Design, fabrication, and static testing of Attached Inflatable Decelerator /AID/ models

Section

Title Page Section H a r d s t r u c t u r e . . . . . . . . . . . . . . . .

LABORATORY TESTING . . . . . . . . . . . . 40

IX General . . . . . . . . . . . . . . . . . . . .

Textile Tensile T e s t s . . . . . . . . . . . . . . 40

F a b r i c Permeability T e s t s . . . . . . . . . . . 41

AeroshellInlet Model Evaluation T e s t s . . . . . .

Deployment and Inflation Test . . . . . . . . . .

X CONCLUSIONS AND RECOMMENDATIONS . . . .

ADDendix 4 8 A DETAILED DESIGN DRAWINGS . . . . . . . . .

5 8 B STRESS ANALYSIS . . . . . . . . . . . . . .

6 0 C DETAILED INFLATION ANALYSIS . . . . . . . .

D AEROSHELL INLET ANALYSIS. . . . . . . . . . 65

7 6 REFERENCES . . . . . . . . . . . . . . . . . . . . . . .

ILLUSTRATIONS Figure Title Page 1 Drag Performance of AID and Disk-Gap-Band Para- chute . . . . . . . . . . . . . . . . . . . . .

2 Variations in Decelerator Mass and Landed-Payload M a s s with Ballistic Coefficient. . . . . . . . . .

Details of -101 Assembly . . . . . . . . . . . . 6

Details of - 103 Assembly . . . . . . . . . . . . .

Details of -105 Assembly . . . . . . . . . . . . .

Details of AID Deployed by Water-Alcohol Solution . 8

Details of Ram-Air Deployed AID . . . . . . . . 8

Cutaway Drawing of a n AID . . . . . . . . . . . . 10

Comparison of Experimental and Analytical P r e s s u r e

Uistributions for Curr?-.lt Canopy Shape (Moo = 3.0) 12

Correlation of Experimental and Predicted Inflation Analysis Results . . . . . . . . . . . . . . . . .

11 15 Predicted Inflation Rate for Specified Inlet Sizes . . .

12 Reference A r e a s of A D . . . . . . . . . . . . . .

Aeroshell Inlet Geometry . . . . . . . . . . . . . 16

14 Thicknesses Associated with Boundary Layer . . . .

15 Aeroshell Inlets P r i o r to Deployment . . . . . . . 19

Schematic of Aeroshell Inlet Restraining System . . . 20

Aeroshell Inlets Deployed (External View) . . . . . 21

18 Aeroshell Inlets Deployed (Internal View) . . . . . .

Relative Angular Displacement of Inlets (Looking Downstream) . . . . . . . . . . . . . . . . . . .

Canopy Forward Attachment Scheme for -101 and -105 Models . . . . . . . . . . . . . . . . . . . . .

Displacement be tween Canopy and Ae r oshell Periphery 24

- vii -

Figure Page 2 2 Rear Attachment Scheme . . . . . . . . . . . . .

2 5 23 Model Gore Patterns . . . . . . . . . . . . . . .

2 4 Initial Stowage and Release Scheme . . . . . . . . . 30

Canopy Stowage and Release Scheme . . . . . . . . 32

26 Folding Sequence . . . . . . . . . . . . . . . . . 33

27 AID Model in Packaged Mode . . . . . . . . . . . 34

Load Cell Attachment To Meridians Profile of A D Showing Meridional Tape Load Cell Locations . . . . . . . . . . . . . . . . . . . .

Serial Numbers and Locations of Load Cells for Each

Assembly . . . . . . . . . . . . . . . . . . . . 36

Water-Alcohol Inflation T e s t Setup . . . . . . . . . 44

AID Differential P r e s s u r e - T i m e History . . . . . 45

Load Cell Force-Time History. . . . . . . . . . .

Fully Inflated AID. . . . . . . . . . . . . . . . . 46

D - 1 F o r c e s Acting on Aeroshell Inlet . . . . . . . . .

D-2 Change in Direction of Flow Entering Aeroshell Inlet D - 3 F o r c e s Acting on Inlet Release Spring . . . . . . .

D - 4 Forces Acting on Release Lever . . . . . . . . .

D - 5 Release Spring Geometry before and after Release . .

7 8 -viii- TABLES Table Page

I AID Design Conditions . . . . . . . . . . . . . . 9

1 1 AID Coordinates . . . . . . . . . . . . . . . . . 27

Lobe Geometry, Radial Growth Due to Unrestrained 1 1 1

T h r e a d Racking and Fabric S t r e s s e s . . . . . . . . 28

IV Index to Attached Inflatable Decelerator Drawings . . 37

V Design F a c t o r s . . . . . . . . . . . . . . . . . 38

V I Tensile Strength Test Results of Canopy F a b r i c and

Meridian Tapes . , . . . . . . . . . . . . . . . 41

VI1 Tensile Strength Test Results of Dacron Canopy Seams 4 2 VI11 Tensile Strength Test Results of Load Cell Attach- ments and Canopy Forward Attachments. . . . . . .

4 2 F a b r i c Permeability T e s t Results . . . . . . . . .

IX 4 3

D -I Release Spring Input F a c t o r s . . . . . . . . . . . 7 2

-ix- DESIGN, FABRICATION, AND STATIC TESTING OF

- ATTACHED - INFLATABLE - DECELERATOR (AID) MODELS

By G. L. Faurote Goodyear Aero space Corpora tion SUMMARY Goodyear Aerospace Corporation (GAC) designed and fabricated wind-tunnel models of a n Attached Inflatable Decelerator (AID) for the NASA Langley Research Center (LRCF An AID-is a flexible canopy, deployed and inflated by ram air, which is attached to the base of the body to be decelerated. The geometry of the canopy can be tailored to the application for which i t is in- tended and to the available attachment locations.

T h r e e models were furnished by Goodyear Aerospace for the c u r r e n t pro- gram: two AID canopies were attached to a simulated payload within a 120- deg conical a e r o s h e l l while the third canopy had a n outer attachment to a 140-deg conical aeroshell and a n inner attachment to a simulated payload.

All models utilized four quick-release a e r o s h e l l inlets for deployment and four canopy inlets to produce final canopy pressurization.

An inflation test was conducted i n a n environmental chamber to investigate packing, deployment, and inflation characteristics of the model p r i o r to wind- tunnel te sting .

I. - INTRODUC TION

The need for a deceleration system with good operational c h a r a c t e r i s t i c s a t supersonic speeds h a s led to the development of a n Attached Inflatable De- c e l e r a t o r (AID). An AID is a flexible canopy normally deployed and inflated by r a m a i r , which is attached to the b a s e of the body decelerated.

A s u m m a r y of the extensive analytical and experimental development of the AID to date is presented in References 1 through 8. The encouraging results of these development efforts have prompted planetary-mission studies (Refer- ences 1 and 9) in which the AID is considered for application as an initial stage of a two- stage deceleration system. Basically, this two- stage approach takes advantage of the high supersonic drag of the AID and the high subsonic drag of a t e r m i n a l stage parachute such a s the disk-gap-band parachute ( s e e Figure 1).

Reference 1 indicates that for entry into the low density atmosphere of M a r s , the AID will r e l a x the stringent deployment conditions on the parachute and will provide significant increases in landed-payload m a s s without increasing To compliment the work the size of the basic entry aeroshell (see F i g u r e 2 ) .

accomplished in Reference 1, Goodyear Aerospace is performing a parametric t h e r m a l and s t r e s s analysis of an AID operating in the most s e v e r e of the The r e s u l t s of this work a r e presently unpub- postulated M a r s atmospheres.

lis hed.

-1- 1.5 1 .o v I- z 0 . 5

u

IL w

I

N O T E : U

C U R V E R E P R O D U C E D FROM R E F E R E N C E 1 I

a 0 1 2 3 4 5 M A C H N U M B E R , Moo

Figure 1. - Drag Performance of AID and Disk-Gap-Band Parachute

N O T E S : 1 , S O L I D L I N E R E P R E S E N T S M A C H N U M B E R D E S I G N C O N S T R A I N T TO ~ E N S U R E T H A T A E R O D Y N A M I C N O C O N S T R A I N T H E A T I N G P R O T E C T I O N I S N O T

I ' I O N M A C H N U M B E R

R E Q U l R E D 2. C U R V E I S R E P R O D U C E D F R O M P A R A C H U T E R E F E R E N C E 1 .

AND A I D MASS 0.6 0.3 0 . 4 0 . 5 E N T R Y B A L L I S T I C C O E F F I C I E N T , B E (SLUGS P E R SQUARE F O O T )

Figure 2. - Variations in Decelerator Mass and Landed-Payload Mass

with Ballistic Coefficient -2- The c u r r e n t investigation is part of a continuing effort by LRC to develop and evaluate the AID. Since the AID i s aerodynamically shaped, its canopy may be tailored to the application for which i t is intended and to the available attachment locations. The detailed design presented i n Reference 3 is for a n AID attached to the periphery of a 120-deg conical aeroshell with inflation initiated by a water vaporization system. Utilizing the s a m e aerodynamic shape, the inflation system was subsequently replaced with r a m air inlets located a t the aeroshell periphery (see Reference 1). Key features of the design presented herein a r e the canopy attachment to a simulated payload in- small r a m - a i r inlets as an integral t e r i o r to the periphery of the aeroshell and Thus, the objectives of the program a r e the design, part of the aeroshell.

AID models in preparation for wind tunnel fabrication, and static testing of (1) demonstrate ram-air deployment of the testing to evaluate the following: AID by quick-opening aeroshell inlets; (2) investigate the supersonic behavior of the canopy in the vicinity of the aeroshell rim where the canopy is unre- strained; ( 3 ) determine the effect of increased inflation time on the deployment integrity of the AID; and (4) measure the transient loads in the meridional tapes of the AID, These objectives a r e discussed in detail a t various points through- out the report.

SYMBOLS A a r e a , sq in.

B ballis tic coefficient, slugs/sq ft drag coefficient based on maximum projected diameter cD C

p r e s s u r e coefficient 9 P L - Poo/9a3

P D diameter, in.; drag, Z b depth of lobe, in.

dg E Young's modulus, p s i circumferential membrane s t r e s s coefficient; force, Zb F factor of safety F. S .

tensile strength, Ib; Zb/in.

Ft f fabric s t r e s s , lb/in.

-

f nondimensional fabric s t r e s s circumferential fabric s t r e s s resultant, lb/in.

f C meridional fabric s t r e s s resultant, lb/in.

fm H compressible form factor h inlet height, in.

-3- k constant dependent upon m a t e r i a l properties 1 length, in.

M Mach number margin of safety M. S .

m m a s s flow, slugs/cu f t N s t r e s s , lb/in.

n number of gores

pi - pb, psf; static p r e s s u r e , psf

P total p r e s s u r e a c r o s s bow shock, psf P t2 dynamic p r e s s u r e , psf q maximum canopy radius excluding burble fence, in. ; gas R constant, ft-lb/lb-deg R r aeroshell base radius, in.

r lobe radius, in.

g S surface distance f r o m apex to base of aeroshell half-arc length of lobe, in.

sg T tension, lb

T nondimensional meridional tape load

U f r e e - s t r e a m velocity, fps V velocity, fps aeroshell volume, cu ft ' a burble fence volume, cu ft vb canopy volume, cu ft VC total enclosed volume, cu f t Ve X inlet width, in.

central half-lobe angle, deg; constant dependent upon

P

mate rial proper ties bias thread s e t angle, deg, ratio of specific heats for air Y elongation, in.

A boundary layer thickness, i n . ; elongation, in. /in.

d displacement thickness, in.

hhk s t r a i n due to thread racking, in. /in.

E e local surface angle relative to direction of free-stream flow, deg; momentum thickness, ft; circumferential direction porosity

x

D Poisson's ratio (T s t r e s s , psi meridional direction

B

Superscript k constant dependent on material properties Subscripts a a e rodynamic ; a e r o shell b direction of bias threads; base; bending br bearing C canopy, closed E entry e external f fabric; final; forward fm maximum friction force, lb internal; initial, inlet i is inlet spring local m meridian, maximum 0 stagnation; nominal porosity, p r e s s u r e , principal P r r e a r rl release lever S surface, s h e a r , spring t to tal U ultimate inlet efficiency Q

e circumferential direction of surface element

meridional direction of surface element

B

00 free s t r e a m -5-

11. - AID MODEL DESCRIPTION

The configuration details of each of the three AID models discussed i n this report a r e shown i n Figures 3 through 5. Additional design details of the models described in Figures 3 through 5 a r e presented i n assembly Drawings 645A000-003-101, -103, and -105, respectively. These assembly numbers will be used in any future reference made to specific details of the models.

A complete s e t of drawings for the models is presented i n Appendix A.

The models shown in Figures 3 through 5 are similar to those previously de- signed and fabricated by Goodyear Aerospace for supersonic wind-tunnel evaluation by LRC. However, important differences that relate to the objec- To point out these differences, the con- tives of this investigation do exist.

figurational details of the models previously tested a r e presented in Figures 6 and 7; additional details of these models a r e presented in References 3 and 1, respectively. The model shown in Figure 6 was deployed by rapid vaporiza- tion of a water-alcohol solution. Vaporization of this solution provided suffi- cient internal p r e s s u r e to e r e c t the ram-air inlets, which then fullyinflate the model. The model shown in Figure 7 was deployed and inflated entirely by r a m a i r .

The forward inlets, packaged behind the aeroshell p r i o r to deploy- ment, a r e released a t deployment and create sufficient internal p r e s s u r e to A model also e r e c t the rear canopy inlets, which then produce full inflation.

was tested successfully that used only the forward inlets.

'I / I / I f

i-

' 4 n-4, I I fi*' LOCATIONS (3 PLACES)

F\ I N L E T

C A N O P Y (4 P L A C E S )

' A T T A C H E D

I T O S I M U L A TE I P A Y L O A D R E S T R A I N I N G I I C O R D S I N F L A T A B L E D E C E L E R A T O R C A N O P Y ( N O M E X F A B R I C ) U N R E S T R A I N E D A T A E R ~ S H E L L A E R O S H E L L INLET (4 P L A C E S ) E D G E E LL ROSH -R = 30.0 I N . ( N O M I N A L )

Figure 3. - Details of -101 Assembly

-6- I B U R B L E F E N C E L O A D C E L L L O C A T I O N S A P I R A M - A I R R I N G I (3 P L A C E S ) Y I N L E T

A ’ (4 P L A C E S )

I I I I

F ‘ I N L E T

R E S T R A I N I N G C O R D S

-----

C A N O P Y /. A T T A C H E D

I N F L A T A B L E F O R W A R D D E C E L E R A T O R A T T A C H M E N T ( N O M E X F A B R I C ) R I N G A E R O S H E L L I N L E T (4 P L A C E S ) A E R O S H E LL IN.

- A - (N OM I NA L) -R = 30.00 IN. ( N O M I N A L )

Figure 4. - Details of -103 Assembly

S U P P O R T / /

\ S L E E , V E 4 - -1 /

A F T R A M - A I R I N L E T

K (4 P L A C E S )

INLET R E S T R A I N I N G C O R D S A T T A C H E D I N F L A T A B L E U N R E S T R A I N E D D E C E L E R A T O R A T A E R O S H E L L ( D A C R O N F A B R I C ) E D G E O S H E L L I N L E T (4 P L A C E S ) I N . - , A E R O S H E L L I

Figure 5. - Details of -105 Assembly

-7- B U R B L E F E N C E S U P P O R T / # / S L E E V E / --1 / A P T R A M - A I R I f I N L E T I 1 (4 PLACES)

A

I $ C A N O P Y R E A R

A T T A C H M E N T

! I R I N G

A T T A C H E D D E C E L E R A T O R ( N O M E X F A B R I C ) F O R W A R D A T T A C H M E N T R I N G N O T E S : 1. C O M P L . E T E D E S I G N D E T A I L S P R E S E :NTED I N R E F E R E N C E 3

- R = 30.0 IN. ( N O M I N A L )

2. WIND-1 rUNNEL T E S T R E S U L T S "-re- r n c a c N T E D IN R E F E R E N C E 5

F i g u r e 6. - Details of AID Deployed by Water-Alcohol Solution

B U R B L E F E N C E I- \ F O R W A R D C A N O P Y A T T A C H M E N T I N L E T C O R D S I . A T T A C H E D I N F L A T A B L E D E C E L E R A T O R (NOMEX FABRIC) / F O R W A R D R A M - A I R I N L E T (4 P L A C E S ) N O T E S : -0 1. A D D I T I O N A L D E T A I L S O F T H E S E I = 12.0 IN.

E R O S H E L L M O D E L S A R E P R E S E N T E D IN f N O M l N A l I . - _._, R E F E R E N C E 1.

C-- R = 30.0 IN. ( N O M I N A L ) 2. W I N D - T U N N E L R E S U L T S A R E P R E - S E N T E D IN R E F E R E N C E 6 .

Figure 7. - Details of R a m - A i r Deployed AID

- 8 - Id pro- The use of r a m a i r to depl y the AID, r a t h e r than water-alcohol, wa vide a weight savings, which is of obvious importance in the design of a flight unit. In addition, the ram-air deployment technique eliminates the potential problem of sterilization associated with using a liquid-vaporization s y s t e m for planetary mission applications.

It may be advantageous in c e r t a i n instances to have available an alternative r a m - a i r deployment scheme to that described above. Therefore, one purpose of this investigation was t o develop a s y s t e m whereby the AID canopies a r e deployed by four symmetrically located aeroshell inlets. A s in the models previously tested, the canopy inlets a r e retained to provide full inflation and final pressurization.

III. - DESIGN CONDITIONS

The aerodynamic and s t r u c t u r a l design conditions for the AID canopies a r e presented in Table I. The h a r d structure was designed to m e e t the require- m e n t s specified in Arnold Engineering Development Center (AEDC) Document QCP-000-21.

TABLE I. - AID DESIGN CONDITIONS

Condition Magnitude 2.5 ft Maximum inflated radius, R P Canopy attachment F o r w a r d x / R = 0. 1538; y/R = -0.745 x/R = 0. 1538; y/R = 0.291 Aft

-

0.44 Nondimensional meridional tape load, T Nondimensional fabric, f 0.09 Number of gores, n Ratio of circumferential fabric s t r e s s e s

< <

0.8 = fc/fm = 2.0 to meridional s t r e s s e s , fc/fm 3 . 0 F r e e - s t r e a m Mach number, Ma, 120 psf F r e e - s t r e a m dynamic p r e s s u r e , 9 , - 9 -

IV. - AERODYNAMIC ANALYSIS

General The AID, which is constructed by overlaying a coated fabric with many load- carrying meridional tapes (see Figure 8), is basically a uniform-stress (isotensoid) structure a s described i n Reference 4.

The canopy shape is maintained by p r e s s u r e recovered f r o m ram-air inlets aligned with the local airflow. F o r subsonic speeds, a burble fence on the AID is necessary to provide stable operation.

Shape Analysis The entire isotensoid analysis as applied to the AID was computerized during a previous program (Reference 3) conducted by Goodyear Aerospace for LRC.

That analysis utilizes a theoretical aerodynamic p r e s s u r e distribution to cal- culate a n initial shape. The initial shape then can be used to obtain a n experi- mental p r e s s u r e distribution which may be inserted in the computerized analy- s i s to obtain an "iterated" shape.

This procedure can be repeated until the desired correlation i s obtained.

The p r e s s u r e distribution and shape (profile) for the c u r r e n t AID models w e r e derived in Reference 3 .

The derivation of the p r e s s u r e distribution used

Vl A E R O S H E L L

Figure 8. - Cutaway Drawing of a n AID

- 10-

I

modified Newtonian aerodynamics over the front surface of the AID and a base p r e s s u r e equivalent to the f r e e - s t r e a m static p r e s s u r e .

F o r the front surface, modified Newtonian aerodynamics relate the p r e s s u r e coefficient, cp, to the stagnation p r e s s u r e coefficient, cpmax by the following relation: -3 L c = c sin 8 , P Pmax where

-

C Pmax - (Po - P,)/9, 9

and 8 = local surface angle to free-stream flow.

I

The internal canopy p r e s s u r e was obtained f r o m the empirical relation: = 2.0

I

The p r e s s u r e distribution derived in Reference 3 and an experimental distri- bution subsequently obtained by LRC (Reference 11) on a rigid wind-tunnel

I p r e s s u r e model i s presented i n Figure 9. As can b e seen, the calculated

and m e a s u r e d p r e s s u r e coefficients a r e in fair agreement over the primary A ID profile.

I

Inflation Analysis AID models previously tested i n the wind tunnel were designed ( s e e Refer- e n c e s 1 and 3) for rapid inflation ( t A s the size of a n AID in- s 0 . 4 sec).

f for instance to a 20-ft-diameter flight test model, the feasibility and c r e a s e s , While the I advisability of inflating the AID within 0 . 4 s e c is questionable.

the inlet a r e a volume of the AID i n c r e a s e s a s a function of its radius cubed, Therefore, to main- i n c r e a s e s only as a function of the inlet radius squared.

tain the same AID volume-to-inlet a r e a ratio, which is essential i f the infla- tion times a r e to be the s a m e for a given s e t of f r e e - s t r e a m conditions, the Additionally, l a r g e local fabric velocities inlet sizes begin to get cumbersome.

(possibly g r e a t enough to cause canopy failure) could occur during the inflation

I

of a 20-ft-diameter AID within 0.4 sec. However, excessively long inflation t i m e s have the apparent disadvantage of permitting the canopy to flag in To investigate the effects the a i r s t r e a m , possibly resulting i n canopy failure.

of longer inflation times on the AID canopy, the r a m - a i r inlets w e r e sized to

I

extend the inflation time in excess of one second.

I

-11-

I

-

2.2 -0.4 -0.3

2.0 --

-0.2

I .e --

I .6 -- -0.1

1.4 -- 0.0

1.2 --

t 0.1

I .o --

0.2

0 . 8 -- 0.3

0.4

0.6 --

P c 0 . 5 0 . 4 - - - V - L h 0.6 w 0.2-.

V W

5 0.0-- 0 . 7

VI VI W a 0 . -0.2-- 2 0 . 8

- - - MODIFIED N E W T O N I A N T H E O R Y

- - - E M P I R I C A L R E L A T I O N

c I I I 0 0.1 0.2 0.3 0.4 0.5 0 . 6 0 . 7 0.8 0.9 1.0 1 . 1 x/ R

Figure 9. - Comparison of Experimental and Analytical Pressure

Distributions for Current Canopy Shape (Moo = 3.0) -12-

I

i

A computerized inflation analysis ( s e e Appendix C ) , similar to the method

I

described in Reference 1, was used to determine inlet size and inflation rate.

T o establish the validity of the method, the inflation times of three previously tested AID wind-tunnel models (Reference 6) w e r e calculated. The correlation

I

obtained f o r those models is presented in F i g u r e 10.

F i g u r e 10B shows the comparison for the AID with dual inlets. Using the analysis of Appendix C the following values of volume-to-inlet a r e a ratios

I

w e r e chosen for the present models.

1. At deployment (aeroshell inlets only),

I

Ve/Af = 8600 in.

I

After deployment (aeroshell and canopy inlets), 2 , Ve/(Af t A r ) = 2860 in.

I

F o r these restrictions on volume-to-inlet a r e a ratio the resulting predicted inflation time is 1. 5 sec a s shown in Figure 11.

I

T o solve for the actual inlet dimensions, the total enclosed volume ( s e e Fig- The canopy volume V is determined f r o m the u r e 12) must be determined.

.c I isotensoid computer analysis described in the previous section.

The total enclosed volume then is:

I

= vc t Vb t va:k

v

e = 23. 95 t 1 . 4 t 0 . 3

I

= 25.65 cu ft ,

I

*

The volume difference between the 120- and 140-deg aeroshell is s m a l l and therefore, only the one value presented was considered.

I

-13- TIME (SECONDS) TIME ISECONDS) ( A I 181 zoo - + U W 4 150 W NOTES v) 1 . E X P E R I M N T A L DATA WERE TAKEN FROM REFERENCE 6.

z loo - 2. SOLID LINE REPRESENTS THEORETICAL W PREDICTION VI E M J z LL W I - z 0 0 5 I O 1.5 2 0 TIME (SECONDS) IC)

Figure 10. - Correlation of Experimental and Predicted

Inflation Analysis Results

- 14-

I

i

I

0.5 1 .o 1.5

i

Figure 11 . - Predicted Inflation Rate for Specified Inlet Sizes

I

u

I "b

I

Figure 12. - Reference Areas of AID

I

-15-

I

F r o m Equation 1

I

*f = 3mtEc

I

= 5. 15 sq in.

I

The required effective area f o r each of the four aeroshell inlets then becomes A f / 4 of 1.285 sq in. F r o m Equation 2

I

V,/(Af t Ar) = 2860 in.

I

Therefore, 2860 Ar = (25.6)(12) - (2860)(5. 15) ,

I

o r A = 10.3 s q i n .

r The requiredopen area for each of the four canopy inlets then becomes A+/4 of 2. 56 sq in.

I

A s shown i n Figure 13, the aeroshell inlets are submerged partially in the

I

R E L E A S E ROD,

I

A E R O S H E L L A E R O S H E L L

I

Figure 13. - Aeroshell Inlet Geometry

I

- 16-

i

I

forebody boundary layer flow. . A s a result the bqundary layer effects

I

m u s t ’be considered in sizing the inlets. In Figure 14, two general types of thicknesses associated with a typical boundary layer a r e depicted. By defi- nition the displacement thickness, 8*, defines the distance through which the body streamline is shifted due to the formation of the boundary layer.

Since

I

the displacement thickness describes a region of zero velocity and, therefore, zero m a s s flow, the height of the inlets ( s e e Figure 13) m u s t be increased by&*.

I

F o r the given wind-tunnel test conditions and based upon an aeroshell diame- ter of two feet, the Reynolds number for the upper one-third of the aeroshell

I

surface is found to be on the order of 10 . Therefore, a turbulent boundary

layer analysis was reformed by using the momentum integral solution of Ref- erence 12. A computer p r o g r a m from Reference 13 was utilized in the solution.

The p r o g r a m yielded the following values for the boundary layer momentum thickness, 8:

I D

1. F o r the 120-deg cone: 0 = 0. 14315 X lo-’ ft ’ 2. F o r the 140-deg cone: 8 = 0.15182 X 10” ft

I

F r o m R e f e r e n c e 1 2 a compressible f o r m factor, H, is given by:

I

Figure 14. - Thicknesses Associated with Boundary Layer

I

-17-

I

i

I

F o r Ma, = 3.0, H = 5.4.

By using the results above for the momentum thick- ne s s, the displacement thicknes s is:

I

= 0.0929 in. (120-deg cone), and 6 * = 0.0977 in. (140-deg cone).

The l a r g e r of these two values was utilized in arriving a t the inlet height. Ad-

I

ditionally, the diameter of the inlet release rod was accounted for in determin- ing the inlet height. F r o m Figure 13 the inlet height is:

I

= 1.34 in., I

where x = 1. 15 in. was chosen as being convenient f r o m a fabrication stand- point .

I

The rear inlets of the AID models shown in Figures 6 and 7 w e r e displaced approximately 0.25 in. from the canopy surface to minimize boundary layer

I

effects. Based on the effective performance of those inlets during the wind- tunnel tests, the r e a r inlets of the c u r r e n t models w e r e displaced a similar distance from the canopy. F r o m the a r e a of each r e a r inlet, which was p r e - viously established a s 2. 56 sq i n . , the diameter is:

I

= 1. 80 in.

I

V. - SYSTEM DESIGN

Aeroshell Inlets

I

The aeroshell inlets a r e forced into the a i r s t r e a m by a sliding flat spring.

When in the extended position, the inlet spring d i r e c t s the a i r flow into the canopy. While other types of springs also w e r e considered for deploying the inlets, the sliding flat spring was chosen p r i m a r i l y because of i t s compactness.

Figure 15 is a photograph of the internal surface of the a e r o s h e l l showing the

I

inlets p r i o r to deployment. The inlets are r e s t r a i n e d in the closed position As shown in Figure 15 the four r e l e a s e l e v e r s are con- by the release lever.

The restraining s y s t e m shown nected mechanically by the restraining wire.

I

- 18-

I

-19-

I

Figure 16 shows the inlet i n Figure 15 is shown schematically in Figure 16.

I

When the electrical pyro- release spring, which is not visible in Figure 15.

technic cutter is activated, the restraining wire is cut, permitting the inlet r e l e a s e spring to rotate the release levers. When the tabs rotate to the dashed position shown i n Figure 15, the inlets a r e deployed into the a i r s t r e a m .

I

Figure 17 i s a photograph of the aeroshell external surface showing the inlets as determined in Section I V , in the deployed condition. The proper inlet height,

I

is achieved when the inlet stops contact the internal surface of the aeroshell as shown in Figure 18. A positive spring force is maintained when the inlet is in the open position to prevent the possibility of inlet flutter.

I

Calculations a r e presented in Appendix D for the (1) aerodynamic forces that a n inlet spring must overcome to deploy an inlet, ( 2 ) frictional f o r c e s that a n inlet release spring m u s t overcome to release a n inlet, and ( 3 ) aerody-

I

namic forces acting on the aeroshell inlets in the open position.

Canopy Inlets The canopy inlets, which a r e identical for all three models, are of the type used on the models of Figures 6 and 7. As determined in Section IV, the

I

canopy inlets have a 1. 8-in. diameter. Figure 19 shows the relative angular displacement of the canopy and aeroshell inlets. While the forward and rear inlets a r e not a t 45 deg relative to each other, this placement should lead to satisfactory performance. If the two s e t s of inlets had been placed a t 45 deg

I

, A E R O S H E L L I N L E T

I

T E N S I O N I N SPRING 9

I N L E T ( I N L E T S C L O S E D ) R E L E A S E

I

ROD R E L E A S E L E V E R

\ P I V O T P O I N T

( I N L E T S C L O S E D )

I n . _

I

U U

\

- \- - /u .

’ SPRING

R E S T R A I N I N G WIRE

/

R E L A X E D

I

( I N L E T S L I N E C U T T E R O P E N ) N O T E : O N L Y TWO I N L E T S SHOWN

I

Figure 16. - Schematic of Aeroshell Inlet Restraining System

I

-20-

I

Figure 17. - Aeroshell Inlets Deployed (External View)

I

relative to each other, two meridional tapes would have been severed where 19).

the canopy attaches to the aeroshell inlets ( s e e Figure This latter place-

I

ment then would be unacceptable from a s t r e s s standpoint.

Canopy Attachment Scheme

I

The forward attachment of the canopy to the h a r d structure of the -101 and -105 assemblies is shown in Figure 20. The canopy is clamped to the a e r o - shell inlet tubes a s shown. Other schemes, such as a flexible fabric connec-

I

tion between the canopy and the inlet tubes, a l s o were considered. However, the approach shown was adopted because of dimensional limitations and sim- A s shown in Figure 20, the canopy will remain unrestrained in the plicity.

I

vicinity of the aeroshell rim.

F o r w a r d attachment of the canopy for the -103 assembly was continuous a t the aeroshell periphery, similar to the attachment of the models shown in

I

F i g u r e s 6 and 7. Attachment to the simulated payload would have resulted in a n undesirable displacement between the canopy and aeroshell periphery ( s e e Figure 21) for the current canopy profile.

I

The r e a r attachment for all three models was made to the simulated payload as shown in Figure 22.

I

-21-

I

I~

i

I

SUPPORT

SLEEVE I

-

I

F 'Y

I

T INLET STOP I N CONTACT W I T H

I

AEROSHELL

I

FORWARO CANOPY A T T AC UMENT CANOPY ATTACHMENT RING TO INLET ( 4 PLACES)

I

oyed ( I n t e r n a l View)

F i g u r e 18. - A e r o s h e l l I n l e t s Dep

I

-22-

I

A F T RAM-AIR I N L E T G O R E NUMBER

I

A L )

I

F i g u r e 19. - Relative Angular Displacement of Inlets (Looking Downstream)

I

- 2 3 -

I

Figure 20. - Canopy Forward Attachment Scheme for - 10 1 and - 105 Models

m

I

0 1 S P L A C EM E N T

I

B E T W E E N C A N O P Y A N 0 A E R O S H E L L \ / 7 0 OEG

I

Figure 21. - Displacement between Canopy and Aeroshell P e r i p h e r y

I

-24-

I

X / R = 0.1538 Y / R = 0.291 C L A M P

Figure 22. - Rear Attachment Scheme

-25-

I

Canopy Gore Pattern

I

While the profile of the current models, presented in Table II, is the s a m e a s the previous wind tunnel models ( s e e Figures 6 and 7 ) , i t was necessary to modify the previous gore patterns to account for the c u r r e n t attachment loca-

I

tions.

The specific equations that define the gore geometry, a s derived in Reference 3, a r e repeated below for convenience.

I

_g R = ( % ) P i

I

2 d = (%)(l - cos p ) ; R

I

c = J ~ s i n y - 1 ;

e

I

c -

- -

f m tan y

I

Table 1 1 1 gives the numerical values of the lobe geometry, radical growth due to unrestrained thread racking, and fabric s t r e s s e s based on the preceding e qua tions .

I

A s shown in F i g u r e 23, the gore patterns, a s constructed, actually comprise three gores, The technique of combining three gores into one pattern to fa- cilitate fabrication w a s applied successfully to the models of Figures 6 and 7.

Due to fabric width limitations, a splice o r c r o s s s e a m in the gore pattern of the current models was necessary. The splice begins a t the burble fence and proceeds to the r e a r postion of the pattern where l e s s aerodynamic loading

I

resulting in 48 takes place. A total of 16 of the patterns shown i n Figure 23, actual gores, are used. Four of the 16 patterns a r e modified for each model to accommodate the inlet a s s e m b l y a s shown i n F i g u r e 23.

I

-26-

I

TABLE II. - AID COORDINATES

Y/R, front y/R, r e a r surface surface 1.00OOOOOOEtOO 0. 0 .

9.99000000E -0 1 -2.35478194E-02 3.40563513E-02 9.75000000E-01 1.62452220E-0 1 -1. 15803332E-01 9.50000000E-01 -1.59982524E-01 2.230866593-01 9.25000000E-01 2.67284617E-0 1 -1.91343665E-01 9.00000000E-0 1 3.03167273E-01 -2. 155876273101 8.75000000E-01 3.3382 1811E -0 1 -2.34951926E-01 8.50000000E-01 3.60836801E-01 -2.505926163-01 8.25000000E-01 3.85244206E-01 -2.63196950E-01 8. O O O O O O O O E -0 1 4.07622681E-01 -2.732093133-01 7.75000000E-01 4.28356524E-01 -2.80932716E-01 7.50000000E-0 1 4.47723896E-0 1 -2.86580582E-01 7.25000000E-01 4.65934253E-01 -2.90305617E-01 7.00000000E-01 4.83150325E-01 -2.92216955E-01 6.75000000E-01

4.99 50 1793E - 0 1 -2.923907333-01

6.50000000E-0 1 5.15094220E-01 -2.908770463-0 1 6.25000000E-01 5.30015103E-01 -2.87703934E-01 6.00000000E-01 5.44338105E-01 ~ -2.8,2880197E-01 5.75000000E-01 5.58126106E-01 -2.763967473-0 1 5.50000000E-0 1 5.71433447E-01 -2.68227011E-01 5.25000000E-01 5.84307618E-01 -2.58326465E-01 5.00000000E-01 5.96790546E-01 -2.46631265E-01 4.75000000E-01 6.08919603E-01 -2.330558823-01 4.50000000E-01 6.207283963-01 -2. 17489501E-01 4.25000000E-01 6.32247084E -0 1 -1.99790778E-01 4.00000000E-01 6.43499407E-01 -1.79780287E-01

3.75000000E-01 6.54505351E-01 - 1.572295863 -0 1

3.50000000E-01 6.65284423E-01

- 1. 3 1845032E-0 1

3.25000000E-01 6.75855781E-01 -1.03243152E-01 3.00000000E-01 6.862383043-01 -7.09115788E-02 2.75000000E-01 -3.414381233-02 6.96450654E-01 2.50000000E-01 7.065113373-01 8.077529323-03 2.25000000E-0 1 7.16438752E-01 5.730988 13E -02 2.00000000E-01 7.26251250E-01 1. 16163698E-01 1.75000000E-01 7.359671763-01 1.89708972E-01 2.91079991E-01 1.50000000E-01 7.45604921E-01 1.25000000E-01 7.551829693-01 0.

1.00000000E-01 7.64719940E-01 0.

7.50000000E-02 7.742346373-01 0.

7.83746097E-01 5.00000000E-02 0 .

7.93273635E-01 2.50000000E-02 0 .

0. 8.02836895E-01 0 .

-27- TABLE In. - LOBE GEOMETRY, RADIAL GROWTH DUE TO UNRESTRAINED THREAD RACKING, AND FABRIC STRESSES v

- -

-

E Min Min Deg dg/R rg/R S g / R Fb e - - 10 55 0.00096 0.05315 0.01013 0. 1538" 44 0.006 1.02 0.1050 39 -0.006 12 0.05337 0.01121 1.03 2 0.001 17 0. 1051 0. 17 44 35 -0.007 0.007 14 5 0.00162 0.05380 0.01322 0. 1054 0. 20 44 -0.010 0.010 1.04 17 25 0.00251 0.05467 0.01662 0.1060 0. 25 44 -0.016 0.015 1.06 20 0.00357 0.05572 0. 30 -0.023 0.022 1.09 38 0.02010 0. 1067 43 43 23 43 0.35 1.12 0.0048 1 0.05695 0.02357 0. 1076 43 20 -0.030 0.029 25 30 0.00563 0.05777 0. 1083 0. 38 43 3 -0.035 0.033 1. 15 0.0 257 1 0 . 4 1 1. 17 27 14 0.00650 0.05864 0.02787 0. 1088 42 47 -0.039 0.038 28 55 0.00743 0.05956 0.03006 0. 1096 0.4400' -0.044 0.043 1. 19 42 29 32 0. 5102 38 0.00978 0.06 193 0.03527 37 0.01326 0.06540 0.04239 0.6032 -0.077 0.062 1.35 8 0.1140 40 44 41 54 0.01791 0.7148 0.07006 0.05123 46 4 0.0230 1 0.07515 0.06042 0. 1229 0.8264 38 -0. 127 0. 113 1.63 49 34 0.02813 0.08044 0.9349 0.06959 50 0.9607 0. 131 20 0.02956 0.08172 0.07179 0.1273 36 48 -0. 152 1.79 51 4 0.03084 0.08299 0.07397 0.9862 51 24 0.03144 0.07500 0.08360 0.9978 51 27 0.03154 0.08370 0.07516 0. 1294 1.0000 21 -0. 163 0. 139 1.85 55 0.0347 1 0.07904 1.0000 0. 159 2.05 0.07709 0.1144 34 57 -0. 190 55 42 0.03433 0.07866 0.07647 0.9933 54 0.03262 0.07696 0.07363 0. 1145 0.9608 35 4 -0. 188 0.158 2.03 53 37 0.03040 0.07473 0.06993 0.9194 51 5 0.02624 0.07057 0.8389 0.138 1.84 0.06292 0.1097 36 26 -0.160 33 0.02264 0.05675 0.7670 48 0.06697 0. 107 1. 58 4 5 0.01871 0.06304 0.04986 0. 1029 0.6850 -0. 119 19 38 29 40 0.05872 0.04200 0.5885 59 0.01439 0.064 1. 31 35 15 0.00996 0.05430 0.03341 0.0958 0.4787 41 11 -0.059 0.00668 0.05101 0.02640 0.3856 29 39 0.04850 0.02022 0.3000 0.029 1. 12 2 3 53 0.00415 0.09 14 43 19 -0.029 0.01585 20 0.00265 0.04698 0.2377 53 0.00176 0.04610 0.01278 0. 1928 44 -0.012 0.013 1 . 0 5 15 0.0898 15 0.04575 0.01 143 0.1730 14 0.00142 1.03 0.00111 0.04603 0.01015 0.0908 0. 1538 44 -0.008 0.008 12 38 32

- -

*

'Indicates x / R value of forward attachment point f o r a s s e m b l y -103.

- 2 8 - 4 (-101 AND -105 ASSEMBLIES)

(-103 ASSEMBLY) - -

REAR INLET MERIDIANS CENTERLI NE AEROSHELL INLET CUTOUT

- -2 -----\I-\---- --

\ GORE CROSS SEAM REAR FRONT .

BURBLE FENCE CENTERLINE

Figure 23. - Model Gore Patterns

Canopy Stowage and Release Two methods of stowing and releasing the canopy w e r e considered. A discus- sion of both is included because, depending upon interface constraints, one The f i r s t method to be discussed concept may offer advantages over the other.

was the approach considered initially for the current wind-tunnel models.

This concept employed 16 small stowage loops, each of which attached to the center meridian of the individual gores. The canopy then is con- strained in the packaged mode, as shown in Figure 24, by a deployment cord passing through the deployment ring and stowage loops on the meridians. The position of the stowage loops along the -103 a s s e m b l y meridians would be the s a m e since the forward attachment i s continuous. However, to prevent loose canopy material in the packaged mode f o r the -101 and -105 assemblies, the This variation results f r o m the position of the stowage loops would vary.

of the continuous attachment a t t h e simulated payload by the attach- interruption m e n t of the canopy a t the aeroshell inlets.

The canopy may be deployed by pyrotechnically cutting the deployment cord.

This permits the p r e s s u r e recovery realized f r o m the a e r o s h e l l inlets to sys- tematically force the cut deployment cord f r o m the stowage and deployment ring loops.

- 2 9 -

I

L I N E C U T T E R C A N O P Y STOWAGE LOOPS ( R E FE R E NC E)

I

D E P L O Y M E N T

I

RING

I

D E P L O Y M E N T CORD

t

I

L I N E C U T T E R 1

I

Figure 2 4 . - Initial Stowage and Release Scheme

I

- 3 0 -

I

Though this canopy r e l e a s e scheme provides simplicity in design, the canopy m a t e r i a l would b e subjected to severe buffeting loads during wind tunnel start.

Consequently, this scheme was discarded for the c o r s e t scheme described below.

The canopy stowage and r e l e a s e scheme adopted for the c u r r e n t models incor- porates a c o r s e t to r e s t r a i n the canopy prior to deployment. At deployment a flexible music w i r e restraining rod is extracted f r o m loops attached to the c o r s e t by the force generated in the stretched bungee cord, permitting the c o r s e t to fall f r e e (see Figure 25). P r i o r to deployment the bungee force is counteracted by a restraining cord, which is pyrotechnically cut a t deployment.

This cutter will be electrically sequenced with the aeroshell inlet r e l e a s e cut- t e r a t the wind tunnel facility s o the aeroshell inlets will b e r e l e a s e d approxi- mately 0.5 sec after r e l e a s e of the canopy. This delay in activating the cutters will permit the release rod sufficient time to withdraw f r o m the c o r s e t loops, A s shown in Figure 25, the release rod is guided into the tube, which houses the bungee, to prevent the possibility of damage to the canopy by the r e l e a s e This method of stowage and r e l e a s e was successfully demonstrated in rod, an inflation t e s t ( s e e Section IX) conducted as part of the c u r r e n t program.

The sequence in which the canopy was folded prior to placing the c o r s e t in The canopy folding procedure illustrated in Fig- place is shown in F i g u r e 26.

Packing of the -103 assembly was u r e 26 was utilized on a l l three models.

the most straightforward since the forward canopy attachment for this model was continuous a t the aeroshell periphery. Packing of the -101 and -105 a s s e m - blies was m o r e difficult due to their noncontinuous forward attachment schemes Figure 2 7 shows the -103 assembly in the packaged as previously described.

mode.

VI. - MODEL INSTRUMENTATION

Each model was instrumented with p r e s s u r e transducers, supplied by AEDC, f o r measuring both aeroshell inlet and canopy internal pressure-time histories.

In addition, six load cells supplied by LRC were installed in each model to m e a s u r e the load-time historv of the meridional tapes.

The method in which the load cells were attached to the meridians 1s shown in Figure 28. Two load cells were installed on each of three meridional tapes p e r model a s shown in Figure 29. The serial number and location of the load cells for the -101, -103, and -105 assemblies a r e presented in F i g u r e 30.

It is expected the load cells, which a r e capable of measuring loads up to 180 lb, will provide good accuracy in detecting the maximum anticipated meridian load of 44.5 lb ( s e e Appendix B).

-31-

I

i

I

W I N D - T U N N E L S T I N G

rn

I

R E L E A S E ROD R E S T R A I N I N G C O R D ( C U T A T C A N O P Y D E P L O Y M E N T SO P Y R O T E C H N B U N G E E W I L L E X T R A C T R E L E A S E C U T T E R R O D F R O M C O R S E T ) C O R S E T LOOPS

I

T H R O U G H W H I C H R E L E A S E R O D PASSES F R O N T VIEW

I

S I D E VIEW

I

L A C I N G

C O R D -

I

W

I

R E A R VIEW

F i g u r e 2 5 . - Canopy Stowage and Release Scheme

I

- 3 2 -

rtl B U R B L E F E N C E

(A) F E N C E F O L D E D INWARD TO R E D U C E EXCESS M A T E R I A L I CLAMPS USED T O RESTRAIN FOLDS DURING PACK1 NG (01 GORES F O L D E D LONGITUDINALLY (LOOKING FORWARD)

A

CORSE IC) CANOPY P R O F I L E FOLDS

Figure 26. - Folding Sequence

- 3 3 -

Figure 27. - AID Model in Packaged Mode

- 34-

E L E C T R I C A L L E A D S L O A D C E L L

-

F A B R I C A T T A C H M E N T , ( I N T E R I O R L O O P SEWN SUR F A C E )

r T H R O U G H TO

/ I I M E R I D I A N M E R I D I A N f

/ M E R I D I A N

F i g u r e 2 8 - Load Cell Attachment to Meridian

X/ R

F i g u r e 29. - Profile of AID Showing Meridional Tape Load C e l l Locations

-35- A F T RAM-AIR INLET , M E R I D I A N ( T Y P I C A L ) \ S E R I A L N U M B E R F R O N T R E A R POSl T I O N S U R F A C E S U R F A C E ASS EM B L Y 1 4 X 2 5 -101 Y 2 3 6 7 10 X 8 1 1 - 1 03 Y 2 9 12 X 13 16 14 17 - 1 05 Y 1s 18

Figure 30. - Serial Numbers and Locations of Load Cells for E a c h Assembly

-36-

Next: Appendix A. →

Appendix A.

VII. - DESIGN DOCUMENTATION

The detailed engineering drawings of each AID a s s e m b l y are presented i n Table IV p r e s e n t s an index delineating the details and subassem- Appendix A.

blies for the three final assemblies.

TABLE 1V.- INDEX T O ATTACHED INFLATABLE DECELERATOR DRAWINGS De tails Subas semblie s F i n a l a s s e m b l y

(645A000- -- - )

(645A000-003- - )

(645A000- -- - 1

-003-3, 5, 7 , 9 , 11, 13, 17, -101 -004-101, 115 19,21 -102-101

- 116-7

-103-103 -117-5, 7, 17, 19

- 116- 101

-119-1,7, 13 -117-101, 103, 105, 111 -118-101 -003-3, 5, 7, 9 , 11, 15, 17, ~ -004-103, 115 -103 ~ -102-101 -116-7

1 -103-103

-117-7, 17, 19

' -116-101

- 119- 1,7, 13,47

-117-101, 105, 107, 109, 111

- 118- 103

-119-105, 107, 109 -003-3, 5, 7, 9 , 11, 13, 17,

- 105 -004- 105, 115

19,21 -103-101 -116-7 -103-103 -117-5,7, 17, 19

- 116-101

-119-1,7, 13 -117-101, 103, 105, 111 -118-101

- 119-103, 107, 109

-37-

Next: SECTION VIII - MATERIALS SELECTION →

SECTION VIII - MATERIALS SELECTION

SECTION VIII - MATERIALS SELECTION 1. FABRIC Two of the AID canopies were fabricated o t Nomex and the other ot Dacron.

The Dacron fabric was chosen for one model because i t r e s u l t s in a fabric ultimate load to developed load m o r e representative of flight model design than does the Nomex fabric. The design factors for both the Nomex fabric and Dacron fabric a r e presented in Table V.

TABLE V. - DESIGN FACTORS

I Nomex I Dacron*

F a b r i c Me r idian s Meridians F a b r i c F a c t o r s

--

Overload (flow breakdown) 1.50 1. 50 1.00 1.00 Dynamic ( f l a g snapping) 1.50 1. 50 1.50 1.50 1.38 1.33 I 1.38 1.33 Temperature

1.25 l . o o i 2.64 1.00 Seam efficiency (see

Section I X )

Racking ! 1.10 I 1.00 1.10 1.00

1. 50 2.00 2.00 1.50 Safety factor

- - - -

Composite de sign factor 8.55 6 . 0 0 9.00 3.00 F r o m Appendix B the expected fabric s t r e s s , for the design conditions given I of this report, i s 3.09 lb/in. Thus, the ultimate room temperature in Table m a t e r i a l strength needed is: = 27.6 lb/in.

The actual Dacron strength given in Section IX is 40.8 lb/in. which r e s u l t s in the following margin of safety: 40.8

(M.S.) = - - 1 . 0

27.9 = 0.46.

-38-

I

The actual Nomex f a b r i c strength given in Section I X i s 96.4 lb/in., which r e s u l t s in the following margin of safety:

I

96.4

- -

1.0 (M. S.) = 27.9

I

2.45.

Thus, the Dacron fabric has a margin of safety which i s approximately a

I

factor of 5 l e s s conservative than the Nomex fabric.

Meridional tapes with an ultimate strength of 548 Lb, which is essentially the

I

strength used on the previous wind-tunnel models, w e r e used on the current models. Appendix B indicates that the actual loads in the meridians will not exceed 44.5 L b for the conditions presented in Table I of this report. B y ap-

I

plying the flight test composite design factor of 3 . 0 f o r the meridians, the required meridian ultimate strength i s = (3)(44.5)

I

= 134 Ib,

I

The reason that the which i s well below the 548-lb strength actually used.

because the basic isotensoid theory 548-lb strength meridians were used i s says that the elongations of the meridians and fabric m u s t be the s a m e i f these

I

two components a r e to c a r r y the proportion of the total drag load predicted by the theory. A comparison of the elongations, which can be expected for the c u r r e n t models, i s shown below. F o r the fabric,

I

- FtU d f - - f b

I

- 40.8 - - 3. 09

I

= 13.2 .

F o r the meridians, - - - dm T

I

m - 548

I

-44.4 = 12.3

I

Therefore, the elongations are nearly the same, which means that the tapes and meridians will c a r r y the proportion of the total load predicted by the isotensoid theory. This observation a s s u m e s that the s t r e s s - s t r a i n charac-

I

t e r i s t i c s of the tapes and fabric a r e the same.

-39-

I

Both the Nomex and Dacron f a b r i c s were calandered and coated to reduce their permeability to a level shuch that the r a m - a i r recovered by the inlets

I

will maintain the design shape and internal p r e s s u r e . Calandering of the fabric, which is a mechanical process for reducing the fabric permeability, i s accomplished by passing a heated m a n d r e l under 85 tons of p r e s s u r e over

I

the cloth. The cloth is restrained a t its perimeter to reduce shrinkage.

The permeability of the fabric is then further reduced ( s e e Section IX) by applying a coating which forms a mechanical bond with the fabric.

I

The Nomex fabric was coated with a Viton fluorelastomer coating that had The Dacron been successfully employed in the model design of Reference 3.

fabric was coated with a single component silicone coating. Both coatings

I

at 150 deg and final cured at 325 deg F.

were machine applied, prediried The m a t e r i a l s for the model hard structure were selected on the basis of

I

meeting the AEDC wind tunnel quality control requirements and a r e not discussed in this report.

IX. - LABORATORY TESTING

D

General Laboratory tests that w e r e conducted include: (1) tensile t e s t s of both the basic textile components and various s e a m s and joints used in the decelerator con- struction; (2) permeability tests of both the coated and uncoated Dacron and Nomex cloths; ( 3 ) a functional evaluation of a full-scale aeroshell inlet a s s e m -

I

bly; and (4) a rapid deployment and inflation t e s t of one of the completed AID assemblies. A discussion of the objectives and r e s u l t s of each of these t e s t s is presented below.

Textile Tensile T e s t s

I

Variations in the tensile strength characteristics of different lots of textile materials woven to the same specification c a n occur. A s a result, i t is neces- s a r y to conduct tensile strength tests to determine the exact strength charac-

I

teristics of the lot being used. Tensile tests w e r e conducted for one-inch raveled s t r i p s of both the coated Nomex and Dacron fabrics in accordance with Federal Specification CCC-T- 191 b on a tensile testing machine. Strength tests were not conducted on the Nomex meridian tapes o r on the various thread

I

types used in the fabrication of the decelerators. These m a t e r i a l s w e r e taken f r o m the same Lot as the m a t e r i a l s used in the fabrication of the models shown in Figures 6 and 7. The test results, which established the thread and tape

I

strengths used on those models, a r e presehted in Reference 3 . However, tensile tests were conducted on the Dacron meridional tapes since previous tests had not been conducted, A s u m m a r y of the results f r o m the above tests

is presented in Table VI. 1

Tensile tests also were performed to v e r i f y the s t r u c t u r a l integrity of the main gore seam, r e a r burble fence s e a m , and inlet s e a m f o r the Dacron canopy.

VII. Similar tests w e r e The results of these tests a r e summarized in Table conducted for the Nomex canopies of the model shown in F i g u r e 6 ( s e e Refer- The s t r u c t u r a l ence 3) and a s a result were not repeated during this program.

I

-40-

I

TABLE VI. - TENSILE STRENGTH TEST RESULTS O F

I

CANOPY FABRIC AND MERIDIAN TAPES Nomex fabric Dacron fabric Dacron meridian ultimate load ultimate load tape ultimate load Specimen (lb/in. ) (lb/in. ) (1b)

I

Warp direction 1 42.00 90.0 . . .

2 41.75 95.0 . . .

I

3 37.50 97.0 . . .

4 102.0 41.75 . . .

5 98. 0 41.00 . . .

Average 96.4 40.80 . . .

Fill direction 1 113.0 42.0 . . .

I

2 108.0 43.0 . . .

3 110.0 43.5 . . .

4 111.0 44.0 . . .

I

5 112.0 43.0 . . .

Average 110.8 43.1 . . .

1 . . . . . . 518

I

2 . . . . . . 56 1

3 540 . . . . . .

Average . . . . . .

I

integrity of theloadcell attachment to the meridians, and the forward inlet

I

and forwara attachment, also were established by tensile tests. The results of these tests a r e presented i n Table VIII.

F a b r i c Permeability T e s t s

I

Results of wind-tunnel tests ( s e e Reference 6) of AID models, similar to that shown in Figure 7, indicate that the r e a r surface permeability has a m a r k e d

I

Models tested with a high rear surface influence on the AID drag coefficients.

permeability (10 cu ft/min-sq f t a t 0. 5 in. of w a t e r p r e s s u r e ) had a drag coefficient on the o r d e r of 60 percent of that of the low permeability model

I

(0.02 cu ft/min-sq f t a t 0 . 5 in. of w a t e r p r e s s u r e ) .

To obtain the high p r e s s u r e recovery needed to produce the design shape and design drag coefficient, both the Nomex and Dacron canopies w e r e calendered

I

and coated to reduce their permeability levels. Permeability tests w e r e con- ducted on the Nomex and Dacron f a b r i c s before and after calendering and The results of these tests a r e presented i n Table IX.

a f t e r coating.

I

-41-

I

TABLE VII. - TENSILE STRENGTH TEST RESULTS O F

I

DACRON CANOPY SEAMS

I

U l t i m a t e I n a d Item Specimen (lb/in. ) Main gore seam 1 15.0 2 15.0 17.0

I

4 15. 5 5 15. 0 Average 15. 5

I

Inlet gore seam 16.0 2 16. 0 3 15. 0

I

4 19.0 5 21.0 17.4 Average

I

Upper burble fence seam 1 19. 5 2 20.0 3 19. 5

I

4 19.5 5 19.2 Average 19.5

I

1 21.0 Lower burble fence seam 2 18. 5 3 21.0

4 21.5 I

5 20.2 20.5 Average

I

TABLE VIII. - TENSILE STRENGTH TEST RESULTS O F

LOAD C E L L ATTACHMENTS AND CANOPY

I

FORWARD ATTACHMENTS

I

Ultimate load Specimen (1b)

I

Load cell attachment Nomex 450 Dacron 425 Nomex 422

I

Canopy forward attachments 468 (average of 3 tests) Nomex 464 (average of 3 t e s t s )

Dacron I

I

TABLE IX. - FABRIC PERMEABILITY TEST RESULTS

I

- -

------" _- - -- I- ._ - - . -- _- -_ - -

Dacron (before calendering) 335 cu ft/sq ft/min a t 0 . 5 in. of H 2 0

I

18 cu ft/sq ft/min a t 0.5 in. of H 2 0 Dacron (after calendering) Dacron (after calendering and

coating) 0 . 2 1 x cu ft/sq ft/min a t 1. o in. of H ~ O

Nomex (after calendering and

coating) cu ft/sq ft/min a t 1, o in, of H ~ O 0 . 2 1 x

I

- _ _ -__ . -

- . - - ----- - _.--_ ~ _ _ _ __I_

Aeroshell Inlet Model Evaluation T e s t s A functional model of a n aeroshell inlet was fabricated for evaluation testing.

The model was constructed of the materials t o be used in the actual wind-

I

The following points w e r e investigated during the model tunnel assemblies.

evaluation: (1) effect of fabrication tolerances on inlet functioning; (2) mag- nitude of friction force created between the inlet r e l e a s e lever and the inlet r e l e a s e rod; and (3) susceptibility of the inlet sliding spring to fatigue.

I

T e s t s w e r e conducted with p r e s s u r e loads on the inlet exceeding these antici- pated in the wind tunnel t e s t s (see Appendix D) to demonstrate the inlet r e -

I

lease mechanism. The force on the inlet release spring (Figure 16) was al- ways sufficient to overcome the friction force f r o m the inlet r e l e a s e rod.

Also, approximately 200 deployments were made with the model to deter-

I

mine if the sliding flat spring was susceptible to fatigue. Additionally, the t i m e period that the wind-tunnel models will be stored, with the inlets in the closed position prior to testing, was duplicated with the inlet model.

Again s e v e r a l deployments were performed without signs of fatigue in the sliding

I

flat spring .

I

Deployment and Inflation Test The -101 assembly was successfully inflation deployed in an environmental

I

chamber a t 2 0 psfa, 103,000 f t pressure altitude, and at approximately 162 F.

Inflation w a s initiated by vaporization of a water -alcohol mixture contained in a r e s e r v o i r within the packaged company. The p r i m a r y objectives of the test w e r e t o demonstrate the adequacy of the packing, deployment, and inflation characteristics of the model. Additionally, load cells attached to the model ( s e e F i g u r e 30) provided a n indication of the loads imposed on the meridians during deployment. Figure 31 shows the inflation test setup.

I

. e

I

-43-

I

D

I

Figure 31,. - Water-ALcohol Inflation T e s t Setup

-44 -

I

‘ I

The spoked guide s e r v e s to cradle the AID when inflated since the absence of aerodynamic f o r c e s would otherwise c a u s e large deformations f r o m the

I

design shape. (See Reference 3 . ) The differential p r e s s u r e - t i m e history of the AID during deployment of the

I

canopy is shown in Figure 32, A p r e s s u r e - t i m e history for a similar test of the model in F i g u r e 6 is shown for reference purposes. F u l l inflation of the canopy o c c u r r e d in approximately 0 . 2 4 sec and the d e s i r e d differential

I

p r e s s u r e of 1 . 0 psi essentially was obtained a s indicated.

Load cell force-history for t h r e e load cells a r e shown in F i g u r e 33. Load 5 , and 6 on the r e a r surface meridians w e r e inoperative during this c e l l s 4,

I

deployment. The peak load occurs just prior to f u l l inflation and is slightly l e s s than the calculated meridian tape load a t that p r e s s u r e . (See Appen- dix B).

I

F i g u r e 34 is a photograph of the AID model shortly after inflation to the fully inflated shape. A review of the film coverage of the test indicated that the

I

packing and deployment scheme w a s suitable for use in the wind tunnel.

I

/

I

C U R R E N T R E S U L T

I

F i g u r e 3 2 . - AID Differential P r e s s u r e - T i m e History

I

-45-

I

T I M E [SECONDS)

I

I I

F i g u r e 33 .- Load C e l l F o r c e - T i m e H i s t o r y I

I

i

I

F i g u r e 34. - F u l l y Inflated AID

I

-46-

I

I1

R

X. - CONCLUSIONS AND RECOMMENDATIONS

I

Conc lu s ions

I

Three attached inflatable decelerator models have been designed and fabri- cated f o r supersonic wind tunnel evaluation. The models incorporate four

I quick-release aeroshell inlets to accomplish deployment and four canopy

inlets to effect final canopy pressurization. The following conclusions can be drawn:

I

1. AID models have been designed and fabricated which deminstrate alternate attachment schemes of the canopy to the aeroshell and payload.

I

2. An alternate means of r a m - a i r deployment of the AID, by use of aeroshell inlets, h a s been developed.

I

3 . A rapid inflation test demonstrated the workability of the AID model's inlet r e l e a s e s y s t e m and packaging and deployment s y s tem.

I

II

I

-47-

I

Next: APPENDIX A - DETAILED DESIGN DRAWINGS →

APPENDIX A - DETAILED DESIGN DRAWINGS

APPENDIX A - DETAILED DESIGN DRAWINGS This Appendix presents the detailed design drawings used in the manufacture of the AID assemblies. Only those drawings that differ significantly from those of Reference 3 have been included.

I

-48-

I

~I

P O T OR B E A D C A N O P Y C M D BANDS, T R A N S D U C E R , CUTTERS F A i R l N G

~I

WITH DOW COUNIUG RTV-131 A I K 5 l y . WHEN USING 1 1 7 V - 7 3 1 ON M E T A L , PRlh(E WlTU now C O I N I N G A - 4 0 9 4 PR~MZR ARC- ’ID%mE W I T H M . 6 3 5 7 5 0 D R I V E R , R E C E S S N.1 , T O IN. L g S .

T O R W E W I T H M S 3 3 7 S O O U l V E l , R S C C S S N . 3, T O E O IN L B S .

~I

TOSQUE W l T U M S 3 5 7 S O DRIVER, R E C E S S N O 4, To 40 IN Lm, TORQUE WITU M S ~ J ~ S O DRIVER, R E C E S S NO h, TO IN LBS.

INSTILL INSEUT P S I MIL- SPEC M S 3 3 6 4 b ( G A C PROCESS S R L M69)

I

. .

I b 4 s ~ o o o - I I P - 1 0 5 c OVER ’ O r

I

I 3 5 (.4rA000-119-101 C W C R I I , I I I I _ 64rAooo-119-7 TU8E 3 3 64rAooo-119-7 TU8E 3 3 -4 I -4 1 4 ” 4 ” 4 4 4 6 4 5 A o O O - 1 1 9 - 1 5 W l Y E L A R M 4 4 4 6 4 5 A o O O - 1 1 9 - 1 5 W l Y E L A R M I I JCRCW I I I t

I

scncw

I

I l l I .. I I I I I b 4 S A 0 0 0 - 1 1 7 - I O I COLLAR ASJY 1-3 I

I WASMCR

I l l I I 64saooo-117-1D7 C L A W Q A J I Y I W A S U L R 4 1 4 1 4 1 I 6+Sno00-117-10S T U U E ASSY I W A S H E R ,*ILIT - L O L L

I I N S C R l -LOCK I I I I

9 1 9 4 N 9 ( . 0 - 1 0 W A S H E 5 4 N R S b L O - L WnSMER 6 4 h A ~ 0 0 - 1 1 1 . - 1 0 l S C O O P A S S Y 4 U I S L 7 9 A D b NUT I I I I I b *SA 0 0 0 - 0 -1 O S I DECLLEPATO R A51 Y I 64SA000-004- 103 zLJ DECELERATOR A S I Y M A I b7 9 A 3 Ir b IC I NUT

I

64SAOOO-004-101 DECLLER&TOR *LIY I z c z Ib Ib I(. N A S I L I 9 - 0 4 - I Z SCREW a 8 8 N A 5 1219-C-11 S C R E W

I

3 2 48 3 2 N A 5 l Z 1 9 - 4 - 1 4 S C R E W N A S I Z l 9 - 4 - I Z S C R L W 8 8 8 5 4 / 4 b c4 N & S I Z \ 9 - 3 - 1 0 SCREW -102 PAR1 NOYENCUlURE l l E Y -103 CODE OR YAlERlAl AND SIZE ZONE F : . ” , IOENl OR

I

DESCRlPllON NO. I ) ! & NO. IDENTIFVING NO D . D l < I ICT

I

< u

~~ ~ ~~ ~ ~~ ~~

I

F i g u r e A-l- Model Assembly with 60-In. Dp (Drawing 645A000-003, Sheet 1)

I

-49-

I

lI

I

8 - 3 2 U C I C - 3 s TWIN c I o L o o o - I o z - I o I S U r P e R l - e U O L E S D E P T H .3a A N Z I S - 8 R L L S C R E W 2 R E P 0

I

-50-

I

/

I

~ 4.20 (REF) -

I

v

I

II n

I

, , b 4 S A O o o - l l 7 - 1 0 1 COLLAR^ m:

I

$ 8 , ." I

I

F i g u r e A - 3 - Model Assembly with 60-In. Dp (Sheet 3 )

I

-51-

I

'I

1 I

I ,-~loAooo-lo+-lo1 S U P P O R T ASJY <I-) I/ i

I

M S Z I Z O I F S - z o INSERT (INSTALLLO IN LIOAOIO-IOZ-IOI SUPPORT A S S Y ) N L S \ L L I - S - Z I CCRCW

I

8 R L P D ( . 1 O A O O O - I D Z - I O I C U P P O R T A S S Y 4 C + S A 0 0 0 - 1 1 7 - I 0 3 C O L L A R T O 0 s M L T C H DRILLLD

I

i

I

b \ E W S!AOW\NG P A C U A G E D DECELERATOR L R E L E R S S MLLHhNIS1.L. ON WtYO T U N N E L ST1YG Y O s c a L r

I

-52-

I

S S - i ? 5 0 - P S O N SHOULDER SCREW 4 REQD A N 9 b 0 - 1 0 W I \ S M I R (4 REOID) ( U N D E R S W I V E L A R W )

I

-53 -

NOTES CONlINULO: F i I. IDENTIFY PER MIL- STD- 130 ( o a c V R O C L S S SPLC 12) 2 . S E W P E R STO 711 3. SEE G R ~ O - 4 - 5 1 Foe S L W I N G S Y M S O L EXPLANATION D P h R T N . - 1 TURU -10, - 2 5 THRU - 4 1 , - 4 9 , - 5 1 , - 5 9 , - L O 6 - 6 9

T O %E e F O U - 1 0 1 e -103 A 5 5 Y L F o u -105 A S S Y

GOoDTLLR A E I O I P h C E C O R P C O D E G X 6 0 1 V O 3 O L , N O M C X / V I T O N FA0RIC (WHITE) Z.G O Z / Y O ~ , 9s a 8s LIS/IN T. s.

e D A C R O N C L O T H , 1.1 07./YD2 ( N C T Y R h L ) , P A l T C R N N . 15573 ~ 35 X 3C L d S / I U 7 . S.

N O W E X TAPE - S I M I L A R T O M I L - T - 5 0 3 0 , T I ? & Ip:, 1.00 W8-E M A K E NDMEX T U B U L A R U R A I D E D C O R 0 F R O M I 2 0 0 D E N I E R 3 ! / a 01 T W l S T - N O M E L I P R N A S F O L L O W S : I(. CARUIERS, 2 E N D S PER C b Q R l E k , 9 TO I 1 P I W S PER I N C H M A T E R I & L FOR - 1 0 1 e -103 TO B E N O M E X T 4 F 1 , 51MIL(\R T O M I L - T - 5 0 3 0 T Y P E P , 9/tU W W L , C O O L I S TEN.

M A T E R I A L F O R -105 A 5 J Y TO BE DACRON T A P E , PATTERN N o 7 9 2 8 .50 W b O E , S O 0 L 3 J TEN. - B A L L Y R l B B O U M I L L S , B A L L Y , P A .

@ NOWEX T h P L , S I M I L A R T O M I L - T - S O 3 0 , T T P L P , 9/tb U l D L , S O 0 L B S . T E N .

L R P D l n S S T R I P O N L I N C H A T S P L I C E S P U T - 1 1 M E R I D I A W W E B UNDCU 5 L B S FOR O N E M I N U T E M A X .

T O M A R K T O EUVELOPE W 4 T L M M P R K S 13- CANOPY S H A P E P A L A M L T E U L : K ( - o . s I 9 9 , K , - 0 . 4 4 2 1 , g = 0. 6 8 5 120. C O N F l G U R A T I O N : F R O N T ATTACHVIEUT - X / r ( r 0.1538, Y / * = - O - i + S R C A U A T T A C H M I N T - 0 . I r 3 8 , Y / R = 0 . 1 9 1 1 4 0 - COb4FIGUUnTIOU: FRONT A T T I C H M I U T - X / R = 0, 4 4 , Y/R: - 0.b253 REAR A T T A C H M E N T - X/R = 0.1538, y / n = 0 . ~ 9 1 O E T E U M I N L -65 R O P E L O C A T I O N B Y F I T T I N G T O C L q M P A T W E S T A S S E M B L Y INSTALLATID* INSTRUCT~ONS F O R - L 9 R E l N F - T A C K E D G E ALONG r O F - 1 1 MERIDIAN W L U , S E W L O W E C E D C L A S SHOWN. SEW A L O N G h L L M E R ) O I ~ N W E B S USING ? ! .

C L D S l N G AT TACKED E D G E W l T H I/L IN L A P 304 5-7 P R l o R T O C L O S I N G A S S Y , COAT A L L S T I T L Y I N G INSIDE E N V E L O P E ( I N C L U D I N G B U R O L E FLHCZ) T O L E P L H C E FOR O l M E N S l O N 0 TO 2.0 IN. 2.04 10 T O 3 0 I N . ? . I 9 Z TO 10.0 IN. t.01. 30 T O L O 1N 2 . a 5 1 . 0 I W . c "? r I70 MCPIOII"' ,LO. * P A L T ) S C H E M A T I C OF INLETS ( L O O K I N G INTO FORWARO E N D ) N O SCALE

Figure A - 6 - Decelerator Assembly 60-In. Dp (Sheet 1 )

i

I !

-54- \

h

Figure A-7- Decelerator Assembly 60-In. Dp (Sheet 2 ) -55- b l N S C R T L Y D INJIDL ColO T O F O R M L O O P e SEW F

I

Figure A - 8 - Decelerator Assembly 60-111. Dp (Sheet 3) i -56- .E OF C O O R D I N A l r c s - S C - L s/r Pn 8NLL F W D

-

GYM

i

:b X DIM 9 . 3 E T A I L -7 REINF S C A L E 2/1 - 9 b/F PANEL (SHOWN) -10 ( O P P ) 31.59

/ ' - &

F i g u r e A - 9 - Decelerator Assembly 60-In. Dp (Sheet 4) -57-

A P P E N D I X B. - STRESS ANALYSIS

Fabric S t r e s s e s The decelerator s t r e s s analysis in Reference 3 presents the relationships, based on the isotensoid method of analysis, f o r determining the Limit Loads for the meridians and fabric. These relationships are: - k p n R 2 ( B - ' l ) Trn - n where from Reference 3 kf = 0 . 5 2 , and F = 0. 1294.

F r o m the relation C = 2.0, (see Reference 3 ) , Pi Pi - P, = 2 . 0 ; 9 , Pi - P , = 2.0 .

0 . 7 M o o P , The refore,

- Pi = 0 . 7 Ma2 ( 2 ) t 1

( B - 3 ) pa3 = 1 3 . 6 for M , = 3 . 0 .

Since p = pi - pb , F r o m the analytical p r e s s u r e distribution of Figure 9 and by substituting Equation B-4 into Equation B - 3 , the following relation i s obtained:

J?- = 1 3 . 6 0 - 0 . 4

pm = 1 3 . 2 -58-

I

F o r q, = 120 psf and Moo = 3.0,

I

= 19.05 psf.

PCQ Therefore, p = (13.2)(19.05)

I

= 252 psf = 1. 75 psi.

I

Therefore, the limit meridian and fabric loads a r e , from Equations B-1 and B -2, respectively, - (0. 52) (1.75) ( n ) (27.3)2 T r n - 48

I

= 44.5 tb,

B

and ( 0 . 1294)(1.75) (27.3)

-

fb - 2

= 3.09 Ib/in.

I

Hard Structure

I

The h a r d structure for the c u r r e n t models was s t r e s s analyzed in accordance with the AEDC wind tunnel quality control procedure (AEDC QCP-000-2 1 ) .

The analysis is not detailed in this report.

I

Next: APPENDIX C. - AID INFLATION ANALYSIS →

APPENDIX C. - AID INFLATION ANALYSIS

In setting up a computer program to predict the inflation rate of an attached inflatable decelerator, a basic opening model m u s t be postulated to charac- terize the events that take place during the inflation process. An effort has been made to keep this model simple while still permitting a n adequate model- ing of the flow a t the inlets. Wind-tunnel movies of the AID have been used to Reasonable agreement generate the inflation model shown in Figure C - 1.

with experimental results should be attainable with this model.

The initiation of the inflation sequence is taken as the time when the front inlets a r e facing the a i r s t r e a m . As this f r e e - s t r e a m airflow approaches the blunted conical forebody, i t passes through a detached shock wave that i s very nearly normal to the flow in the nose region of the forebody. The flow that e n t e r s the inlets p a s s e s through this normal portion of the shock.

The f r e e - s t r e a m total p r e s s u r e and temperature of the flow a r e given by the following r e l a tions , r e spe c tive 1 y: 00 a3 The total temperature of the air remains constant as i t passes through the shock wave; however, the total p r e s s u r e will decrease a c r o s s the shock wave.

F o r the flow passing through the normal portion of the detached shock, the new total p r e s s u r e is: The flow, after passing through the shock, slows briefly in the stagnation region, turns along the surface of the blunted nose and expands onto the conical surface and subsequently e n t e r s the forward inlets. If the cone angle is small o r the Mach number high, the flow a t the forward inlets could be supersonic.

The static p r e s s u r e at the forward inlet is found, assuming Newtonian flow over the cone, f r o m the following relation:

- 2 sin2

Pf - Y P,Mm Z

The Mach number in the inlet region is given by the relation: -60-

I

O W a 0 J 9 n

0 W I

11 n .- E a

I

II

I

I- v) I - a I- UI

I

a a I 1 .- a

I

n o o a

I

Figure C - 1. - AID Inflation Model

I

-61-

I

Of course, if Mf = 1, the front inlet total p r e s s u r e , ptf, is still pt. However,

i f Mf > 1, t h e n the total p r e s s u r e behind the inlet shock is given by the rela-

tion: At the time the canopy s t a r t s to inflate i t is assumed that the internal p r e s - s u r e , pi, of the undeployed decelerator is equal to the base p r e s s u r e of the The base p r e s s u r e is based on the empirical correlations: vehicle.

p b - p c o - ( M - 2 1 )"p_M_ 2 .

t 0 . 7 The total p r e s s u r e ratio a c r o s s the inlet is now known and the m a s s flow rate into the inlet can be calculated. The m a s s flow rate i s a s s u m e d to be propor- tional to the isentropic flow rate, where the proportionality constant is the inlet flow coefficient, This coefficient is treated in the manner suggested in Reference 1. The coefficient is expressed a s a combination of two factors: the f i r s t i s the flow coefficient of a sharp-edged orifice. This coefficient is e x p r e s s e d a s a function of the total p r e s s u r e ratio a c r o s s the orifice. The second factor is a constant inlet efficiency to account for the inlet geometry and i t s fabric construction.

The mass flow rate at the forward inlet then, is given by the following relation: w h e r e Cf = f (ptf/Pi) the As the decelerator s t a r t s to inflate behind the aeroshell ( s e e Figure C - 1 ) , volume will increase with time, while the internal p r e s s u r e remains at pb according to the following relation: m t f

V ( t ) = -

P -62- air a t a p r e s s u r e p At some nominal volume, the canopy will be filled with b is and from that time on, the p r e s s u r e will r i s e until the final p r e s s u r e reached. F o r the model in Figure C-1, the nominal volume was chosen as Vo, which is equivalent to the total canopy volume. After V o is reached, the internal p r e s s u r e rises as a function of time according to the following relation: t Pi =/ RTt .

Once the p r e s s u r e in the canopy exceeds that in the base region, the air will s t a r t to flow through the fabric. The rate of this flow will depend upon the a heavily coated cloth this loss porosity o r permeability of the fabric. F o r will be neglegible, but for some lightly coated cloths, this l o s s m a y be s i g - nificant enough to modify the inflation p r o c e s s . Another factor that m u s t be considered in calculating the air flow from the rear surface fabric is the in- c r e a s e in effective porosity of the r e a r surface due to an increasing p r e s s u r e differential a c r o s s the r e a r surface as the model inflates. A relationship between the porosity and p r e s s u r e ratio of the following form is used.

The m a s s flow rate due to this porosity is given by the following relation: Now, the mass flow into the system is given by:

= mf - mP

The canopy cannot inflate beyond the edge of the conical forebody until the internal p r e s s u r e exceeds the f r e e - s t r e a m static p r e s s u r e . When the p r e s - s u r e rises above the free-stream value, the f a b r i c adjacent to the h a r d cone a s s u m e s the form and angle 8 shown in Figure C - 1 relative to the aeroshell longitudinal axis; in this case the assumed Newtonian p r e s s u r e distribution is: 8 = a r c sin As this angle increases, the AID continues to inflate as a r e s u l t of the air flow into the front inlets only, until the r e a r inlets become effective. At the point where the r e a r inlets become effective, defined as the position where the -63- r e a r inlets become parallel with the flow, the mass flow equation m u s t be

modified. The angle by which the inlet l a g s behind 0 is defined by 8, there-

-

the angle between the f r e e - s t r e a m flow and the r e a r inlet is 8 - 0.

fore, The static p r e s s u r e in the r e a r inlet region is: 2 2 sin ( 0 - e’> , Pr = YMco P , and the Mach number a t the r e a r inlet is: y-l r =&%)[($) 41.

<

However, i f M ’> 1, Again, if M r = 1, then the r e a r inlet total p r e s s u r e is p the value will be given by t’ r L J A s before a dual factor flow coefficient is used and the m a s s flow into the r e a r inlets is given by: t Y - 1 Y

mr = Q r r r C A 4 * ) ( & ) ( 2 F ”

-(e)

The mass flow into the system is now:

. 0 e e

m = m f - m P + m r This flow is allowed to continue until the p r e s s u r e ratio a c r o s s theminletsn e a r s unity. I f the r e a r p r e s s u r e ratio p./p exceeds 99 percent, then mr is t r If the front inlet b r e s s u r e ratio Pi/Ptf exceeds 99 percent, equated to zero.

the computation is concluded and the system is considered to be fully inflated.

The integration of the m a s s flow rate with time is done by simply summing If a f t e r 200At intervals, the system has not fully inflated, the &At values.

The summation of 200 At intervals can occur computation is terminated.

The inflow through the inlets will be e a s i l y f o r systems with porous fabrics.

equalized by the outflow through the fabric and the p r e s s u r e will stabilize a t a value well below 0.99 ptf.

- 64-

Next: APPENDIX D. - AEROSHELL INLET ANALYSIS →

APPENDIX D. - AEROSHELL INLET ANALYSIS

General This appendix discusses the analysis conducted to determine the forces acting of the inlet spring, and the fric- on the aeroshell inlets, the characteristics tional forces that the deployment spring m u s t overcome to release the aero- shell inlets .

Inlet F o r c e s When the door is in the closed position, the net force acting on the inlet i s P r e s s u r e measurement from the wind-tunnel tests of models shown in Fig- pi, of the model prior to u r e s 6 and 7 indicate that the internal p r e s s u r e , deployment i s relatively small when compared to pa . The external p r e s s u r e , pe, however, is approximately one order of magnitude greater than pa a s will be shown subsequently. Therefore, assuming that the internal p r e s s u r e is negligible a s compared to the external p r e s s u r e , the maximum differential p r e s s u r e acting on the inlet is equivalent to the external p r e s s u r e . The maxi- m u m force acting on the inlet in the closed position then is: F = peAi .

m To determine the magnitude of the external p r e s s u r e when the door is in the closed position, modified Newtonian theory is used to predict an average p r e s - F r o m the Newtonian theory, this ratio is defined as: s u r e ratio of pe/pa.

- Pe = 1 . 0 t ("' - - 1.0) s i n 2 8 .

pa3 Po0 F r o m Reference 14, a t Ma = 3 . 0 : = 1.21 F o r the 120ddeg conical aeroshell, 0 = 60 deg. Therefore, Pe

- = 1.0 (12. 1 - 1.0) sin2 60 deg

pa0 = 9 . 3 2 .

-65- F o r the 140-deg aeroshell, Pe

-- - 1.0 t (12. I - 1 . 0 ) sin2 70 deg

pa3 = 1 0 . 8 .

Using the l a r g e r of the p r e s s u r e ratios, the maximum force acting on the inlet is: Pe F = - m Ai = 5.36 , where A. is as d e t e r m a e d in Figure 13.

A s stated i n Section V , the inlet spring was designed to ensure a positive force F o r design in the open orientation to present the possibility of inlet flutter.

purposes the spring was considered to maintain a force equivalent to that act- ing on the inlet i n the closed position, which was previously determined as The subsequent analysis will show that there is actually a small a e r o - 5. 36 lb.

dynamic force tending to maintain the inlet open. Therefore, the aeroshell inlets should deploy r a t h e r positively and maintain the deployed position with- out unfavorable dynamics o r flutter. The f o r c e s acting on the aeroshell inlet ( s e e Figure D-1) a t the beginning of inflation only a r e represented by the fol- lowing relationships: and The Larce resulting from the cliange in momentum in F,gure D-2 m u s t be equal T h e r e f o r e , to the resultant of the three forces shown in Figure D-1.

- A 2 d F - F1 - F2 = x(m’;) .

By summing forces i n the x direction, then: - 2 A

CF = - F t F1 cos (90 - 0 ) -t F2 C O S 0

X 4 d = - P V A [ V 1 cos ( 9 0 - a ) ] - 1 1 1 - 6 6 - X

F \

I \ I

F i g u r e D-1. - F o r c e s Acting on Aeroshell Inlet

I

I I I A 1 V I 1 I, I I I

i

I

F i g u r e 0 - 2 . - Change in Direction of Flow Entering Aeroshell Inlet

-6.7- The ref0 re, A 2 A --. -L

F = F1 cos (90 - CY) t F2 cos cy t PIVIA1[V1 cos (90 - a)] t

A F A 1 p2V2A2 [V2 cos (90 - 8 + c y ) J .

The force of the fluid on the inlet is computed below for both the 120- and 140-deg aeroshells at: (1) s t a r t of inflation and ( 2 ) full inflation.

The follow- ing p a r a m e t e r s do not vary for e i t h e r configuration.

= 190.5 psf, pal m = 3 . 0 , A1 = 0.00892 sq f t , A2 = 0.0211 s q f t , 8 = 6 5 deg, and CY = 23.6 deg.

The inlet Location is 0.64s, where s is the distance f r o m the aeroshell apex to the base.

F o r the 120-deg configuration, pL/p, = 9.97, and ML = 0.49 a t 0.64s f r o m the apex according to Reference 17.

The ref0 r e , pL = (9.97)(19.05) = 190 psf, and - Y p m q L - 2 L L = (0.7) (190) (0.49) = 32 psf.

At the s t a r t of inflation, P2v2 = P I V l = 2qL 9 where T o determine the resultant of the forces shown in Figure D-1, the values of the above p a r a m e t e r s are substituted into Equation D-1: A

F = (190)(0.00892)(cos 66.4) + (190)(0.02ll)(cos 23.6) t

2(32)(0.00892)(cos 66.4) t 2(32) (O.O2ll)(cos 48.6) = 5.45 Lb.

A acts on the back surface of the inlet in a negative The recirculation force, x direction. It is where A. is determined from Figure 13.

The refore, A = (190) (0.0236) FB = 4.48 .

The resulting aerodynamic force acting on the inlet is: F = F - F B a = 5.45 - 4.48 = 0 . 9 7 .

At full inflation, Fa = qCDA1, where the CD of a flat plate is used.

Therefore, Fa = (32)( 1. 17)(0. 00892) = 0.884 Lb.

F o r the 140-deg configuration, p /pa = 10.28 and ML = 0.441 a t 0.64s f r o m the apex according to Reference 1 2 .

Therefore, PL = (10.28)(19.05) = 196 psf, and

- ( 0 . 7) (196) (0.441)’

qL -

= 26.7 psf.

-69-

I

Using the same calculation procedure as for the 120-deg cone, the resulting aerodynamic force a t the beginning of inflation is

I

F = 0.92 Ib, a and a t full inflation, the resulting aerodynamic force is F = 0. 88 Lb.

I

Therefore the fact has been established that the resultant aerodynamic force acting on the inlets tends to maintain the inlet in the open position.

I

Release Spring F o r c e s a force g r e a t e r than the fric- The release spring must be capable of producing tion force generated between the four restraining rods and the release levers.

B y summing moments about the rotational axis of the inlet shown in FigureD-3, then o r (0.65) (25) F = 2.55 rl

I

= 6.38 Lb.

i

I

1 / R E L E A S E L E V E R

I F . I N L E T IS R E L E A S E R O D R O T A T I O N A L AXIS

I

~~ Figure D - 3 . - F o r c e s Acting on Inlet Release Spring

I

- 70-

I

The maximum friction force occurs as the inlet release lever s t a r t s to slide over the release rod. F o r steel on steel, the coefficient of static friction is

I

p = 0.28; therefore, Ffm = (0.28)(638)

I

= 1. 78 Lb.

I

F r o m the geometry in Figure D-4, Frs is a minimum i f l is made as l a r g e as possible and l2 as small a s possible.

On the basis of the physical limitations involved, the maximum 11 is 2.06 in. and

I

the minimum 12 is 0.90 in. Before the r e l e a s e spring can open inlet release, the following inequality m u s t be satisfied:

I

Therefore, >, (1.78) ( 0 . 9 0 ) F - 2.06 r s

I

2 0.78 lb.

I

Ffrn

t

R E L E A S E L E V E R R E L E A S E L E V E R R E L E A S E L E V E R R E L E A S E L E V E R

I

P I V O T P O I N T P I V O T P O I N T I N L E T / I R E L E A S E I R O D

\

I

i

I

Figure D-4. - F o r c e s Acting on Release Lever

I

-71-

I

Since the f o r c e developed by the release spring is transmitted about a circu- lar plane of the cone ( s e e Figure 13), a force input factor a t the various loca- tions on the circle m u s t be determined. Table D-I s u m m a r i z e s these factors f o r the inlet locations (relative to the release spring) and the resulting release spring forces required.

By summing the forces in Table D-I and applying a safety factor of 2.0, the r e l e a s e spring must provide a n 8.6-lb force to release the inlets. In addition the spring m u s t possess the physical characteristic ( s e e Figure D-5) to satisfy the following inequality:

>

i i - if = i 2 e .

TABLE D-I. - RELEASE SPRING INPUT FACTORS*

Inlet position relative to release spring Input factor, rs H (deg)

I HF

I I 20 1.24 0.97 110 1. 27 0.99 20 0 1 . 4 0 1.09 1. 25 29 0 1. 60

*

Source: Reference 19 Since 1 = 0.90 in.,

> 3

1. - lf = (0.90) -

1 n

>

= 0.86 in.

A helical spring with a n li = 3.0 in. and an lf = 1. 60 in. satisfies this con- The spring chosen has a spring constant, K , of 14. 5 lb/in. and will straint.

deliver a force f o r an elongation of 0.86 in.

F = K x rs = (14. 5) ( 0 . 8 6 ) = 12. 5 lb, which is g r e a t e r than the 8.6 l b previously determined as being required.

-72-

I

I L

f

I

( A ) B E F O R E R E L E A S E

I

------I

F6V

t' D

I

1 .--&

I

( 8 ) A F T E R R E L E A S E

If ___c(

~ Figure D-5. - Release Spring Geometry before and after Release Figure D-5. - Release Spring Geometry before and after Release

I

Inlet Spring F o r c e s

I

The f o r c e that the sliding flat spring was designed to exert to maintain the inlet in the deployed condition is computed by summing moments about the inlet rotational axis, f r o m Figure D-6.

I

M = 0 . 0 = (0.5)(Fi) - xFs .

I

There- The f o r c e acting on the inlet, as determined previously, is 5.36 lb.

f o r e f r o m Equation D-1, for an inlet t h r e e inches in length,

I

xFs = 8 . 0 4 in. -1b.

(D-2)

I

The bending s t r e s s and deflection in the spring a r e defined by the following relations, r e spec tively , 6 F sl - - (D-3) fb - wt

I

and

I

-73-

i

I

F I X E D END

I

S P R I N G P O S I T I O N , I N L E T C L O S E D

I

I I I I t

I

I / SPRING P O S I T I O N ,

I

I N L E T O P E N R O T A TI ON A L C E N T E R

I

N O T E : THIS F I G U R E N O T DRAWN TO S C A L E

I

Figure D-6. - Aeroshell Inlet and Spring Geometry

I

F l 3 S

d = - I

( 0 - 4 ) s 3EI Due to the number of variables involved in Equations D-2 through D-4, a

I

trial-and-error method was employed to determine the maximum bending s t r e s s . A solution was sought that produced a bending s t r e s s below the ulti- < mate of available steel alloys (fb = 200, 000 psi). A value of x = 0.83 in.

satisfies this constraint a s will be shown below. F r o m Equation D-2, 1

8 . 0 4 F = - S O 0.83 = 9.68 Ib.

I

F o r a spring 2. 0 in. long, 1. 1 in. wide, and 0 . 0 4 in. thick, the resulting de- flection required f o r the f r e e end of the spring to move from the f r e e position to the inlet open position ( s e e Figure 0 - 6 ) is:

B

6 FSol ' 1 = 3EI

I

- - (9.68) (2) (12) ( 3 ) ( 2 9 X l o 6 ) (1.03) (0. 04)3

I

= 0. 162 in.

I

The additional deflection, dS2, that the spring must experience to permit the close i s found f r o m Figure D-6 (when drawn to scale) to be 0.26 in.

inlet to

I

i n the spring, generated when the inlet is in Therefore, the maximum force the closed position, f r o m Equation D-4 is:

I

- - (3)(29 X 106)(l.03)(0.04)3 (0.422) (2. 0 l 3 (12)

I

= 25.0 lb.

I F r o m Equation D-3

- (6)(25) (2.0)

fb -

I

(1.03)(0.04 in.)

= 182,000 psi < 200,000 psi

I

Type 1090 spring steel tempered to 250,000 psi was chosen f o r this application.

The loading for the hard structure components making up the -103 assembly

I

a r e either loaded l e s s severely than the -101 and -105 assemblies o r were previously shown to be adequate in Reference 3 .

I

-75-

I

RE FE RENC ES

I

Bohon, H. L. ; and Miserentino, R. : Attached Inflatable Decelerator 1.

Performance Evaluation and Mission-Application Study. AIAA Paper No. 70-1 163, AIAA Aerodynamic Deceleration Systems Conference

I

(Dayton, Ohio), Sept. 1970.

2. Mikulas, M. M. , Jr. ; and Bohon, H. L.: Summary of the Development

I

Status of Attached Inflatable Decelerators.

AIAA P a p e r No. 68-929, AIAA Second Aerodynamic Decelerator Sys tems Conference (El Centro, California), Sept. 1968.

I

Barton, R. R. : Development of Attached Inflatable Decelerators f o r 3 .

Supersonic Application. NASA CR-66613, 1968.

I

Houtz, N. : Optimization of Inflatable Drag Devices by Isotensoid Design.

4 .

AIAA Paper No. 64-437, F i r s t Annual AIAA Meeting (Washington, D. C.), 29 June through 2 July 1964.

I

Baker, D. C. : Investigation of a n Inflatable Decelerator Attached to a 5.

120-Deg Conical E n t r y Capsule a t Mach Numbers f r o m 2. 55 to 4.40.

AEDC-TR-68-227, U. S. Air F o r c e , October 1968.

I

Baker, D. C. : Investigation of a n Attached Inflatable Decelerator with 6 .

Mechanically Deployed Inlets a t Mach Numbers f r o m 2.25 to 4. 75. AEDC-

I

TR-69- 132, U. S. Air Force. June 1969.

Reichenau, David E. A. : Investigation of a n Attached Inflatable Deceler- 7 .

I

ator System for D r a g Augmentation of the Voyage Entry Capsule a t Super- AEDC TR-68-71, U. s. Air F o r c e , April 1968.

sonic Speeds.

8. Bohon, Herman L. , and Miserentino, R. : Deployment and Performance I

Characteristics of 5-Foot- Diameter Attached Inflatable Decelerators F r o m Med Number 2 . 2 to 4.4. NASA TND-5840, August 1970.

I

Gillis, C. L. : Aerodynamic Deceleration Systems for Space Missions.

Pa.), AIAA Paper No. 68-1081, AiAA Fifth Annual Meeting (Philadelphia, Oct. 1968.

I

10. Burgess, J. L. ; Sowa, W. W. ; and Houmard, J. E.: Flight Model - Attached Inflatable Decelerator (AID) P r o g r a m . GER- 14939. Goodyear

I

Aerospace Corporazon, Nov. 1970.

11. Deveikis, W. D. ; and Sawyer, J. W . : Static Aerodynamic Characteristics, P r e s s u r e Distributions, and Ram-Air Inflation of Attached Inflatable

I

Decelerator Models a t Mach 3.0. NASA TN D-5816, May 1970.

12. Reshotko, E. ; and Tucker, M. : Approximate Calculation of the Compres-

I

sible Turbulent Boundary Layer with Heat T r a n s f e r and A r b i t r a r y P r e s - s u r e Gradient. NACA TN 4154, December, 1957.

I

-76-

I

~~ 13. User's Manual for the Boundary Weinberg, S. A.; and Tucker, D. W.: GER-14534. Goodyear Aerospace L a y e r and Wake Computer P r o g r a m .

li

Corporation, September, 1969.

14. Barton, R. R. : Development of InfLatable Attached Supersonic Decelera-

t o r s , Design Report - Phase I. GER-13429. Goodyear Aerospace

Corporation, 18 August 1967.

15. Roark, R. J.: F o r m u l a s for S t r e s s and Strain. Third ed. McGraw-Hill

'I

Book Co., Inc., 1954.

16. A m e s R e s e a r c h Staff: Equations, Tables, a n d C h a r t s for Compressible

I

Flow. NACA Report 1135, 1953.

17. Stallings, R. L. ; and Tudor, D. H. : Experimental Pressure Distribution on a 120° Cone a t Mach Numbers f r o m 2.96 to 4.63 and Angles of Attack f r o m 0 to 20. NASA TN D-5054, March 1969.

P r e s s u r e Distribution on 140°, 160°, Campbell, J. R. ; and Tudor, D. H. : 18.

I

and 180' Cones a t Mach Numbers f r o m 2.30 to 4.63 and Angles of Attack f r o m O o to 20'. NASA T N D-5204, May 1969.

I

19. Anon. : Tru-Lay Push-pull Controls. Folder DH-219-B, American Chain and Cable Company, Inc. (Detroit, Michigan).

Overview

Section View in PDF