C H CONTRIBUTIONS NASA research on the special problems of supersonic - cruise air - planes began in connection with the B - 58 and B-70 programs. A number of significant contributions of direct interest for the supersonic com - mercial air transport (SCAT) have been made since this time and are listed as follows: Airframe and engine configuration concepts: As emphasized in the initial papers, a basic part of the NASA program has been the evolution of advanced airframe and engine configuration concepts and the submittal of these concepts to the industry for study from the viewpoint of.the practical airplane designer. This activity is quite different from the traditional NASA research effort but has been considered necessary in this case because many of the special problems of the supersonic trans - port are directly dependent on the overall airplane configuration and on the compromises that the designer must make to arrive at a practical airplane.

Performance aerodynamics: A large amount of researcn has been accomplished recently in the field of performance aerodynamics. For example, a method for calculating and optimizing the pressure drag has been evolved, validated, and adapted for machine computing. As men - ) tioned in preceding papers, an advanced machine computing program developed by and obtained through courtesy of The Boeing Company w i l l be made available to the industry by NASA.

Additional research on boundary layers has resulted in verification of the applicability of the T ' method of extrapolating the turbulent skin friction of smooth surfaces from model - scale Reynolds number to full - scale supersonic - transport Reynolds numbers. It is hoped that these new data will settle the long - standing controversy on this subject.

The drag'contributions of a variety of additional types of surface roughness also have been correlated and quantitized.

Studies of the use of wing warp for reducing the drag due to lift and trim have been continued with successful results. Overall improve - ments in lift - drag ratio of about 1 . 0 have been obtained for wing plan - In addition, special forms of interest for the supersonic transport.

calculation methods have been evolved for handling arbitrary planforms such as cranked, ogee, and M wings.

Finally, a much better understanding has been obtained of the nature of both drag and lift interference effects at supersonic Mach This understanding has led to.the capability of using these numbers.

interference effects to obtain overall configuration improvements.

SCAT 15 provides an outstanding example along this line inasmuch as addition of the engine nacelles to the basic wing - body combination was accomplished with no loss in cruise lift - drag ratio.

NASA stability and control research has Stability and control: defined methods for minimizing the aerodynamic - center shift and pitch- up of variable - sweep - wing configurations, has demonstrated the feasi - bility of high - lift systems for both the variable - sweep - wing and the SCAT 17 type, fixed - wing configurations, and has evolved highly swept wing planforms which are regarded as promising successors to the SCAT 4 concept. In addition, the lateral - directional and control problems of a number of candidate and general research configurations have been explored, and extensive simulation research has been conducted to help define problems and requirements in the areas of handling qualities and flight operations.

In the area of propulsion, studies conducted in 1959 Pro ulsion:

and 1 9 + 0 provided an early definition of the basic requirements for the

supersonic - transport engine. Promising engine cycles were subsequently evolved in a joint program with Federal Aviation Agency (FAA), U.S. Air Force (USAF), and industry. Major air - inlet problems have been explored, and promising jet - exit concepts have been both evolved and - studied experimentally. This latter item is regarded as particularly important inasmuch as a 1 - percent loss in nozzle efficiency in cruise is M = 3 equivalent to a 4 - to 5 - percent increase in specific fuel consumption.

j Hence, jet - exit research is being given attention in the current SCAT program comparable to that devoted to reduction of aircraft drag.

Materials: In the area of materials, a preliminary screening of structural materials of interest for the Mach 3 class of supersonic transport was accomplished by a special NASA committee composed of rep - resentatives of the airframe companies, alloy producers, and interested government agencies. In a follow - on program, the mechanical properties of the more promising candidate materials are being determined after prolonged exposure to M = 3 tenrperatures. At the present time, some of the materials already have passed the 14,000 - hour mark. Initial results also have been obtained in research on other important areas such as the susceptibility of the materials to salt stress corrosion and the determination of practical methods of fabrication.

Sonic boom: Because of its dominant effect on the overall feasi - bility of the airplane, a great deal of research has been devoted to the sonic boom. In this research, basic sonic - boom theory was verified, measurements of intensity were obtained, and exploratory studies of public acceptance were conducted in a joint program with the F A A and USAF. Calculation methods and wind - tunnel techniques also have been developed by NASA f o r studying configuration effects. Because of their pirst-order influence on configuration geometry and aircraft weight, it 'is anticipated that these methods and techniques w i l l take their place along with classical drag - estimation methods as primary tools of the designer, Finally, this total effort has led to the establishment of approximate lower bounds of sonic - boom intensity from the configuration - effects viewpoint.

BROAD CONCLUSIONS OF CONTRACT STUDIES OF SCAT CONFIGURATIOMS The purpose of the contracts let with The Boeing Company and Lockheed California Company was fourfold: To obtain the input of the industry in the areas in which it is uniquely qualified; to determine the feasibility of the four concepts studied as candidate concepts for the U . S , supersonic transport; to establish the technological state of the art for the guidance of the national program; and to uncover problem areas needing further research attention by NASA. This section of the paper summarizes the broad conclusions of these contract studies.

First, airplanes developed from the SCAT 1 6 and 17 configuration

concepts appear to meet essentially all of the specified mission requirements. They are competitive with regard to overall performance in this mission but are larger and heavier than corresponding subsonic ) j e t s .

Airplanes based on the initial SCAT 4 and 15 configuration concepts are ruled out unless solutions can be found to their stability or weight problems. The fact that they were dropped from further consideration by the contractors after midterm review of the contracts does not mean that hope is abandoned for these concepts, but rather that they were consid - ered noncompetitive with their stability and weight problems outstanding.

Solutions to these problems were not forseen in time to permit proper study by the contractors during the remaining period of the contracts.

Variable sweep provides improvements in mission versatility, low - speed handling qualities, and community noise.

Delta - type wings provide low structural weight and near - linear pitching - moment characteristics for most operating conditions.

The sonic boom has become a dominant factor in design with effects on the gross weight and range comparable to lift - drag ratio, specific fuel consumption, and structural weight. It is clear that great care must be exercised both in specifying the sonic - boom requirements and in configuring the airplane so as to obtain a minimum sonic - boom " form factor. " New and advanced engines are required for the airplane. An associ - ated major research and development effort w i l l be required in the areas of air inlets, jet exits, and propulsion system controls.

Major reductions in gross weight and sonic - boom intensity compa - rable to the gains obtained by the use of advanced engines are obtain - able by the use of titanium as the basic structural material and also by substantial reduction of the fuel reserves such as might become feasible following improvements in air traffic control and landing systems. It appears mandatory that these avenues of improvement be exploited.

Within the ground rules of the study, M = 3 airplanes with advanced engines and titanium structures have lower gross weights and lower levels of sonic boom than lower - speed aluminum airplanes. The lower - speed airplane probably would have been more nearly competitive However, the if more advanced engines had been permitted in the study.

use of highly advanced engine technology was considered to be inconsis - tent with the assumption of an aluminum structure.

Finally, the overall implication of the results of the studies is that highly advanced airframe and engine configurations are needed to meet the objectives of the national program.

ASSES- OF SCAT CAPABILITIES On the basis of the NASA research results and the results of the SCAT feasibility studies by Boeing and Lockheed, it is possible to arrive at an assessment of the state of the art. It is useful in such an assessment to begin with a look at the Breguet range equation: efficiency (Symbols used are defined in the appendix at the end of this paper.)

As indicated by the grouping of terms, the two basic factors affecting air - craft range are the flight, efficiency (range factor) and the fuel-to- initial weight ratio or, in the case of a complete mission, the fuel - to - gross weight ratio. The state of the art with regard to these two factors w i l l be reviewed briefly.

i

The flight efficiency expressed in the form Mm is plotted in

SFC figure 1 as a function of Mach number for representative U.S. subsonic jet transports and for supersonic airplane designs developed from the four SCAT configuration concepts by the NASA SCAT feasibility study contractors. Each of the supersonic airplanes was assumed to be powered by the highly advanced D-type turbofan engines, and its engines and wing area were sized to satisfy the mission rules spelled out in the SCAT feasibility study work statement.

At supersonic speeds, the SCAT 16 and 17 airplanes in their present form provide approximately equal flight efficiencies.

The level of these efficiencies is significantly lower than those for the SCAT 4 and 15 class of airplanes which featured highly advanced aerodynamics but were found to be impractical for other reasons.

The present SCAT 1 6 and 17 airplanes provide only about three - quarters of the cruise flight efficiencies of the best present subsonic jets.

Hence, it is regarded as extremely important that research be continued to attain a higher level of supersonic cruise efficiency for practical configurations. The flight efficiencies of the SCAT 4 and 15 class of configurations are considered to establish a realistic goal for this upgrading of the state of the art.

At subsonic speeds, the SCAT 16 type of airplane characteristically exhibits marked superiority over the SCAT 17 type and provides flight ifficiencies approaching those of the present subsonic jets.

This good subsonic efficiency is of advantage in subsonic cruise and hold opera - tion but may or may not be a determining factor in overall configuration selection, depending on the degree of mission flexibility demanded of the airplane.

In figure 2 the disposable - load ratio, which is the ratio of the weight of the usable fuel plus payload to the take - off gross weight, is plotted as a function of airplane gross weight.

The disposable - load ratio is a significant par s it defines the upper limit of the fuel - carrying capability of the airplane and recognizes, to some extent, the fuel - payload trade. The disposable - load gross weight char - acteristics of a number of U.S. subsonic jet transports and of the SCAT 1 6 and 17 airplanes (as determined by Boeing and Lockheed) are shown by the circular, square, and diamond syrdbols, respectively.

For the subsonic jets, the points plotted include newly all the Boeing, Douglas, and Convair series and represent a consistent family of carefully optimized designs beginning with small ranges and payloads at the lower left of the figure and extending to large payloads and very long ranges at the upper right. The Boeing 707-120 airplane is identified because it has approximately 3,200 - nautical - mile range capa - bility with the 26,125 - pound payload. In other words, it meets all of the performance requirements of the NASA SCAT mission with the exception of cruise speed.

The SCAT 1 6 and 17 airplanes require a disposable - load ratio of 0.02 to 0.05 greater than that for the 707-120 for the design SCAT mission.

This difference is caused by the flight efficiency of the SCAT airplanes being lower than that for the 707-120 both in cruise (as shown in fig. 1 ) and in off - design flight so that more fuel is needed. The requirement for a higher disposable - load ratio plus the inherent weight penalties of the supersonic airplane associated with greater overall slenderness, higher - thrust engines, more complex inlets and exits, and so forth,

results in the SCAT 1 6 and 17 airplanes being 130,000 to 200,000 pounds

heavier than the 707 - 120. The present objective of NASA advanced - concepts research is to evolve designs which w i l l provide major reduc - tions in this penalty through improvements in aerodynamics, propulsion, and structural weight.

The lines plotted in the vicinity of the points for the SCAT 1 6 show estimates of the way in which the disposable - load ratio attainable with this configuration with a payload of 26,125 pounds varies over a range of design gross weights. Two acceleration sonic - boom limits are 2 . 5 and 2 . 0 lb/sq ft.

considered: At gross weights in excess of the plotted points, the slope of the curve for is low = 2.5 1b/sq ft compared vith the trend curve through the family of subsonic jets. Thus, it is indicated that with even a modest boom limitation the gross weight of the supersonic airplane is much more sensitive to changes in payload and range than is the case for its subsonic counterpart.

With a boom limit of 2 . 0 lb/sq ft, the slope of the trend curve is further reduced and the curve actually hooks down at gross weights exceeding about 480,000 pounds. This hooking is caused by the necessity for increasing the acceleration altitude progressively as the gross weight is increased in order to maintain a constant level of sonic - boom overpressure. Each increase in altitude requires increases in both engine size and wing size so that eventually the operating weight empty i begins t o increase a t a f a s t e r r a t e than the g r e s u l t i s that, i n contrast t o t h e subsonic case, it i s no longer pos - s i b l e t o increase range - payload performance by merely f u r t h e r increasing the gross w e i g h t . A t this point f u r t h e r performance gains can be obtained only through basic improvements.in flight efficiency, reduc - tions i n t h e sonic - boom " form factor, " or reductions i n t h e operating w e i g h t empty.

From the preceding discussion, it i s clear that the sonic boom has introduced another f a c t o r i n t o t h e design process and t h a t this f a c t o r has e f f e c t s on gross weight and range f u l l y comparable t o the f l i g h t - efficiency and weight - ratio terms of t h e Breguet equation. This s i t u a t i o n as i n the case of j e t noise, a compromise must be means t h a t once more, jade between the desire of t h e designer and operator t o develop t h e most economical airplane and the desire of t h e public t o be subjected t o a minimum amount of noise. The s i t u a t i o n i s complicated by the f a c t t h a t the l e v e l of sonic boom acceptable t o the public cannot be pinned down exactly a t t h e present t i m e . A great deal of additional thought and study will have t o be devoted i n the near future t o the question, " What should the specified sonic - boom requirement be? " i f unnecessary design penalties are t o be avoided.

The state of t h e art with regard t o mission performance i s pre - sented i n figure 3 i n which t h e gross weight and the payload - to - gross The mission weight r a t i o a r e plotted as a function of design range.

requirements considered are those specified i n the N A S A SCAT f e a s i b i l i t y study work statement and include a 26,125-pound payload and appropriate operating r e s t r i c t i o n s . The cruise Mach number i s 3; the maximum allow - able sonic - boom overpressure i s 2.0 lb/sq ft; and the airplane i s assumed t o have titanium structures and the D-type turbofan engines incorporated.

It should be recalled t h a t t h e D engine i s highly advanced and i s approaching the l i m i t of present development capabilities. The band of values shown covers the spread between the minimum - weight configurations determined by the two contractors i n t h e i r NASA SCAT studies.

A t a 3,200 - nautical - mile design range, t h e state - of - the - art minimum gross w e i g h t s as defined i n figure 3 range from 380,000 t o 430,000 pounds with a corresponding spread of payload - to - gross w e i g h t r a t i o s from 6.9 '-0 6.1 percent. When t h e range i s increased t o 4,000 s t a t u t e miles ,3,475 n. m i . ) , a value of current i n t e r e s t , the required gross weights increase by increments of 40,000 t o 80,000 pounds.

Figures 4 and 5 present the r e s u l t s of extensive computer studies

conducted by NASA, u t i l i z i n g a complete mission optimization process, t o establish t h e importance of t h e major design trades f o r SCAT'S 16 and 1 7 .

The values shown generally a r e i n good agreement with values reported t o date by t h e contractors.

The following"'qua1ifications should be made regarding the data pre - sented in figures 4 and 5 to permit proper interpretation: (1) To correspond with the results presented by the con - tractors, the sonic - boom calcu$ations of the upper left plots were made with the reference - pressure term in the Whitham equation defined as i p - instead of the mean - altitude pressure which now appears to give somewhat better correlation between measurements and theory.

(2) In the plots at the upper rigbt, the range or weight increments caused by changes in passenger - type payload take into account changes in fuselage and accommodation weights and changes in the fuselage drag in addition to the change in the payload weight itself.

(3) In the plots at the lower right, the engine and structure weight increments were assumed to be reflected directly without growth factors in changes in the operating weight empty.

The results of these trade studies have been used to determine a projection of the present state of the art (fig. 6) because of the belief that modest or significant improvements can be made in a number of areas during the airplane development process. As a starting point, it was assumed that the correct weights for the SCAT 1 6 and 17 airplanes are each the average of the values reported by the contractors. The particular improvements listed are considered to be individually attain - able. (The reductions in engine and structure weight are assumed to be reflected in equal changes in the operating weight empty.)

As shown in figure 6, if all of these gains can be achieved simultaneously, it appears that a 4,000 - statute - mile range can be attained with the standard payload of 26,125 pounds with a gross weight in the neighborhood of 360,000 pounds. Further, if the payload is increased to 35,000 pounds, the predicted gross weight f o r the sane range increases only to about 430,000 pounds. These values represent very substantial improvements over the current capabilities and provide a much more attractive airplane. How soon can The main questions are: .i such improvements be realized? and Can they all be realized in one design?

NASA RESEARCH PROGRAM The NASA SCAT research reported at this conference is being con - tinued with the objective of solving outstanding specific problems and bringing about a general upgrading of ~ aircraft capabilities. The over - all program includes extensive research in the following areas: (1) Advanced configuration concepts such as the improved versions of the present configurations illustrated in figure 7 (2) Basic and configuration aerodynamics ( 3 ) Propulsion (engine cycles, air inlets, jet exits) (4) Sonic boom and noise ( 3 ) Materials and their fabrication problems (6) Fabricated and fail - safe structures (7) Handling - qualities problems (8) Operating and airworthiness problems It is hoped that this research will make major contributions to the first U.S. supersonic transport.

SYMBOLS V velocity SFC specific fuel consumption L/D lift - drag ratio

w weight

M Mach number sonic - boom overpressure at sea level *%Ax ambient pressure at altitude P a l t ambient pressure at sea level Psl Subscripts: rnG engine STEWCT structural

STATE OF THE ART - FLIGHT EFFICIENCY

D ENGINES, NASA MISSION RULES U. S. SUBSONIC JET TRANSPORTS S M (LID)

- -

S FC - 1 0 i I I I I 0 I 2 3 M F i g u r e 1 STATE OF THE ART - DISPOSABLE LOAD ESTIMATED TRENDS 1 6 WITH 26,125 LB T I I I I I 0 100 200 300 400 500 6AOXlO3 WGROSS LB F i g u r e 2 PERFORMANCE

STATE OF THE ART - MISSION

WGROSS, LB WPAYLOAD 550 xi03

r WGROSS

.05 r

- Ti STRUCTURES - .06 - .07

I

- .08 - 2,500 3,000 3,500 4,000 DESIGN RANGE, N. MI.

Figure 3 MAJOR DESIGN TRADES BASE - LINE CONFIGURATION: SCAT 1 6 WITH D ENGINES AW, LB L B I F T ~ ~p CARGO OR RANGE, N.MI. 2 5 PASSENGERS RESERVES 3,500- 3,000 - I I I I I 2,500 L : I Figure 4 MAJOR DESIGN TRADES BASE - LINE CONFIGURATION: SCAT 1 7 WITH D ENGINES AW. L B " M A X ; . LB/FT* 25 CARGO OR RANGE, N.MI.

3,500 3,000

I