Flight test report of the NASA icing research airplane: Performance, stability, and control after flight through natural icing conditions

J NASA Contractor Report 17951"5

Flight Test Report of the NASA

Icing Research Airplane

Performance, Stability, and Control After

Flight Through Natural Icing Conditions

(NASA-CR-1795151 FLIGHT _ESI _._ORT OF THE N87-1 17c7 NASA ICING RESEAFCH AI_LAN_: _]_5_OI_kNCE, ST_BILI[[Y, AND CCI_RCL AFteR _IIGHT THHOUGH NATURAL ICING CCNDITICNS _inal Report U nc la s [Kohlman Systems Research, Inc.) ]59 p 4_924 G 3/0 8 Jerry L. Jordan, Stewart J. P|atz, and William C. Schinstock Kohlman Systems Research, Inc.

Lawrence, Kansas October 1986 Prepared for Lewis Research Center under Contract NAS3-24547 National Aeronautics and Space Administration

REVISIONS

DATE APPI%OVAL RE'V DESCRIPTION T ,l 14 MAR Or J gJ nal

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I_6 R£POe_ NO. 8(,-o] ,,fe_, ,4 DE HAVTL/.AND TW_ OTTER

KOHLMAN SYSTEMS RESEARCH

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HAV'TLLAND TWIN OTTER 86-01 APe,, A KOHLHAN SYSTEMS RESEARCH PAGE i J i DHC-6-_4 _:._r.D_iG FAGE B_K NuI: _X._P aL

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KOHLMAN SYSTEMS RESEARCH

PAGE Iv DHC-.6--II_ TABLE OF CONTENTS PAGE # REVISIONS ......................................................

LIST OF ACTIVE PAGES ........................................... iii TABLE OF CONTENTS .............................................. v LIST OF FIGURES ................................................ vii 1.0: INTRODUCTION ............................................. i. 1 2.0: LIST OF SYMBOLS .......................................... 2.1 3.0: AIRPLANE ................................................. 3.1.1 3.1 GENERAL AIRPLANE DESCRIPTION ........................ 3.1.1 3.2 TEST AIRPLANE ....................................... 3._. 1 3.3 WEIGHTS, CENTERS OF GRAVITY AND INERTIAS ............ 3.3. 1 4.0: FLIGHT TEST OPERATION .................................... 4.1 5.0: DATA ACQUISITION SYSTEM DESCRIPTION ...................... 5.1.1 5.1 DATA ACQUISITION SYSTEM DESCRIPTION ................. 5.1.1 5.2 SPECIAL SENSORS ..................................... 5. 2. i 5.3 DATA ANALYSIS ....................................... 5.2.1 6.0: AXIS SYSTEM AND SIGN CONVENTIONS ......................... 6.1 7.0: FLIGHT TEST ANALYSIS TECHNIQUES .......................... 7. i.I 7. 1 STABILITY AND CONTROL DERIVATIVE METHODS ............ 7. i. i 7.2 PERFORMANCE MODELING METHODS ........................ 7.2. i 7.3 AIR DATA CALIBRATION TECHNIQUES ..................... 7.3.1 8.0: RESULTS .................................................. 8.1.1 _. ! LONGITLiDINAL DERIVATIVES ............................ 8. I. i 9, ()_ REFERENCES ............................................... 9. 1 ADPE_D!:X A: FL ]GH r ITS! DA'IA BASE FORi_A_ ...................... A. I THIS PAGE INTENTIONALLY LEFT BLANK vi LIST OF FIGURES ¢ b PAGE # FIG. # FIGURE TITLE 3.1.1 AIRPLANE OPERATING ENVELOPE 3.1.2 AIRPLANE GEOMETRY 3.3.1 FUSELAGE STATION AND WATERLINE REFERENCES 3.3.4 MASS MODEL FOR NASA FLIGHTS 6-11 3.3.2 _.II MASS MODEL FOR NASA FLIGHTS 12-25 5.1.5 5.1.1 KSR DATA ACQUISITION SYSTEM 5.2.1 KSR FORCE WHEEL 5. d._ KSR FORCE PEDALS 5.2.2 5. -_ • ,-t 3 ROSEMOUNT MODEL 858AJ TYPE FLOW ANGLE SENSOR 5. 3.2 5.3. ia AND CALIBRATION 5.3.3 KSR DATA MANAGEMENT SYSTEM - DATA ANALYSIS 5.3. Ib 6.2 6.1 AIRPLANE AXIS SYSTEM 6.3 6.2 CONTROL DEFLECTION SIGN CONVENTIONS 7. l.Z_ 7. i.I TEST DATA AND MMLE MATCH FOR LONGITUDINAL DOUBLET (FLAPS O) 7 :' 5 l u-.

1 "='• 1 FORCES AND MOMENTS USED FOR PERFORMANCE ANALYSIS i_ POSITION CORRECTION FOR THE DE HAVILLAND 7.=' 1 TWIN OTTER 7.3.5 ROSEM[]UNT PROBE ALPHA CALIBRATION FOR THE 7.3.2 DE HAVILLAND TWIN OTTER 7.3.6 7. 3. 3 ROSEMOUN] PROBE BETA CALIBRATION FOR THE DE HAVILLAND TWIN OTTER 7.3.8 7.5:.4 AL_A BCfOiv, PENDING CORREC-[IO'_ FC*_ THE DE HAVILI_AND TWIN OTTER 7. 3. 9 7. 3. 5, BE I(i BOOm BENDING CORRECTION FOR THE DE HAVII LAND TW;N OTTER vii e_ING PAGE BLANK NQT FiLMF,D FIG.

FIGURE TITLE , ° CHANGE IN AIRCRAFT LIF] DUE TO rICE 8. i.4 8. i.8 EFFECT OF ICE ON STATE COEFFICIENTS (C + C ) L L.

q 8. I. I0 8.1.3 CHANGE IN ELEVATOR LIFT DUE TO ICE 8.1.16 8.1.4 CHANGE IN AIRCRAFT DRAG DUE TO ICE 8. 1.22 , 8.1.5 CHANGE IN AIRCRAFT ITCHING MOMENT DUE TO ICE 8. 1.24 8.1.6 EFFECT OF ICE ON STATE COEFFICIENTS (C + C h3 rll .

q 8.1.7 8.1.26 CHA_GE IN ELEVATOR CONTROL POWER DUE TO ICE 8.1.8 TIME HISTORY FOR NIII+36A 8.1.31 8.1.9 TIME HISTORY FOR NIII+37A 8. i.38 8. i. I0 TIME HISTORY FOR NI23÷23A 8.1.45 8.1.11 TIME HISTORY FOR NI23+24A 8. i. 52 8.1.12 TIME HISTORY FOR N123+25A 8. 1.59 8.1.13 TIME HISTORY FOR NI23+26A 8.1.66 8.1.14 PLOT FOR RUNS NIIIm36A AND NIII+37A 8.1.73 C /C h h r B 8.1.15 C /C 8.1.74 PLOT FOR RUNS NI23+24A, NI23+25A AND h h NI23+26A r B viii THIS PAGE INTENTIONALLY LEFT BLANK i. INTRODUCE iO_x ..................... OF POOR QU_LiT_ This reoc, r= c,-,rDtalr;s the res,Jits ,-,f filDnt tes;t- _ ,:,f tr_E- lxH_F4 Lewis De Havillar_d Twir, Otter_ Serial Nurnbe_- (-)4. The f_s_ s_x sectlorss cor_tair, ir;formatic, rl abc.ut the test airDlar, e, the techr, iQues use0 for _ata gat her ir,_c, a i ist ,:,f syrnbc, i s, ar, o axes/si or, cot;verst ior, s. A brief descriDt i,_-,r_ of the flight test tecnr, iQues useO by KSR are preser;ted it, Sectior_ 7. The c,:,efficier, ts ar;d c_erivatives for the lor_gitudir, al axis are cor;tair, ed irJ Secti,_-,r_ 8.

Due to c,:,rltract argo bueDetacy cor_strair_ts, or, ly a port i,:,r_ of the data acouirerJ could be ar_alyzed. Thus, priority was Qiver, tc, ar; ar, aiysis of aircraft oerforn_ar, ce, i,_,r,_it udir,al etabl I ity, ar,o selected asyrmnetric power sidesl i D rnar, euve_- showir, n _-udder fi,-,a_ characteristics with ar, d withc, ut ice. It should be r,,-,teo that or, eli stability ar, d cc, r,t ro ] flignts i at era l-d i rect i,_-,r,a i date _as acquired, h,-,wever, these data were r,ot ar,alyzed f,-,r this reD,-,rt for the reas,z,r,s orevic, usly nler, t i,-,r, ed. It is our hope at KSR that fv.'_uce cr, r, tracts will aliow us to ar, alyze this e_Citior, al data ar, c verify it with sirnulati,-,r, techr, iques.

,_E, OEDING PAGE BLAHK NOT FtLMED 1.1 THIS PAGF INTEN]IONALkY LEFT BLANK 1.2 2.0 LI ST OF SYMBOLS DEFINITION SYMBOL longitudinal acce]eration (body axis) A X lateral acceleration (body axis) Ay vertical acceleration (body axis) A Z positive downward b wingspan b aileron span a b elevator span e b rudder span r c wing mean aerodynamic chord (MAC) c dvei d_e aJ it_l Ull t_tva_ u a c average elevator chord e c average rudder chord r center of gravity c.g.

drag coefficient, D/½pV2Sw C D 8Ch/86 r Ch 6 r ¢h O C h/Sr g r 6 lift coefficient, L/½pV2S C L W 0 CL/86 e C L e dc 8CL/8 (_7) C L .

OCL/O (_--_) C L q 2".1 LIST OF SYMBOLS (Cont'd) DEFINITION SYMBOL C pitching mom,'nt coefficient m q_ C OCm/O ('-_-) m q C OCm/O ( -_-_ ) m.

ac /06 C m e

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e P/pn 3 D 5 P Cp F/pn D 4 P C T drag D propeller diameter P data acquisition system DAS thrust wheel force due to ailerons F a wheel force due to elevators F e gross thrust (engine axis) F g net thrust (engine axis) F n rudder pedal force or ram drag (stability axis) F r F aerodynamic force along the Y-axis Y acceleration due to gravity g moment of inertia about X axis l XX moment of inertia about Y axis Iyy LIST OF SYMBOLS (Cont'd) SYMBOL DEFINITI ON I moment of inertia about Z axis ZZ XZ product of inertia Ixz ITT turbine temperature V/n D J P l_ft or rolling moment M pitching moment or Mach number MAC mean aerodynamic chord normal load factor, g's, n -= -A Z or propeller speed n w_nd axis load factor, x-dlrectJorl x wind n W]rld axis i £)acl I ac ct_L _ t. i i-&c t] OFf z wind N yawing moment or correction exponent for the thrust model G gas generator RPM P pressure.

P propel ler rol 1 rate P P power dynamic preusure or pitch rate q r yaw ra Le wi ng area S W T tempera ture V _a] ibrated _i i speed C 2f.3 LIST OF SYMBOLS (Cont'd) SYMBOLS DEFINITION V ground speed g true airspeed Vt, V W weight W air flow rate a fuel flow rate Wf aircraft coordinate axes - defined in Section 6 X_ y_ Z t_ angle of attack of longitudinal body angle of attack of horizontal stabilizer n B sideslip angle 6 control or trim surface deflection normalized ambient pressure ratio differential element (generic) downwash angle at horizontal stabilizer inclination angle E downwash angle at horizontal stabilizer for _ = 0 air density P e pitch angle roll (bank) angle yaw angle normalized total pressure ratio at the t 2 compressor face normalized total temperature ratio at the t 2 compressor face flight path angle P.4 LIST OF SYMBOLS (Cont'd SUBSCRIPT DEFI NIT] ON aileron a A aerod ynam_ aircraft A/C C control ¢olumn elevator e f flap left L rudder r R right test t t tr_m total T control whee] W Z.B THIS PAGE INTENtIONALlY LEF_ BLANK 2.6 3. AIRPLANE 3.1. General AirQlane Desciptio D The De Havilland Twin Otter is a 13 to 20 passenger, one ,z,r two crew, turboprop airplane used primarily for commuter trar, s- portatic°n. Maximum range is 700 n.m (at 10,000 ft. ) with a 45 minute reserve. Maximum cruise speed at 10,000 feet is 165 KTAS.

The airplane operating envelope is shown in Figure 3. 1. I, and the geometry and operating weights are shown in Figure 3.1.2.

The airplane is powered by two Pratt and Whitney (UACL) PT6A-2C) engines_ _ach flat rated at 55(] shaft horsepower and a Hartzell three bladed, driving constar, t speed, reversible, blade T10173+I.

[netal propeller, model H3-B3TN-3, The primary flight control system is fully reversible. A gear tab is fitted to the rudder to lighten control forces and a tab fitted to the starboard elevator is linked to the flaps to control longitudinal trim during flap retraction and extension.

The horizontal tail has fixed incidence. Geared trirn tabs are located on both ailerons.

The flaps are hydraulically actuated and incorporate a hand pump in the crew compartment to provide emergency pressure. No speed brakes or spoilers are instal leo.

Figure 3. 1. 2 shows the locatic0n of the flight control syster_ layout. Stall warning is visual and aural. No stick shaker is instal led.

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.... "t - - +.............. KOHLMAN SYSTEMS RESEARCH- 3.1.3 LAWlIE NC| R&N|&S _R-.,I ! r 1- THIS PAGF INTENTIONALLY LEFT BLANK 3.1.4 3.2 Test A_Qlane The test airplane (N607, SN 04) instrumentation include_ control surface positions, pilot forces, engine data, air data, and inertial data. The KSR Data Acquisition System (DAS) was mounted inside the cabin near the aircraft, c.g. A Rosemount differential pressure angle of attack and sideslip sensor was mounted on a nose boom. This sensor is detailed in Section 5.2, •Special Sensors'.

Several geometric constants were used to process the flight test data. These are listed Jr, Table 3.2.1. The wir, g area_ chord (MAC) were used to nor+- span, and mean aerodynamic dimensionalize the aerodynamic coefficients. The standard c.g.

location is the point where all aerodynamic forces are assumed to act. A standard moment transformation is then used to transfer the moments to the actual c.g. The propeller and r,ozzle locations are where the propeller and jet forces act.

3.2.1 TABLE 3.2.1 DE HAVILLAND TWIN OTTER BASIC CONSTANTS LIST VALUE UNITS PARAMETER 420.03 I SQ. FEET WING AREA INCHES 780.000 2 WING SPAN I0.00 3 ASPECT RATIO 4 INCHES 78.000 M.A.C.

188.240 5 LEADING EDGE STATION INCHES INCHES -123.5 6 F.S. OF ALPHA/BETA SENSORS 14.68 7 AILERON AREA (EACH) SQ. FEET 8 ELEVATOR AREA (EACH SIDE) 21.61 SQ. FEET 9 RUDDER AREA 33.55 SQ. FEET 10.500 ]0 AVERAGE AILERON CHORD INCHES 26. 600 Ii AVERAGE ELEVATOR CHORD INCHES INCHES 29.000 12 AVERAGE RUDDER CHORD INCHES 223.058 13 STANDARD LONGITUDINAL C.G.

12.84 14 STANDARD VERTICAL C.G. INCHES INCHES 131.44 15 F.S. INLET 29.6800 16 W'L. INLET INCHES ll0.000 17 B.L. INLET INCHES 141.8 ]8 FS. NOZZLE INCHES 43.43 19 WL. NOZZLE INCHES 20 INCHES ll0.000 B.L. NOZZLE 21 INCHES 124.49 FS. PROPELLER 22 INCHES 43.43 WL. PROPELLER 23 INCHES ll0.000 B.L. _PROPELLER 24 DEG 0.0 ENGINE TOE-OUT ANGLE 25 ENGINE INCIDENCE ANGLE DEG 0.0 26 I007o NG RPM 37468.

1.00 27 RAM RECOVERY FACTOR 2.0 28 NUMBER OF ENGINES 511.93 29 X LOCATION H-TAIL (25°/°MAC) INCHES 3O 54.00 Z LOCATION H-TAIL (25% MAC) INCHES 31 519.500 X LOCATION V-TAlL (25% MAC) INCHES 32 8 ]. 000 V LOCATION V-TAIL (25°/°MAC) INCHES 33 l0.0 STANDARD FLAP SETTING l DEG 34 STANDARD FLAP SETTING 2 20.0 DEG 35 STANDARD FLAP SETTING 3 37.5 DEG FORCE WHEEL RADIUS INCHES 5.25 GROUND DECK ANGLE DEG -1.5 D.A.S. SAMPLE RATE SPS 8.567 PROP. DIAMETER INCHES 102.0 3.2.2 The weight, center of gravity (c.g.) and ir, ertias for the aircraft were determined from a combination of manufacturer's data and from experimental tests, measurements, and computations by KSR. The weight of the aircraft was measured in flight by subtracting the measured value of fuel used from the initial weight of the aircraft with full fuel, crew and flight test equipment. The zero fuel weight was determined by weighing the aircraft. The fuel used measurement was confirmed by weighing the aircraft before and after a number of the early flights. The fuel used measurement was found to be accurate to within 5(:) pounds. Between NASA flights Ii and 12, (KSR flights 6 and 7 ), additional equipment was installed ir, the aircraft. Therefore, two sets of plots for aircraft weight, c.g., and inertias are used. One set is valid for flights 6 through 11 and a second for flights 12 and or,.

Since weight is the primary variable for deterr,1ir, ing the center of gravity and inertias during any given flight, changes Jr, their values have beer, expressed in terms of fuel used. The same fuel burning sequence was used for all test flights. The initial weight, c.g. and inertias of the aircraft for each flight were adjusted for the addition of observers and test equipmer, t.

The center of gravity position in flight was determined as a function of the fuel used. The longitudinal position of the 3.3.1 center of gravity was established based on infornlation contained in the Pilot's Operating Handbook (P.O.H.) and data from the zero fuel weighing. The vertical center of gravity was determined fron_ specially provided manufacturer's data. No corrections were made for fuel slosh with change in attitude of the aircraft.

The coordinates used to measure the X and Z c.g. location are shown in Figure 3.3. 1. The origin is located 21 inches aft of the nose, in the forward baggage compartment. Waterline and fuselage station are measured positive upward and aft fron_ this point. Because this sign convention ls contrary to the convent ional aerodynamic body axis coordinates (poslt ive X forward and Z downward), the signs have been changed in the KSR fl i ght test data and are expressed in feet for consistency of the MAC would be of that data set. Thus, the leading edge listed as -15.69 feet in the KSR data base.

The lateral (Y-axis) location of the c.g. is assumed to be zero (i.e., lies in the X-Z plane).

Moments of inertia (Ixx, lyy, Izz, Ixz) were calculated based on manufacturer's data for the aircraft with empty tanks and with eight fuel Ioadings. The aircraft c.g. and inertia data for flights 6 through 11 are presented in Figures 3.3.2a thru 3.3.2g. For Flight 12 and on, this data is presented in Figures 3.3.3a thru 3.3.3g. The effects of crew and test equipment were accounted for, however, the dynamic effects of fuel slosh were not. This could cause variat ions of damping and momentary aircraft response more violent to control inputs during maneuvers.

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3.3.18 4. FLIGHT TEST OPERATION Data for this report were gathered in 18 data flights totaling 38.9 hours. The test airplane was the NASA Lewis Research Center modified De Havilland Twin Otter which has been equipped for icing research.

A summary of the total flight test program is given ir, Table 4.1. Aircraft instrumentation started in September 1985 and data system installation and calibration were completed by the end of the month.

A typical flight began with a ground transducer check.

Every parameter was checked for operation and its counts value hand recorded. The flight test englneer briefed the pilots on the maneuvers to be flown. Before takeoff, three preflight runs were recorded on the DAS. A "zero" run data record was taker_ with the airplane static, engines at idle, and the controls at a pre-determined position. Then a data record of a cc°ntrol sweep in each axis was taken. An engine run-up was used to check engine parameters. At this time, the cockpit engine ir_dications were hand recorded. After the flight, these steps were repeated.

The values from the "zero" runs were used to correct bias errors in some parameters, such as angular rates, air data, and forces.

4.1 TABLE 4.1 TWIN OTTER FLIGHT PROGRAM SUMMARY FLIGHT TIME PERF. S & C PERF. S & C MISC. REMARKS DATE FLT.

(DRY AIR) (DRY AIR) (ICE) (ICE) ( 198 5 ) NO. * TOTAL 1.4 System Check Out 27 Sep I (6) 1.4 2.4 Airspeed Cal.

30 Sep 2 (7) 2.4 1.0 1.0 Finished Airspeed I Oct 3 (8) 2.0 Cal.

2.9 3 Oct 4 (9) 2.9 2.1 4 Oct 5 (I0) 2.1 23 Oct 6 (ii) 1.9 1.9 I.]

]9 Nov 7 (12) i.] System Check Out 2.6 Airspeed Failure 21 Nov 8 (13) 2.6 4 Dec I0 (15) 2.8 2.8 2.2 4 Dec i] (16) 2.2 1.9 5 Dec 12 (17) 1.9 12 Dec 19 (19) 1.8 1.8 ].4 .5 12 Dec 20 (20) 1.9 2.2 13 Dec 21 (21) 2.2 ]4 Dec 22 (22) 3.0 l.O ! 16 Dec 23 (23) 2.5 i 18 Dec 24 (24) 1.8 1.8 2.4 I 19 Dec 25 (25) 2.4 Total Hours 38.9 3.7 5.9 8.4 9.0 Ii .9 PERF. - Performance DRY AIR Baseline Flights ICE S & C - Stability and Control Icing Data Flights

L

MISC. - Miscellaneous * Numbers in parenthesis represent the NASA flight test numbering system.

4.2 5. DATA ACQUISITION AND ANALYSIS 5. 1. Data Acquisition The data for this report were recorded on the KSR DAS. The DAS records data with 12 bits of resolution at approximately 8.6 samples per second. All channels are sampled within a one-millisecond time interval and then recorded on tape. This effect ively el irninates time skews due to sequential sampling.

The parameters recorded by this system are listed zn Table 5.1.1.

Figure 5. 1. I shows the DAS which includes the computer, computer controls and display, tape recorder, and signal conditioning. The KSR inertial transducer package which c,:,ntair, s three linear accelerorneters, three rate gyros, and an attitude gyro was rnour, ted near the longitudlnal and vertical c.g. but tc, the left of the lateral c.g.

5.1.1

TABLE 5.1.1

RECORDED DATA

NASA TWIN OTTER DHC-6-O04

VARIABLE # VARIABLE VARIABLE NAME UNITS

FILE COUNTER

1 FILCNT

BLOCK COUNTER

2 BLKCNT

ENGINEER'S STATUS BYTE

3 ESB

AIRCRAFT STATUS BYTE

4 ASB

PAUSE EVENT ROSEMOUNT HEAT GEAR POSITION GYRO ERECTION MARKER PILOT EVENT (SEC) TIME 5 TIME (G) LONGITUDINAL ACCELERATION 6 AX (G) 7 AY LATERAL ACCELERATION (G) Z-DIRECTION ACCELERATION 8 AZ (DEG/SEC) PITCH RATE 9 PITCH_RATE (DEGISEC) ROLL RATE 10 ROLL_RATE (DEGISEC) YAW RATE 11 YAW_RATE <DEG) PITCH ATTITUDE 12 PITCH_ATT (DEG) ROLL ATTITUDE 13 ROLL_ATT (PSF) PRESSURE ALPHA 14 DELP ALPHA (PSF) 15 DELP_BETA PRESSURE BETA (PSF) PRESSURE REFERENCE 16 DELP_REF (DEG) AILERON DEFLECTION 17 DELTA_A_L

i

EMPTY 5.1.2 19 ELEVATOR DEFLECTION DELTA_E (DEG) 2O RUDDER DEFLECTION DELTA_R (DEG) 21 FLAP FLAP POSITION (DEG) EMPTY EMPTY EMPTY EMPTY EMPTY EMPTY EMPTY EMPTY EMPTY i EMPTY DIFFERENTIAL PRESSURE DIFF_PRESS (PSF) INDICATED TOTAL TEMPERATURE AIR TEMP (DEG_K) INERT_VREF VERTICAL GYRO REFERENCE VOLTAGE (VOLT) STATIC PRESSURE STAT_PRESS (PSF) REFERENCE VOLTAGE (VOLT) CPT_VREF 37 REFERENCE BATTERY BOARD 1 (VOLT) BATTERY_A EMPTY FUEL USED (LB) FUEL_USED 40 PAF PILOT AILERON FORCE (LB) PEF PILOT ELEVATOR FORCE (LB) PILOT RUDDER FORCE LEFT PRF_L (LB) PILOT RUDDER FORCE RIGHT PRF_R (LB) EMPTY EMPTY 5.1.3 EMPTY EMPTY EMPTY (%) GAS GENERATOR RPM-N1 LEFT 49 N1 _L (%) GAS GENERATOR RPM-N1 RIGHT N1 _R (%) PROPELLER RPM LEFT PROP_RPM_L (%) PROPELLER RPM RIGHT PROP_RPM R (PSI) ENGINE TORQUE PRESSURE LEFT 53 TORQUE_L (PSI) ENGINE TORQUE PRESSURE RIGHT TORGUE_R (LB/HR) INDICATED FUEL FLOW LEFT ENGINE FUEL-FLO_L INDICATED FUEL FLOW RIGHT ENGINE (LB/HR) FUEL_FLO_R FUEL TEMPERATURE LEFT ENGINE (DEG_K) FUELTEMP_L 58 FUEL TEMPERATURE RIGHT ENGINE (DEG_K) FUELTEMP_R EMPTY EMPTY EMPTY EMPTY EMPTY (VOLT) 64 LIQUID WATER CONTENT INDICATOR JOHNS_WILL (VOLT) 65 LEIGH ICE DETECTOR UNIT (VOLT) ICE DETECTOR UNIT 66 ROSEMOUNT (VOLT) 67 DEW POINT HYGROMETER GEN_EAST EMPTY EMPTY EMPTY (VOLT) 71 REFERENCE BATTERY BOARD 2 BATTERY_B 5.1.4 THIS PAGE INTENTIONALIY LEFT BLANK 5.1.6 5.2. _R_ie!__D_r_ Special sensors were installed on the test airplane as described in the following paragraphs.

5.2.1. Traili_g__D_ To calibrate the airplane's airspeed system, a trailing cone was used to measure atmospheric pressure. The cone was trailed approximately 100 feet behind the airplane on a flexible tube which entered the aircraft through a metal tube connected to the aircraft at a_i openir,u created by removing the tail skid. A previous inflight inspection confirmed that the cone trailed well in all configurations. Static ports were located approximately 5 cone diameters forward of the cone where the pressure is uninfluenced by either the cone or the airplane. Cone static pressure was sensed by at, Calibration methods are independent pressure transducer.

described in Section 7.3.

5.2.2.

Primary Control Force Measurement KSR has developed a system to measure column, wheel, and toe brake forces. Elevator and aileron control pedal, forces are measured with the force wheel shown in Figure 5.2.1. Rudder and toe brake forces are measured with the pedal system shown in Figure 5.2.2.

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K b h S 4 5 ) R A W V ( L b W R f h i t 5.2. 3. A_ngle of Attack and SidesliQ Sensors A Rosemount Model 858AJ Type Flow Angle Sensor to measure angle of attack arid sideslip was installed on the aircraft noseboom. The system consisted of five pressure ports; or,e located at most forward point and four others located to the left, right, above and below the forward port. The geoti_etric layout of the system is illustrated in Figure 5.2.3. A char, ge in angle of attack (alpha) or sideslip (beta) resulted In a change in differential pressure between the resoectlve pressure ports. Calibration methods are described in Sect1,_, 7.3.

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KOHLMAN SYSTEMS RESEARCH.

5.2.5 LAWW_C_ K&NS&S D_AWI_! ! ;i J THIS PAGE INTENTIONALLY LEF_ BLANK

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5.2.6

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5. 3.

Data Analysis The data were processed or, a Cadmus 90C)C_ microcomputer and a Gould SEL 32/77 minicomputer both located at KSR in Lawrer, ce, Kansas. A block diagram of the data management system is shown in Figures 5.3. la and 5. 1.3b. This data management system was designed to function with the KSR DAS.

The initial phase of the processing involved transferring the flight tapes to the Cadmus in a raw data format. The data are then converted to engineering units, and the air data are correct ed for position error. Also computed are the weight and balance, inert ias, linear accelerations at the center of gravity, correct ions to alpha and beta due to angular rates and boorn bending, thrust, and other parameters required for subsequent analysis programs.

The flight test data base was then transferred to the SEL where it is saved on magnetic tape for subsequent data analysis.

The data analysis techniques are discussed in Section 7. The flight test data base was stored in unformatted, 200-word blocks (each word is 32 bits). Each block represents c.ne time-slice of flight test data, and every measured or derived parameter occupies a specific word as listed in Table A. 1 ir, Appendix A.

5.3.1 RECORDED I FL]GH'r DATA

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DHC-6-O0& KOHLMAN SYSTEMS RESEARCH

- I 5.3.2

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5.3.!_ DHC-6-O0/, _ 5.3.3 ' THIS PAGE INTENTIONALLY LEFT BLARK 5.3.4 6. AXIS SYSTEM AND SIGN CONVENTIONS The axis system and sign conventions used in this report are defined in Figures 6.1 and 6.2. Figure 6.1 shows the sign con- ventions used for body motion and Figure 6.2 shows the sign con- vent ions for the control surface deflect ions. Rudder and ai leron trim have the same conventions as rudder and aileron deflectic0n.

Posit ive control forces move pi lot control posit ions in a positive direction.

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6.3

THIS PAG_INTENTIONALLY LEFTBLANK

6.4

7. E_I_!_!E_!_Q_ALYSIS TECHNIQUES techr, i- This section gives a general outline of the various ques used by Kohlman Systems Research to obtain the results presented in Section 8.

7. 1. Stabilit_ and Control Derivative Methods The primary method used to determine stability derivatives is the Modified Maximum Likelihood Estimation technique (MMLE).

The MMLE method is based on an assumed mathematical model of the airplane's aerodynamic and inertial characteristics. Initial conditions and dynamic control inputs measured in flight are applied to the model and the response of the model is compared to that of the airplane. The difference is a response error. The MMLE program then changes the aerodynamic derivatives by a compu- tational algorithm to reduce the response error. The new derivatives are used in the math model to compute a new response error. This iteration procedure continues until a specified convergence criterion is met. The final derivatives represent airplane aerodynamic characteristics that maximize the likelihood that the error between airplane and mathematical model responses will be minimized.

7.1.1

The accuracy and consistency of results obtained by MMLE

methods are dependent on many different factors. First, the model must _eamonably represent the airplane. Since the model is linear, any nonlinearities during a maneuver will affect the results. Good results achieved during relatively linear, small perturbation maneuvers will represent only local slopes.

It Nas found on the lowest speed flight test runs, buffet due to ice accretion caused the MMLE analysis to converge poorly or not at all. Therefore, in general, the lowest speed data points for all configurations were not included in the analysls.

Since the MMLE program will attempt to match _h_tever actual response is presented to it, it is important to minimize the sources of error in the recorded data. Air turbulence can be minimized by doing tests only when the air is smooth. Data system noise minimization requires a data system having a _ery errors in calibrat_o_ curves high signal to noise ratio. Bias used to produce the data do not affect the aria L ys _ _ Howev_ of the c_olzbr_ior : _irvt÷_ d, changes in sensitivity or slope tc_c, slow ._ chann_l_ _t Time skews, caused by sampling between undetected G_ eat c,_,c rate, produce subtle errors that can go has been taken to eliminate these problems.

The final non-dimensional derivatives obtained from MMLE are dependent on the initial conditions of the maneuver (airspeed and altitude) and the airplane weight and moments of inertia. The MMLE program will attempt to correct any errors in these para- REV a 7.].2 meters by adjusting the derivatives. For example, if the roll inertia is calculated to be higher than it actually is, the MMLE derived roll power C_6 will be higher than it actually is. This a illustrates the importance of obtaining accurate airspeed cali- brations, weight and inertia data, and accurate calibration of the data acquisition system.

An illustration of the ability of the MMLE program to match the model response to the airplane response is provided in Figure 7.1.1 which presents time histories for an elevator doublet. The quality of these matches provide confidence in the accuracy of the derived stability derivatives. However, these results are collected _or many runs over a wide range of flight conditions, thus some judgement is involved in arriving at the stability derivative curves.

Careful interpretation was required for each data point produced by MMLE. Important characteristics included the level of excitation of particular parameters, the quality of the time history match provided by the MMLE program, and the nature of the dynamic maneuver. Full documentation of each individual MMLE data point is available for inspection and further interpre- tation as required.

The MMLE results, combined with theoretical aerodynamics and were used to quality of individual maneuvers, judgements on the believed to represent the form derivative curves that are airplane as accurately as possible.

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[ ,*$=PD DHC- 6- ()04 i APPD , _ RESEARCH 7.1.4 KOH_MAN SYS'EM _ 14:5_:24 _ME 7.2 _f_£[]_D_ ModeliDg Methods Performance modeling is a techr, ioue developed by KSR through which a complete model of the propulsior, system and the static longitudinal aerodynamic coefficients of the airplar, e are defir_ed. This is done by performing a series of level flight full and partial power accelerations over the flight er, velope.

The first step ir, perforrnance rnodel ing is to define a propulsion system model from the engir,e deck provided by the er_gine manufacturer and the propeller characteristics frc°m the prope I !er manufacturer, mh_ _=+ ,=+ +_ .... + ca _''!ated fronl the deck is used direct ly, while ir, fl i ght measured torque and propeller rpm are used to calculate shaft horsepower.

Once the propulsion model is defined, the aerodynarnic lift and drag models are determir, ed Jr, terms c,f: VS. (_ a) CLA/c b ) C D vs. CLA/C c) c VS. Ci m untrimmed 7.2.1 The forces and mc, mer, ts pertir, ent to the steady state analysis are shown in Figure 7.2.1. The airplane lift coefficient is generated with the following equations.

C L = [ W'n - Fpsin(a + X p) - FjsJn(e_ + _.j) ]/qS A/C Zwind where : C L = C L - C 6 untrimmed A/C L 6 e e from MMLE data reduction CL 6 _-p propeller inclination an?to X iet exhaust inclination angle ] Fj net jet thrust CD 6 The airplane trim drag car, r,,:,t be det.=r_ rl, .._n=_- ÷:.-c,r_ e the maneuvers performed.

Therefore a trimmed drag coefficier_t is generated from: C D = [Fpcos(_ + ),p) Fjcos(_ + X j) - W'_ x ]/qS wi nd 7.E.2 Ur,t r i mined ( 6 = O) pitchir, g momer, t coefficient data were e also derived from the performance maneuvers. The governir, g equations are: c .= c u a_- c_ a - c - CDA_ muntrimmed unt 6 e m T where: AX : (X - X ) Cgs td Cgts t A_. (Z - Z ) Cgst d Cgts t CmT= -(FjZTj + FpZTp) _SE C ZTp and ZTj are m6 and CL 6 , are predicted from MMLE.

e e a function of airplane geometry and c.g. posit ior,.

A summary of the performance flight tests is presented it, Table 7.2.1. These tests consist of a series of accelerations and decelerations at selected torque settings and weight-pressure ratios < W/6).

The various airplane cor, figuratior, s evaluated durir, g these maneuvers are listed below: FLAPS GEAR CONFIGURATION Cruise 0 Down Approach 10 Dowr, Approach 2(1) Dowr, Takeoff 0 Dowr, Takeoff 10 Dowr, Landing 37.5 Dowr, 7.2.3

At the time the data was being analyzed, it was felt that

the KSR torque readings were in error when compared to values

recorded in previous flight test by NASA. The evidence of this

effect was most noticeable in the drag polar. After much

discussion it was decided to adjust the torque readings so that

the KSR drag polar matched the NASA drag polar shown in the NASA

re0ort given at the AIAA 23rd Aerospace Sciences Meeting (paper

number AIAA-85-0468). To match this drag polar 45 foot-pounds of

torque was subtracted from the recorded torque value of each

engine. The flight test data base mag tapes that were made for

NASA do not have this adjustment made for the engine torque so

that the user could make any changes as deemed necessary.

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7.;_.6 7.3. _Ai_r Dat, a _alibration Technigues This section presents the techniques used to calibrate the air data measurement system. The static pressure system is corrected for static port position error. The delta-p alpha and beta systems are calibrated directly to degrees, which includes correction for upwash and sidewash.

Angle of attack and sideslip data were measured using a differential pressure systenl mounted on the nose boom of the air- plane. The sensors used are detailed it, Section 5.2, "Special Se_so_s _ .

7.3.1

The boom mounted forward of the nose of the aircraft con-

tained static pressure ports located several diameters aft on the cylindrical shaft of the Rosemount Flow Angle Sensor. The ship's standard system Cco-pilot's) pitot probe located on the right side of the aircraft nose was also utilized.

The airspeed calibration used a trailinD cone extending more than one fuselage length behind the airplane in undisturbed air. A static pressure port located on the tube leading to the cone, well forward (at least 5 cone diameters) of the cone, was used to determine the true ambient pressure. The ambient pressure was compared to the indicated static pressure measured on the boom to determine the position error.

stabilized at selected speeds and The airplane was alt it udes. Data were taken under stabilized conditions to prevent pneumatic lag from influencing the measurements.

contained in These data were analyzed using standard methods Reference 4 and are presented in Figure 7.3.1.

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APPD _ _ _" DE HAV'ILLAND T_3_I OTTE_ APPD ..L,J

KOHLMAN SYSTEMS RESEARCH 7.3.3

T_MI_ t6 :S_:47 7.3.2. B_gle of Attack and SidesliQ Calibrati¢2_ The angle of attack and sideslip were measured using a 858AJ is Rosemount 858AJ probe (see Section 5.2). The supplied with a manufacturer calibration. These calibratic, ns were used following a process of validation using actual flight test data.

The angle of attack calibration was verified by the analysis of stable points. With the aircraft stabilized at a constant altitude and airspeed the angle of attack is equal to the inverse tangent of the ratio of the longitudinal and The Rosemc, unt calibratic, n was normal body axis accelerations.

verified using this method to validate its use during this flight test program.

The validation process was repeated on the angle of side- slip calibration using the dynamic calibratior, technioue_ described in Reference 6. This method involves the analysis of rudder doublets and compares the measured sideslip angle to a sideslip angle calculated fror0 the inertial data. Ir, addition to the validation of the Rosemount calibration this analysis procedure also facilitated the investigatiors of dynamic lag in sideslip measuring system. This lag was found to be insignificant for the flight conditions experienced Jr, this flight program. The calibrations for angle of attack and sideslip are shown in Figures 7.3.2 and 7.3.3.

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I ' "ALC DATE Fig. 7.3.2 RO_EHOUNT PROBE: ALPHA CAL / _PD DE HAV3::I.I.AND "E_Z:N OTTER _PD

KOHLMAN SYSTEMS RESEARCH 7.3.5

"r_Mg m .... i........ i........ i......i........ i ..i ......... i........ i......... i........ i........ i........ i......... i..... i .... i........ i.... l.... iT_ .... i.... i.... il .... il .... i........ !.... I.... i.... I.... i.... I.... i.... l.... i.... I.... i.... .

.... i........ i.... I .... i....l .... i..._! .... I .... i--..l .... _---I .... _ ........ _.-I .... i.... I .... i---I .... i.... I .... i.... l.... _ .... " .... !I .... i"I .... i'I iI .... .... ....i'i ' .... .... .... .... .... iI .... .... i' .... .... i""I .... i'I iIi - .... .... .... ....

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Fig. 7.3.3 ROSEHOUNT PROBE BETA CAL DE HAVZLLAND I_]:N 01"i'E]_

I 'I'l "°

TZH[ !tl :S8:22 KOHLMAN SYSTEMS RESEARCH 7.s.6

An additional boom bending calibration was performed to correct for errors caused by bending due to dynamic rnaneuvers.

The calibration shown in Figure 7. 3.4 was accomplished by loading the boom to represent the load experienced in dynamic maneuvers and then measuring the deflection of the boom.

Since the boom is symmetrical in both axes, the beta bc0or,_ bending correction shown in Figure 7.3.5 was derived from the alpha boom bending.

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030 - NOT.LO3EIBO0 VH'Id'Tf

I

ALOHA BOOM BENDING CORR.

DE: I-_VZI.L_WD TWIN OTTER

KOHLMAN SYSTEMS RESEARCH 7.3.B

I : • : : : : : • : - . :_ : : ,,11.

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_ALC 83/04/88 R_VZ_IED DATIr .,'1 F_g. 7.3.5 BETA BOCIq BENDZNG C0R_.

CHECK _/_'/_

DE HAV'£L.I.AND "I'WZNOTTER DHC-O-O_ AP'PD ,.Jt.,J

RESEARCH

KOHLMAN SYSTEMS ?.3,9

"l-ZMnr 10:15:31 THIS PAGE INTENTIONALLY LEFT BLANK

h

7.3.10 This section contains the results of all analyzed flight test data.

This includes stability and control, performance, and special tests and calibrations.

All flight test data in this report assume a center of gravity position of 26_ MAC. Standard transformations can be used for other e.g. positions. The coefficients and derivatives in this report are based on the stability axis system.

Flights 20 and 21 were selected for performance analysis.

Flight 20 represented a moderate to heavy mixed icing condition and 21 represented a moderate glaze icing condition. Flights 16 and 17 represented mixed icing conditions and were used to analyze the longitudinal stability. Flight 16 compared between the baseline and all iced, and then baseline and wing only iced. Flight 17 compared the baseline with all iced, and then the baseline with the tail only iced. Flight 23 represented rime ice conditions and provided an analysis of the directional derivative ratio Ch6 / C hB for the r baseline the vertical tail iced configuration. Icing cloud data by NASA during natural icing test is shown in Table 8.1.

are in sub- The longitudinal derivatives and coefficients sections representing lift, drag, and pitching moment.

H.].] REV A 8.1.1. _ifl E_!ffi_it_, _ The lift decre_nts that were broken out by selectively deicing the aircraft, remain relatively the same percentage--wise at both high and low power settings. The total degradation in lift coefficient at 6 degree angle of attack is approximately 7X with 40 psi torque and 8_ with 15 psi torque. The individual contributions of each iced component are nearly the same at both power settings.

with the aircraft in the The elevator effectiveness, C L 6 e all-iced configuration showed an average 10% degradation.

When the tail only was iced this degradation averaged around 8.5%, and when the wing only was iced a 2% degradation was found. When flaps were lowered to 10 degrees the all-iced degradation in these coefficients increased to 15-16% and the wing ice accounted for approximately 9% or nearly equiva- lent in magnitude to the tail ice contribution.

It was found that the state coefficient (C L + CL. ) q due to ice in the flaps-up case, however, showed no change when the flaps were lowered 10 degrees and there was ice on the wings, tail, or both, this coefficient was degraded approximately 23 percent.

REV A 8.1.2a TABLE 8.1 SUMMARY OF AVERAGED ICING CLOUD DATAFOR PERFORMANCE AND STABILITY AND CONTROL Flt Start End Dew Aft., TAS, AOA, Static Average Avg No. Time Time Point LWC MVD PA, Kts deg Temp -3 ft °C °C g-m Em__ 16 14:32:58 15:12:18 8060 138 6 1.4 -u.u 19 O. 25 17 10:08:28 10:27:18 7309 135.0 1.6 -7.2 -8.1 .33 21 2O 12:42:48 13:36:58 6163 127.3 1.8 -9.5 -10.1 .46 14 21 10:40:38 4315 130.8 1.6 -5.0 -4.8 09:55:38 -. 20 15 23 10:13:38 11:11:48 4330 136.4 0.5 -10.7 -10.0 • 33 10 REV A

THIS PAGEINTENTIONALLY LEFT BLANK

REVA 8.1.2e

The aerodynamic lift force coefficients investigated

for the NASA Twin Otter are presented below. The lines

designated as "baseline" represent the dry air baseline testing. The iced configurations are defined using the letters E thru J, with multiple designations for a given curve indicating that there was no discernable difference between results for the different configurations.

Letter Code Aircraft Confi|lur_at io_n E Cruise configuration, aircraft all iced F Cruise configuration, aircraft complete tail and struts iced G Cruise configuration, aircraft vertical tail and struts iced H Cruise configuration, aircraft horizontal tail and struts iced I Cruise configuration, aircraft struts iced J Cruise configuration, residual ice only Ei_ure # 8.1.1a - 8.1.1d CLA/C 8.1.2a - 8.1.2b (C L + CL. ) q c_ 8.1.3a - 8.1.3d C L6 e

I F.C DING PAC BLANK I 1" FILM

8.1.3 REV A T .................. a ..... m bJ Z ... ..... i .... ..... i:.....

r-- ....................

7 .....

i --J D 7 o CALC REVISED DATE CHANGE IN AIRCRAFT LIFT DUE TO ICE, FJ g. 8.1.3 CHECK TORQUE 40 PSI, FLIGHT 20 DHC-6-004 f I_PPD j APPD 8.1.4 KOHLMAN SYSTEMS RESEARCH- DRAWN LAWRENCE I&NIA| z,c, Rm,_ W Z _1 W mL_LL W bJ ..J Z ILl :3 O0 I-0 "rl- I U.O..

if) !

"1:) CALC I i REVISE° DATE Fig.

CHAN(,E IN AIRCRAFT LIFT DUE TO ICE, _ _ _8.2__L]_ TORQUE ] 5 PSI , FLIGHT 20 DtlC- 6-004 ..... 4 - iAPPD I i KOHLMAN SYSTEMS RESEARCH .. 8. 1.5 t AWR! hie ! M ANSAS m m Ld Z (/.)

(.J F-- I- h O UJ ..I Z I if-) t¢') i "1_)

_ALC li_EV,S_r,i OA_E

CHECK r- CHANGE IN AJRc-RAFT LIFT DUE TO ICE, TORQUE 40 PS], DilC-6-O04 FLI(,II7 21

....... I T!- -_ ...... APPL]

KOHLMAN SYSTEMS RESEARCH- _.1.6 LAWN| NCf KANS&5 LO Z .J LO U_ I- J ({ h i,i / W (D Z Nn _ T I !

7 o CALC I REVISED DATE Fig.

C|tANCf_ IN AIRCRAF] LIFT DUE J'O ]CE, 8.1 .]d CHECK TORQUE ]9 PSI, FLIGtt] 21 DHC-6-O04

APP

A P PD 8.1.7 KOHLMAN SYSTEMS RESEARCH.

LAWRENCE K AN_*A$ DRAw. Z_ I_, t FLIGHT 16 _ 17 C T = .06--.08 (CL.q+ CL_)boseline = 10.8 rod -I S F = 0 ° z_(CLq+ CL _) ice = 0 Fig.

AL_C !

EFFECI (-)I ICE ON S1A]t:: COEFF 1 C 1 ENTS 8.1 .2a ( (:1. _ C ) {)_ 1._ ' 6F (I Fl.l(;lllg I(_ _, 17 : -]]REVIS[D T_]

8:1:B

KOHLMAN SYSTEMS RESEARCH.

L AWRf NC! KANSAS CT= .05 - .07 -I = 5.9 rad +C L ) [ CLq &, baseline 6F=IO ° FLIGHTS 16 E_ 17 -3.0 I qo o -2.0 t== v (p , ALL ICED"_ o TAIL ICED /_ J WING I C Ep_.)

U -I.0 4- .J

.(o .15 .£o

.05 MACH • T.

F_g.

CALC i REVISED COEFFICIEN] DATE EFFECT OF ICE ON SLATE 8.] .2b (CL + CL- )' 6F = ]0°' q _' DI1C-6-004 APPD _ IY l FLI (;ttTS ] 6 & ] 7 APPD KOHLMAN SYSTEMS RESEARCH . E_. 1.9 LAWRENCE KANSAS C T = .06 - .08 = .0069 deg -I • CL6e baseline 6F=O" FLIGHT 16 -.0015 -.0010 ALL ICED-- I @ TAIL ICED-- "1o o -.0005 o_ ID --!

¢O

<3

.,g .2"o .z'5

MACH Fig.

CALC _ _REVISED DATE 8.] .3a CtIANCE IN ELEVA]-OF L]FT DUE ']'0 ICE, ])HC-6-004 0 F : 0 °, FI, IGHT ]6 APPD i "i - -'_ -- - - . 8.1.10 KOHLMAN SYSTEMS RESEARCH ' " 't I LAWI_E NC_ KANSAS C T • .06- .08

.... CLse baseline= .0069 deg- I

6 F = 0 ° FLIGHT 17 -.0015 -.-oo-.0010 "i ALL ICED

d

-.0005 • WING ICED

.,'5 .2'o .z'5

MACH Fig.

CALC REVISED DATE 8.1.3b CHANGE ]N ELEVATOR LIFT DUE TO ICE, CHECK _ -- 6 F = 0 °, FLIGHT ]7 DIIC-6-O04 APPD KOHLMAN SYSTEMS RESEARCH_ 8.1.11 LAWR|NC_ KANSAS DRAWN KE-i_ _ _;_ F.._ CT=.05-.07 -I = .0056 deg CL 6ebaseline SF= I0" FLIGHT 16 -.0015 ===, !

-.0010 ALL ICED-

=o

d

TAIL ICED-- (_ -.0005 .05 .10 .1'5 .20 MACH ] I ]i I DATE Fig.

CALC I t ii REVISED CHANGE ]N ELEVA'IOR L1F'I DUE "I0 ICE, 8.1.3_ DttC-6-004 _W/-q- ..... _.......

6l. 1()" FI.I GtlT 16 KOHLMAN SYSTEMS RESEARCH.

8.1.12 LAWM| NC| KAN,_A_ ................... :...ct. =..OS: .07 ....

...... ° !_ ......................... :_ ................... :SF.._'_._: ...... • ...............

i: :: : _ ! :.eL..:_: ..... :ZOosS=,.-_, ...... : ...............

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L:.......... : ........................... :. . - ......

-.0015.

'T

-Oh .:0 v Q - .... .,.y -.0010 , ALL ICED- QI) {---- a_ -

- -- .0005 ,, • • -WING ICED --

i : ........... O = • ) ...()5 .lO .15 .20 . .0 r ._.:,,_. __.:. ...........

................ MACH Fig.

CALC REVISED DATE CHANGE IN ELEVATOR LIFT DUE TO ICE, 8.l .3c CHECK DHC-6-00L A PPD/_ 6 F = 10 °, FLIGHT ]7 A PPD KOHLMAN SYSTEMS RESEARCH_ B.1.13 DRAWN LAWRE NCE I_ANSAS As it was four_ with lift, the drag decrements remain relatively the same percentage-wise at both high and low power settings. Within the linear portion of the fliQht envelope the relative magnitude of lift and drag decrements due to ice appear independent of power. However, the absolute magnitude of these decrements are dependent on power effects.

FliDht 21 points out that the shape of aircraft ice accretions on both lifting and nonlifting surfaces is the most important factor influencing performance. For example, the icing encounter on flight 20 lasted 54 minutes at an approximate average LWC of 0.20 gm/ 3. Though the average MVD on each flight was approximately 14 to 15 microns, the temperature differed by 4.5 degrees C. This difference in temperature resulted in a glazed-type ice formation on flight 21. The encounter on flight 21 was nine minutes less than on flight 20 and the LWC less by a factor of greater than two, yet the overall drag increase was about 15-20% more than that of flight 20.

The aerodynamic drag force coefficient investigated are presented below. The lines designated as "baseline" repre- sent the dry air baseline testing. The iced configurations are defined using the letters E thru J, with multiple REV A 8.1.14 indicating that there was no designations for a given curve results for the different discernable differmnce between configurations.

Aircraft GgD_in_E_iRD _£__9_ Cruise configuration, aircraft all iced E Cruise configuration, aircraft complete tail F and struts iced Cruise configuration, aircraft vertical tail G and struts iced H Cruise configuration, aircraft horizontal tail and struts iced Cruise configuration, aircraft struts iced I Cruise configuration, residual ice only J E_g_re # 8.1.4a - 8.1.4d C D 8.].15 REV A r ................

i i :1,-1 .... • :o'> i ! ; _. ............................ ,_ ........ _ ............. i........ i ...... i....... i-1 ...... till ! b. i(,D _ "'ID " ' • : .... : • 0,,I •

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.CO ioj _1 O .¢D ! i_N.__O !.'1 ¸ L ..... - ...... " .... : ...... " . .

[ ...........

L__I............... _. i..............

r '_" N 0 0 q o O3 Fig.

CALC REVISED DATE CHANGE IN AIRCRAFT DRAG DUE TO ICE, 8.1.4a TORQUE 40 PSI, FLIGH] 20 o Ec, 2Y.. g.

!DI!C- 6- 004 A_PD_' -f APPD KOHLMAN SYSTEMS RESEARCH.

8.1.16 KF-JB C>RAWN Z_ Frq_l_,, LAW_|NC| KANIA$ cD N N _1 bJ O_ NO I- o XI-- o,1 0 CO _ 0 O. 0 ; T Fig.

(_: ALL. 1 | li REVISED I DATE CttANGE ]N A1RCRAF] DRAG DUE TO ICE 8.1.4b TORQLIE ]5 PSI, FL](;HT 20 DtIC-6-O0_ B.I.I?

KOHLMAN SYSTEMS RESEARCH.

...... t - t' ....

L AWlqlE NEE KANSAS _D ! • I

I

bJ Z / hi <{ m L N.

i CO

¥

._1 U ) _O '_" 0 E) _ o. 0 0 O0 Fig.

CALC i _ I;rEVISE b I DALE" CHANGE ]N AIRCRAF'I DRAG DUE TO ICE, 8.1.4¢ CH_CK "I'ORQLTE 4C) PS1, FI. IC,,H] 2] DH(:- 6- O0/-, APPD APP_ 8. 1. lt:l KOHLMAN SYSTEMS RESEARCH- " i ..... t DRAWN LAWRt NL! KANSAS Z ,.j .... 2.- ............. _ ..............................

e!, .... -i, m b ...%. Q q o OO Fig.

CALC REVISED DATE CHANGE IN AIRCRAFT DRAG DUE TO ICE, 8.1.4d TORQUE ]5 PSI, FLIGHT 2] APPD _V /" APPD WOHLMAN SYST'i;MS RESEARCH.

LAWRENCE KANSAS DRAWN Z_ l_._e6 KF-.E_ 8.1.3. Eii_hinn__gm_D___Q_ffi_i_Df, Q_ The effect of aircraft icing on pitching moment at the 15 psi torque setting was not seen in data from either flight aO or 21. At the 40 psi torque setting small reductions were measured in pitching moment on both flights. However, both the glazed and mixed type ice shape gave the same results.

The elevator power, Cm6 with the aircraft in the e all-iced configuration showed an 10_ degradation. When the tail only was iced this degradation averaged around 8.5%, and when the wing only was iced a 2% degradation was found. Whet, flaps were lowered to 10 degrees the all-iced degradation in these coefficients increased to 15-16% and the wing ice accounted for approximately 9% or nearly equivalent in magnitude to the tail ice contribution.

It was found that the pitch damping state coefficient showed no chan_e due to ice in the flaps-up case, however when and there was ice on the the flaps were lowered 10 degrees coefficient wings, tai I, or both this was degraded approximately 23%.

The aerodynamic pitching moment coefficients investigated are presented below. The lines designated as "baseline" re- present the dry air baseline testing. The iced configura- REV A 8.] .20 letters E thru J, with multiple

tions are defined using the

that there was no designations fop a given curve indicating for the different discernable difference between results configurations.

E Cruise configuration, aircraft all iced F Cruise configuration, aircraft complete tail and struts iced G Cruise configuration, aircraft vertical tail and struts iced H Cruise configuration, aircraft horizontal tail and struts iced I Cruise configuration, aircraft struts iced J Cruise configuration, residual ice only 8. 1.5a - 8. 1.5b Cm 8. 1.6a - 8. 1.6b (Cm + Cm. ) q c_ 8. 1.7a - 8. 1.7d _.] .21 REV A - . ........ : ....... .. • • : i-......... :- .............. -: .......................... "; ............ :" ....... ";...........

T" : " ":_:'-?-" L ;..::-.:":-."__ • _ -'":" ....... _T ........... .......: - ..... "r"..... ". ........... .e-.. --- L:.::_ .... :---:-: ......... : ....... i.-;:;:-- -.... :- _ .. i.: _._.:...... .: .....

!:: :!; i :¸ : ._ L '.. _.i-. _.L ...... : .....i.L::,, i. ....... : ..............

i---:--==--- i ...... f ....... :............. _ ......... "

,,.,J

_ ............................... .,, ...... :,_ _,_

__:.__ ......................... _. :, \../III . . ... .

___:_ :_o:_ ........... :_:.... :...... .\ _,,-\ _ .......... '.-' -_

/ : ..3 ' l/,/' " " " :,.o /// _ .

. , ,l.## !

",,---,.'---- .........

|.-_ I I_I -- 0 ._ e,,l I I u.l:_ Fig.

CALC REVISED DATE CHANGE IN AIRCRAFT PITCHING 8.1.5a MOMENT DUE TO ICE, TORQUE 40 PSI, CHECK L_.

FLIGHTS 20 & 21 APPD APPD I_'E;_ KOHLMAN SYSTEMS RESEARCH- LIWItINCE IANIAI I Z rr tO Z _J J D Z W z_ _J v W 0 i- tiJ II1 F- _D !

O o T U- I N" I-- i 0

'/j

W rr t--- _J TO') LL O n Z -JU_ LL-- I if) I T t_ IN -- 0 -- CM I I m 0 i Fig.

CHANGE IN AIRCRAFT PITCHING 8.1.5b MOMENT DUE TO ICE, TORQUE 15 PSI, DIIC-6-O0 FLIGHTS 20 & 2] KOHLMAN SYSTEMS RESEARCH- B.1.23 LAWRE NCI[ KANSAS FLIGHTS 16 8, 17 C T = .06 -.08 (Cmq + Cm,_ )baseline = - 48 rod -I SF= 0 ° Z_ (Cmq+Cm_.)ice = 0 Fig.

CALC REVISED DATE EFFECT OF ICE ON S'IATE COEFFICIENTS 8.1 .6a (c + c ),6F =0° rrl m° q e_ FI.lgUTS 16 _., ] 7 APPD KOHLMAN SYSTEMS RESEARCH..

DRA* , KEe

t AWR! NEE R&NSAS CT= .05- .07' (Cmq4- Cm_) baseline =-26 rod -I 8 F = I0" FLIGHTS 16 _ 17 ALL ICE D--_ T 6 TAIL ICEDP- "0 WING ICED3 t.

Q) U • -- 4 E ÷ E o o .o5 .io .15 .2o MACH / F_g.

L REVISED t CALC DATE_ CHANGE IN P]TCH DAMPING STATE APPD ' "'1 l() ° , l.'l.t(;tilS 16 (, 17 DIIC-6-004 A PPL) f]

..... ] l

' _ / KOHLMAN sYSTEMS -R-ESEAR6i_ _- e. I .zs C T = .06- .08 =-.0285 deg -I ' Cm $e baseline 6 F = 0 ° FLIGHT 16 .003 • ALL ICED- TAIL ICED '-" .002 t_ E .001

<3

_V ' ' 0 .10 .15 .20 .25 MACH !1 CALC DATE CONTROL POWER DUE CHANGE IN ELEVATOR TO ICE, 6 = 0 ° CHECK FLIGHT 16 F APPD APPD KOHLMAN SYSTEMS RESEARCH- DRAWNI LAWRENCE KANSAS CT=.05-.07 Cm6e .0232 deg-I b0seline = SF =10" FLIGHT 16 .004, ALL ICED i o= "o TAIL ICED- .002 • II) (J o_ q) (0 E .001.

(.;

<3

• 1'5 .2"0 .10 0 .05 MACH Fig.

CHANGE ]N ELEVATOR CONTROL POWER 8.]. 7_ DUE TO ]CE, DHC-6-O0_ 6 F = ]0 °, FLIGHT ]6 KOHLMAN SYSTEMS RESEARCH,. B.1 ._7 LAWRENCE KANSAS C T -- .O6--.08 -I Cruse = -.0285 deg baseline S F = 0 ° FLIGHT 17 .OO3 ALL ICED- m "1 O_ ¢D "10 .002 a) O °m Q) .001 E C)

<3

WING ICED-- I .20 .:25 0 .I0 .15 MACH ; Ii [ Fig.

CALC I _ REVISED I DATE (:ONTROL POWER DUE CHANGE iN ELEVAI'()R 8.1.7_ ' ;4 TO ICE, 6 = 0 ° , CHECK DHC- 6-004 FL]GH3" ] 7 F APPDd ...... _L APPD KOHLMAN SYSTEMS RESEARCH.

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Time history plots are presented to show the effect of ice on the vertical fin. These are presented for six runs from two flights in Figures 8.1.8a thru 8.1.13g. Perhaps the most interesting time history of this series is run 25, flight 23, which shows a decrease in vertical fin power due to ice.

For five selected runs the ratio of C_/C_was calculated and u _ r plotted in Figures 8.1.14 and 8.1.15.

Time history for N111+36A 8.1.8a - 8.1.8g Time history for N111+37A 8.1.9a - 8.1.9g Time history for N123+23A 8.1.10a - 8.1.10g Time history for N123+24A 8.1.11a - 8.1.11g Time history for N123+25A 8.1.12a - 8.1.12g Time history for N123-26A 8. 1.13a - 8.1.13g - 8.1.14 Ch 0 /Ch B plot for runs N111+36A & N111+37A r - 8.1.15 plot for run N123+24A, N123+25A & N123+26A Ch 6 /Ch 6 r r REV A 8. ]. 30 O ¢'g _ _10 - -- ....

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.J Wn l | 0 i . 0 GD _ 0 "T i i i/q_) / Jgq 9 S/q:) / J8 q3 CALC REVISED DATE Ch6 RATIO OF ]() ChB FLIGHT ]6 r DIIC-6-004 APPD APPD KOitLMAN SYSTEMS RESEARCH., 8.1.73 LAWRI.; NC I! KANSAS 1%1 Ix) ILl I"- I'M (JO

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CI,6 RATIO OF TO ChB FL]GHI' 23 r DIIC-6-O0_ APPD t t I-- . J .... -<+ - -- i .... T-- + l APPD _ I I I" _,. 1.74 KOHLMAN SYSTEMS RESEARCH- _AWN 21 _(- E LAWlq_ NCf _ANSA_ 91 REFERENCES 1. Flight Manual DHC-6 Twin Otter Series 200 (PSM 1-62-1A) Schweikhard, W.G. "A Method of In-Flight Measurement of Ground Effects on a Fixed-Wing Aircraft," Journal of Aircraft, Aprll 1967.

m Schweikhard, W.G. and Kohlman, D.L., "Flight Test Principles and Practices," University of Kansas Extension Divislon, 1982.

4. Hoak, D.E. et al: USAF Stability and Control DATCOM, Wright Patterson Air Force Base, Ohio, September 1970.

5. Larson, T.J. and Flechner, S.A. , "Wind Tunnel Invest igat ion of an All Flush Orifice Air Data Syster_1 for a Large Subsor_ic Aircraft, " NASA TN 1642, 1980.

6m Gilyard, G.B. and Belte, D., "Flight Determined Lag of Angle of Attack and Angle of Sideslip Sensors in the YF-12A Airplane from Analysis of Dynamic Maneuvers," NASA TN D-7819, 1974.

9.1

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I

I

9.;' A_ENDIX A: FLIGQI I_E_SI_ _D_A___A BASE F__O_R_M_A]_ This Appendix cor, tairts a table c0f the flight test data base (]able A. I) used ir_ KSR flight test ar_alysis. The table cor_tair, s the pertiner_t variables with their cc, rrespor_dir_g cori10uter r_arnes and ur_its.

The fii0ht test data base is saved or_ MAG TAPE in the followir_g forr_lat. The data are stored as 5 recr0rds oer ohysical block. Each record of 2c')0 words (32 bit floating poir_t) represer_ts one time slice of flight test data. Table A. 2 defines the floating represent at ior, used on the Cadmus 9000. The r0urnber of each variable indicates the correspor_dir, g word irm each block where the value is stc, red. The ta_e is partioned ir_to files separated by orme EOF mark with the final file er_dir_g with two EOF marks.

For each file, valid data begir_s with the first block after the block wherJ BLKCNT equals -300. It ends wherJ BLKCNT equals -600. Ir_ the block wher_ BLKCNT equals -300, variable #4 cor, tair_s the aDprc, xi- mate size of the file ir_ blocks of 200 words.

Nornla!ly, data are sampled cor, tir, uously, however, the KSR DAS has the capability to "pause" A pause of arbitrary ler, gth is sianifiee by Erlgir, eerir, g Status Byte equal t0-, I00 fc, r the first block after tree _use. The DAS also can sample ira a "burst "" rnc, de, where, it wiil a]terr, a%ely record eight sarnples a_ 8.5 Hz armd ther, pause for" one _.: i5 secc, r_c1_. For these cases ESBI : i01 to 115 for the firs_ san, r-,i_- _fte_." _he pa,_s-.e. T_,e _ast two oiQits of ESBI are approximately e___,ul _,:, tree ler, n_th of the o_uEe ir_ secor_ds.

A.I N1 FTDB VARIABLE LIST TABLE A.] LeRC Twin Otter Icing Fltm Date: 03Mar86 Twln Otter Time : 15:40 DHC-6-004 FTDB NUMBER VARIABLE NAME CHANNEL I I F ILCNT (DIGITAL) 2 BLKCNT (DIGITAL ) 3 ESB (DIGITAL ) 4 ASB (DIGITAL ) 5 TIME (SEC) 6 AX (G) 2 AY (G) 8 AZ (G) 9 PITCH RATE (DEG/SEC) 10 ROLL RATE (DEG/SEC) II ii YAW RATE (DEG/SEC) 12 PITCH ATT (DEC) 13 ROLL ATT (DEC) 14 DELP_ALPHA (PSF) 15 DELP_BETA (PSF ) 16 DELP_REF (PSF ) 17 DELTA_A_L (DEG) e 19 DELTA_E (DEC) 20 DELTA_R (DEG) 21 FLAP (DEC ) e 23 .....

Z'q 35 R}'N_I A.2

TABLE A.1, CONT'D

26 RPN02 26 2_ 29 .....

32 DIFF_PRESS (PSF) 32 33 AIR_TEMP (DEG K ) 33 34 INERT_VREF (VOLT) 34 35 STAT_PRESS (PSF) 35 36 CPT VREF (VOLT) 36 37 BATTERY_A (VOLT) 37 39 FUEL_USED (LB ) 39 40 PAF (LB ) 40 41 PEF (LB) 41 42 PRF__L ( LB ) 43 PRF_R (LB ) 44 PRF_NET (LB ) 173 48 0 49 Ni_L (%) 50 NI__R (M) 5i PROP._RPM_.L (Y.)

52 PROP_RPM_.R (_) 52 53 T ORQUE_L (PSI) 54 TORQUE_R (PSI) 55 FUEL_FLO__L (LB/HR) A.3 TABLE A.I, CONT'D FUEL_FLO R (LB/HR) 52 FUELTEMP_L (DEG_K) 58 FUELTEMP_R (DEG_K) 64 JOHNS_WILL (VOLT) LEIGH (VOLT) ROSEMOUNT (VOLT) 67 GEN_EAST (VOLT) BATTERY__B (VOLT) 72 PAUSE_EVNT (0 : NO ) 73 PIRAM_HEAT (0 : OFF ) 24 ROSEM_HEAT (0 = OFF ) P5 76 GEAR_POS (0 = UP ) GYRO_ERECT (0 = YES ) 78 PILOT_EVNT (0 : NO ) 81 P_CORRECTD (DEG/SEC) 82 Q_CORRECTD (DEG/SEC) 83 R_CORRECTD (DEG/SEC) 84 ROLL__ACCEL (DEG/S/S) 85 PITCH_ACCL (DEG/S/S) A.4 TABLE A.I, CONT'D 86 YAW__ACCEL (DEG/S/S ) 86 87 AX_CORRCTD (G) 87 88 AY_CORRCTD (G ) 89 AZ_CORRCTD (G) 90 EULER__ROLL (DEG) 91 EULR_PITCH (DEG) 92 EULER_YAW °" (DEG) E ULER_P ( DEG/SEC ) E ULER__Q (DEG/SEC) 94 E ULER__R (DEG/SEC) 95 ALT_INDCID (FEET) AIRSPD_IND (KNOT) 97 MACH_IND 98 DYN_PRS_IN (PSF) 99 CLIC 101 NCLIC P OS_COR_P (PSF) 102 TOTL_PRESS (PSF) AMB_PRESS (PSF ) 104 QC (PSF) PRES_RATIO 10& PRES_RAT_T 107 ALT_CAL ( FEET ) 188 AIRSPD_CAL (KNOT) 109 MACH_CAL 110 DYN._PRS_CA (PSF) 111 112 CLT 113 NCLT P OS_COR_V (KNOT ) 114 P OS_COR_H (FEET ) 115 A.5 TABLE A.I, CONT'D li6 POS_COR__ 117 A I RSPD__EQ (KNOT) TEMP__AMB (DEGK ) 119 TEMP_TOTAL (DEGK) THETA i21 12i THETA_TOTL !22 AIRSPD_TRU (KNOT) DENSITY (SLG/CUFT) DENS_RATIO ALT_DENS (FEET) TTL_PRS_ST (PSF) 127 AMB_PRS_ST (PSF) 128 QC_STD (PSF) PRS_RAT_ST PS RT T ST PROP_THR_ (LB) PROP_THR_R (LB) i33 i33 FGJ_L (LB) FGJ_R (LB) WA_L (LB/SEC) WA_R (LB/EEC) i32 RAM_DRAG_L (LB) 138 RAM__DRAG_R (LB) THR_COEF__L THR_COEF__R CT_CALC_L CT_CALC__R PWR_COEF_L 144 PWR_COEF_R 145 145 ADV_RAT_L A.6

CONT ' D

ADV_RAT_R 146 (DEG ) PRP_BETA_L (DEG ) PRP_BETA_R 148 (KNOT ) 149 D E L_VE (DEGK ) DEL_T._AHB 158 (KNOT ) DEL_VT 151 DEL_DENS (SLG/CUFT) 152 (DEG ALPHA_CAL I 153 @ (DEG BETA_CAL 156 ALPHA_TRUI (DEG ) 157 BETA_TRUE (DEG ) NLOAD (G ) 161 CG_LONG ( FEET ) CG_LATERAL (FEET) CG_VERT (FEET ) 164 I XX_BODY (SLG-SQFT) I YY_BODY (SLG-SQFT ) I ZZ_BODY (SLG-SQFT) IXZ.BODY (SLG-SQFT ) CG_PCT_tt_C (Y.)

FUEL_USED (LB ) AC_WE I GHT (LB) FUEL_.WT (LB ) 172 ALPHA_ROSI (DEG) 174 ALPHA_ROS2 (DEG) 175 A.?

TABLE A.I, CONT'D 176 BETA_ROSI (DEG) BETA_.ROS2 (DEG) TEMP 179 DPAOQ 180 ATAN_AXAZ (DEG) 181 TORQUE P L (PSI) TORQUE P R (PSI) TEMP_AIR ( ( ) _90 g A.8

TABLE A. 2

Cadmus Floatinq Pointing Format

Real*4:

s eeee eeee mmmm mmmm mmmm mmmm mmmm mrnm

where s = sign (O=pos, l=neg) e = exponent (unsigned binary integer, excess 127 code m = mantissa (binary fraction, normalized with leftmc, st bit implied) Example: exp-127 +/-2 * !._antissa A. 9 Yw THIS PAGF INTENTIONALIY LEFT BLANK A.IO Government Accession No. 3. Recipient's Catalog No.

1. Report No. 2. NASA CR-179515 4. Title and Subtitle 5. Report Date October 1986 Flight Test Report of the NASA Icing Research Alr- Plane. Performance, Stability, and Control After 6. Performing Organization Code Flight Through Natural Icing Conditions 8. Performing Organization Report No.

7. Author(s) KSR 86-0l Jerry L. Jordan, Stewart O. Platz, and William C. Schlnstock 10. Work Unit No.

505-68-II 9. Performing Organization Name and Address _i-ContntractorGrantNo.

Kohlman Systems Research, Inc.

NAS3-24547 319 Perry Street Lawrence, Kansas 66044 13. Type of Report and Period Covered Contractor Report 12 Sponsoring Agency Name and Address Final National Aeronautics and Space Administration 14. Sponsoring Agency Code Lewis Research Center Cleveland, Ohlo 44135 15. Supplementa_ Notes Project Manager, Kevln L. Mlkkelsen, Propulsion Systems Division, NASA Lewis Research Center.

16. Abstract Flight test results are presented documenting the effect of airframe icing on performance and stability and control of a NASA DHC-6 icing research alrcraft.

Kohlman Systems Research, Inc., provided the data acquisition system and data analysis under contract to NASA. Performance modeling methods and MMLE Tech- niques were used to determine the effects of natural ice on the aircraft.

Results showed that ice had a significant effect on the drag coefficient of the aircraft and a modest effect on the MMLE derived longitudinal stability coefft- clents (code version "MMLE"). Data is also presented on asymmetric power si,{_'- slip maneuvers showing rudder floating characterlstics w_th and wittlout ;_ .,, the vertical stabilizer.

17. Key Words (Suggested by Author(s)) Unclassified - unllmlted Aircraft icing STAR Category 08 Aircraft performance Stability and control 18. Distribution Statement 20. Security Classif. (of this page) 21. No. of pages 19. Security Classif. (of this report) Unclassified 159 A08 Unclassified 22. Price* *For sate by the National Technical Information Service, Springfield, Virginia 22161