c3 - apps.dtic.mil · ype of report &period covered erodynamiccharacteristics of a 10 sharp c...
TRANSCRIPT
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NAVAL SHIP RESEARCH AND DEVELOPMENT CENTER • tBethesda, Md. 20084
*-No AERODYNAMIC CHARACTERISTICS OF A 110 SHARP CONE
"AT HYPERSONIC SPEEDS At,;" HIGH ANGLES OF ATTACK
by
z< •.ieorge S. PickSzSamuel E. Dawson' •Robert L. Walker
2
0
0-
APPROVED FOR PUBLIC RELEASE: DISTRIBUTIC I UNLIMITED
REPRINT FROM THE 9TH NAVY SYMPOSIUM ON . ROSALLISTICS.APPLIED PHYSICS LABORATORY. THE JOHNS HOPý !NS UNIVERSITY.
9-11 May 1972
U.0
AVIATION AND SURFACE EFFECTS DEPA ITMENTcc< RESEARCH AND DEVELOPMENT REP()RT
I
0
1! June 1975 "1 Ll• Report 4692
UU C3
The Naval Ship Research~ and Development Center Is.a U. S.Navy center for laboratoryIeffort directed at achieving improved see and air vehicles. It warn formed in March 1967 by
merging the David Taylor Model Basin at Caadtarock,,Mwyland with the Idarine EngineeringL~boratowy atAnnapolis, Metyland.
Naval Ship, Research and Development Center
Bethesda, Md. 20084
MAJOR NSRDC ORGANIZATIONAL COMPONENTS171...COMMANDER
*REPORT ORIGINATOR TECHNICAL DIRECTOI~
OFFICER-IN.CHARGE OFFICER-IN-CHARGEI[ C:AIDEROCK L ANNAPOLISI
SYSTEMSDEVELOPMENT
y DEPARTMENT
SHIP PERFORMANCEEFCT
15 DEPARTVENT
STRUCURESCOMPUTATIONDEPARMENTAND MATHEMATICS
1DEDEPARTMEN
jSHIP USTICS AUXILIARY SYSTEMSARMn~19 DEPARTMENT
"'MATERIALS5ETA
,..Q.EPARTMENT 28DPRMN
NDW-NSRDC 3960/43b (Rov. 3-12)
......... GPO 920-106
UNCLASSIFPIED
41u.MI~MiI BEFORE COMPLETING FORMSECURITY ~ ~ ~ ~ ~~ 2 CLSSFIATO OFTI6AE(~w aeC;ESSION N3.RECIPIENT'S CATALOG NUMBER
E n IS 5. YPE OF REPORT &PERIOD COVERED
( ERODYNAMICCHARACTERISTICS OF A 10 SHARP
C GOrgES ck ATIYESamuel Finaawlo RoerpLfwlk
I. ERFR~iG RGAIZAION PAEADS ANDES HIGH ANGLORMELMNTSROETTS"NFavTal Ship ReerhanEPOmetee ' RA&W R N T .N MBERSBeheda Marylan 2084 ,
Naval Ahir Systemsc Cmand Deve RpmentC) Jur
WsigoD.C. 20361 13. RMER OF PA3ES
2514. MONITORING AGENCY NAME 6 ADDRESS(If different from Controlling Office) 13. SECURITY CLASS. (of this report)
UNCLASSIFIED
I$&. DECL ASSI FICATION/ DOWNGRADING
III. DISTRIBUTION STATEMENT (of this Report)
APPROVED FOR PUBLIC RELEASE: DISTRIBUTION UNLIMITED
5. 17. DISTRIBUTION STATEMENT (of the abstract mnisteed in Block 20, If different from Report)
;.I5PresntMed t thOMN avy Symposium on Aero ball isticsv'A pplied Physics Laboratory,The Johns Hopkins University, 9--Il May 1972.
It. KEY WORDS (Continuil On rereree Old* U neoeeee07an d Identity by block number)(~Pf Hypersonic Flow
High Angle of Attack AerodynamicsFree Flight DataLift, Drag Charac.,:ristics
10. AGSTRACT ýConflnue an reverse aide It roes&7y and Identityp by block numbs?)Afe-flying instrule iot oe a ese nahp onic facility at M 5.30, 6.34
ýJand 9.94 and Re xl~to l.2x10 6/ft Aerodynamic lift, total drag and base dragcoefficients were obtained together with detailed base pressure distribution measuremenits'atangles of attack from 0 to 60'. The measured lift and drag coefficients showed good agree-ment with the modified Ncwtonian theory. The ceneli P istributonwsnal
the~ ~ 'nelný~(ontiasnuedrlindependent of teangle of attack from - I0 toC0 n 4'd rapidly o ees) ~A
D M173 :TO FINV1I BOEE- UNCLASSIFIED 2__A 4S/N012-04-601 Z.'- SECURITY CLASSFICATION Of THIS PAGS (Wh. 0Due 81411e44)H~
*"'?4A6 7'41 ~I k4j$
IYI
~ .4 -UNCLASSIFIEDI I(-LU qITY CLASSIFICATION Of THIS PAGE(W7,ef Date &Kt...E)
Block 20 (Continued) Ifor all test conditions beyo~nd a -3O. For similar conditions at angle of attack andReynolds number, an increase in' Mach number resulted in an increase of the base
- pressure ratio. The measurements indicated a highly complex base flow region, andtentative explanation of the observed results was offered.
I 1, 1)1
UNLASIIE
89yRI' CLII.1INO HS A 9Wo aaEt a
...... .
TABLE OF CONTENTS
Page
ABSTRACT.................. . . ..... . . ....... .. .. .. .. .. .....
NOMENCLATURE................. . . ......... .. .. .. .. .. ....
INTRODUCTION .. ..... .... .......... .............. ...... ..... 2
*BACKGROUND................. . . ... . . ..... .. .. .. .. .. ....TEST APPARATUS, EXPERIMENTAL TECHNIQUES,TEST CONDITIONS. .. ...... .......... .......... .......... ..... 4
WIND-TUNNEL FACILITY. .. ........ .......... .......... ..... 4MODEL DESCRIPTION. .. .. .......... .......... .............. 4INSTRUMENTATION AND DATA ACQUISITIONSYSTEM. .. .... .......... .......... .......... ............ 5
Pressure Sensors. .. .... ........... .......... ............ 5Antenna Geometry. .. .... .......... .......... ...........Drop Mechanism .. .. ...... .......... ............ ..... 5Signal Processing and Data Acquisition .. .. .......... ............ 6
PRELIMINARY SYSTEM STUDIES. .. .... .......... .............. 6Prcssure-Tinle Response .. .. .......... .......... ............ 6Transducer Interactions and Calibration .. .. ...... .......... ..... 6
TEST CONDITIONS .. .. .......... .......... ............ ..... 7
TEST PROGRAM. .. ........ .......... .......... .......... ..... 7
DATA REDUCTION. .. ...... .......... .......... .......... ..... 7REDUCTION OF HIGH-SPEED MOTION PICTURE DATA. .. ............ 7BASE PRESSURE AND FORCE REDUCTION PROGRAM .. .. ...... ..... 9
DISCUSSION OF RESULTS. .. ...... .......... .......... .......... 10BASE PRESSURE DISTRIBUTION. .. ........ .......... .......... 10AERODYNAMIC COEFFICIENTS AT TIHI MODEL BASE. .... .......... 12TOTAL DRAG AND LIFT COEFFICIENTS ... .. .. .. .. . ......... )MODEL TRAJECTORY RESULTS. .. ........ .......... .......... 13
CONCLUSIONS. .. .... .......... .......... ............ ........ 13
REFERENCES .. .... ............ ........ ............ .......... 13
0I
LIST OF FIGURES
Page
AN1 - Exploded View of Inscrumented 100 Half-Angle Cone BasePressure Model ......... ................ .... ..... .. 15
2 - Schematics of Model Interior ........ ..... ..... ..... 15
3 - instrumented Cone Model and Antenna Installed inHypersonic Tunnel ....... ........................ .... 15
4 - Terminal Release Velocity and Trigger Switch Locationas Functions of Model Drop Height . . ..... ... ................ 15
5 - Block Diagram of 100 Cone Model and its Instrumentation in theHypersonic Wind Tunnel ....... ............... ..... 15
6 - Sample Oscillograph Record ...... .................... .... 15-7 - Tie Response Characteristics of the Transducer System . . . . . . . . . 16
8 - Centerline Base Pressure Ratio as Function of Angle of Attackand Mach Number at Re = 1 x 106/ft ....... ..... ....... ..... 16
9 - Centerline Base Pressure Ratio as Function of Angle of Attackand Mach Number at Re = 6 x 1O5 /ft.16
4. I 10 - Centerline Base Pressure Ratio as Function of Angle of Attackat M =6.34 and Re= 3 x 105/ft ......... .................. 16
I1 - Base Pressure Distribution in the Vertical Meridian Plane atII1 M 6.34 ........ ..... ..... ..... ..... ..... ..... ..... 17
12 - Base Pressure Distribution in the Vertical Meridian Plane atM 5.30 ........ ..... ..... ..... ..... ..... ..... ..... 17
13 - Base Pressure Distribution in the Vcrtical Meridian Plane atM =9.94 ........ ..... ..... ..... ..... ..... ..... ..... 17
14- Base Pressure Distribution at M 6.34 and Re 6.2 x 105/ft .... ....... 17
15 - Base Pressure Distribution at M 5.30 and Re = 5,9 x 105/ft .... ....... 1716 - Base Pressure Distribution at M = 9.94 ,'sd Pe = 6.4 x 105 /ft ........... 17
17 - Base Pressure Coefficients as Functions of Angle of Attackand Mach Number at Re % 6 x 105 /ft ..... ................ .... 17
18 - Base Pressure Coefficients as Functions of Angle of Attackand Mach Number at Re 1 x 106 /ft ....... ................ .188
19 - Base Drag Coefficients as Functions of Angle of Attack andMach Number at Re • 6 x 105 /ft ........ .................. 18
20- Base Drag Coefficients as Functions of Angle cf Attack andMach Number at Re I x 106 /ft ........ ..... ..... ......... 19
ivF
Page
21 - Base Lift Coefficients as Functions of Angle of Attackand Mach Number at Re -t 6 x 105 /ft .................... ... 19
22- Base Lift Coefficients as Functions of Angle of Attackand Mach Number at Re I x 106/ft .................... ... 19
23 - Total Drag Coefficient as Function of Angle of Attackand Reynolds Number at M = 5.30 ........ ..... ..... ....... 19
24 - Totrd Drag Coefficient as Function of Angle of Attack2 and Reynolds Number at M = 6.34 .. .................. 19
25 - Total Drag Coefficient as Function of Angle of Attackand Reynolds Number at M = 9.94 ..... ................. .. 20
26 - Total Lift Coefficient as Function of Angle of Attack.and Reynolds Number at M = 5.30 ........ ..... ..... ....... 20
27 - Total Lift Coefficient as Function of Angle of Attack
and Reynolds Number atM = 6.34 ..... ........... ...... .20
28- Total Lift Coefficient as Function of Angle of Attackand Reynolds Number at M 9.94 ..... ................. .. 20
29 - Measured and Computed Modcl Trajectories at M = 5.30and Re = 5.9 x 10 /ft . . . . . . . . . . . . . . . . . . . . . . 21
30 - Measured and Computed Model Trajectories at M =6.34and Re=6.2x 10 /ft .......... ...................... 21
31 - Measured and Computed Model Trajectories at M = 9.94and Re= 11.4x 105 /ft ...... ...................... ... 21
Table 1- Test Program ........... ......................... 7 7
Lv
MW 04V
L -] -.. ,
j•- 1 .,9th Navy Symposium on Aeroballistics [.
AERODYNAMIC CHARACTERISTICS OF A 100 SHARP CONE ATHYPERSONIC SPEEDS AND HIGH ANGLES OF ATTACK
(Paper UNCLASSIFIED)
by
George S. Pick, Samuel E. Dpwson, and Robert L. WalkerAviation and Surface Effects Department
. Naval Ship Research and Development Center
.- "Bethesda, Maryland 20034
ABSTRACT. A free-flying instrumented C N Normal fý,rce coefficient" I model was tested in a hypersonic facility at
M = 5.30 6.34 and 9.94 and Re = 3 X 10 to C Base pressure coefficient1.2 X 10 /ft. Aerodynamic lift, total drag Band base drag coefficients were obtained to- D Base diameter, in.gether with detailed base pressure distribu-tion measurements at angles of attack from 0 F Total base force, lbs
to 600. The measured life and drag coeffi- a
cients showed good agreemenc with the modified 9 Gravitational acceleration, ft/sec
Newtonian theory. The centerline K Model geometry and m~iss functiondistribution was nearly independent of the L Model length, in.
angle of attack fro.n -10 to -30° but increasedrapidly for all test conditions beyond M Free-stream Macai number
S= -30Q. For similar conditions at angle ofattack and Reynolds number, an increase in M- Mach number 'ased on local propertiesMach number resoulted in an increase of the ps su s
base pressure ratio. The measurements indi- B
cated a highly complex base flow region, aid ' B sun
tentative explanation of the observed results B mounted model, psia~was offered.wPB Average base pressure, psij
171 OMECLAUREpA Local pressure, psia
a Horizontal accelera-ion component ofmodel center of gravity, ft/sec p Total pressure, psa
Vertical acceleration component of p, Free-stream static pressure, psiay model center of gravity, ft/sec•
q,• Free-stream dynamic pressure, psiaSB Equation (18)
CA Axial force coefficient r Radial d&.stance from the centerline
C Total drag coefficientRe Unit Reynolds number/ft
C Base drag c oefficient ReD Reynolds number based on the base
l Dra coefficient at opo 00 diameter
Do Re Reynolds number based on localC, Total lift coefficient properties
Lif coffiien du tomodl bseReL Reynolds number based on length
B force Transition Reynolds number at angio
................................................t,,.
- -.., - .: ." .
9th Navy Symposium on Aeroballistics
Sof attack •jetted into the flow field at predetermined
angles of attack and roll and to be releasedRe Transition Reynolds number at zeroRoe an tn with adequate vertical velocity to fly
t .. eol nub atzthrough to the fSo fection. a r me': : SB Base area, in \;
B
S Horizontal component of model center BACKGROUNDof gravity displacement measured fromthe model release position. in. Free-flight techniques to investigate
Vtthe aerodynamic characteristics of models inS Sy Vertical component of model center of hypersonic wind-tunnel facilities have been"gravity displacement measured from the in-use for the past 7-8 years. However,
9th Navy Symposium on Aeroballistics
into account. wake of a iW. half-angle cone with variousbluntness ratios at M = 8, 10, and 12.3 and
Practically all of the experimental a 0° with a special compression strutinformation reported in the literature deals system to minimize support interference.with axisymmetric geometries at zero angle of Their measurements showed both laminar andattack. The Todisco and Pallone5 compre- transitional flow conditions. The surveyed
•,•hensive summary of experimental work includes regions included the cone surface and basenear-wake data for a variety of vehicle con- areas, the cone shoulder boundary layer and
figurations, free-stream Mac't numbers, shock layers, the recirculation region, theReynolds numbers, wall temperature ratios, shear and mixing layers, and the inviscid
S. etc. supersonic weak-wake region. Axial and radial jeti". pressures and temperature profiles were takenA few investigators have obtained some with various probes. The base pressure
data in the base region of slender cones at results showed that a slight bluntness had no 'Aangles of attack up to 100. Schlesinger and effect but that a greater bluntness (Rn/Rb =Martellucci 6 tested a 100 half-angle cone at 0,15) increased the base pressure level above
4 Mw = 6.0 and at a 100 angle of attack. The values for the sharp leading edge. The studyfree-stream Reynolds number was sufficiently showed that pB/p decreased with increasinglarge so that fully turbulent flow was Reynolds numbers and increased with Machachieved both on the cone surface and in the numbers for a given Reynolds number. Thenear wake. Significant changes in the flow variation in radial pressure for tne basefield were observed when a configuration was was maximum at the centerline and decreasedstudied at angles of attack other than 00. toward the edges. As the Reynolds number
reieandhfoincreased, the nonuniformity decreased; atSchmidt and Crsieaie h lwturbulent conditions, the base pressure dils-
characteristics in the near wake '.f a 100 tribution was uniform.*•: •half-angle cone at M = 7.7 and at 100 angle 10
of attack for laminar flow. They obtained Ward and Choate0 used a free-flightradial vwriations of pitot and static pres- technique to measure base pressure and basesures at selected axial location in the near heat transfer cn a 100 half-angle cone atwake. Their measurements indicated that the M= 10 and 00 angle of attack while theangle of attack produced a region on the lee- shoulder boundary layer and base region"ward side of the cone surface wherein the conditions were laminar. The absolute valuesboundary layer which was originally laminar of the base pressure at Re. = 0.5 X 106 wereon the windward side became transitional or in the range of 0.004 psia ±0.001. Theturbulent at lower Reynolds numbers than ex- model base pressure ratio results showedpected from the axisynmnetric flow case. This pB/pm s= 0.58 at Rej 0.5 X 103 with a nearlyaffected the mixing processes and therefore linearly decline to 0.35 at Rey = 2 X 10s.the behavior of the local flow conditions inthe wake. In addition, it appeared that Cassanto et a111 correlated 0 angle of
vertical inviscid flow above the cone surface attack free-flight centerline base pressurecaused a large-scale mixing for the angle of data for both sharp and blunt 10 'vIlf-angle
attack configuration. Picot pressure pro- cones at the speed range of M N 4 tc M = 19.
files showed that the symmetry axis in the Both laminar and turbulent flow conditionsviscous core was displaced toward the leeward were tested and analyzed. The sharp coneside of the cone by about one-tenth of the data showed strong Reynolds number dependence
. base diameter. The stagnation pressure ratio whereas no such dependence was exhibited forincreased by a factor of five above the the blunt cone. The level of base pressurelaminar axisymmetric flow conditions, for the blunt configuration was higher than
Mreccfor the sharp cone for comparable free-streamMartellucci and Schlesinger 8 utilized a conditions of both the laminar and turbulent
Wi 100 half-angle cone at 0° angle of attack and flows. Laminar flow centerline base pressureTM = 12.3 to obtain laminar near-wake and correlation exhibited a maximum dlue where
base pressure data. They employed a special pB/pt was a single-valued functLion ofsupport system to minimize support inter- M2 • Re This function appeared t.) relateference and measured centerline base pressure, Zdividing velocity line, and static and total much of the data in the test program. Base
dividing~ veoiylnadsaih an oa pressure correlation for turbulent flow con-pressure distributions together with static dtesshoe rio suc iumbutand total temperature distributions in the detions showed .o Such maximum buw pB/peinear wake. The results showed that decreased with increasing 14V
B OD ±Much of the published experimental in-9 formation concerning base flow has been
Martellucci and Ranlet examined the near clouded by the uncertainty introduced by
i -i
9th Navy Symposium on Aeroballistics
model suppcrt interference. The sting model attack increased; at c = 40$, pB/PB / 1.75.support system, in common usage in wind-tunnel For the free-flight model at Mach number 9.9,measurements, is bound to distort the flow the magnitude of the measured base pressurefield to some extent; consequently, the re- ratios at corresponding flow conditions andliability of the resulting data might be a.ngles of attack was about 70% above the sting-questionable. mouifted model. Beyond about 40°, no steady
base pressure value could be reached with theConsiderable experimental work on the sting-mounted model in either of the tested
"problem of support interference in supersonic Mach numbers because of the severe effects of
speeds has been done at 00 angle of attack. sting interference.It has confirmed that the base pressure isstror.gly influenced by the support inter-ference and can serve as the first indication TEST APPARATUS, EXPERIMENTAL TECHNIQUES,of flow distortion caused by, the presence of .EST CONDITIONS
•i•- model support. 12-21WIND-TUNNEL FACILITY
On the basis of some data, 1 5 1 8 it A
appears that for laminar flow in the range of All experiments were conducted in theM- 1.5 to 5.0, there are critical ratios for NAVSHIPRANDCEN hypersonic tunnel, a facilityboth length and diameter. However, whereas equipped with a series of axisymmetric con-the sting length/model diameter ratio increased toured nozzles with exit velocities rangingfrom 3.0 to 6.0 between M = 1.5 and 5.0, the from M = 5 to M = 10.0 and an open-jet testcritical sting diameter/model diameter ratio section, The outside diametet of the jet is Ahad a maximum value of 0.5 at M P 4.0, and de- approximately 13.5 in. with a uniform (flow,creased with both increasing and decreasing uniformity ± 1.75%) core of approximatelyMach number. Whitfield20 states that the sting 6 in. The unit Reynolds number can be varieddiameter effects may be important when attempt- from 8 X( 10 td10V/ft by varying the supply
K. ing to correlate data with free-flight results. pressure over the range from 15 to 600 psiaand varying the supply air temperature from
Peckham2 2 conducted a qualitative explora- ambient to 25000F. A more detailed descrip-tory study at M = 6.8 where transition occurred tion is available in Reference 24.upstream of the model base so that a turbulentwake was formed. He found that at a = 200,the flow pattern on a delta wing model was not E SIaffected by sting diameter in the range of The 100 half-angle cone models were0.4 < d/D < 0.6. designed to be injected into the flow field
and released to fall freely through the hyper-In their experimental study with a sharp56 thalf-anger conrenataM stu14,Steo and ssonic test section. They were 6 in. long withi95.60 half-angle cone at M =14, Stetson and
a nose radius of 0.003 in. and a base diametercomunicatfound etwee there wase no subsonc of approximately 2 in. The skin was machinedcommunication between the base and leeward from corrosion-reslstant steel that was polish-regions and, consequently, little interaction ed and chrome plated to provide abrasion re-up to an angle of attack of approximately sistanc. These models were capable ofthree times the cone half-angle. However, Smeasuring base pressure distri bution at vari-heyond this incidence angle, the leesideminimum pressure region was reduced to base ous roll angles, angles of attack, andminium ressre egin wa reucedto ase Reynolds and Mach numbers. Figure 1 showsa
pressure level and thus there was subsonican exploded view of the cone model, and Fig. 2
communication between these two flow fields. is the schematic drawing of the interior.This means that at large angles of attack,any type of mechanical support system will A conical brass weight was attached todisturb the flow f'eld upstream as well as
the forward dnterlor to provide for grosscenter of gray y adjustment. Threaded slugs
r n within the weight were used for fine trim..Pick2 3 Four differential pressure transducer telem-effects on meesured base pressures at H = 6.3and 9.9 for a unit Re: holds number of eters were housed in a split, filled Teflon
: insert; the power supply leads were ý,'oughtforward and connected in parallel to three
free-flight half-angle sharp cone models wereuse bewee ~ 0 nd . ,*~surrnets mercury cells. The battery pack was clampedused between a- 0 and a - 4no. Measurements
on the front half of the insert. The trans-showed that the sting-interference effectsw o s e6 w1 ducer measuring ports were connected through
i, thick-walled flexible plastic tubing to brass(I < p/ .2), but that the flow became nipples on the base orifice plate. A softprogressively more distorted as the angle of rubber plug was mounted in the Micarta base
4
_10
,, 9th Navy Symposium on Aeroballistics
plate, and in conjunction wit'l a vacuum• pump arrived at by trial and error. After an ex-and hypodermic needle, served as a self-seal- tensive pretrial period during which several :.,Ing valve for evacuation of the model in- antenna geometries were tested, satisfac;ytearor (transducer reference pressuze volume), performance was achieved by using two elet..
threaded bezel provided clamping force for trically parallel fclded dipoles with thethe 0-ring seal located between the orifice elements arranged approximately in the sL:-' ,:plate and model skin. of a parallelepiped.
By using this antenna geometry, sigi --L-•i
INSTRUMENTATION AND DATA ACQUISITION SYSTEM s y usi s antenna ugeom ery , .i.i'istrength versus roll angle studies were ir.Afor three (x,z) coordinate points wi-'"- .2
P expected flight envelope at a = 0, - , and-60 ,In general, signal strength :7.iues
As previously mentioned, the free-f U*ght I eeasgasr"imoeI~ws~evius~eulpe wlhmnt~nd'ourdifeeniathefre-liht decreased below an acceptable !•:-1only over •.._.•
model was equipped with four differentialp r s t t kisolated areas amounting to 5 t.: ;.0 of roll
These units were based on a Harrison design2 5 angle each. It was further fou•. that thetransducer center frequency was independent
• j . and were fabricated by JPL. The telemeters of pitch and roll angles. Fiure 3 depictsutilized ultrastable Colpitts oscillators the final version of the antenna system, thewhich were operated at center frequencies drop mechanism, and the model installed inbetween 106 and 137 MHz. The oscillators tehprnc unlJ. the hypersonic tunnel. :-. . :were frequency modulated by the variablecapacitance-type pressure transducers. The ,ropMechanismtelemetry assembly consisted of four major Drop Me
components: a printed circuit inductor, a Prior to its injection into the hyper-variable capacitance-type differential sonic jet stream, the mouel was guided by apressure transducer, a battery package, and specially constructed drop mechanism thata microminiatire package of pellet-type con- held it at a pructede d pitch angle andt
struction for all other circuit components. hl ta rdtrie ic nl n
The transducer sensing element was a 0.00025- gimin.-thick prestressed metal diaphragm solder- was completely submerged into the inviscidcore of the flow, the restraining arms opened <•ed over the reference port case and separated corelof the f ow, thes i a ms opne
elecricaly rom he ressre ort. Theand released the model smoothly without im-I electrically from the pressure port. The-parting any side force or yawing moment. The
stress level of the diaphragm determined the vertical speed of the model was determinedsensitivity of the uni.t. The diaphragm and by the release height of the carriage. Fourthe inner surface of the pressure port formed ttension springs ensured that the restrainingthe variable parallel plate capacitor. The 'owivarious components of the unit were rigidly arms moved out of the flow field within 10encapsulated nto msec after model release so that they didthermal isotion e y potn providne, not disturb the flow around it. The drop
temlisolation and sho(. resistance.sequence had been so automatized that oncethe circuitry was energizeo, a multicam timer
The rise time of the pressure telemeters took over. The drop mechanism was calibratedwas about 0.5 msec without tubing, and their to determine the correct location of the 4thermal stability was good. Harrison found trigger switch so that the model was releasedthat a temperature increasi of 100 F at at-mospheric pressure produced a 2.7" increase at 0.25 (± 0.1) in. above the point where the
g in the oscillator frequency and a 0.11% chanje carriage was mechanically restrained. This.in the full-scale sensltivi' calibration, ensured that the model was released smoothly
The temperature response time was aboeat 3 min. and very nenrly at the same posion in thestream. Calibration data showe-d that spring-
performance over a temperature rangp from 30 mass system of the restraining arms andor attached coil springs had a natural frequencyto 200 F. Details on the construction of of about 20 msec/cycle and the motion was
these units and their performance character- dam out w0thin 3cyclea h.damped out within 3 cycles,.
istics are given in the Harrison report.25
After model release into the stream,Antenna Geometry the carr[i.e impacted ou a rubber bumper
Tfewhich causv,, it to bounce upward about 3 in.The signals from the telemetry units Tentrifeunyo hsmto aThe natural frequency of this motion was
were intercepted by a complex antenna system such that the carriage and the release armsthat. surrounded but was completely outsidethe iypersonic jet flow. Since the system were completely removed fromtheoyperad ier fi nmsec after release and only 200 msec later
opertedin nar ieldin he cnries o a well after the model had left the stream)steel housiag, which contained the open jet, (wed iter ah i fod he secn le ofuid it return again for the second cycle of 0:the optimum antenna geometry had to be its damped oscillation.
s d "scllai'n
; :i :;' •a • •.. " - -• . .... ..... ... .. "'-:".. " .. • ' : . •'•:(..:• A
9th Navy Symposi,;i on Aeroballistics _-_....___
A series of drops at varying carriage digt) r.• and reduced to useful Zorw by data ""-heights (between 0 and 20 in. above the flow rediction procedures detailed later.
centerline), corresponding to a range of"terminal release velocities, provided data During the drop, the tunnel cab pressurefor the determination of (a) terminal re- was continuously monitored by a Datametricslease velocities as a function of drop height Type 1014 Electronic Manometer and a-Typean' (L) trigger switch location as a function 511-3 Barocel capable of measuring pressures 4
of drop height; see Fig. 4. between 0 and 1 psi on seven consecutive
scales ranging from 0.001 to 1.0 psi full
"Signal Processing and Data Acquisition scale. The instrument had been calibrated
Systems and certified to an accuracy of ± 0.1% oneach scale by the Bureau of Standards. The
""As mentioned earlier, the signals from accurscy was very important since this measure-the telemetry oscillator were beamed to a ment provided the reference pressure vilues to
. complex antenna system. This antenna system which. the base pressure signals were Telated
was matched to a coaxial -cable by using a in order to determine their absolute levels.I Blonder-Tongue Model MT-283 matching trans-
* former and a signal splitter. Two prearnpli- Since the data acquisition aed processingfiers (Jerrold Model '440) were then used in system was rather complex an,' hjmi to be syn-
v. : - - cascade to obtain at least a 40 dB gain for chronized with the model carriage release anddriving the signal tapoffs (Jerrold Model solenoid release signals, a timer mechanism
UT-22W).. Four tapoffs wer3 used to. couple took over all functions as soon as wind-tunnel"the input signals to the four FM receivers. free-sLream conditions were- established and a"The resulting. isolation was at lease 31 dB timer starter switch was manually activated;-and feed-through losses were less tbhan 0.25 see -ig. 5.dB for the frequency range from 100 to
" .'., •140 MHz.14.M}z- PRELIMINARY SYSTEM STUDIES
The RF and oscillator stages -f thetuners (McIntosh Model Type MR71) were modi- Pressure-Time Response
fled to provide reception in the bani from
100 to 140 tA{z. An a-c vnltage regulator The available ti~me for physical measure-(Sorensen Model ACRIO00) was used to main- mentE is very limited for all free-flighttain tuner line voltages within 1 0.1%. model experiments; it is on the order of 10
Tuner outputs were then fed into four amplex- to 50 msec and seldom exceeds 0.1 sec.er amplifiers (Beckman Model C-44) which in Trajectory predictions showed an availabletur.i. drove the 600-Hz Type 7-323 galvanom- flight time of about 65 msec in the hyper-
eters in a recording oscillograph (C.E.C. sonic stream. As a final check on the in-
• - Type 5-124). Figure 5 is a block diagram strument design, pressure response data wer.'
of the signal processing network, obtained in che expected pressure range forshort connecting-tube transducer systems and
Ten channels recorded the iLncoming data correlated with some analytical formulas2 6
.on the oscillograph. Four processed the Figure 7 shows the key results for the design
pressure signals, four measured the signal confz...jration. Measurements indicated that
strength of each transducer channel network, the time response within the expected pressure!r.u recorded the solenoid trigger release range did not exceed 5 msec.
signal, and one recorded the output signal
of a 1000-Hz reference timer. Figure 6 Transducer Interactions and Calibrationshows a iample cf the oscillograph record.
Pick and Dawson2 4
systematically studiedThe motion of the model was recorded by the effects of close proximity on the simul-
two high-speed motion picture cameras: (1) taneously operation of transducer telemetev
"a Photosonics camera recorded the yawing and uthits. They evaluated such parameters asrolling mocions of the models at the rate of interaction as a function of transducer-to-1000 frames/sec and (2) a Hycam model (in- transducer distances, center frequency spread.;Lall.ed at right angle to the viewing window and its effect on tie interaction, change of
of the wind tu,ýnel) recorded the pitching sensitivity in relf.tion to center frequency,motions of the cone at about 2000 frames/sec. center frequency F,hifts as functions ofThe output of the 1000-Hz reference timer, transducer-to-transducer distances, and the
. rocord:id on the oscillograph, was &lso re- effects of com'son power supply and shielding.corded on film so that the film and oscillo-graph data could be synchrunized. BoLh the The results showed that the interactionfilm and oscillograph daza were manually increase was approximately proportional to
6
"_ __ 9th Navy Symposium on Aeroballistics
the inverse 0.65 power of the edge-to-edge estinaate, to be within 5%.distance and that the center frequei)cy spreadhad the most significant effect aa the inter- TEST PROGRAMaction.
Table I outlines the test program andSThe frequent interaction calibration the main parameters investigated. For each
measurements showed that ý.he interaction model dIrop, four data points were obtainederrors did not exceed ± 1.3% for press,-res at three different radii of the base area and
up to 3.0 mm1 Hg. In most C-ases, the irLter- at centerline. For a given flow condition andaction error was within ±0.5%. The instru- angie of attack, a series of 12 drops at 30
. mented model was recalibrated daily with roll-angle increments provided good defi•i.tionknown pressure inputs. Corrections to com- of the complete base pressure distribution.pensate for interaction errors were contained More than 450 data drops were conducted atin the calibration. The daily calibration seven separate flow conditions, And more Lhanresults showed that all the transducers were 90% of them were satisfactory.
C linear withir ± 1.5% and day-to-day repeat-ability was good. .TBES- TABLE 1
TEST CONDITIONS . Ts rga
Thorough surveys of Mach number distri- sbutILn and temperatnre distribution were Re/ft •0.undertaken for each test speed. The result3showed that in the jet core (11 in. horizon- 6.34 ±0 . il 6.2 ± 0.5 0 - 60tat icngth, 6 in. vertical height. and. 4 in.width) where the model trajectories were 6.34 ± 0.11 11.2 ± 0.7 0 - 54
located, the Mach number values were defined 6.34 ± 0.11 3.0 ± 0.2 0 " 80-as M -. 30 1 0.11, 6.34 ± 0.11, and 9.94 9.± '9.94 ± 0. 6.4 ± 0.3 0 80
•-- •9.94 ±L 0.11 11.4 :k 0.6 0 -" 70
Measurements indicated that steady flow 5.94±0.11 1.4±0.6 0-50)was established within 5 sec after startup.To ensure steady conditions and provide for 5.30 ± 0.11 10.7 L 0.7 0 -45time lag in the cab pressure sensor system,all model drops were initiated at Least 15sec after tune. start.
I' DATA REDUCTIONThe settling chamber .emperature, up-stream ot the nozzle throat, was continuously Two digital computer programs were
monitored during runs by using a Chromel- developed to reduce data in useful forms. The"•Alumel thermocouple. Prliminary measure- first converted film data to model pitch angle;m"dents establised th temperature drop in the trajectory and aerodynamic coefficient (CL andnozzle between the monitoring thermocouple!•:. ;CDý information as functions of time. The.and the test section for each stagnation second converted information from the oscillo-"condition. This information was used in graph records to base pressurv ratios pB/'pcomputing the tree-stream Reynolds number. an as funcions of time
The temperature differential between thesettling chamber and test section rangedbetween 30 and 80 0F depending on time and Reduction of High-Speed Motion Picture Date
•I• • stagnat ion pressure.
V Tie aodel motion and trajectory informa-Variations in actual unit Reynolds tiott ftr each run, obtained by the Hycam
numbers from run to run due to particular camera, was converted to useful form. Threetunnel conditions were withii, + 8% of the poinis on the cone were digitized by ujing theaverage values in all tested cases (the apex and two selected points on the windward
W -average unit Rsy;ý.aIds number values were 6.2, and lee•,ard generaturs. The horizontal and6.4, 6.9, 3.0, Il 2, 11.4, 10.7 X lOb/ft) and vertical locations of each of these pointscontributed little to the overall inaccuracies were defined relative to a fixed fiducialoi the data. coordinate system built into the camera. This
information was recorded on punched cards for
baseIn all, the total error in the measured every fifth or tunth frame (a time interval ot•'•: i • base ir' .. ata diet n•ruealn about 2.5 to 5,0 reset). The digitized frame -
interaction, &'6 time response errors was position wah defined in a time coordinate
-. 7•:
LEE I" "'! '. ! ""-
A Jaý
9th Navy Symposium on Ainrobalfis'tics o
system using timing marks which were recorded SubstitutIng Equations (5) and (6) intc.on the edges of the film. In this way, thma Equations (3) and (4) yields:position of the model could be defined with =*
respect to aisy event (Such as the opening of, ,y (CW B~.the releage arm) and thus the measured base -
-pressure data could be correlated with the and .
model trajectory at any instant.= g* (W t 2 i + S.t(a
The pirogram converted he raw digitized .frame image:.:tc full-scale. "o rdiznat.as and V
~2§ aplid crrctinsforopt~. tes istr-According to the modified Newtoniantions. With the cone converted to a 4u1 tertenra n xa oc oftsize, . he. center of pressure lo.zation and et ofa on a aya l fatc cwangle of attack were then determined fromb epese s
o-the digitized point locations using the ~F+-z
~~~P~~ 3'" c+ -cosO, (cotatanOc + 2tan~ 9
Jbut in the present experimentad
'~.1X =X .(2) [2s-Clsd (1cg cp CA L nc s .
This procedure was repeated for each 3sin e,) j + f2-. cosý i~ l 2 ~ (0digitized frame in a particular drop and the Jinformation was stored in the memory bau.k of whethle comouter for further use. To compensate ~hn~
7"for snail human errors in the manual process si (1of digitization, a second order smoothingu\ nafunction was applied at this point to thestored center of pressure and angle of attack whendata as functions of time. The availableinformation was sufficient for the calcula- a> ction of the model. terminal velocity (Vu) be-fore release, and thi's was uised as initialanvertical trajectory velocity i~nput for compu- andtation of the equationa of motion. Tu - (12)
Two-dimensional curvilinear equations ofmotion were utilized to describe the trajec- wetory of the center Ir 'ravity. These equa- a-tions are expressed in rectilinear componentform as:.1~Finally, the total drag and lift coeffi-
x 2 .& t (3) cients in terms of normnal and axial forceX 2 Xcoefficients are givcn as:
and~and a 2 +y (4) C0 CN $in o + CA Cos a- (13)
C L CN Cos 01 CA asino'14The horizontal component of the accelerationin the present investigation -is given as:-
In the computer program, Equations (7)9 C~~~
a~ gC% 8 (5) to (14) were combined in functional forms'as-
S, f(K4B,C&)L (15)and the vertical component of Lthe arcolerationis: and
(W C~~Sg h(K,B,a-)t4 + Vtr. (16) I --
8
*-5- -- - - - ---. .
-' . . -
-- - - - - - -- - - - - - - -- - - - - - -_mg* -'.-
where nubr Mach nube an Thfe en rsue
and the end of the run. A linear interpolation
- Equations (15) t (1)wr tenused to The total force on the model base is aobtain theoretical trajectory information function of the base pressure distribution.f'rom the measured time and angle of attack Generally, this is in the form:data based on point-to-point predictor-corrector techniques. Total drag and liftcoefficients were also-obtained using theR Tmeasur-'d trajectory, temnlvelocity, andr
time information based on the film data in FB dJ (9coajuuction with Equations (7) and (8).
0 0The final output provided angle of attack
and trajectory results as functions of timeand information on how they compared withTofcltenuria cmpainstheory together with measured and ýheoretical Equation (19) was approximatod by:aerodynamic coefficients as functions of time.Prt..:±siol- were also made for plotting all the(da in graphicAl form. 0.41 B p'm.~
Base-Pressure and Force Reduction Program 0.39)
Four oscillograph trace3 were recorded + 0.783 ( /" t.for each drop, one foy :-ach transdorer, A B P. , --
sample output is shown in Fig. C.. Themeasuý -d cab reference p'reitsure and the free- + 2.29 * (20)
4tream total pressure and Mach number values ( 7 mvere noted fo~ Lach zurt togethsrr with a start--ing titre corre-s pond.'ng, to :ie time when the wjhere PB /p., etc., were obtained by graphi-
* retraiingcal integration of the plots of pB/p. versuswhe retr.!.ig ams ~mpetey rrratedroll angle (ýp) (see Fig. 14 to 16) at the
and interfereni-e free data are ohtainted and vrosnniesoa aerdif~nll, telenigth of th, u-;ful running
time. By using the reference prescsure as Th vrl vrg bs rsuetezern line, the pressure readirgs were aken Teoealaeaebs rsuete
K. ~~~~~in counts (100 2uLunts/in.) along the vertical i opte0rmEutin(01saxis -together with the corresponding timeFreadimgs along the hori.;ontal axis, again in PB B (21)terms %if couoB
TheaJ readii~g onthe vertical d~rection
mesue tedlifercnce b'.twiqon the' refer- From thi, the base pressure coefficient and
converted into pnunds per in. by the trans-
ducer calibratlor, Linstartrs. The re-iu-zing P .Pprress~cre was then either added to or sub- C (22)tractec from the -efe.en-,e prt..;ure to yieldabsolute pr-essure v'alues. fhe free-stream and j:static: pzessure war. detprmined from the Machn umber and toLai pressure taput data. The *B os
A -time readings were converted into milli- CDB -q23
-seconds using the reference timing triarks onthe scilogeph. n~~mumof ightprosure-time readings could be taken along any At Fngle of attack, the base force alsoone trace. has a vertical component. This represents a
lift force. In coefficiotit form this is'¾ -The program output contained the run expressed'as:
9 I
tl' *,A *l. ..- . .
9th Navy Symposium on Aeroballistics
made elsewhere 5 , 11. Both the preliminary•M..' Cq_ ~~(24) and present results indicate that conditions .. : :.:.
•..i- qW at the bane at M u 6.34 and 5.30 ranged fromlaminar a0 Ge= 00, to tranoitionall at mader-ate angles of attack, to turbulert a- high
DISCUSSION OF RESULTS angles of attack.
According to the measurements of Pate35,
S BASE PRESSURE DYSTRIBUTI)N transition at a = 0 occurs at Reynoldsnumbers based on local inviscid conditions of
An importnnt aspect of the experiteatal•i : ~about 4 X 106. At angles of attack, the.' !•iprogram was to measure the pressure distribu- t n ue e a.;.:• . ~transition Reynolds number decrease.a rapidly ,•"t ion of the complete base area. This wasncsr foteopaino thbeprincipally because of cross-flow formation.: .necessary for the €ompuL.tion Of the base"
cfier anbs ialin the boundary layer from the windward to• pressure coefficients and base drag in allpre-ue dagthe leeward side and vortex shedding actiontest conditions. Furthermore, these measure-m p eg i t ton the leeside. Both of these phenomena are• ~~~meats provided an insight into the external cne~dwt nl fatc•te is
"flow field and recirculation region adjacent a aty,•-:: appear at a -•0,8 UC (where 6c is the semi- .to the ba,.e:. This in turn offers the passi- apa taO8~(hr ci h ei'.1 t to formulate a inew turn 0fr te possic--l vertex angle of the cone), as was shown bybility to formulate a new phenomenological Moore 3 l and Tracy 3 2 , and inc,'ease in strengthmodel to serve as a more realistic basis for•i•: ~and effect as the angle of att~ack increases. "Snear wake computations. The boundary layer cross flow end, more
Sparticularly, the leeward vortev: shedding,SFigure• 8 to 10 show the variaition of the causes insta'-ilities in both the viscous and?,•:•base pressure, normalized by free =, treambe rs , o ie b r- rainviscid portion of the flow around the cone
static pressure, as a fuiction of angle of and in the near wake. Feldhuhn et a12 8"-attack, Mach number and Reynolds number. Note mrs?•:• measured the variation of transition Reynolds•'Y t~~hat in each curve the base pressure at .iith't° in eachtluyveither basenpre at =- o number with angle of aitdck and found that at
a 0 0 is slIightly higher than at 'x = -10 or 6.an a 30, et/eo .1whr0. M = 6.0 and a _? 30'), Retof/Reto• 0. 15 where --20o. At or around a =1 0 ° p /p® was a the quantity represents ratio of the transi- ",.:, minimum. There was relatively little yak'i s-mionimum. Therewasrrelatbveeypressure vaw-ia- lion Reynolds number at angle of attack rela-do ng in athe bcwterie base prssr an ith0 tive to the transition Reynolds number atangle of attack between a =-10 and -30 .a~ rlmnr ne~gtowt•:•• "• =0 . A preliminary investigation, withSchlesinger and Martellucci 6 found similar 00-, a 9half-angle sting-mounted cone, under--results for a 100 half-angle cone at M =6.0, a9hl-nl tn-one oe ne'ucenterlinefr a0baha -resre values we Mtaken prior to the main experiments showed
0e.g, 0ethat the recirculation region in the base a t
lower at a = -10 than at 00. However, be- M = 6.3 and Re = I x 10G/ft was tra-isitioningSyond cy= -30°, the base pressure ratio in- 0'y3srwhen the angle of attack exceeded 20 andcreases quite rapidly with angle of attack, became turbulent at around 30..A direct relationship was found between
pand Mach number; for the same a and In the light of the aforesaid, it can beSReynolds number, the base pressure ratio was concluded that in the base region of the modellower for lower Mach number. This is in and at M = 5.30 and 6.34, transition and
• : ~agreement with the results given by Reference:;Iia n w t r t eb rc possible turbulent flow conditions prevailed!• ":'.1 1 , 2 8 , a n d 2 9 .,, 8 ad2.when Re = 6 X 10/ft and the angle of attack
u e th uexceeded 200, whereas conditions were laminar•:'" , Further examination of the results in
Fu 8n 9h ht r 94aat lower angles of attack. This is further." i./ ~Figures 8 and 9 show that for M - 9.94 and sbtnitdo lsreaiainosubstantiated on closer examination ofidentical angle of attack conditions, in- Figures 11 and 12. These show the base
creasing the Reynolds number caused a de-ca pressure distribution along the vertical
crease in r/p valumeridian for M = 6.34 and 5,30 whte the• ..... conditions. The effect of Reynolds ,numbercondition. Te fo f Reynolds e lowest point on the vertical scale corresponds
seemed negligiblethe wrdward generator and the highestresults provide some indications as to the t the l eeward g enerator Reynolds .. 4s: point to tie leeward generator. Reynoldsstate of the cone boundary layer as well as number and angle of attack ere Independentthe flow conditions immediately downstream of variables In each graph.the base. In addition to these re.Sults,temperature measurements and varioLs flowvisualization studies performed previously At a - J0, the base preosure distributionin a preliminary program indicated that the 4n the vertical meridian in all measuredcone boundary layer and the base flow at conditions showed a maximum value at theH = 9.94 were laminar in the unit Reynolds centerline and tapered off toward the edges.number and angle of attack range tested. This is typical of the laminar conditions :-...This was in good agreement with measurements found In several investigations11 ,32 ,33 .
•,•: 10'
•t::• " :.: • '..-•::"•:::,.;"-. " .... • "• ' "....' "........."•:"'" ""'"""....
7,11
K. 9th Navy Symposium on Aeroballiaties
As the angle of attack increased from Schlieren phbtographs taken during thet 40 to -200, the general level of base preliminary investigation Qhowed that the
V pressure along the base region decreased incipient shock on the windward siic of therelat•ve to the zero angle of attack and the cone became highly curved downstreem of thscenterline pressure more particularly de- model at high angles of attack. The shockcreased even further than the edge. valu"s so appeared to turn toward the base centerlineK' that fzm tnP stagtuation line to the center- downstream oi the rec.rculation regio.a. Itline, basc pressure distribution was nearly is conceivable chat the incipieut shock on
uniform (as was the cast- of Re = 3 x lO•/ft, the windward side deformed the inner shear14 - 6.34 or M = 5.9 at both Reynolds numbers) layer causing a loral expansion and conse-
.[or decreased toward the c.rterline (s.e the quent drop in locat pressure. This drop of. two higher Reynolds numbers at M = 6.s4). In local pressure would be felt throvghout thecontrast, the base pressura ratio increased recirculation region and consequently, tl,ebeyond the centerlines, toward the leeward base pressure close to the windward sideside in the .meridian plane. This tendency within a * 200 region would exerience a dropwas also noticeable at higher angles of toward the centerline. The combination of
attack. Cassanto3 3 showed that the peak these effects would be a plausible explana-value of p./p. shifted to the leeward side tion for the shape of the base pressurefrom the centerline ana that the radial base dittribution -at high angles of attack.pressure ratio also increased toward the lee-ward side when a _• 50. Although Cassanto'sresults were obtained at somewhat h4 gher Mach At M 9.94, the cone boundary layer andnumbers, it seems to apply qualitatively to base flow were laminar for the complete test
Y, the present data as well. The general uni- range- Examination of Figure 13 reveals thatI formity of base pressure along the meridian at both Re = 6.4 x lOs/ft and Re = 1.1 x 10'/ft
plane as well as in the entire base region at there was a maximum value for pB/p at the cen--10 to -300 was usually within ± 15%; this is terline for 00 angle of attack. AS the angleconsidered to indicate transition or curbu- of attack increased to c = -300 and above,lent flow conditions. Martellucci and Ranlet 9 the general level of base pressure increasedreported also that in the transition and tur- but in a very nonuniform way. From the edgebulent flow regimes, .J- nonuniformity of the to the centerline starting at the windwardbase pressure distribution decreased and ulti- generator, the base pressures decreased ormately diminished due to large mixing. were nearly uniform toward the centerline. ?
There they reached a minimum point and thenV When the angle of attack increased increased rather abruptly toward the leeward
further, the general level of the base pres- generator where the base pressure level wassure increased but the characteristic distri- 70 to 80% higher than at the centerline.
P!. bution stayed the same as indicated by de- This peculiar pB/p distribution was apparent-creasing base pressure from the most windward ly caused by the sTrong interaction betweengenerator to a minimum point close to the the base flow field and the leeward vortexcenterllue and thereafter an increase as the flow field combined with the highly curvedleeward side generator was approached. The incipient shock wave that acted downstreamlevel of base pressure was, in most cases, of the base on the windward side and deformedhigher at the leeward side that at the wind- the shear layer. Similar, but not identical,ward side. Stetson and Friberg2 2 explain behavior was noted for the other two Machthis phenomenon by the fact that at large numbers.
* •angles of attack, strong interactions occurbetween the base flow field and the leewardvortex flow field. They observed that the Flgures 14 to 16 show examples of the"base pressure increased locally in regions base pressure distributions at three differ-adjacent to the minimum pressure region on ent radii (r/R = 0.24, 0.47, and 0.71) alongthe cone and attributed the effect to weak the roll angle range of 0 to 3600 for alllocal shock wave systems. The pressures in three Mach numbers at Re = 6.0 X 10/ft.the separated regAon on the leeward side were These figures were partially analyzed in thealways larger than the base pressure, but the preceding sections, However, there areinteraction between the two regions apparently some interesting points which have not been4zausod a local increase around the neighbor- discussed yet. Notice in Figure 14 thehood of the leeward meridian * 300. This is minimum areas at roll angles of 30 and 240Ufurther confirmed by the analysis of the in the vicinity of the base centerline (r/Rcomplete base gressure distributions at and 0.24) at o -40". Adjacent to these areasbeyond o = -30 , in particular when r/R 0.71
(see Fig. 14 and 15).
W I
-M-7
9th Navy Symposium on Aeroballistlcs .
local maximum goir.ts occurred at roll angles Examination of the base drag to total Aof 120 and-310 . this same phenomenon was drag ratio (if comparison is made between Arepeated at angleb of -10, -20, and -1300. At Fig. 19, 20, and 23 to 25) showed that at 2 AH =-9.94 (Fig. 16), there were minimum zones ak 00, t~he base drag was about 15 to 181% ofii't roughly 9Q0 from the vertical meridian the total drag ac M = 5.30 and 6.34 and wasplane and maximum zones adjacent to Lhe lee- in the neighborhood of 0 to 7% at M = 9.94.ward meridian. The fact that local maxima Results of other investigators corroborateand minima were close to the cnter area but these measurements. At ce--10 , the baseoccurred only at certain spots indicate local drag to total drag ratio dropped to aboutflow reactachments due to the iuteraction of 4-to 5% for all three Mach numbers. Atrolled up vortex sheets from the leeward side -= 20" and beyond, CD/CDB gradually --1and the recircul..cion region. Apparently decreased from about 1.7 to 1.1%.
- this interaction resulted in an intensive "local mixing and secondary reattarhment. The When the model flew at an angle of
Macph"nomenon is noasymmetric. At the lowest attack, a base lift force was generated inMach number tested (see base pressure distil- addition to the base drag since the totalbution in Fig. 15), this phenomenon was much force at the base was no longer parallel with -1
:-less pronounced than for the higher Mach the free-stream direction. Fig. 21 and 22
numbers. A possible explanation is that the show the base lift coefficients as functionspressure gradients withit, the base region were of M, Re and a. Apparently, CLB is nearly ,much smaller than in the other cases because independent of Hand Re and chiefly dependentintensive turbulent mixing alleviated the on the angle of attack. The ratio of CLBsecondary reattachment. to the total tift was in the order of I to
- 2% (if one compares Fig. 21, 22, and 26 to28).
- AERODYNAMIC COEFFICIENTSAT THE MODEL BASE TOTAL DRAG AND LIFT COEFFICIENTS
Bise pressure coefficients, as defined Figures 23 to 25 show the variation ofby Equation (22) and evaluated by Equations the total drag coeffcicnts with angle of(20) and (21), are shown in Fig. 17 and 18. attack for all test conditions. In additionStrong Mach number Jependen'3e was evident to the measured values, the inviscid dragfrom the results. At ao= 00, the ICPBJ value coefficients were computed utilizing thewas about 50% larger at M = 5.30 than at modified Newtonian theory. Equations (9) to
s. M = 6.34 but 4 to 7 times as large as the (14) were used for these computations and thebase pressure coefficient at M = 9.94. The results are also shown on the graphs. Thedifferences decreased as the angle of attack close agreement between measurements andincreased. Around ot= -10 to -200, CPB was theory is readily apparent, although someminimum for all test conditions and thereafter experimental scatter shows up in Fig. 24.increased with increasing angle of attack. Results of the investigation conducted byThe base ?tessure-coefficients did not depend Ward et a13 showed that at relatively lowsignificantly on Reynolds number throughout Mach numbers (below about 7) the viscousthe test envelope. The only exception was at drag was itsignuii..arit and the total drag wasM = 9.94 and a = -10h where Cp, at Re the sum of the inviscid drag- and base drag.
/6 x 10f/t was only 50% of the base pressure But the base drag was only about 1.5% of the
oefficient value measured at Re :- I 10o/ft. total drag at high angles of attack for theThe Reynolds number effect is evident when present investigation, as was discussed in theone compares Fig. 17 and 18. previous section, so the good agreemunt be-tween measured and calculated drag coefficients
She base drag coefficients for was not surprising. At HMi 9.94 the viscous
Re - 6 x 105 and i X lOS/ft are shown in drag became significant at a- 0 and conse-
Fig. 19 and 20, respectively. Mach and quently, the measured tot~al drag coefficientsReynolds number effects were quite insignifi- deviated sonewhat from the computed ones. Thecant throughout the test envelope. The base measured values were between CD = 0.1 and 0.12.drag coefficient decreased between a = 0 and These values were in good agreement with theabout o = -15° and increased with increasing measurements of Welsh et al and Ward et a1 3 .
-- = angle of attack thereafter. The measurementsof base drag at a = 00 were in good agreement Figures 26 tc 28 show the total liftwiLh data presented by Ward and Choateo10 . coefficients as functions of M, Re, and a forThey found that at M = 10 and zeý--3 to6X 10l, all test conditions. Lift coefficients, basedthe base drag was close to 0.008. The present on the Newtonian inviscid theory, were alsomeasurements showed very similar results in computed using Equations (9) to (14). ;'he
"CDB values at M = 9.94 (see Fig. 19 and 20). computational results are shown as solid .i
12 ;
""* . , . , ." ' '-"
-- ".-..
9th Navy Symposium on Aeroballistics
curves on the graphs for comparison. Again, (5) Base drag at . 00 = 0 represented athe measered and computed va ces ftre in good . considerable fraction of the total drag butagreement. deczeascr. tv around 1.5% above the angle of
4track at -7ýOo.
MomE Tw rA (f RF.LTS (6) Clooe agreement was found between-
I e a j o dthe measured and predicted values of total:!il;-i " i ~ ~In the great -,'•ajorty of th,• m'odel drops, lf n rgcefcet n hnte•
* the model departure in the yaw plane was lift and drag coefficients and when thegl b a tr e e t- &modified Newtonian theory was used.!•i •aegligible, and therefore, use cf two-di;::,tý .
-sional curvilinear, equations of tion was (7) The computed and measured trajector-"justified. Consequently, Equations (15) to(18), programmed for computer calculations,provided good approximations for se-mi-theoret- .1lcal trajectory predictions in conjunction REFERENCES"'th predictor and corrector techniques.Figures 29 to 31 show that the measured andcomputed model trajectories were in good . 1 rfih .. adLG ur"Comparisons of Free-Flight and Wind-Tunnelagreement throughout the Mach number and angle Data on Slender Cones," Arnold Engineeringof attack renge covered by the test program. Development CenteNote that At lower angles of attack (10 j<20°), r AEDC-TDR-64-272 (Dec 1964).
the agreement is relatively poor at M w 5.301and 6.34. The steepness of the trajectories 2. Welsh, C.G., G.L. Winchenbach and
may account for this disagreement where the A.N. Madagan, "Free-Flig:.t Tnvesitgation of.ay ithe Aerodynamic Characterisitcs of a 10-Degsmal nSemiangle Cone at Mach Numbers from 6 to 16,"procedure become relatively more significant Arnold Engineering Development Center, AEDC-than at trajectories where the horizontal TR-69-63 (Apr 1969).distance is longer (i.e., trajectories corre-sponding to higher angles of attack). As afinal comment on the trajectory calculations, 3. Ward, i.K., A.E. Hodapp and R.H. Choate,it is worthwhile to mention that even when the "Description of a Model Launcher and Techni-
•;:. ques used for Obtaining Model Free-Flightmodel angle of attack changed significantly rteasurements in the VIF Continuous Flow Wind
pon" p , tTunnels at Mach Number from 1.5 through 10,".•: ! bAP!,! on local angle of attack was in closeb o o a oa k i lArnold Engineering Development Center, AEDC-
a•,rreement with the measured path. TR-66-112 (Aug 1966).
4. Carpenter, 2.W. and W. Tabakoff, "SurveySCONCLUSIONS and Evaluation of Supersonic Base Flow
Theories," University of Cincinnati,The results can be summarized as: Cincinnati, Ohio, NASA CR-97129 (Oct 1968).
(I) The centerline base pressure normal- 5. Todisco, A. and A. Pallone, "Measurementsized by free-stream static pressure, was of Laminar and Turbulent Near Wakes," AVCO,relatively constant at angles of attack Everett, Mass., AIAA Paper No. 67-30between -10 and -300. Above -300, pB/p in- (Jan 1967).creased with increasing angle of attack.
( , 6, Schlesinger, A. and A. Martellucci, "Wind-(2) An increase in Mach number caused Tunnel Invesrigations of the Turbulent Near
an increase of the base pressure ratio. Wake of a Cone at Angle of Attack,"GASL, N.Y.,Tech Rpt 581 (Mar 1966).
(3) At M- 9.94, laminar conditions pro-vailed throughout the entire test envelope. 7. Schmidt, E.M. and RJ. Cresci, "NearHowever, at H~ - 6.34 and 5.30, there was Wake of a Slender Cone in Hypersonic Flow,"transition and turbulent flow in the no,..r Polytechnic Institute of Brooklyn, Brooklyn,
wake at angles greater than -200. N.Y., AGARD Conference on Fluid Physics ofHypersonic Wakes, Vol. I (May 1967).
(4) Measurements indicated a highlycomplex base flow region. A tentative ex- 8. Martellucci, A. and A. Schlesinger,planation was offered, but much more flow "Measurements in the Near Wake of a Cone atfield data are required to define all aspects Mach 12," GASL, N.Y., Tech Rpt AD803-892 Iof the base flow at high anglas of attack in (Mar 1966)the hypersonic flow regime.
"1A • 13
"-":.'-
9th Navy Symposium on Aeroballistics
9. Martellucci, A. And J. Ranlet, "Experi- 21. Peckham, D.H., "Exptoratory Tents on• I~i. i mental Study of Near Wakes," GASL, N.Y.,$- Sting interference at a Mich Number of 6.8," i+ ... •
Tech Progress Rpt (Oct 1966). RAE Te'..h Note Aero 2721 (Oct 1960). 110. Ward, L.K. and R.H. Choate, "A Model 22. Stetson, K.F. and E.G. Friberg,Drop Technique for Free-Flight Measurements "Communication Between Base and Leeward
'i I in Hypersonic Wind Tunnels Using Telemetry," Region of a Cone at Angle of Attack in aAEDC-TR-66-77 (May 1966), Hypersonic Flow," W-P AFB, ARL 69-0015
(DDC AD 696 584)(Jul 1969)..11. Cassanto, J.M., N.S. Rasmussen andJ.D. Coats, "Correlation of Measured Free- 23. Pick, G.S., "Sting Effects in Hypersonic -1Flight Base Pressure Data for M = 4.to M - 19 Ba':e Pressure Measurements," NSRDC Test Rpt,in Lamir and Turbulent Flow," AIAA Paper AL-85 (Dec 1971).S..... ~No. 68-699 (Jun 1968) .': "t.No. 8-69 (Ju4. Ziegler, N.G., "The David Taylor Model
12. Potter, J.L. and J.D. Whitfield, Basin Wind-Tunnel Facility," DTMB Aero Rpt"Preliminary Study of the Effect of Unit 1027 (Jul 1963).
,4 Reynolds Number on Transition Sensitive Data,"W AEDC-TN-57-37 (AD 135338)(Feb 1957). 25. Harrison, R.G. Jr., "A Pressure Telemeter
for W•ind-Tunnel Free-Flight Pressure Measure-13. Bogdonoff, S.M., "A Preliminary Study ment," Jet Propulsion Laboratory Tech Rptof Reynolds Number Effect on Base Pressnre No. 32-763 (Feb 1966).a, M = 2.95," Journal of Aeronautical SciencesVol. 19, No. 3 (Mar 1952). ?6. Pick, G.S., "A Study on Short-Time, Low
-Pressure Response in a Transducer System,"14. Whitfield, J.D., "Critical Discussion Wash., NSRDC Test Rpt AL-74 (DDC AD 711 191)•:• of Experiments on Support Interference at (May 1970). .
Supersonic Speeds," Arnold EngineeringDevelopment Center, AEDC TN-58-30 (Aug 1958). 27. Picki G.S. and S.E. Dawson, "Multi-
Channel Pressure Telemetry System for Base15. Love, E.S., "A Summary of Support Pressure Measurements on Small Models,"Interference at Transonic and Supersonic Proceedings, IEEE 3rd International CongressSpeeds," NACA RM L53KI2 (Jan 1954). on Instrumentation in Aerospace Facilities "
(May 1969).16. Kavanau, L.L., "Results of Some BasePressure Experiments at Intermediate Reynolds 28. Feldhuhn, R., A. Winkelmann and L. Pasiuk,Numbers with M - 2.84," University of "An Experimental Investigation of the Flow 4California, Institute of Engineering Research, Field Around A Yawed Cone," AIAA JournalRpt No. HE-150-117 (Oct 1953). (Jun 1971).
17. Kavanau, L.L., "Base Pressure Studies 29. Soffley and Graber, "An Experimental Iin Rarefield Supersonic Flows," University of Study of the Pressure and Heat Transfer onCalifornia, Institute of Engineering Research, the Base of Cones in Hypersonic Flow," AGARDRpt HE-150-125 (Nov 1954). Proceedings, No. 19 (May 1967).
18. Reller, J.O. and F.M. Hamaker, "An 30. Pate, S.R., "Measurements and Correla-Experimental Investigation of the Base tions of Transition Reynolds Numbers on Sharp"Pressure Characteristics of Non-Lifting Slender Coaes at High Speeds," AEDC-TR-69-172Bodies of Revolution at Mach Numbers from (Dec 1966).2.73 to 4.98," NACA TN-3393 (Mar 1955).
31. Moore, F.K., "Laminar Boundary Layer on19. Sivier, K.R. and S.M. Bogdonoff, "The Cone in Supersonic Flow at Large Anglo ofEffect of Support Interference on the Base Attack," NACA TN 2844 (Mar 1963).Pressures of a Body of Revolution at HighReynolds Numbers," Princeton University Rpt 32. Tracy, R.R., "Hypersonic Flow over YawedNo. 332, AFOSR TN-55-301 (Oct 1955). Circular Cones," Hypersonic Project Memo 69,2.e ,WR "efe o t California Institute of Technology (May 1963).i 20. Steling, W.R., "The Effect of Sting •
Diameter and Length on Base Pressure at 33. CassanLo, J.M., "Free-Flight Base3.88," The Aeronautical Quarterly, Vol. Pressure Results for a Sharp Cone at Angle
19 (Nov 1968). of Attack," G.E. Aerodynamics Laboratory Data
Memo 70-88 (Mar 1970).
14*i _ _.oo .o 4< ii
: h .': . + . .:- , . . . .. ••.. ,• •..•••,.•,:+ . ... . y+•,.,- -• •.••, -. ,-• ,. .+,• . . .. ... .. ... .. ..... . .. ... . . •... . ... ' 4 •
MM.,. .
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9th NayfmoimonArblltc~Z~~A1 (~ -t
-. *
74~-
2.~~L -t'ay~ p~no - - mRUN3 2.0w
-a-'
ofSMBL MoeMDo-Hih
~ .2.6J~~*U 1 _______
Figue 2 cheatic of ode IntrioIoFiur 5 lc ga f10 CoNe oe
an IM IntUNnato in t .0eHyesoi Wid une
Figure 4 TermSaplRe O leogase Vel corad.
3~~~~o Moe Dntroepe Hoeeodlanghtet
4 av
jk.ý
A - ~ Ali-
-. c
Fiue2 Shntc fMde Interior-~
* 9th Navy Sympetum on Aeroballistlcs
.4.
Z4-, ~~* M .Re
2.2 x 05)5t 1 05/;T
CORRELATION C~9.94 6.4 x 1 5lP
1.4L
--" . t! -- ~~--.-
2 3 4 56 8 10 15 fTIME RESPONSE (MSEC)01
kFigure 7 -Time Response Characteristics of 0.4- 4-the Transducer System02---
(lIN DEGREES)Figure 9 -Centerline Base Pressure Ratio as
1.4 Function of Angle of Attack and Mach Number
0 &.30 10.7 a 1 5IFT 18 a e 6~ 0/e
1..*
ow-39 1051Fi
1.2---.
F ~ ~~~~~o-----S -- - -
0 -10 -n -30 -040 -U 10 -W0040-0 - -40 -50 .40 -70 40 (INoauu-~Figure 10 - Centerline Base Pressure Ratio &aq
Function of Angle of Attack at N 6.34Figure 8 - Centerline Base Pressure Ratio as and Re 3 x 106/ftFunction of Angle of Attack and Hach Number
at Re-I X 10 0 /ft
iv,, R
~ V.
01
ri 0 7 .1 - -l
US? : i- r&L
(L -- SO*
0 40 80.1@10US3W~t4 0.8 0~~~* 04,08 0.4 OZ'-*~A~tER
-Fiur 1- asePrssreDisriutoni -Figr 14 'B~ase Pressure Distribution-, at.Verti"a ~iia'ln tNm63 -6.:34= 10 'R '. OIft
it .. to~' 11.0R..5.%OSFT Re I. OPIFT
K a4~. -~FA'~-~T+-0.2
-- '
C0.4}~if 04 I I_
0. G 08 0 A 0.47u.0n~e5915 f
1.0
o-- 9a--3 A 0~0~~ 044 0.0 12 1.0 NO 20 006 0.8 3 0ue1 rsueDitiuina
0 .4 as 0 .4 0. M-5.304and Re 5.9 XlOr/ft
figure 132 Base Pressure Distribution ini theVertial. Meridian Plane at N 5 .94
j~Z)~ ~4Van, I i§
OR X1 Re 11'440
- --~~- - ---~ -
.. ...." . ...-,
jr>'
fth Navy Symposium on Aeroballis tics .
a- ~309 0.01 M R
0.8 05.30 10.7 x 105/FTCLA0. 06.34 11.2 x105/FT
KU1.2
04 viA -- 0.01
0* L1 120 14020 20 326 NO 3
~~< Figure 16' (Continued)-02
0 -0.03040 IC 12rift0.1240 -0 -0 44O LA4 L OGR E D G E S
0 5. 20.9 160 2 2o/*3Mn
0 5.30 _W9x-40F0.02~~~ a(634G.2EES)/F
00.01
0.0
0 -130 -20 -30 -4 -5
Fuptsn ofAAl0f2tac n
0ahNme tR ~6~l~fA03 -10 -20 -30 -40 _0
a4 (DEGREES)
Figuree 17 Bas Drsg Coefessur aFunctions of Angle of Attack and
Mach N~umber at Rei 6 X LOF,/ft
--.. '. T%
~~~~~~~~~~~~~~~~~ M&~ , ~ ,~ . 4**\* 1 ~4 ¶* 4
M 9th Navy SymposumOn' Airoblitc
K- ft~4 '~ ~ ~
0 .3 11.2x 105/0T '
______ 1.~~6: 0 5.30 -. 1 5 F7!W -1 -2O-r 5~.30 10.7,x405/FT..
Figre 0 -Base Drag Coefficients as '- ,ss*~~
unios of Angle Of Attack and-- NEWTONIAN .. ..-
- M01 a (DEGREE)
Figure 23 -Total Drag Coefficient as Function ,,¾
0 5.30 5.9.x 1 5/FT of Angle Of Attack and Reynolds Number ,
6.34 6.2 x105/F at M 5.30u
-10 -20 -0 -0 -
a(DEGREES)" M Re
J ~ ~~Figure 21 -Base Lift Coefficients s16 a634 .20/FFunctions of Angle of Attack and 1x1/T
Macb Number at Re 6 x 10"/ft
44* t7 4 -0.0
-020 ..10 -20 -30 -40 ~ -
i7 ea (DEGREE)105.30 10.7 x 105/FT
0 6.34 11.2 x105/FT09.93 11A41 5 P Figure 24 -Total Drag Coefficient as Function
'0 -10 -20 Of Angie Of Attack and Reynolds Nwnbur
a (DEGREES) .4~-.
*Figure 22 -aUse Lift Coefficients as . ~Functions of Angie of Attack and
.sk umber at' Re I x 1 0lt .
-. 4-
I 9th Navy Symposium on Aerobatlistics
2.0 __1.0
M Re
1.6 * 9.94 6A x 10 5 /FT ____.8_
* 9.94 11.4x105/FTNET IA
1.2 -0.
0. 8 -- -0.4 _
NEWTONMAN MR
0.4 - HOY-0.2 036.34 6.2 x105 /FT
86.34 11.2 x10 5/FT
0 -10 -20 -30 -40 -50 -60 0 -10 -20 -30 -40 -50 -60
a (DEGREE) a (DEGREE)
Figure 25 - Total Drag Coefficient as I unctioflof ngl ofAttck nd eynldsNumer Figure 27 - Total Lift coefficient as Function
at1 .4of Angle of Attack and Reynolds N'umber
at 14=6.34
-0.-0-8
-36 0.HE_ THEORY
~ -0.4:C -0.4
M ReM Re-0.2 0 5.30 5.9 x105/FT -0.2 0~9.94 6.4 x105/FT
05.30 10.7xl05/FT *9.94 11.4xl05 /F-
0 -10 -20 -30 -40 -to 6 0 -10 .20 -30 -40 -50 -60
a (DEGREE) a (DEGREE)
kFigure 26 - Toial Lift Coefficient as Fun..tion 1j.A.re 28 -Total Lift. Coefficient as FUni.Li~n
of Angle of Attack and Reynolds Number of Angle of Attack and Reynolds Nume)ec
at M4 5.30atM 94
20 1
- ..-. -----..---- '----.--.-.----
9th Navy Symposium on Aerobailistics
-COMPUTEDI--MEASURED
(IN.)
8 AVE -32.9V
"1 \AVF1. CAE~4O0~~OMPUTED00 1 2 3 4 MEASURED
S. (IN.) 2
Figure 29 -Measured and Com~puted ModeýIrajectories at M -).30 and Re =5.9 x 105/ft4
SY(IN.) aAVE -59.70
-COMPUTED
- - -MEASURED
2
=.-14.2'
(I. 4" ~ 1100 AV 2 3 4 5
sy N ~AV, (I N.)
(IN.)aAVE 46.3'
6 NFigure 31 - Measured and Computed Model\NNTrajectories at M =9.94 and Re 11.4 x IO05 /ft.
aAVE ~11
101. AVE=_-19-0'10 1 2 3 4 S, (I N.
Figure 30 -Measured and Computed ModelI
Trajectories at M 6.34 and Re 6.2 x l0f/ft
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