apollo 15 aerodynamic stability characteristics of the apollo launch escape vehicle by nasa

8/13/2019 Apollo 15 Aerodynamic Stability Characteristics Of The Apollo Launch Escape Vehicle By NASA

1/134

N A S A T E C H N I C A L N O T E TN -D-3964-

AERODYNAMIC STABILITY CHARACTERISTICSOF THE APOLLO LAUNCH ESCAPE VEHICLE

ATIONAL AERONAUTICS AND SPACE ADMINISTRATION0 WASHINGTON,D. C. 0 J U N E 1967


2/134

TECH LIBRARYKAFB, NM

AERODYNAMIC ST AB ILIT Y CHARACT ERISTICS OF

THE AP OL LO LAUNCH ESCAPE VEHICLE

By Wi l l i a m C . M o s e l e y, Jr., a nd J a m e s G. H o n d r o s

M a nn e d S p a c e c r a f t C e n t e rHo u s to n , Texas

NATIONAL AERONAUTICS AND SPACE ADMINISTRATION~

Fo r sa l e b y t h e C l ea r i n g h o u se fo r Fe d e ra l Sc i e n t i f i c a nd Te c h n i c a l In fo rmat i o nSp r i n g f i e l d , Vi rg i n i o 22151 - CFSTI p r i c e $3.00


3/134

ABSTRACT

A program of wind-tunnel tests was conductedto deter min e the stat ic and dynamic stability of theApollo launch esc ap e vehicle. Static stabil ity studiesincluded the effects of escape rocket firing. Result sof static tests showed that the vehicle wa s s table i nthe no rm al angle-of-attack operat ing range and thatrock et thrusting generally de cre ase d the longitudinal

stability at all Mach num bers tested . Dynamic testdata indicated that the vehicle had positive dampingover the normal angle-of-attack operating rangeexcept for a very s mal l angle range (*2 ) near thetr im angle of attack at subsonic speeds.

ii


4/134


5/134

Section Page

CONCLUDING REMARKS . . . . . . . . . . . . . . . . . . . . . . . . . . . . .REFERENCES . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19

17

iv


6/134


7/134

FIGURES

Figure Page

1 Body system of axes. Force and moment coefficients on modelincluding roc ke t th ru st component. Ar row s indicate positived i r e c t i o n . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

2 Sketch of Apollo launch es ca pe vehicle. Dimensions are forfull-scale vehicle . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26

3. Photographs of test models and components

(a) Apollo LEV model inst all ed in the 8- by 'I-foot test sec tion of

(b) Cataly st packs used in esca pe motor for hot-jet thrustin g

(c) Breakdown of the hot-j et model of the Apollo LEV showing

(d) Hot-jet model tower showing hollow leg s for conducting H 2 0 Z

(e) View of the Apollo LEV model mounted on the t ra ns ve rs e- ro ddynamic stability test setup in the NAA Tri son ic WindTunnel . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29

the Ames Unitary Plan Wind Tunnel . . . . . . . . . . . . . . . 27simulation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27

component par ts . . . . . . . . . . . . . . . . . . . . . . . . . . 28to catal yst pack in escape motor . . . . . . . . . . . . . . . . . 29

4 Model drawing showing propell ant l in es . . . . . . . . . . . . . . . . . 305 Sketch of ra di al flow decomp ositio n cha mb er . . . . . . . . . . . . . . 306 Sketch showing cold-jet simul ation technique. Air sys tem is

physically isolated from model . . . . . . . . . . . . . . . . . . . . 317 A comparison of computer si mul ated and wind-tunnel positio n-time

hi sto rie s (diverging from heat-shield-forward tr im position) . . . . 31

8 Static longitudinal stability characteristics for the Apollo launchescape vehicle at Mach num bers 0.25 and 0.5 as determinedin the Ames 12-Foot Tunnel . . . . . . . . . . . . . . . . . . . . . . 32

9 Static longitudinal stability cha rac ter ist ics fo r the Apollo launchescape vehicle at Mach numbers from 0.5 to 6.0 as determinedin the Am es UPWT and the AEDC-A Tunnel

(a) Pitching-moment coefficient, apex, M = 0. 5 to 1. 35 . . . . . . . . 33(b) Normal-force coefficient, M = 0. 5 to 1. 35 . . . . . . . . . . . . . 34(c) Axial-force coefficient, M = 0. 5 to 1. 35 . . . . . . . . . . . . . . 34(e) Normal-force coefficient, M = 1. 55 to 3.4 . . . . . . . . . . . . . 36(d) Pitching-moment coefficient, apex, M = 1. 55 to 3.4 . . . . . . . . 35

v i


8/134

Figure Page

(f) Axi al-fo rce coefficient, M = 1.55 to 3 . 4 . . . . . . . . . . . . . . 37(g) P itching -momen t coefficient, apex, M = 4.0 to 6.0 . . . . . . . . 37h) Normal-force coefficient, M = 4.0 to 6.0 . . . . . . . . . . . . . 38

(i) Axi al-fo rce coefficient, M = 4.0 to 6.0 . . . . . . . . . . . . . . 3810 Static, thrusting longitudinal stability chara cte ris ti cs for the

Apollo launch escape vehicle at Mach numbers 0 .5 to 1.3as determined in the Langley l6-Foot Wind Tunnel

(a) Vari atio n of pitching- moment coefficient with thr us t

(b) Variation of normal-force coefficient with thrust

(c) Variatio n of axia l-fo rce coefficient with thr us t coefficient

(d) Variation of pitchin g-moment coefficient with th ru st

(e) Variation of norm al-fo rce coefficient with thr us t

(f) Variation of axia l-f orc e coefficient with th ru st coefficient

(g) Variation of pitching-mom ent coefficient with th ru st

(h) Variation of no rmal- for ce coefficient with th ru st

(i) Variation of axi al -fo rce coefficient with th ru st coefficient

(j) Variation of pitching-moment coefficient with thrust

(k) Variation of normal-force coefficient with thrust

(1) Variation of axi al -fo rce coeffic ient with th ru st coefficient

(m) Variation of pitchin g-moment coefficient with th ru st

(n) Variatio n of norm al-fo rce coefficient with thr us t

(0) Variation of ax ial -fo rce coefficient with th ru st coefficient

(p) Var iati on of pitching-moment coefficient with th ru st

(9) Variati on of no rmal -forc e coefficient with thr us t

(r) Variation of axi al- for ce coefficient with th ru st coefficient

(s) Variation of pitching-moment coefficient with thrust

(t) Variation of no rma l-for ce coefficient with th ru st

coefficient at M = 0 .5 . . . . . . . . . . . . . . . . . . . . . . 39coefficient at M = 0.5 . . . . . . . . . . . . . . . . . . . . . . 40a t M = 0 . 5 . . . . . . . . . . . . . . . . . . . . . . . . . . . . 41coefficient at M = 0.7 . . . . . . . . . . . . . . . . . . . . . . 42coefficient at M = 0.7 . . . . . . . . . . . . . . . . . . . . . . 43a t M = 0 . 7 . . . . . . . . . . . . . . . . . . . . . . . . . . . 44coefficient at M = 0.9 . . . . . . . . . . . . . . . . . . . . . . 45coefficient at M = 0.9 . . . . . . . . . . . . . . . . . . . . . . 46a t M = 0 . 9 . . . . . . . . . . . . . . . . . . . . . . . . . . . . 47coefficient at M = 1.0 . . . . . . . . . . . . . . . . . . . . . . 48coefficient at M = 1.0 . . . . . . . . . . . . . . . . . . . . . . 49a t M = 1 . 0 . . . . . . . . . . . . . . . . . . . . . . . . . . . . 50coefficient at M = 1.1 . . . . . . . . . . . . . . . . . . . . . . 51coefficient at M = 1.1 . . . . . . . . . . . . . . . . . . . . . . 52a t M = l . l . . . . . . . . . . . . . . . . . . . . . . . . . . . . 53coefficient at M = 1 . 2 . . . . . . . . . . . . . . . . . . . . . . 54coefficient at M

=1.2 . . . . . . . . . . . . . . . . . . . . . . 55a t M = 1 . 2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . 56

coefficient at M = 1 .3 . . . . . . . . . . . . . . . . . . . . . . 57coefficient at M = 1 .3 . . . . . . . . . . . . . . . . . . . . . . 58

vii


9/134

Figure Page

(u) Var iat ion of axial -forc e coefficient with th ru st coeffici entat M = 1 . 3 . . . . . . . . . . . . . . . . . . . . . . . . . . . 59

11 Variation of aerodynamic char act eri st ics with angle of attack forthe Apollo launch esc ape veh icle as determined in theLangley 16-Foo t Tunnel

. . . . . . . . . . . . . . . . . . . . . . . . . .a) Cm,a M = 0 . 5 6 0

(c) C A , M = 0 . 5 . . . . . . . . . . . . . . . . . . . . . . . . . . . 62(4 Cm,a M= O. 9 63(e) C N , M = 0 . 9 64

(f) CA, M = 0 . 9 . . . . . . . . . . . . . . . . . . . . . . . . . . . 65M = l . l . . . . . . . . . . . . . . . . . . . . . . . . . . 66

(b) CN, M = 0 . 5 . . . . . . . . . . . . . . . . . . . . . . . . . . . 6 1

. . . . . . . . . . . . . . . . . . . . . . . . . .. . . . . . . . . . . . . . . . . . . . . . . . . . .

(g) cm, (h) CN, M = l . l . . . . . . . . . . . . . . . . . . . . . . . . . . . 67

ti) Cm,a(k) CN, M = l . 3 . . . . . . . . . . . . . . . . . . . . . . . . . . . 70

. . . . . . . . . . . . . . . . . . . . . . . . . . .i) CA, M = l . l 68M = 1 . 3 . . . . . . . . . . . . . . . . . . . . . . . . . . 69

(1) CA, M = l . 3 . . . . . . . . . . . . . . . . . . . . . . . . . . . 71

1 2

13

Static, thrusting longitudinal stability characterist ics for theApollo launch escape vehicle determined in the Langleyl6-F oot Tunnel at M = 0.9 (total coefficients, inc lud esth rus t components)

(a) Variation of total pitching-moment coefficient with ang le of

(b) Variation of total normal-force coefficient with angle of

(c) Variat ion of tot al axia l-forc e coefficient with angle of

attack . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 72attack . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 73attack . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 74

Static, thru stin g longitudinal stability ch ar ac te ri st ic s of theApollo launch e sca pe vehicle de term ine d in the AEDC-A Tunnelat Mach numbers fro m 0.7 to 5.97

(a) Var iati on of pitching-moment coefficient with th ru st

(b) Variation of no rmal- for ce coefficient with th ru st

(c) Variati on of a xial-forc e coefficient with th ru st coefficient

coefficient at M = 0.7 . . . . . . . . . . . . . . . . . . . . . .7coefficient at M = 0.7 . . . . . . . . . . . . . . . . . . . . . 76at M = 0 . 7 . . . . . . . . . . . . . . . . . . . . . . . . . . . 77

viii

. . . . .


10/134

Figure Page

(d) Variation of pitching-moment coefficient with th ru st

(e) Variation of nor ma l-fo rce coe fficient with th ru st

(f) Variation of axial-for ce coefficient with th ru st coefficient

(g) Variation of pitching-moment coefficient with th ru st

(h) var iati on of norm al-forc e coefficient with thrus t

(i) Variation of axia l-forc e coefficient with th ru st coefficient

j) Variation of pitching-moment coefficient with th ru st

(k) Variation of normal -force coefficient with th ru st

(1) Variation of axi al-for ce coefficient with th ru st coefficient

(m) Variation of pitching-moment coefficient with th ru st(n) Variation of norm al-f orc e coefficient with thrus t

(0) Variation of axia l-forc e coefficient with th ru st coefficient

p) Variation of pitching-moment coefficient with th ru st

(9) Variation of norm al-f orc e coefficient with thrust

(r) Variation of axi al-for ce coefficient with th ru st coefficient

(s) Varia tion of pitching-mom ent coeff icien t with thrus t

(t) Variation of norm al- for ce coefficient with th ru st

(u) Variation of axial- force coefficient with th ru st coefficient

coefficient at M = 1 . 4 8 . . . . . . . . . . . . . . . . . . . . . 7 8coefficient at M = 1 . 4 8 . . . . . . . . . . . . . . . . . . . . . 79at M = l . 4 8 . .

. . . . . . . . . . . . . . . . . . . . . . . . .80

coefficient at M = 1 . 9 8 . . . . . . . . . . . . . . . . . . . . . 8 1coefficient at M = 1 . 9 8 . . . . . . . . . . . . . . . . . . . . . 8 2at M = 1 . 9 8 . . . . . . . . . . . . . . . . . . . . . . . . . . . 8 3coefficient at M = 2 . 9 9 . . . . . . . . . . . . . . . . . . . . . 8 4coefficient at M = 2 . 9 9 . . . . . . . . . . . . . . . . . . . . . 8 5at M = 2 . 9 9 . . . . . . . . . . . . . . . . . . . . . . . . . . . 8 6coefficient at M = 3 . 9 9 . . . . . . . . . . . . . . . . . . . . . 8 7coefficient at M = 3 . 9 9 . . . . . . . . . . . . . . . . . . . . . 8 8at M = 3 . 9 9 . . . . . . . . . . . . . . . . . . . . . . . . . . . 8 9coefficient at M = 4.99 . . . . . . . . . . . . . . . . . . . . . 9 0coefficient at M = 4 . 9 9 . . . . . . . . . . . . . . . . . . . . . 9 1at M = 4 . 9 9 . . . . . . . . . . . . . . . . . . . . . . . . . . . 9 2

coefficient at M = 5 . 9 7 . . . . . . . . . . . . . . . . . . . . . 9 3coefficient at M = 5 . 9 7 . . . . . . . . . . . . . . . . . . . . . 9 4at M = 5 . 9 7 . . . . . . . . . . . . . . . . . . . . . . . . . . . 95

14 Static, thrusting longitudinal stability ch ar ac te ris tic s of theApollo launch escape vehicle at M = 1 . 9 8 to 5 . 9 7 asdetermined in the AEDC-A Tunnel

(a) Varia tion of pitching-mom ent coeff icien t with angle of

(b) Vari ation of no rm al -f or ce coeff icien t with angl e of

(c) Varia tion of ax ial -fo rce coeff icien t with angle of

(d) Varia tion of pitching-moment coeff icien t with angle of

attack at M = 1 . 9 8 . . . . . . . . . . . . . . . . . . . . . . . 96attack at M = 1 . 9 8 . . . . . . . . . . . . . . . . . . . . . . . 9 7attack at M = 1 . 9 8 . . . . . . . . . . . . . . . . . . . . . . . 98a t t a c k a t M = 2 . 9 9 . . . . . . . . . . . . . . . . . . . . . . . 9 9

ix


11/134


12/134

Figure Page

18 Dynamic longitudinal osci lla tor y stabil ity deri vati ves fo rthe Apollo launch escape vehicle as determined usingthe free -to- tum ble technique in the NAA-TWT andLeRC facilit ies at Mach numb ers from 0. 5 to 1.98

(a) Variat ion of damping-in-pitch pa ramet er with angle ofattack at Mach numbers 0. 5, 0.7, and 0.8.

Im

attack at M = 1.59 . . . . . . . . . . . . . . . . . . . . . . . 116attack at M = 1.98 . . . . . . . . . . . . . . . . . . . . . . . 117

2= 0.251 slug-ft . . . . . . . . . . . . . . . . . . . . . . . 115

(b) Variation of damping-in-pitch pa ramet er with angle of

( c ) Variat ion of damping-in-pitch pa ramet er with angle of

19 Comparison of st ati c pitching-moment coefficient dat aobtained with sting-mounted and tra nsv ers e rod-mounted model s. (x/ d = -0.104, z/d = 0) . . . . . . . . . . . . 118

20 Variation of summ ary param eter s, A, a=O0 N Ly

and C with Mach number . . . . . . . . . . . . . . . . . . . 119m, aLy

xi


13/134

AERODYNAMIC STABILITY CHARACTERISTICS OF

THE APOLLO LAUNCH ESCAPE VEHICLE

By Will iam C . Moseley, J r. , and Jame s G. HondrosManned Spacecraft C ente r

SUMMARY

Wind-tunnel tests were made at several facilities to determine the static and thedynamic stability ch ar ac te ri st ic s of the Apollo launch esca pe vehicl e at Mach numbersfrom 0.25 to 6 0 Static stability tests included the effects of launch esca pe rock etthrusting. Dynamic stability der iv at ive s were determ ined using both forced - and free -oscillation te st techniques.

Re su lts of the t es ts indicate the Apollo launch esc ape vehicle is statically stable,both power on and power off, ov er the no rm al angle-of-attack oper atin g ran ge near thetr im angle of attack. The effect of ro ck et thrusti ng is to reduce the stability, part icu -l a r ly at the hig hest va lu es of th ru st coefficient investigated. Also, the Apollo launchescape vehicl e generall y ha s positive damping ove r the normal angle-of-attack oper -ating range except for a very small angle range (*2 ) near the trim angle of attack atsubsonic speeds.

INTRODUCTION

The Apollo Sp aae cra ft Pr og ra m, with the eventual goal of a manned lunar landing,w a s initiated by the National Aeronautics and Space Administration (NASA) as p art ofthe continuing pro gram of sp ace expl oration following Pr oj ec t Mercu ry and the GeminiProg ram . Initial study contra cts, NASA Space Tas k Group stud ies, and other non-funded stud ies est abl ishe d the design requ ire me nts for the Apollo configuration, usingthe separ ab le module concept. Some of the initial wind-tunnel stud ies use d subsequentlyto support and verify this selection can be found in references 1 to 4 .

A s p a r t of the design and development pro gra m initiated in su ppo rt of the ApolloSpacecra ft Pro gra m, the Apollo wind-tunnel pro gra m w a s established. (See tabl e I for alist of test facilities and capabilities. ) The total program, discussed in more detai l inreference 5, w a s planned to yield design data on static and dynamic stability, aerody-namic heating, and aerodynami c loads; and the pro gra m was planned t o evaluate thor-oughly su ch specific problem s as interact ions between separating bodies during normalor ab ort operations, jet-plume interactions or effects, and launch vehicle compatibility.The stability characteristics of the Apollo command module (CM) are presented in refer-ence 6. The program had to provide the e xperim ental data necess ary f or efficient


14/134

1111111111111111111111111111111111111 Ill I I I

space craft design, as well as data fo r stud ies of all phase s of the flight prog rams . Oneprimary area of concern was the prelaunch o r at mosp heric portion of the flight whereany fai lur e would likely resu lt in a launch vehicle explosion. In the event of any mal-function requiring a pr em at ur e termin atio n of the flight, the Apollo Launch esc apevehicle (LEV) a s designed to remove the CM and its occupants a safe distance fromthe launch vehicle. The LEV comprises the CM plus the launch escape system, which

consi sts primari ly of the escape tower, the escape rocket, the jettison rocket, the pitchcontrol motor, and the can ard surfaces. As stated previously, the system is designedf o r us e during the prelaunch and/or atm osphe ric portion of the flight and is jettisonedwhen no longer needed, since other propulsion systems provide a means of escape inthe latter st ag es of flight. In or de r to design and fully evaluate the LEV, nvestigationswere made to determine the aerodynamic charac ter ist ics of the vehicle.

The purpose of this paper is to present the aerodynamic cha rac ter ist ics of theApollo LEV as dete rmined by wind-tunnel tests. Thrusting s tat ic stability data havebeen determined at Mach numbers from 0.5 to 6.0 over an angle-of-attack rangefrom -5" to 60". Nonthrusting s tati c stability data were obtained over the same angle-of-attack range at Mach numbers fro m 0 . 2 5 to 6.0. Dynamic stability studies weremade at Mach numbers from 0. 5 to 6.0, using forced-oscillation, limite d free-oscillation, and fr ee -to-tumble te st techniques to dete rmin e the damping-in-pitchderivative.

SYMBOLS

The force and moment coefficients presented in this paper a r e referenced aboutthe body system of axes as shown in figure 1.

nozzle exit ar eaAe

nozzle throat ar eaAt

cA axial-force coefficient,axial force

qa3s

axial-force coefficient at a, = 0"'A, a=O0

C pitching-moment coefficient computed about th eoreti cal apex,m, a pitching moment

q,Sd

pitching-moment coefficient computed about a nominal center of gravity,(x/d = -0.104, z/d = 0)

, c. g.

C slope of the pitching-moment coefficient measured at a, = 0"

2


15/134


16/134

d

(Yt r im

Y

P

u

Subscripts:

C

fs

j

m

T

00

time rate of change of an gl e of attack

t ri m angle of att ack

rat io of specific heats

m as s de nsity of air

freq uenc y of oscill ation

chamber

full scale

jet

model

total coefficients, includes aerodynamic and thrusti ng for ces

fre e-s tre am conditions

The aerodynamic coefficients presente d for the thrust ing tests, unless otherwise noted,include the interf eren ce or aerodynamic effects of the roc ke t exhaust, but with thethrusting components removed.

MODELS AND TEST TECHNIQUES

Tes t models varying in s ize from 0.045- to 0.105-s cale we re used in conductingthese te sts . The geomet ric simila rity of the model to the full-scale vehicle w a s main-tained except fo r the thrust ing models. In the cold-jet t hrus tin g model, a high-pressure air line was rou ted through the ce nte r of the tower to the esca pe motor andthe hot-jet model had tower legs which wer e slightly lar ge r than sc ale in ord er to ac-commodate the pro pell ant lines . A sket ch of the Apollo LEV is presented in figure 2,and photographs of t es t models mounted in test faciliti es and model components a r eincluded in figure 3.

The ba si c mod els of the CM did not sim ulat e the protuberances and cavities suchas antennas, umbil ical fairing, air vent, windows, and tower le g wel ls which we reincluded on the mode ls used in later tests. T es t s of the CM with th ese su rf ac e modifi-cations a re reported in mo re detail in reference 6.

Selective model-mounting techniques made testing possible through the completeangle-of-attack rang e fr om 0 o 360". Available balances w ere not r eadily adaptable tothe l arge angle ran ge s tested and in many c as es were not ideally suited for sp acec raft

4


17/134

testing; consequently, they w e r e selected for gross overall loading. Models w e r e de-signed for use in more than one facility to allow for the efficient and economical testingover the full range of Mach num bers and Reynolds num bers. A list of the models usedand ranges over which they were tested is included (table II).

Static Stability Te sts

Power off. - Ea rl y' te st s wer e made with 0.045- and 0.105-scale models of theApollo LEV to define the power-off static stability cha ract eris tics . Data are includedwhich were used to extend the available data to M = 6 . 0 for a configuration that isslightly different fro m the ba sic flight configuration. The alt ern ate configuration isidentical to the bas ic flight configuration except for the tubular bracing in the most for-ward se ction of the e scap e tower.

Power on, hot jet. - Tests, using a hydrogen peroxide gas generator and high-pressure cold air, were made to obtain the effects of r ock et thrusting on the stabilitycharacteristics of the LEV. A detailed descrip tion of the H202 ystem is given in

reference 7; therefor e, only cert ain pertinent details will be discuss ed here. Forsimulating thrust (hot jet), liquid H202 as brought through the sting-support sys teminto the model by two stainless steel propellant lines. Within the CM, the lines wereformed into concentric right- and left-hand helices around the strain-gage forcebalance. From the helical coils, the H 0 w a s fed into a torus-shaped plenum cham-

be r, and then through the four tower l eg s into the gas g ene rato r within the es cap e motorcasing (fig. 4) . The escape motor contained a radial flow decomposition chamber,which utilized a catalyst pack as shown in figure s 3 and 5. Silver scre en, coated withsamarium oxide, w a s cut into discs that w e r e compressed around a perforated distri-bution pipe between thin spa ce r disc s. The pack was enclosed in a 30-percent openperforated stainless -steel cylinder. This motor w a s designed to decompose 90-percenthydrogen peroxide at a maximum rate of 15 lb/ sec.

2 2

Pro duc ts of decomposition (super-h eated ste am and fre e oxygen) flowed along thepack to the re a r of the motor and out the four nozzles shown in detail in fig ure 5. Thenozzle s had divergence half-angles of 1 7.5". The two nozzles in the yaw plane and theupper nozzle in the pitch plane had equal throat areas and equal ex it- are a ra ti os of 10;the lower nozzle in the pitch plane had a larger throat area and a smaller exit-arearatio of 7.62. The asy mm etr ic th rus t of the nozzle s in the pitch plane provided an off-set thru st vector 2 O45').

A brief r6 sum6 of the da ta cycle is presen ted for clarity. When the escape rocketjets were not being simulated, the model balance syst em rec orded the aerodynamicch ara cte ris tic s of the Apollo LEV. With the roc ke ts thrusting , the balance sys temrecorded the rocket thrus t fo rces and moments as wel l as the aerodynamic for ces andmoments for the vehicle. These data are designated as total force and moment coeffi-cients. Static rocket thru st calibratio ns w e r e made with a shroud around the model toprovide the th rus t components. By removing the applicable components of the jet thr ust,the aerodynamic c oefficien ts are obtained. These coefficients rep res en t the aerody-namic ch ara cte ris tic s of the Apollo LEV in the prese nce of the free s t ream as alteredby the rocket exhaust plumes.

5


18/134

I IIIIlllllIIlllIllIIlI I I

Power on, cold jet. - Unheated compressed air was also used to simulate jetthru stin g (cold jet). The air sy st em of the model, as shown in figure 6, extended fromthe sting, through the CM, to the esc ap e mot or. Since this air syst em was physicallyisol ated fr om the balance, the CM, and the tower, the aerodynami c coefficientsmeas ured re pr es en t those ch ara cte ris ti cs of the Apollo LEV in the presen ce of thefree st re am alt ere d by the rocke t exhaust. The air, f ro m a 4000-psi supply, was

brought to the model support s tru ctu re and then ducted through two pipes, located oneit her si de of the balance, to a manifold in the CM. A sing le pipe, concentr ic withthe model cent erli ne, extended fr om the manifold, through the tower, to the nozzleadapter which contained the simulat ed LEV. In ord er to achieve an acceptable sim-ulation, these nozzles we re not constructed to scale except for the exit area, the exitlocation, and the ir ce nter lin e inclination to the model centerline . Thi s simulation w i l lbe discussed later.

Dynamic Stability Tests

Prov isio n was made, in the wind-tunnel prog ram , fo r extensiv e evaluation of thedynamic stability c ha ra ct er is ti cs of the Apollo LEV. Dynamic stability data wereacq ui red through the us e of the following test techniques:

(1) Forced oscillation

(2) Limited free oscillation

(3) Free to tumble

A detailed di scuss ion of the techniques and appa rat us used fo r measu ring dynamicstability parameters for a rigidly-forced oscillation system is found in ref ere nc e 8,and a method for reducing these aerodynamic char act eri sti cs to coefficient form isdescribed in reference 9. Limited free-oscill ation t es t techniques, apparatus, and

data reduction methods are discussed in reference 10.

The free-to-tumble technique is a method which allows statically balancedmodels, mounted on a tr an sv er se rod through the cen ter of gravity, to tumble freelythrough an angle -of-attack rang e of 360 . A photograph of a typical model installationfor thi s technique is shown in figure 3. Some prob lem s w ere encountered in designinga method of mount ing the model on a system where min imal friction and interferencewere required. A gas-bearing support, sim il ar to one used successfully in limitedfree-oscillation tests, failed becau se of galling under high-loading conditions. Amethod of mounting, using preci sio n ball bearin gs, w a s developed and proved satis-factory. It w a s determi ned that friction o r tare damping generally contributed a frac-tional part (less than Cm

Bench te st s we re made to determ ine the friction damping under load, and tare correc-tions were applied to the data.

+ C m = *O. 5) of the aerod ynam ic damping of the sys tem.q c

6


19/134

FACILITIES

The broad range of expect ed flight conditions (Mach numb er, Reynolds number,and angle of attack), and the limita tions of any single wind tunnel to sim ula te all theseconditions, dictated the utiliz ation of a number of test facilities.facilities used to acqui re sta ti c and dynamic stability data on the Apollo LEV, along

with tunnel size and capability, a r e listed in table I.

The wind-tunnel-test

TEST CONDITIONS AND ACCURACIES

Test C onditions

Table III lists the test conditions by facility . Complete data we re not available,which accounts for the absence of information in portions of this table.

Accuracies

Standard statistical analys is of balance calibration data and data repeatabilityindicated cert ain accuracy tole ranc es of the sta ti c forc e and moment coefficients. Thi sinformation, where available, has been compiled in table IV.

For the forced-oscillation dynamic testing, stru ctu ral damping values w ere ob-tained at vacuum conditions, before tunnel te st s w e r e made, to evaluate the still-airdamping contribution. Using known displ aceme nts and moments, tran sdu cer ca li br a-tion fact ors were obtained for the model displa cemen t and input torque. Resultinger ro rs in ei ther parameter were within +O. 75 and * l . 0 percent of the maximum valuesof the range in which eac h para me te r w a s calibrated. Considering the uncert ainties inthe system and the fact that the aerodynamic damping-in-pitch parameter is a function

of the difference between wind-on and vacuum conditions, the es tima ted maximum un-certainties in Cm + Cm were to. 50 for the Apollo LEV.q d

In the limi ted free-oscillat ion technique of dynamic testing, two so urc es of er r o rwere of pr im ary im portan ce in evaluating the accurac y of the data: the me as ure -ments from the angular displacem ent trans duc er and the determ inatio n of the taredamping of the ball-bearing support system.

+ CmmThe estimated maximum uncertainties in

were *O. 50 for the Apollo LEV.q &

Dynamic stability data obtained in the free-to-tumble technique cannot be said toThe accuracy of these data isave a certain accuracy at a particular angle of attack.

determined by the quality of the match between computer simulation and measuredwind-tunnel position-time histories. Themeasur ed and computed time-motion hist ori es indicate an excellent match except for aslight variat ion in the frequency of oscillation. Thi s frequency difference, however,is negligible when converted to full-scale values.

An indication of the matching is shown in figu re 7.

7


20/134

SIMULATION

Static Stability Tests

Close simu lation of ac tual flight conditions is nec ess ary to achieve usablefro m wind-tunnel tests.geo met ry, (2) Reynolds number, and (3) Mach number. If model geometry is similarto the prototype, if Reynolds number of model and prototy pe are matched, and if Machnumber of the tunnel airfl ow equals flight Mach number, good simulation is assumed.

Numerous fact ors m ust be cons idered including (1) model

Thrusting jets exhausting near the CM su rf ac e affect vehicle flight characteristicsin two ways:men t of the jet on the su rfa ce of the CM. The first affects the aerodynamic character-ist ics and stabilit y of the vehicle by al teri ng the free-stream flow around the CM. Thesecond incr eases surfa ce pr es su re s and temperatu res which in tu rn affect the stabilityof the vehicle . Clo se simula tion of jet thrust ing require s a similarity of parameterswhich affect the jet penetratin g into the fr ee -s tr ea m flow and the j et mixing with thefree st re am . Some of the fa ct or s which mu st be cons ide red include Reynolds number,Mach number, velocity, tem pera ture , density rati o, m a s s flow, r ati o of specific heats,and scaled jet thrust.

(1) blockage of the fr ee -s tr ea m flow by the jet plumes and (2) impinge-

tot-jet simulation. - The hot-jet model had scaled rocket nozzle throat areas Aand exit areas Ae and the sa m e expansion ra ti o Ae/At as the full-scale rocket

nozzles. The equation for the thru st coefficient is

where s, in k


21/134

boundary shape is, therefore, desirable. If the model has the sam e nozzle divergenceangle as the full-sc ale vehicle, the exit static press ure, along with y. and M deter-

mines the initial plume shape similarity. The jet shaping downstream of the nozzle isalso important if the jet effects are to be pro perl y defined. Matching of the mixingboundary shape may be possible with the simulati on of gas-con stant tot al-t em pera tur evalues for the full-sc ale rocket (ref. 8 ) . Although the decomposition pro duc ts of H 0

have gas-constant total-temperature values of less than 40 percent of the high-temperature full-scale vehicle, the downstream simulation is still considered accept-able. Compromises in nozzle geometry are necessa ry to obtain the required simulationof both the initi al and the downstream sha pes using air. The simulation is discussedbelow.

3 1'

2 2

Cold- jet simulation. - The cold- jet model used unheated h igh-pr essure air as thefluid for simulat ion of the launch es cap e rock et exhaust produc ts. Since the two fluids(gases) a r e considerably different in physical chara cter isti cs, a brief discussion ofthe analyses which supports the simulation technique is presented. As stated previ-ously, the simulation para me te rs which were determined to be the most important arejet momentum and jet plume shape. Initially, let us assume that the nozzle efficienciesof the full-scale vehicle and the model are equal. The ideal momentum rat io is defined

as the ideal thrust coefficient CT where

where Is the m as s flow rate. Thi s can be expanded to

in the tunnel and Pw , m and y 0, m a r e set equal to M m , f s and Y f since Mequals P . fo r an isentropically expanded jet, this equation redu ces to

9


22/134


23/134

frequency param ete r is considered. This paramet er represent s a ra ti o of the cha rac -ter i st ic length of the body to the wavelength of a disturbance. Since every point of abody disturbs the flow, the reduced frequency cha ra ct er iz es the mutual influence be-tween the motion at va rio us points on the body. Therefore, matching the reduced fre-quency pa ram et er i ns ure s the sim ila rity of the two syst em s when investigating unsteadyaerodynamics. Dynamic sim ila rity between the full -sca le vehicle and the te st wascorr ela ted by using the reduced frequency par am et er k.

=($)m = (%)fs

where w is the frequen cy of oscill ation , d is the ref ere nc e length, and V is thevelocity.

(G)m ( fs

where

and

I static operating temperature s a re equal, o r Tm = Tfs, then Vm = Vfs and theabove identity becomes

11


24/134

However, if the velocities are not equal, the identity becomes

Therefore, the reduced frequency parameter may be managed by varying thedynamic pr es su re o r changing the moment of iner ti a by ballasting. The approach useddepended on the conditions, which wer e determ ined by the wind-tunnel test used.

DISCUSSION

Presentation of Results

The results of this repor t are summarized in the following figures .(1) Figures 8 and 9 presen t the static longitudinal stab ility c hara cte rist ics of the

Apollo LEV.

(2) Figures 10 o 14 present the longitudinal stability c har act eri sti cs of the LEVin the pre sen ce of thrusting.

(a) Figure 10 presents the data from hot-jet simulation tests.

(b) Figure 1 1 presents the data from hot-jet simulation tests against the angleof attack at Mach numbers 0.5, 0.9, 1.1, and 1.3.

(c) Figure 12 presents the total aero dyn ami c ch ar ac te ri st ic s of the Apollo LEVa t Mach number 0.9 plotted ag ains t the angle of attack.

(d) Figures 13 and 14 present the results of the co ld-jet method of simulation.Figure 13 presents the aerodynamic coefficient data plotted against the thrust coeffi-cient, and figure 14 pres ents the aerodynamic coefficient data plotted agains t the angleof attack at Mach numbers 1.98, 2.99, 3.99, and 5.97.

(3) Figure 15 presents a comparison of data obtained by the hot- and cold-jetmethods of simulation at Mach number 0.7.

(4) Figure 16 presen ts the dynamic stability characteristics determined by aforced-oscillation technique.

(5) Figure 17 presents the dynamic stability characteristics, as determined bythe limi ted free-osci llatio n technique at three ce nt er s of gravity.

(6) Figure 18 presents the dynamic stability characteristics, as determined bythe free-to-t umble technique, fro m Mach num bers 0.5 to 1.98.

12


25/134

(7) Figure 19 pre se nt s a comparison of pitching-moment ch ara ct eri sti cs obtainedwith sting-mounted and transver se-rod-m ounted models.

(8) Figure 20 presents summary data.

Static Stability Characteristics

Static stability cha ra ct er is ti cs of the Apollo LEV (nonthrusting) have been d et er-mined at Mach numbers from 0.25 to 1.35 at the Ames Unitary Plan and 12-Foot WindTunnels. Additional data ar ep re se nt ed to extend the data to Mach number 6.0; how-ever, the configuration tested had a slightly different tower bracing than the final con-figuration. Stati c stability data for the Apollo LEV with simu lated LEV rocket t hru sthave been determined at Mach numbers from 0.5 to 1.30 at the Langley l6 -Foot Tra n-s o n i c Tunnel using a H 0 gas generator to simul ate the rocket thrust. Data forM = 0.7 and M = 1.48 to 5.97 were determined at the Arnold Engin eering DevelopmentCorporation Tunnel A using high p re ss ur e air to simul ate the rocket thrust. The aer o-dynamic coefficients, un le ss otherwise noted, include the aerodynamic or i nter fere nceeffects of the thrus tin g rock et, but have had the components of the roc ke t th ru st re-moved or isola ted fro m the balance during testing. Detai ls of the test procedures forthe thrusting tests can be found in refer enc es 7, 15, and 16.

Static, nonthrusting. - The static stability data (nonthrusting), given in fig ur es 8and 9 and summ arized in. fi gu re 20, indicate that the gene ral varia tio n in the pitching-moment coefficient with the angle of a tt ack is only slightly changed with the in cr ea sein Mach number up to M = 1.55. The pitching-moment curve slope C de cr ea se s

with Mach number to near M = 0.90 and then inc rea ses above M = 0.90. The favo r-able incre ase in the overall stability at Mach numbers above M = 1.55 is attributedto a shock interaction that re sul ts in an i ncre ase in press ure on the lower surface of theCM at mode rate angles of at tack (a = 10" to 40'). Note the co mp ar iso n of the pitching-moment curve slo pes determine d fro m the nonthrusting and thrusting te st installations

(fig. 20). Th er e are som e diff eren ces in the pitching-moment curv e slope (power off)as determined from the nonthrusting and thrusti ng te st installations. Note that thepitching-moment coefficient va ria ti on with angle of at tac k is generally nonlinear andany slope para me te r pr ese nt ed would be indicative only over a limited angle range n earthe point of me asu rem en t. Data of the pitch ing-momen t coefficient plotted agai nst theangle of attack at M = 3.0 (figs. 9(d) and 14(d)) indicate th er e ar e only sm al l differ-ences in the data as determ ined by the two te st installations. The slope-p aramet er data(fig. 20) indicate the difference could be much gr ea te r. Th ere are also some differ-ences in model installation, the majo r one being the inc reas ed sting diame ter ne cess aryto accommodate the rock et exhaust syst em. The variati on of the normal-force coeffi-cient with the angle of attack (figs. 8, 9, and 20) in di cat es th at CN dec reas es slightly

with the Mach number up to M = 0.80. There are rather abrupt changes in thenormal-force cur ve slope in the t ransoni c region (M = 0.9 to 1.2). Except for a de-crease in C near M = 3.0, the normal-force curve slope is generally constant

through the sup ersonic speed range (M = 1 .2 to 6.0). The power-off norm al-f orcecurve slope from both nonthrusting and thrusti ng test installations com pare s favorably

2 2

ma

(Y

N a

13


26/134

except fo r so me differen ces indicated at M = 1.0, 3.0, and 6.0. The power-offforce coefficient data (figs. 8, 9, and 20) ind ica te an in cr ea se in CA, up to

M = 1.1 with a sl ight decrease in CA,above M = 1.1. The comp arison between the power-off and the nonthrusting andthrusting installations should again be noted.

-with furthe r i nc rea se i n the Mach number-

A compariso n of the static stability data for the alt erna te configuration is alsopresented in figu re 9 for Mach numbers 0.7 to 1.35. The se dat a are in very goodagreement with those for the basic configuration for the Mach number and the angle-of-attack rang e tested.

Static, thrusting. - The exhausting of jet st re am s near adjacent surfaces canappreciably alter flow patterns, surf ace pre ssu res and temperatures, and the aero-dynamic stability. Detailed stu die s were made to evaluate the effects of rocket exhauston the static stability ch arac teri sti cs of the Apollo LEV. Again, two diff eren t testtechniques were used to determine the jet effects. A comparison of the data obtainedby using the two techniques is presented in figure 15 for C - 0, 1.34, and 2.40. Onlysm al l vari ati ons i n the pitching-moment and norm al-f orce coefficients (both power offand power on) a re indicated by the mea sure d data. The axi al-f orce data comparefavorably at CT = 0, but are higher for the cold-jet data with the jets thrusting. ThefavorabIe comparison of the summ ary dat a for the hot-jet data a t M = 1.3 with thesummary data for the cold-jet data at M = 1.48 (fig. 20) should be noted, sin ce thecold-jet model w a s designed for the supersonic Mach number range (M = 1.48 to 5.97).

T -

A typical rep rese nta tio n of the total aerod ynam ic forc e and moment coefficientsmeasured at M = 0.9, using the hot-jet simu lation method, is presented in figure 12.The total pitching-moment co efficient C

becomes more positive with increase in thrust coefficient at both the lower and higherang les of att ack while it is slightly m ore negative in the angle-of-attack range be-tween 25" and 40". Th is same tre nd is indicated for other Mach numbers not presented

increases asere, but discus sed in refe renc e 7. The normal-force coefficient C

the thru st coefficient is increased at angle of attack = O o , but remains generally con-sta nt with in cr ea se in angle of attack. Jet-off values show positi ve axial-force coeffi-cient for all angles of attack tested as shown in fi gu re 12(c). With the jet on, thethrust overcomes the aerodynamic axial force and at the higher values of t hru st coeffi-cient results in a forw ard accel erat ion of the vehicle.

as shown on figure 12(a) generallym, a, T

N, T

The variation of force and moment coefficients w i t h th rus t coefficient for Machnumbers 0.5 to 1.30 is pres ent ed in figur e 10 and was obtained using the hot-jet tech-nique. Like data, obtained from tests using the high-pressure cold air system, orMach numbers 0.7 and 1.48 to 5.97

aregiven in figu re 13. The pitching-moment

coefficient varies irregularly with thrust coefficient as a result of the rocket-exhaustfree-stream flow-field interaction. At the lower Mach numbers (0.7 to 1.98) andangles of attack (0 o lo ) , small i rreg ular changes in pitching-moment coefficientoccurred with increase in thr ust coefficient. At angles of attack above about 20,moderate increases in thrust coefficient resulted in higher negative values of pitchingmoment. Further inc reas es in thrust coefficient generally resulted in a trend toward

14


27/134

more positive values of pitching moment.13(g). ) A t the supe rson ic Mach num ber s between 2.99 and 5.97, the effect of increas-ing rocket thru st was a gene ral inc rea se in pitching moment toward mo re positivevalues at all angles of attack as shown in fi gu res 13(j) and 13(m).

(See figs. lO(a), lO(d), lO(s), 13(d), and

The norma l-force coef fici ent vari ed irr eg ul arl y with Mach number and angle ofattack as the thrust coefficient w a s increase d. Generally, the overall trends weresmall and resulted in a dec reas e in normal-force coefficient with in crea se in thrus tcoefficient, although so me exce pti ons are evident at the higher ang les of attack. F orexample, at subsonic Mach numb ers as shown in fig ures l o b ) , lO(e), lob), and13(b), CN generally decreased at the lower angles of attack with a n inc rea se inthrust coefficient. At the higher an gle s of attack, CN showed a general increase asthe thru st coefficient was increased . Increasin g the th rus t coefficient generally ha slittle effect on CN in the tra nsonic and low superso nic Mach number range (up to

M = 1.98) as shown on fi gure s lO(k), lO(n), lO(q), 13(e), and 13(h). At the hig hersupersonic Mach num bers, M = 1.9 9 and above, increas ing the t hru st coefficient r e -sulted in a decrease of the normal-force coefficient at most an gles of a ttack as givenin fig ures 13(k) and 13(n).

The effect of roc ke t th ru st on the axial- force coefficients, pres ent ed in figures 10and 13, is i rregular. The data of f ig ur e 10 indicate t hat the effect of increasing thrustcoefficient is to increase axial-force coefficient at all angles of a ttack teste d at Machnumbers up to 1.3. Except for M = 0.7, the data of f ig ure 13 indicate a generally re-versing trend until M = 1.98 is reached. A t M = 1.9 8 and above, inc reas ing thethrust coefficient results in a dec rea se in axial-force coefficient at som e an gles ofattack. A t the highest Mach numbers tested (M = 4.99 and 5.97) there w a s a decreasein axial-force coefficient with roc ket thr us t at the initial th rus t coefficient t ested andan increase in axial-f orce coefficient a t the highest th rus t coefficient tested.increase is the r es ul t of jet plume impingement. These data are discussed additionallyin refer ence s 7, 15, and 16.

This

The effect of th rus t coefficient on the aerodynamic cha rac ter is ti cs with varia tionsin angle of attack is presented in figures 11 and 14. The pitching-moment coefficientdata vary nonlinearly with the angle of a tt ack at CT = 0. Thi s nonlinearity, as notedin refe renc e 17, r es ul ts fr om the wake of the tower ro ck et disturbin g the flow over theupper su rf ac e of the CM as the angle of attack is inc reas ed. In addition, at Mach num-b ers above 1.98, the bow shock and the CM surface shock in te rac t to produce higherpr es su re s on the lower su rfa ce of the CM result ing in an increase in the power-offstability. The angle- of-atta ck range of stability with jets off is about 40" for M = 0.50and decreases as Mach number is increased, until at M = 1.3, the vehicle is stableonly to about 11". At Mach numbers above M = 1.3, the vehicle tends to be stabl e ove ran inc reas ing angle-of-attack range, until at Mach num ber 5.97, the vehicle is stable upto an angle of at ta ck of 20". The effect of ro cke t exhau st on the pitching-moment coef-ficient vari es with th ru st coefficient and Mach number over the angle-of-attack ran getested. These data indicate that the basi c, o r jet-off, aerodyn amic s have the predomi-nant effect on the st abi lit y of the vehicle. At Mach nu mb ers up to M = 1.98, the effectOf the rocket thrust is small, although it does a lt er the pitching-moment coefficientvariati on with angl e of a tt ack by changing the d eg re e of s tab ili ty and the location of th eunstable angle-of-attack ran ge. At M = 2.99, the effect of rock et th ru st is to

15


28/134

significantly red uce the stab ili ty of the vehicle. At Mach nu mb ers above M = 2-99, heeffects of jet shieldin g and impingement reduce the stabi lity of the vehicle to near zero.

At Mach numb ers up to M = 1.98, the effect of i ncr eas in g rock et thru st on thevari atio n of the nor ma l for ce with angle of a ttac k was generally sm all except for smallirregular changes in slope at som e Mach numbers. (See figs. 11 and 1400). ) AtM = 2.99, the effect of rocket thrust is to reduce the norma l-for ce curve slope result-

ing primarily from the jet shieldin g on the C M and flow sep ara ti on on the rocke t in-duced by jet expansion. This effect becomes increasing ly mo re significant a t Machnumbers above M = 2.99 until a t M = 5.97 the normal-forc e curve slope is near zero.

Data, prese nted in figu res 1 1 and 14 fo r jet-off conditions, indicate that theaxia l-for ce coefficient in cre ase s slightly with the angle of attack up to 10'. At anglesof attack gre at er than lo', the axial-force coefficient decreases with increases in &eangle of atta ck except a t the highest super soni c Mach numbers. At Mach numbers from0.5 to 1.98, the effect of rocket thru st is to in cr ea se the axi al- for ce coefficient. AtMach numbers from 1.98 to 4.99, the effect of rocket thrust is to reduce generally themagnitude of the axi al-fo rce coefficient as a result of the jet exhaust shielding the CM.At M = 5.97, the rocket exhaust impinges on the CM surface and results in an increase

in the axial-force coefficient.The axia l-for ce coefficient CA, and the slope par am ete rs Cm and C N

(Y (Y

are presented in figure 20. Thes e data show the effect of rock et exhaust on theseparameters as a function of Mach number.

Dynamic Stability Characteristics

Dynamic stability cha rac ter ist ics , determine d by thr ee different test methods,

data of figure 16 (ref. 18) at Machfor ced oscillation, limite d free oscillation, and free to tumble, are presented infigures 16 to 18, respectively.

qnumbers 2.0 to 6.0 are fo r an of fs et center-of-grav ity location. For purposes ofcomparison, similar data at Mach numbers 2 . 0 to 6.0 are presented for a centerlinecenter-of-grav ity location. Th ese data were determined utilizing the forced-oscillationtest technique and indic ate that the vehicle generally has positive damping except forM = 6.0 fo r the off set center-of-gravity location where a highly irregular variation inthe damping par am ete r near the tri m angle of atta ck is indicated. Additional studiesusing a forced-oscillation technique are reported in references 19 to 23. The datafrom these studies a r e not presen ted h er e s inc e the configurations t este d differedslightly fr om the production Apollo LEV.

The cm + c m &

The vari atio n of the avera ge damping-in-pitch pa ra me te r + s deter-q m &mined from t es ts using the limi ted free-oscillation tes t technique is presented in fig-

u r e 17 for Mach numbers from 1.5 to 4.0 (ref. 10). Three different center-of-graviwlocations wer e investiga ted and, for som e conditions, data were taken for more thanone Reynolds number. The dat a of fi gur e 17(a) ndicate the vehicle generally haspositiv e damping fo r the osc il lat io n amp litu des shown. The low Reynolds number dataat Mach number 3.0 indicate a trend toward negative damping near the trim angle of

16


29/134

attack which would indicate a li mi t cycl e of about 3.5". The dat a of fig ure 17(b) arevery si mi la r to those of fi gure 17(a) with a trend toward generally negative dampingindicated with decreases in Reynolds number and oscillat ion amplitude. Fo r the alter-nate center of gravity (x/d = -0.316, z/d = 0.0442), the data of figure 17(c) indicate thevehicle has less positive damping than the vehicle with the bas ic ce nte r of g ravity(x/d = -0.208, z/d = 0.037). Th is reduced damping would indicate a limit-cycle oscil-lation of about 4.5" for M = 3.0. As previously noted, the low Reynolds number dat aindicate a tren d toward negative damping.

To further evaluate the damping-in-pitch chara cteri stic s, a free-to-tumble testtechnique was developed. The results of tests using t his technique are given in fig-ure 18. A compari son of static pitching-moment coefficient data obtained using thefree-to-tumble, transverse-rod test technique with data obtained using the conventionalsting-support technique is given in figure 19. The sting-support data were measuredat the Ames Unitary Pla n Wind Tunnel fo r alm ost identical Mach numbers. The dataindicate that there are som e differences in the pitching-moment coefficient nea r themaximum values of pitching-mo ment and that the pitching-moment cur ve slop e throughthe tri m angle of at tack is slightly different. The se difference s are attributable tomodel support interferenc e. Positio n-time h ist ori es were computed using both sets ofpitching-moment coefficient data. Com pari sons indicat ed the only noticeable differenceto be frequency of oscillation which is quite small when converted from model to fu l l -scal e frequencies. The pitching-moment coefficient data, measured with thetransverse-ro d support system, wer e used in the reduction of the dynamic stabilityparameter.

The variation of the dynamic stabil ity der iva tive with the angle of a ttack(fig. 18(a)) indicate s that th e vehicle ha s positive damping for the n orm al opera tingangle-of-attack range at the subsonic Mach numbers of 0.50 to 0.80. Note that verysmall values of negative damping exist at ze ro angle of a ttack. The vehicle al so showsan ar ea of unstabl e damping ne ar an angle of at tack @f 180". The dashed lin es indicatear ea s of insufficient data to evalu ate the damping and are extrapolations of existingdata. The data of fi gu re s 18(b) and 18(c) indicat e that the Apollo L E V has positivedamping over mos t of the angle-of-a ttack rang e test ed. At M = 1.59, there is a rangeof negative damping ne ar an ang le of attack of 180". At Mach number 1.9 8, th er e is asmall range of negative damping nea r the angles of at tack at 170" and 190" for two ofthe reduced frequ enci es tested . The moment of in er ti a values (figs. 18(b) and 18(c))indicate a variation in reduced freauencv. The higher the moment of inerti a. the lowerthe reduced frequency par am ete r. -Th e "damping par am ete r Cm + C u k a l ly i n-

a m&cre ase s with a dec rea se in reduced frequency.

CONCLUDING REMARKS

Wind-tunnel tests to determine the stability c harac teri stic s of the Apollo LEVwere made using sev era l wind-tunnel facilities. Static stability char acte rist ics, boththrusting and nonthrusting, w e r e determined at Mach numbers from 0.25 to 6.0.

17


30/134

Dynamic stability par am ete rs were determined using forced-oscillation, limited free-oscillation, and free-to-tum ble te st techniques at Mach numbers from 0.5 to 6.0.Analyses of the re su lt s indicate:

(1) The Apollo LEV s gene rall y sta ble , both power on and power off, over thenormal operating angle-of-attack range near trim angle of attack.

(2) The effect of thrusting rockets is generally destabilizing, especially at thehighest th rus t coefficients and Mach numbers tested.

(3) The Apollo LEV generally has positive damping over the normal operatingangle-of-attack range except for a very small angle range (+2 ) near the trim angle ofattack at subsonic speeds.

Manned Spacecr aft C enterNational Aeronautics and Space Administration

Houston, Texas, February 9, 1967914-50-10-03-72

.

18


31/134

REFERENCES.1. Morgan, James R. ; and Fournier, Roger H. : Static Longitudinal Chara cteri stic s

of a 0.07-Scale Model of a Propo sed Apollo Spacec raft at Mach Numbers of 1.57to 4.65. NASA TM X-603, 1961.

2. Pearson , Albin 0. : Wind-Tunnel Investigation of the S tati c Longitudinal Aerody-namic Ch ara cte ris ti cs of Models of Reentry and Atmospheric-Ab ort Configura-tions of a Proposed Apollo Spacecraft at Mach Numbers from 0.30 to 1.20.NASA TM X-604, 1961.

3. Pearson, Albin 0. : Wind-Tunnel Invest igation of the Sta tic Longitudinal Aerody-namic C harac teri stics of a Modified Model of a Proposed Apollo Atmosp heric-Abort Configuration at Mach Numbers from 0.30 to 1.20. NASA TM X-686,1962.

4. Fournier, Roger H. ; and Corlett, William A. : Aerodynamic Characteristics inPitch of Sev era l Models of the Apollo Abort Syst em from Mach 1. 57 to 2. 16.NASA TM X-910, 1964.

5. Moseley, William C . , Jr. ; nd Martino, Joseph C. : Apollo Wind Tunnel P ro gr amDevelopment of General Configurations. NASA TN D-3748, Dec. 1966.

6. Moseley, William C., Jr. ; Moore, Robert H., Jr. ; and Hughes, Jack E. : Stabil-i ty Ch arac te ri st ic s of the Apollo Command Module. NASA TN D-3890, 1967.

7. Runckel, Jack F.; Schmeer, Ja me s W.; and Pendergraft, Odis C., Jr. : StaticLongitudinal Aerodynamic Characteristics of a Powered Model of the ApolloLaunch Escape Vehicle fro m Mach Numb ers 0.50 to 1.30. NASA TM X-1215,March 1966.

8. Welsh, C. J.; Hance, Q. P.; and Ward, L. K.: A F orced Oscillation BalanceSystem fo r the von K arma n Facility 40- by 50-Inch Su perso nic Tunnel.AEDC-TN-61-63, May 1961.

9. Campbell, John P. ; Johnson, Joseph L., Jr.; and Hewes, Donald E. : Low-SpeedStudy of the Effec t of Fre quency on the Stability Der iv at iv es of Wings Osci llati ngin Yaw with Part ic ul ar Ref ere nce to High Angle-of-Attack Conditions. NACARM L55H05, 1955.

10. Hodapp, A. E. , Jr. : Free-Oscillation Dynamic Stability Tests of a 0.05-ScaleApollo Command Module and a 0.059-Scale Apollo Launch Escap e Vehicle atSupersonic Speeds. AEDC-TR-63-186, 1963.

11. Runckel, Ja ck F.; and Swihart , John M. : A Hydrogen Peroxide Hot Jet Simulatorfo r Wind-Tunnel Tes ts of Turboje t-Exi t Models. NASA Memo 1-10-59L, 1959.

12. Davis, N. S. , Jr. ; and McCormick, Ja me s C. : Design of Catalyst P ack s fo r theDecomposition of Hydrogen Peroxi de (Pr epr int ) 1246-60 Ame rican RocketSociety, July 1960.

19


32/134

13.

14.

15.

16.

17.

18.

19.

20.

21.

22.

23.

Huff, Ronald G. ; and Abdalla, Kaleel L. : Mixing Cha rac ter ist ics Downsbeam ofCore Re@on of High Tem perature h i s y m m e t ri c Jets Exhausting into Transonicand Sup ersoni c St ream s. NASA TM X-151, 1960.

Pindzola, M. :Nov. 1963.

Jet Simulation in Ground Tes t Facilities. Agardograph 79,

Jones, Je r ry H. ; and Hi lli ard , E. E. : Jet Effects Tests on the Apollo LaunchEscape Vehicle at Mach 1.5 through 6.0. AEDC-TDR-63-240, 1963.

Jones, Jerry H. : Jet Effects Te st s on the Apollo Launch Escape Vehicle atMach 1.5 through 4. AEDC-TDR-64-92, 1964.

Babcock, D. L. ; and Wi l t s e , P. D. : Motor Vehicle Interfaces in the ApolloLaunch Es cape Syst em. Pa per No. 65-152, Am. Inst . Aeron. Astronaut.,Feb. 1965.

Ward, L. K.; and Hodapp, A. E. , Jr.: Dynamic Stability Tests of a 0.059-ScaleApollo Launch E sca pe Vehicle Model at Mach Numbers 1.5 through 6,Ap ri l 1963. AEDC-TDR-65-63, 1965.

Boisseau, Peter C. : Low-Speed Static and Osci llat ory Stability Chara cteri stic sof a Model of the Apollo Laun ch-Escape Vehicle and Command Module. NASATM X-894, 1963.

Kilgore, Ro ber t A. ; and Averett, Benjamin T. : Wind-Tunnel Measurement s ofSome Dynamic Stability Ch arac te ri st ic s of 0.055-Scale Models of ProposedApollo Command Module and Launch-Escape Configurations at Mach Numbersfr om 2.40 to 4.65. NASA TM X-769, 1963.

Averett, Benjamin T. ; and Kilgore , Robe rt A. : Dynamic-Stability Characteris-ti cs of P rop ose d Apollo Configurations at Mach Numbers from 0.30 to 1.20.NASA TM X-912, 1964.

Averett, Benjamin T. ; and Wright, Bruce R. : Some Dynamic-Stability Charac-teristics of Models of Proposed Apollo Configurations at Mach Numbers from1.60 to 2.75. NASA TM X-971, 1964.

Averett, Benjamin T. : Dynamic-Stability Ch arac te ri st ic s in Pit ch of Models ofProposed Apollo Configurations at Mach Nu mb ers from 0.30 to 4.63. NASAT M X-1127, 1965.

20


33/134

TABLE I. - TEST FACILITIES AND CAPABILITIES

_ _

Size of Mach num bertest section rangeest facility

Reynolds numberrange, x

_ __

Ames Unitary Plan Wind Tunnel 8 X 7 t

11 x 11 f t

2.4 to 3.5(Ames- UPWT) 9 X 7 f t 1.5 to 2.6

0.7 to 1.4

Ames 12-Foot Pr es su re Tunnel 12 f t diameter 0.0 to 0.95(Ames 12-ft)

Arnold Engineering DevelopmentCenter, Tunnel A (AEDC-A)

Langley Research Center 16-ft

40 X 40 in. 1. 5 to 6.0

16 t (octagonal) 0. 5 to 1.3Transo nic Tunne l (LRC 16-ft)

21

0. 5 to 5 per f t1 to 7 p e r f t1 to 10 per f t

0. 5 to 9. 0 p erf t

0.3 to 9 pe r f t

1. 2 to 3. 7 pe r f l

Lewis Research CenterSupersonic Wind Tunnel(LeRC -SWT)

8 X 6 f t


34/134

NN

Scale

0 . 1 0 5

TABLE II. - MODELS AND TEST RANGES

Tes t typeeynolds number range, a range, degacilit y Mach number rangex

Static, non-

thrusting

Ames-UPWT . 7 to 3 . 4 3 . 4 to 5 . 2 - 3 to 55 Static, non-

Ames 1 2 - f t 0 . 2 5 and 0 . 5 0 5 . 4 2 t o 1 0 . 3 4 - 1 0 to 4 0

thrusting

Ames-UPWT . 5 to 1 . 3 5 3 . 5 to 6 . 8 - 1 6 to 35 Static, non-thrusting

Model

0 . 0 4 5

0 . 0 8 5

FS-2

3 . 5 to 0 . 4 0 Static, non-thrusting

Static, thrusting,(hot jet)

AEDC-A

LRC 16-ft

FS-3

FSJ-1

0 . 0 4 5S 5-3 static, thrusting,O t o 5 0 (cold jet)

AEDC-A ; 0.7 to 5 . 9 7 0 . 6 to 2 . 0 5

FD-3

FD-5

LeRC -SWT

FD-9

2 . 0 to 6 . 0 4 . 0 to 6 . 0 -5 to 15 Dynamic, forced

1 . 5 to 6 . 0 0 . 3 1 to 5.3 18 ' oscill ation Dynamic, limit ed

oscillation

amplitude free oscillation

0 . 5, 0. 7 , and 0.8

1 . 5 9 and 1 . 9 8

3 . 1 4 to 4 . 8 6

3 . 5 6 to 3 . 8 8

Free to tumble Dynamic


35/134

Facility

_ _

Ames 12-ft

x(a)

Ame s-UPWT

Stagnation Dynamicpressure, pressure,

PSfsia

Ames-UPWT(alternateconfiguration)

AEDC -A(alternateconfiguration)

LRC 16-ft

TABLE III. - TEST CONDITIONS

Mach number

0.25.25.50

0.5. 7

. 91.11.21.35

0.7. 9

1.11.21.351.552.02.43.03.4

4.05.06.0

0.5. 7. 9

1.01.1

1.2I . 3

R

5. 5210.34

5.42

6.875.2

4.323.853.573.57

5.224.343.853.703.603.493.433. 653.903.60

3.5. 7. 4

3 . 13.94.34.44.44.34.2

~

65.025.229.8

12.410.6

8.77.87.16.15.3

Stagnationtemperature,

OF

540540

540540540540

540540540540540540540540540540

691115

72

116152148

3 175 1867777783 5878907

aBased on model diameter.

23


36/134

Facility

AEDC-A

AEDC-A

AEDC-A

NAA-TWT

LeRC-SWT

TABLE III.

Mach number

0.71.481.982.993.994.995.97

2.03.04.06.06.0

1.52.02.03.03.04.04.0

~

_ _ - - - -

0.5

.7.8

1.591.98

TEST CONDITIONS - Concluded

R X(a)

2.051.421.421.701.29

.695

.659

6.04.05.0

4.84.7

3.25.65

3.565.26

. 55.07

.42

4.80

3.984.31

3.563.75

~~

psia

13.808.219.69

19.5625.1322.1531.9

32.536.874.0

2QO. 0195.8

. _ _

-

Stagnationpressure,

~

. .

Dynamicpressure,

PSf

493508505487269.5106.47 50

1675908.8779.5200456.2

868165.8927.5

1089.2102.1794.0

65.1- _ _

640

705838

11601295

-

_ _Stagnation

temperatureOF _ _

8689907 0969686

101108

98180102

Based on model diameter.

24


37/134


38/134


39/134

(a) Apollo LEV model installed in the 8- by 'I-foot test sectionof the Ames Unitary Pl an Wind Tunnel.

(b) Catalyst packs used in escape motor for hot-jetthrusting simulation.

Figure 3. - Photographs of test models and components.

27


40/134


41/134


42/134

Concentric right andleft hand helixes

(4 pipes supplyingpropellant to rocket)

LPropellant l inesfrom supply

Figure 4. - Model drawing showing propellant lines.

,37"41'

Gas f low passage

Se c t i o n 6-6Li q u i d f low passage

Figure 5 . - Sketch of ra di al flow decomp osition chamb er.30


43/134

A i r e x h a u sk t h r o u g h n o z z le sPipe conduct ing a i rto escipe motor

Air exhausts th rough nozz les

F l ex i b l e a i r l i n e\-Two s ta in less s tee l p ipes conduct ing a i r

to command module m an ifo ld

Figure 6. - Sketch showing cold-jet sim ulation technique. Air sy ste m is physicallyisolated from model.

320

a

a

0

0 4 I L2 L6 2 0 2 4 28 3.2 3.6 4 0 4 4 4 8 5.2 5 6

Tim. sa

Figure 7. - A com par ison of comp uter sim ulated and wind-tunnel position-timehist orie s (diverging fro m heat-shield-forward tr im position).

31


44/134

M

Cm ,a

M

M

cN

M

M

cA

M

-0.25

-0.5

= 0.25

-0.5

=O.25

= 0.5

. 2

0 0

0 0

-.2

. 4

0 0

0 0

-.

. 4

0 0

0 0

7 4

-10 -5 0 10 15 20 25 30 35

A n g l e o fattack,a. deg

Figure 8. - Static longitudinal stability characteristics for the Apollo launchescape vehicle at Mach numbers 0.2 5 and 0 .5 a s determined in the Ames12-Foot Tunnel.

32


45/134

ww

m,

I

M-0.5 0 0

M - k P 0 0

, a

M = L 1A 0

M - L 3 5 DO

Angle of attack, a, deg

(a) Pitching-moment coefficient.

Figure 9. - Static longitudinal stability char acte rist ics f or the Apollo launch escap e vehicle at Mach numbe rsfrom 0.5 to 6.0 as determined in the Ames UPWT and the AEDC-A Tunnel. (Flagged symbols denote dat afo r an alt ernate configuration. )


46/134

(b) Normal-force coefficient,

(c) Axial-force coefficient.

Figure 9. - Continued.

34


47/134

M ~2.0

M . 2 4

m, a

M = 3.0

M =3.4

0 0

A 0

0 0

-.

-.

(d) Pitching-moment coefficient.



48/134

WQ

2.0

1.0

M = 1.55 0

M.2.0 0

M s 2 . 4 0

M - 3 . 0 A 0

M.3.4 b 0

-1.0-5 0 5 10 15 20 25 30 35 40 45 50 55 60

A n g l e o fattack, a, deg

e) Nor mal - or c e co efficient .Figure 9. - Continued.


49/134


50/134

M

M

CN

M

-4 0

-5.0

-6.0

20

L O

0 0

0 0

0 0

-1.02 0 2 5 30 35 m 4 5 55-10 -5 0 5 10 15

Angle of attack. a deg

(h) Normal-force coefficient.

Angle of attack. a, deg

(i) Axial-force coefficient.

Figure 9. - Concluded.


51/134


52/134

. 2

a - o o 0 0

a - 11 0

a - 2 0 0 0

0-31' A 0

cN

a = 4 ) ' b O

a - 5 0 " o O

a - 61 00 .5 1.0 1.5 2.0 2.5

Thrust coefficient. C,

3 .0 3.5 4.0 4.5 5.0

(b) Variation of normal-force coefficient with thrust coefficientat M = 0.5.


40


53/134


54/134

I .1 I11 I I I

.05

a =O o 0

a = 11 0

a - 2 0 ' 0 0

cm,

a = 31 a

a.40 bO

a-50 DO

a = 61 0

-. 5

-. 00 . 4 .8 1.2 1.6 2.0 2.4 2.8 3.2

Thrust coefficient, CT

(d) Variation of pitching-moment coefficient with thrust coefficient at M = 0.7.


42


55/134


56/134

a -

a =

a =

a =

a =

a =

a =

0"

11

20"

31"

40"

50

61

1.0

.8

. 6

. 4

. 2

0 0

0 0

0 0

h O

n o

0 0

0 . 4 1.2 1 .6 2 .4 2 .8

Thru st coefficient, CT

(f) Variation of axia l-force coefficient with thrust coefficient atM = 0. 7.

Figure 10. - Continued.44


57/134

.05

a - 0 o 0

a = 11' o 0

a = 2 0 0 0

a = 31 A 0

%,a

a = & b 0

a = 61" 0

-.05

-.100 . 4 . a 1.2 1.6 2.0 2.4

Thrust coeff ic ien t ,CT

(g) Variation of pitching-moment coefficient w i th thrust coefficientat M = 0.9.


45


58/134

IIIIIIIIIII I II I I

a - 0 o

a = 11

a = 31 A

a = m ob

a = % " o

a = 61 o0 . 4 .a 1. : 1.6

Thrust coefficient, C

2.0

jInn

II:II;

I

I

II

2.4

h) Variation of normal-force coefficient withthrust coefficient at M = 0.9.

Figure 10.- Continued.

46


59/134

cA

1.2

1.0

.8

I

. 6

. 4

apollo 15 aerodynamic stability characteristics of the apollo launch escape vehicle by nasa

Documents