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Approved for public release; distribution isunlimited.
Distribution authorized to U.S. Gov't. agenciesonly; Test and Evaluation; NOV 1967. Otherrequests shall be referred to AeronauticalSystems Div., Wright-Patterson AFB, OH 45433.
ASD ltr 28 Nov 1972 ; ASD ltr 16 Aug 1977
^AEDC-TR-67-252 DECLASSIFIED / UNCLASSIFIED
ARCHIVE COPY DO NOT LOAN
ARO, INC. DOCUMENT CONTROL N0 IG-681-343
rnpv / ne 41
SERIES _JL_ PAGES _i*_
WIND TUNNEL INVESTIGATION OF A 1/8-SCALE AMSA AIRCRAFT-INLET MODEL AT
TRANSONIC AND SUPERSONIC MACH NUMBERS (U)
5 g l< * ' a. 88
"^4 Lawrence L Galigher
'*si
JAN^'1968 JUN 4^1968
|uÄ59
* ' ^40/600^1^00 November 1967 JUN ^1971
1
«."«■«hBEHTYOFUfi
PROPULSION WIND TUNNEL FACILITY
ARNOLD ENGINEERING DEVELOPMENT CENTER
AIR FORCE SYSTEMS COMMAND
ARNOLD AIR FORCE STATION, TENNESSEE DECLASSIFIED/UNCLASSIFIBD
GROUP 3 PPA?!7n"rV Or- I I A . ._ Downgraded at 12 year Intervals 1 IV^i _,., , i,,.- |J £ Aip FORCE Not automatically declassified.
> f\/ DOD DIR S200 10 ■AY
AF 40(600)1200
AEDC TECHNICAL LIBRARY
S D72G DD031 b32S
mm When U. S. Government drawings specifications, or other data are used for any purpose other than a definitely related Government procurement operation, the Government thereby ineurs no responsibility nor any obligation whatsoever, and the fact that the Government may have formulated, furnished, or in any way supplied the said drawings, specifications, or other data, is not to be regarded by implication or otherwise, or in any manner licensing the holder or any other person or corporation, or conveying any rights or permission to manufacture, use, or sell any patented invention that may in any way be related thereto.
Qualified users may obtain copies of this report from the Defense Documentation Center.
Hcfoionci's to named commercial products in this report are not to be considered in any sense as an endorsement of the product by the United Stales Air Force or the Government.
Do not return this copy. When not needed, destroy in accordance with pertinent security regulations.
AEDC-TR-67-252
DECLASSIFIED / UNCLASSIFIED
WIND TUNNEL INVESTIGATION OF A 1/8-SCALE
AMSA AIRCRAFT-INLET MODEL AT
TRANSONIC AND SUPERSONIC MACH NUMBERS (U)
Lawrence L. Galigher
ARO, Inc.
a? o
overnments 6r iv witi&pn^fi" approval AFB, Ohio.
.«ft nladdÜÜftn to security (requirements Bra document andlnust(he^SM TBmit|alio\ ■ re». cia^^BPW^^Vv * k» ■"" — -TP^PBPTense must ha^e prior ap-
ZBtf ffJffejK-PaVtfKm AFB, Ohio.
This material contains information affecting the national defense of the United States within the meaning of the Espionage Laws (Title 18, U.S.C., sections 793 and 794) the transmission or revelation of which in any manner to an unauthorized person is prohibited by law.
Distribution lint onlyj Test request to Q
rrf:CLASSlFfl E0 / UNCLASP«**
AEDC-TR-67-252
DECLASSIFIED I UNCLASSIFIED
FOREWORD
(U) The work reported herein was done at the request of the Aeronautical Systems Division (ASD), Air Force Systems Command (AFSC), for the General Dynamics Corporation, Fort Worth, Texas, under Program Element 6340683F, Program Area 139A.
(U) The results of tests presented were obtained by ARO, Inc. (a subsidiary of Sverdrup & Parcel and Associates, Inc. >, contract operator of the Arnold Engineering Development Center (AEDC), AFSC, Arnold Air Force Station, Tennessee, under Contract AF40(600)-1200. The test was conducted from August 7 through September 9, 1967, under ARO Project No. PT0741, and the manu- script was submitted for publication on October 27, 1967.
(U) This report contains no classified information extracted from other classified documents.
(U) This technical report has been reviewed and is approved.
# Richard W. Bradley Leonard T. Glaser Lt Col/ USAF Colonel, USAF AF Representative, PWT Director of Test Directorate of Test
3
DECLASS.F«ED/UNCL».SSffffiO
This page is Unclassified
UNCLASSIFIED AEDC-TR-67-252
UNCLASSIFIED ABSTRACT
(U) Test results are presented for a 1/8-scale inlet model of a proposed AMSA air induction system at transonic and supersonic Mach numbers. The effects of spike bleed pattern and inlet orienta- tion are presented. Compressor face total-pressure recovery and flow distortion data are presented as a function of spike position, compressor-face mass-flow ratio, spike boundary-layer bleed mass- flow ratio, angle of attack, angle of sideslip, and free-stream Mach number.
Distribution limited to U. S^Goj^^^^cies only; Test and Evaluatioi^rf^^P^^^ther requests for th^^^^S^^nust be referred to Cammandeja^^^Mautical Systems Div., Attn^^jj^^^Jright-Patterson AFB, Ohio ^5*M^?er TAB 73-2, 15 January, 1973.
V_
111
UNCLASSIFIED
AEDC-TR-67-252
DECLASSIFIED / UNCLASSIFIED CONTENTS
ABSTRACT . . . NOMENCLATURE
I. INTRODUCTION II. APPARATUS
2. 1 Test Facilities 2. 2 Test Article 2. 3 Instrumentation
III. PROCEDURE 3. 1 General 3. 2 Accuracy of Measurements
IV. RESULTS AND DISCUSSION 4. 1 Isolated Inlet Tests . . . . 4. 2 Inlet-Airplane Tests . . .
V. CONCLUSIONS
Page
iii vii
1
1 2 3
3 4
5 6 3
APPENDIXES
1. ILLUSTRATIONS
Figure
1.
2.
3.
4.
5.
6.
7.
8.
Location of AMSA 1/8-Scale Inlet Model in Tunnel 16T and Tunnel 16S Test Sections 11
Sketch of AMSA 1/8-Scale Inlet Model .'». . 12
Sketch of the Proposed AMSA ,' . . 13
Inlet Airplane Model Installation in Test Section a. Tunnel 16T < ■ 14
b. Tunnel 16S . . 15
Isolated Inlet Installation in Test Section a. Single Inlet 16 b. Dual Inlets: Nonstaggered Arrangement .... 17 c. Dual Inlets: Staggered Arrangement 18
Dimensioned Sketch of Model Nacelle 19
Spike Boundary-Layer Bleed Pattern 20
Inlet Spike Contours . . . . '. 21
v
This^pa'ge is--Unclassified
DECLASSIFIED / UNCLASSIFIED
AEDC-TR-67-252 «^nw* >:$
DECLASSIFIED/UNCiAS£IFtlD * ,,. Figure Page
9. Compressor-Face Station Instrumentation 22
10. (58 Effect of Spike Bleed Flow on Inlet Performance; 27
ll Spike, Ma = 2. 06, a = 0 deg, 0 = 0 deg,
X/R = 1.59, W J0/62 = 185 lbm/sec 23
11. & Effect of Spike Position on Inlet Performance; 27
I. Spike, Ma = 2. 06, a = 0 deg, ß = 0 deg,
w7e/62 = 180 1bm/sec 24 12. (i) Effect of Spike Bleed Pattern on Inlet Performance;
27 l1 Spike, M. = 2. 06, a = 0 deg, 3 = 0 deg,
X/R =1.63 25
13. $) Mutual Interference Effects of Nonstaggered, Dual-Inlet Orientation on Inlet Performance;
I^7 Spike, M«, = 2. 06, a = 0 deg, 8 = 0 deg,
X/R = 1.63 26
14. <i) Mutual Interference Effect of Staggered Dual-Inlet 27
Orientation on Inlet Performance; I- Spike,
M. = 2. 06, a = 0 deg, J3 = 0 deg, (X/R)Inb-d = 1. 63, (X/R)0utb,d = 1.67 27
15. ($ (U) Effect of Spike Position on Inlet Performance; ö = 5 deg, ß = 0 deg a. (<) M. = 0. 85, wJd/52 = 260 lbm/sec 28 b. (#)M. = 1.20, WN/9/62 = 260 lbm/sec 29 c. ($)Mm = 1.75, W>/0/62 = 220 1bm/sec 30 d. (DM«, = 2.05, Wv
/9/62 = 196 lbm/sec 31 e. (|)MW = 2,20, wJe/62 = 180 lbm/sec 32
16. if) Effect of Spike Bleed Flow on Inlet Performance;
I^7 Spike, M,,, = 2. 20, a = +5 deg, ß = 0 deg,
W^e/62 = 180 lbm/sec 33
17. Effect of Angle of Attack on Inlet Performance; 6 = 0 deg 34
18. Effect of Angle of Sideslip on Inlet Performance; a = 6.2 deg 35
19. Optimum Inlet Performance; a = +5 deg, 8 = 0 deg ... 36
r if VI'
.OECIAWFIED i UMcCHSSPlED
UNCLASSIFIED AEDC-TR-67-252
II. TABLE
I. Accuracy of Measurements.
Page
37
NOMENCLATURE
BL
M
n^/mi
msb/mi
N2
Pt R
Tt
WL
w
X
a
ß
Buttock line
Compressor-face total-pressure distortion,
Wmax " (Pt2) min avg
Inlet spike with 15-deg first-cone half-angle
Mach number
Ratio of compressor-face station mass flow to inlet capture mass flow. The inlet capture mass flow is defined as the mass flow passing through an area equivalent to the cowl area when projected on a plane normal to the spike centerline
Ratio of spike bleed mass flow to inlet capture mass flow
Compressor-face total-pressure recovery,
<Pt2>avg/Pt»
Total pressure, psfa
Inlet cowl radius at cowl leading edge
Total temperature, °R
Waterline
Simulated full-scale weight flow at compressor-face station, lbm/sec
Distance from cowl leading edge to spike tip, in.
Angle of attack of model wing chord line, deg
Angle of sideslip, positive nose left, deg
Ratio of compressor-face average total pressure to standard sea-level pressure, (p£2)avg/2116
Vll
UNCLASSIFIED
AEDC-TR-67-252 UNCLASSIFIED
0 Ratio of free-stream total temperature to standard sea-level temperature, T+ /519
vQD
SUBSCRIPTS
1 Inlet lip station
2 Compressor-face station
avg Average
Inb'd Inboard inlet
max Maximum
min Minimum
Outb'd Outboard inlet
OB Free-stream conditions
SUPERSCRIPTS
10, 15, 20, Inlet spike second-cone d 24, and27
viii
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AEDC-TR-67-252
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SECTION I INTRODUCTION
(U) Tests were conducted in the Propulsion Wind Tunnel, Tran- sonic (16T) and Supersonic (16S) to determine inlet performance characteristics of a 1/8-scale, partial span model of the Advanced Manned Strategic Aircraft (AMSA) as proposed by the General Dynamics Corporation, Fort Worth Division.
(£) The objectives of the tests were to evaluate some spike and cowl design variations and nacelle position effects (spacing and stagger) from both the isolated inlet and inlet-airplane standpoint.
(t) This report is concerned with the significant test results obtained for the basic cowl configuration at free-stream Mach numbers from 0. 85 to 2. 20. The complete test data were forwarded to the test user and are available at AEDC.
SECTION II APPARATUS
2.1 TEST FACILITIES
(U) Tunnel 16T is a closed-circuit, continuous flow wind tunnel capable of operating at Mach numbers from 0. 55 to 1. 60. The tunnel can be operated over a stagnation pressure range from approximately 160 to 4000 psfa and over a stagnation temperature range from 80 to 160°F. The tunnel specific humidity is controlled by removing tunnel air and supplying conditioned makeup air from an atmospheric dryer. Perforated walls in the test section allow continuous operation through the Mach number range with a minimum of wall interference.
(U) Tunnel 16S is a closed-circuit, continuous flow wind tunnel that currently can be operated at Mach numbers from 1. 65 to 3. 10. The tunnel can be operated over a stagnation pressure range from approximately 100 to 1800 psfa and over a stagnation temperature range from 150 to 435°F. The tunnel specific humidity is controlled by removing tunnel air and supplying conditioned makeup air from an atmospheric dryer.
1
ncri AäSIFIEO / UNCLASSIFIED
I« AEDC-TR-67-252
DECLASSIFIED / UNCLASSIFIED
(U) A more extensive description of each tunnel and its operating characteristics is contained in the Test Facilities Handbook. *■ A dimensioned sketch showing the model location and the sting support arrangement in both the Tunnel 16T and Tunnel 16S test sections is presented in Fig. 1 (Appendix I).
2.2 TEST ARTICLE
2.2.1 Wing-Fuselage Configuration
($) The model is a 1/8-scale general configuration of the basic AMSA airplane with the variable swept wings in the fully retracted position. The wing tips, aft of the outboard inlet, cowl lip station, and horizontal stabilizer were removed to reduce aerodynamic loads on the wind tunnel model. Dimensioned sketches of the model and the proposed full-scale AMSA are shown in Figs. 2 and 3, respectively. Photographs of the model installation in Tunnel 16T and Tunnel 16S are presented in Fig. 4.
2.2.2 Inlet Configuration
CfcJ The air induction system for the proposed AMSA, as shown previously in Fig. 3, consists of four external compression, axisym- metric double-cone inlets located beneath the wings and attached to the fuselage in dual-inlet arrangement. Only the dual-inlet arrange- ment for the left side of the fuselage was installed on the model for the inlet-airplane phase of testing. During the isolated inlet phase of testing, a mounting bracket was used to position both a single-inlet installation and a dual-inlet installation ahead of the sting. For the dual-inlet installation, separation distances of 0. 70, 0.84, and 1.00 inlet diameters were investigated for a nonstaggered arrangement, and a separation distance of 0. 84 inlet diameters was investigated for a staggered arrangement. In the staggered arrangement the outboard inlet was located 3. 5 inlet diameters downstream of the inboard inlet. Photographs of the isolated single-inlet installation and the isolated dual-inlet installation in both the nonstaggered and staggered arrange- ments are shown in Fig. 5.
JTest Facilities Handbook (6th Edition). "Propulsion Wind Tunnel Facility, Vol. 5. " Arnold Engineering Development Center, November 1966.
ragetASSIFIED i IÜNCLASS1F1E&
AEDC-TR-67-252 ^nOr4MrS«
(e) The inlets had 9. 5-deg cowl lip angles and inlet capture areas of 14. 875 in. 2. They were geometrically similar, both internally and externally, to full-scale inlets from the cowl lip to the compressor- face station (nacelle station 14.375). Compressor-face airflow control was obtained with a variable position plug valve downstream of the compressor face. A sketch of the model nacelle, including tables of the cowl internal and external contours, is shown in Fig. 6.
($ Five interchangeable double-cone inlet spikes were provided for each inlet. The spikes had a 15-deg first-cone half angle with second-cone deflection angles of 10, 15, 20, 24, and 27 deg. They were remotely translatable and, with the exception of the two spikes having second-cone deflection angles of 10 and 15 deg, had boundary- layer bleed holes on the second-cone surface. Spike boundary-layer bleed pattern variations were accomplished by opening selected rows of bleed holes that were initially closed with a filling compound prior to testing. Spike bleed flow was ducted internally through the spike centerbody and vented overboard. Spike bleed control was obtained with a variable position plug valve located within the inlet actuator housing. The two bleed patterns investigated during this test are shown in Fig. 7. A sketch of the spike, including tables of the spike contours, is shown in Fig. 8.
2.3 INSTRUMENTATION
(U) Inlet performance in terms of total-pressure recovery and flow distortion were obtained from compressor-face total-pressure meas- urements; the orifice locations are shown in Fig. 9. A dynamic trans- ducer was mounted in the duct wall at the compressor face of each inlet to monitor inlet stability. Spike boundary-layer bleed and compressor- face mass flows were measured using calibrated metering plugs in conjunction with static and total-pressure measurements and free- stream total temperature measurements.
(U) Linear potentiometers were used to measure spike position, spike bleed plug position, and compressor-face flow plug position.
SECTION III
PROCEDURE
3.1 GENERAL
üß The AMSA model was tested in Tunnel 16T at nominal free- stream Mach numbers of 0. 85 and 1. 20. Free-stream total tempera- ture was maintained at 120°F.;: and free ^stream total pressure was
SECLASSOFHE® ' UNCLASSIFIED
AEDC-TR-67-252
DECLASSIFIED I UNCLASSIFIED
maintained at 1560 and 1200 psfa at Mach numbers 0. 85 and 1. 20* respectively. In Tunnel 16S the model was tested at Mach numbers from 1. 75 to 2. 20. Free-stream total temperature and total pressure were maintained at 200°F and 1100 psfa, respectively, through the Mach number range. The model angle of attack and angle of sideslip were varied from 0 to +14 deg and from -8 to +8 deg, respectively. The majority of the data taken with the model in a yawed attitude was taken at a nominal 6. 2-deg angle of attack.
(?) After tunnel free-stream total pressure and Mach number were established, the model was positioned to the desired angle of attack and/or angle of sideslip. Steady-state pressure measurements were obtained by setting the desired spike position and spike bleed flow and setting selected compressor-face corrected weight flow by varying the primary duct plug position. Except when determining the influence of spike bleed flow rate on inlet performance, the spike bleed flow was gradually increased until further increases in the spike bleed flow had no effect on inlet recovery. Inlet stability was determined by monitor- ing the output of the dynamic pressure transducer mounted near the compressor face. The initial indication of pressure fluctuation at the compressor-face station defined the inlet stability limit as presented in this report.
3.2 ACCURACY OF MEASUREMENTS
(U) The probable errors associated with the various tunnel and model parameters are listed in Table I (Appendix II). Since the data presented in this report were obtained from single sample measure- ments, the errors listed in Table I are the estimated uncertainties of the data based upon the accuracy of the instrumentation components and tunnel calibration data.
SECTION IV RESULTS AND DISCUSSION
(J>) Inlet performance in terms of compressor-face total-pressure recovery and flow distortion is presented as a function of spike posi- tion, compressor-face mass-flow ratio, spike boundary-layer bleed mass-flow ratio, angle of attack, angle of sideslip, and free-stream Mach number.
06CLASSffieO.UNCtASS.ret»
AEDC-TR-67-252
DECLASSIFIED / UNCLASSIFIED
(B) The predicted cruise attitude of the AMSA model, as discussed in this report, is assumed to be +5 deg angle of attack and 0 deg angle of sideslip through the Mach number range from 0. 85 to 2. 20. The total-pressure recovery and flow-distortion data presented for the inlet-airplane phase of testing are those values that were obtained for nominal compressor-face corrected "airflows of 260, 260, 220, 196, and 180 lbm/sec at Mach numbers of 0.85, 1.20, 1.75, 2.05, and 2. 20, respectively. During the isolated inlet phase of testing, the majority of data was obtained for a nominal compressor-face corrected airflow of 180 lbm/sec and at Mach number 2. 06, which was estimated as the local Mach number in the inlet flow field generated by the wing- fuselage combination at free-stream Mach number 2. 20.
4.1 ISOLATED INLET TESTS
4.1.1 Spike Bleed Flow
(£) The effect of spike boundary-layer bleed flow on inlet perform- ance at Mach number 2. 06 is presented in Figs. 10 through 12. »The two spike bleed patterns investigated (bleed patterns 5 and 6) provided the same improvement in total-pressure recovery. As shown in Fig. 10, recovery increased and distortion remained essentially unchanged as spike bleed flow was increased to the maximum obtainable value. At the optimum spike position of X/R = 1.63, as determined from Fig. 11, maximum spike bleed flow increased maximum recovery approximately 1 percent for a nominal corrected airflow of 180 lbm/sec. The data presented in Fig. 11 also show that spike position at maximum recovery was less critical with spike bleed flow than without spike bleed flow. As can be interpreted from Fig. 12, spike bleed flow was more effective at below-nominal values of corrected airflow. The data also show that spike bleed flow decreased the inlet stability margin.
(&) Subsequent isolated inlet testing was done with spike bleed pat- tern 6 operating at maximum bleed flow rate. The optimum spike posi- tion was established as X/R = 1. 63 for a corrected airflow of 180 lbm/sec at Mach number 2. 06.
4.1.2 Mutual Interference Effects
iff) The mutual interference effects of the nonstaggered dual-inlet installations are shown in Fig. 13. These data were obtained by setting a desired performance for one of the inlets and varying the mass flow of the other inlet. For clarity the inlets-were identified, as they were
DECLÄSSIF.EO / UNCLASSIFIED
|*Sh
AEDC-TR-67-252
DECLASSIFIED / UNCLASSIFIED commonly referred to when attached to the AMSA model, as either the inboard inlet or the outboard inlet. For the three inlet spacings investi- gated the inboard inlet performance was not degraded as the outboard inlet mass flow was throttled to the stability limit. As the outboard inlet mass flow was throttled past the stability limit, the inboard inlet performance became slightly sensitive to inlet spacing only when the inboard inlet was set for a below-nominal corrected airflow of 170 lbm/sec. Flow distortion was unaffected at all test conditions.
iji) The interference effects of staggering the outboard inlet 3. 5 inlet diameters downstream of the inboard inlet are shown in Fig. 14. Varying the outboard mass flow had no effect on inboard in- let performance; however, varying the inboard mass flow caused out- board inlet recovery degradation of approximately 2 percent at nominal corrected airflow and approximately 4. 5 percent at a below-nominal corrected airflow of 165 lbm/sec. Flow distortion was unaffected for the inboard inlet and was increased slightly for the outboard inlet when the inboard inlet mass flow was throttled past the stability limit.
(§) The dual-inlet configuration selected as the best compromise installation for the inlet-airplane phase of testing was the nonstaggered dual-inlet configuration having the 0. 84-inlet-diameter spacing.
4.2 INLET-AIRPLANE TESTS
4.2.1 Effect of Spike Position
<f>) The effect of spike position on both the inboard and outboard inlet performance at Mach numbers from 0. 85 to 2. 20 at an angle of attack of 5 deg is shown in Fig. 15. The data, presented at constant corrected airflows, were within ±2 lbm/sec of the previously defined nominal corrected airflow for the respective Mach numbers. The data were used to select the spike positions for optimum inlet performance at the assumed cruise angle of attack of 5 deg. The effect of second- cone deflection angle on inlet performance at Mach numbers of 0. 85 and 1. 20 is also shown. Collapsing the second-cone deflection angle from 15 to 10 deg, at the spike position for optimum inlet perform- ance, improved recovery approximately 0. 3 percent at M,,, = 0. 85 and 0. 6 percent at M,,, = 1. 20. At Mach numbers from 1. 75 to 2. 20 the spike position for optimum outboard inlet performance was at a slightly more extended position than that for the inboard inlet. The discrepancy is believed to be caused by slightly different spike and cowl contours of the two inlets since the discrepancy also appeared during the isolated inlet phase of testing.
„EC..AS5.F.ED/ UNCLASSIFIED
AEDCTR-67-252
DECLASSIFIED t Mdltfs&FIED 4.2.2 Effect of Spike Bleed Flow
(0) Limited data were obtained for determining the effect of spike bleed flow on inlet performance. At Mach numbers of 0. 85 and 1. 20,-the spikes contained no perforations to bleed the boundary layer. In general the spike bleed plugs were fully open at the supersonic Mach number conditions investigated. However, the effect of spike bleed was investi- gated at M0 = 2. 20 at the assumed cruise angle of attack of 5 deg. The data presented in Fig. 16 show that recovery was unaffected for spike bleed flow greater than approximately 2 percent of the inlet capture flow. Spike bleed mass-flow ratio above 0.02 increased recovery approximately 2 percent at M,,, = 2. 20 at the cruise angle of attack of 5 deg.
4.2.3 Effects of Angle of Attack and Sideslip
(0) The effects of angle of attack and sideslip on inlet performance, for previously defined corrected airflows at optimum spike settings, are shown in Figs. 17 and 18, respectively. The inboard inlet performance was more sensitive to model attitude than was the outboard inlet. . At Mach numbers of 1. 75, 2.05, and 2.20, the maximum recovery and minimum distortion occurred at angles of attack greater than the 5-deg cruise angle of attack. To move the peak recovery points closer to the 5-deg cruise angle of attack, the data indicate that the inlets should be pitched up to align the inlets to the local flow angularity. As would be expected at positive angles of sideslip, recovery decreased and distor- tion increased because the inlets were in the flow field wake generated by the fuselage.
4.2.4 Optimum Inlet Performance
{$) Figure 19 shows a summary of the optimum inboard and out- board inlet performance for a 5-deg cruise angle of attack over the Mach number range investigated. Optimum recovery varied from 0. 99 at Ma = 0. 85 to 0. 88 at M,,, = 2. 20 while total-pressure distortion remained at or below 0. 1 through the Mach number range.
(rf) Also shown in Fig. 19 is the isolated inlet performance at the same corrected airflows and spike settings as were determined at the respective Mach numbers for the inlet-airplane phase of testing. The increase in performance shown for the inlet-airplane configuration at Mach numbers greater than 1. 20 results from a decrease in the local Mach number ahead of the inlets which was generated by the weak oblique shock system of the wing-fuselage combination, thereby decreas- ing the total-pressure losses associated with the inlet shock system.
DECLASSIFIED / UNCLASSIFIED
AEDC-TR-67-252
DECLASSIFIED!UNCLASSIFIED SECTION V
CONCLUSIONS
(U) A summary of the significant test results of the wind tunnel investigation of a 1/8-scale AMSA inlet model in Tunnel 16T and Tunnel 16S is as follows:
(JO 1. The nonstaggered, dual-isolated inlet configurations had no mutual inlet interaction effects at nominal corrected airflow at free-stream Mach number 2. 06.
ffi) 2. The outboard inlet performance was adversely affected by" the inboard inlet performance for the staggered, dual-isolated inlet configuration at free-stream Mach number 2.06.
Iß) 3. Inboard inlet performance was more sensitive to airplane attitude than was the outboard inlet.
If) 4. Maximum inlet performance did not occur at pre- dicted airplane cruise attitude for the Mach number range from 1. 75 to 2. 20.
oeciASSff«*' .UNCUASSff«*
UNCLASSIFIED r . , AEDCTR-67-252
APPENDIXES I. ILLUSTRATIONS
II. TABLE
UNCLASSIFIED
TUNNEL I6S STATIONS STA. STA ■ 3.17 0.0
STA 26.90
TUNNEL I6T STATIONS NOTE: TUNNEL STATION IN FEET
Fig. 1 Location of AMSA 1/8-Scale Inlet Model in Tunnel 16T and Tunnel 16S Test Sections
a m o r
■>
c/> Cff nfi rn O
c z o r- >
fti a
> m o n
o •
Kl 1/1
o n
C AIRPLANE BLOOOO(RER)
WL 1.042 (REF.)-
WL 1.406
— 0.8891— I •—1.552- *|
WL 1323 —
I* INCIDENCE
BL 00 (REF.)
WL 1.406 WL l.042(REF.)
o m o r
-n
|
C z o
0» (0
o
o I
ro ►O
DIMENSIONS ARE IN FEET
Fig. 2 Sketch of AMSA 1/8-Scale Inlet Model
<n
& a
33.00
119.97- WING SPAN (EXTENDED)
-58.83- WING SPAN (SWEPT)
BL 0.00
*W ■ SW ,K»
WL 11.25
124.92
I WL 17.83 WLI6.I3
Fig. 3 Sketch of the Proposed AMSA
O m o r- > C/J w
m o
> (A W ■n m O
> m o o I H i
ro In KJ
r5
*.
E D C 11573-67 IG-3329-32
a. Tunnel 16T
Fig. 4 Inlet Airplane Model Installation in Test Section
o
o m o
W
m o
o r- >
(A
m o
A E D C 10885-67 IG-3286-32
O rn O
m o c z o r-
rn o
b. Tunnel 16S
Fig- 4 Concluded
> m o n i H JO i
©.
• o>
> m
n
ro
a. Single Inlet
Fig. 5 Isolated Inlet Installation in Test Section
o m o r- >
CO
o c s (A CO
fti O
A E D C 10759-67 IG-3250-32
b. Dual Inlets: Nonstaggered Arrangement
Fig. 5 Continued
n
a 8 (A
o n
a m o
to (A
m D
C Z o 00 00
rn O
c. Dual Inlets: Staggered Arrangement
Fig. 5 Concluded
NACELLE STA
o m o r >
g»
fn o c z o r- > to CO =n Fn o
14.373 (COMPRESSOR FACE)
48.193
NACELLE GEOMETRY
NACELLE INTERNAL EXTERNAL STATION RAD RAO 00 2.176 2 176 0.281 2 161 2 249 0.625 2 209 2 293 1 290 2 263 2 364 1.875 2.301 2 423 2 S00 2.334 2 478 3 125 2 359 2 325 3 750 2 381 2.566 4.375 2 403 2.601 S.OOO 2.421 2 635 5.6 25 2.440 2.664 6 250 2 456 2.694 6 87S 2 471 2.72 1 7.500 2.484 2 730 8.125 2.496 2781 B.7S0 2.503 2.806 9.375 2 513 2 835
10.000 2.528 2.864 10.625 2.535 2.891 1 1 .250 2. 544* 2.920 1 I.87S 2.949 1 2.000 2 794 13.125 3.003 13.750 3.029 14.375 i 3066 15.375 2.54 4* __ 18.375 2.646*
2 2.075 2.646 * 3.395* 29.445 2.646 • 3.393*
NACELLE LIP (VIEW A-A) X T| T2
0 0 ..0.0 00 0013 0.0 12 0.020 0025 0.01« 9.029 0.050 0022 0 039 0 075 0 025' 0.046 0 123 0023* 0.054 0.138 0.025* _ 0.188 0.056 0 250 0.058* 0 270 0.058*
INDICATES POINTS ON STRAIGHT LINE
DIMENSIONS ARE IN INCHES
o m o to to
m o
z o > to to
m O
Fig. 6 Dimensioned Sketch of Model Nacelle
o n
to
SPIKE BOUNDARY-LAYER BLEED ROW NO. 12 3 4 5
*—T T T -r ■*
D fll a
I w
m o BLEED PATTERN
NO BLEED NO. 5 NO. 6
ROWS OPEN
NONE
4, 5, 6 3, 4,5,6
ROWS CLOSED
ALL (SMOOTH)
1,2.3 1,2
o n
I
0.0625 BLEED HOLES 30° TO SPIKE SURFACE
Fig. 7 Spike Boundary-Layer Bleed Pattern
1.005 RAD-
I5"(REF.)
STA. 0.000
a m o r > CO CO
ni o
c z o r > CO CO
rn o
1 0 15 20 24 27
II I| I| I| I| STA RAD STA RAD STA RAO STA RAD STA RAD
+o.o 0.0 +00 0.0 +0.0 0.0 +0.0 0.0 +0.0 0.0 +2.134 0.572 +2.134 0.572 +2.134 0.572 +2.134 0.572 +2.134 0.572 +3.806 0.867 +3.775 1.012 +3.731 1 .191 +3.686 1.263 +3.648 1 .343 *4.046 0.881 »4.018 1 .045 * 3.960 1 .205 «•3.915 1.333 «3.873 1 .429
4.106 0.898 4.106 1.066 4.106 1.241 4.106 1 .374 4.106 1.479 4.710 1.019 4.710 1 .171 4.710 1.325 4.710 1 .441 4.710 1.533 5.335 1.084 5.335 1.224 5.335 1.360 5.335 1 .466 5.335 1.551 5.960 1.121 5.960 1.250 5.960 1 .373 5.960 1 .470 5.960 1.544 6.585 1.141 6.585 1 256 6.585 1.366 6.585 1 .451 6.585 1.516 7.210 1.143 7.210 1.246 7.210 1.341 7.210 1.416 7.210 1.475 7.835 1 .130 7.835 1.219 7.835 1.303 7.835 1.370 7.835 1.416 8.460 1.109 8.460 1.184 8.460 1.254 8.460 1.310 8.460 1.350 9.085 1.075 9.085 1.136 9.085 1.196 9.085 1.244 9.085 1.271 9.710 1.035 9.710 1.084 9.710 1.129 9.710 1.169 9.710 1.189
10.335 0.988 10.335 1.025 10.335 1.059 10.335 1 .086 10.335 1 .100
10.960 0.940 10.960 0.966 10.960 0.987 10.960 1.000 10.960 1.011 11.585 0.894 11.585 0.905 11.585 0.921 11.585 0.916 11.585 0.924 11.951 0.864 11.975 0.864 11.918 0.864 11.980 0.864 12.043 0.864
.OTE: STRAIGHT LINE BETWEEN POINTS MARKED ++. 1.005 RAD. BETWEEN POINTS MARKED + #, SHARP BREAK AT POINT
o m o f
s ni O
■c 3 CO to
ffi o
MARKED * EXCEPT ON 11
DIMENSIONS ARE IN INCHES
27
Fig. 8 Inlet Spike Contours
> m a n
1
Or
AEDC-TR-67-252
DECLASSIFIED:/ UNCLASSIFIED
COMPRESSOR-FACE INSTRUMENTATION
NOTE: DIMENSIONS IN INCHES
COMPRESSOR FACE 40 TOTAL-PRESSURE PROBES
8 STATIC-PRESSURE TAPS
1.875 INCHES FORWARD OF C. F. ON DUCT WALL I BUZZ TRANSDUCER AT 157.5°
Fig. 9 Compressor-Face Station Instrumentation
DECLASSIFIED / UNCLASSIFIED
'■'-'"-'■ , . 22
UNCLASSIFIED
This paj|2^Un£jajjified
JHfi£t-"' -.
AEDCTR-67-252
DECLASSIFIED1UNCLASSIFIED
er UJ
6* or m oc 3 (/) w m o: 0. i -I
P
0.92
CM 0.68
0.84
SPIKE BLEED PATTERN
O NO BLEED D 5 O 6
ISOLATED-SINGLE
z o H
I o UJ
(/) UJ cr 0. I _l
o
0.2
0 0.01 0.02 0.03 0.04
SPIKE BLEED MASS-FLOW RATIO, (mSb/m|>
Fig. 10 ßf Effect of Spike Bleed Flow on Inlet Performance, If27 Spike, M«, = 2.06, a = Odeg, ß = 0 deg, WVfl/S, = 185 lbm/sec
^•."Äi*"»**
DECLASSIFIED / UNCLASSIFIED
23
AEDC-TR-67-252
DECLASSIFIED / UNCLASSIFIED
0.92
SPIKE BLEED PATTERN
O NO BLEED D 5
O 6
0.76
0.4
CM Q
2 0.3
UJ tr (/> c/J UJ tr Q.
O r-
0.2
0.1 ^*ca^w»-^
1.5 1.6 1.7
SPIKE POSITION, (X/R)
Fig. 11 i8j Effect of Spike Position on Inlet Performance;
I,27 Spike, MM = 2.06, a = Odeg, ß = 0 deg,
■ Wv^«2 = 180 lbm/sec
~^S^
ulsSIFIED / UNCLASSIFIED
AEOC-TR-67-252
DECLASSED I U»CLAS$lflfP
SPIKE 8LEED PATTERN
O NO BLEED D 5
O 6 DARK SYMBOL-INLET STABILITY LIMIT
>- ft UJ > O u UJ tr. ui
w z vt UJ (£ Q.
< I- o
0.92
0.88
0.84
0.80
z o h- o r-
UI £C CM D O (/> to UJ rr a.
0.2
0.1
< i- o
^^t 0.4 0.5 0.6 0.7 0.8 0.9
MASS-FLOW RATIO, (m2/m|)
Fig. 12 (8? Effect of Spike Bleed Pattern on Inlet Performance; lj27 Spike, MM = 2.06, a = deg, ß = 0 deg, X/R = 1.63
DECLASSIFIED/ UNCLASSIFIED
25
n
0.96
Q m ?092
CM
^ 0.88
1 i r ISOLATED-DUAL UNSTABLE OPERATION
L 0.2
0.1
(wvV»2)|NB'D O 170 lbm/sec G 180 lbm/sec
DARK SYMBOL-OUTBOARD INLET STABILITY LIMIT
0.96
Q O 5092
CM
0.88
I I ISOLATED-DUAL
UNSTABLE I OPERATION _1
0.96
Q
CO 50.92
CM
0.88
1 1 ISOLATED-DUAL
UNSTABLE OPERATION
"0= H
UNSTABLE —I OPERATION
JU-CKJD" CK>CD
UNSTABLE i OPERATION *1
T~F [DO DO
UNSTABLE 1 OPERATION"
0.2
0.1 CM Q
0.5 0.6 0.7 0.8 0.9 1.0 OUTBOARD MASS-FLOW RATIO,(m2/«n|)oUTB'D
a- 0.70-lnlet-Diameter Spacing
%
UNSTABLE i OPERATION *1
-CDOC D—'
< § UNSTABLE J OPERATION *1
0.2
a m 2 0.1
CM O
1 UNSTABLE 1 OPERATION *i
B8S ^O | J s UNSTABLE 1 OPERATION "*l
0.6 0.7 0 8 0.9 1.0 OUTBOARD MASS-FLOW RATIO.dn^miJouTB'D
b. 0.84-lnlet-Diameter Spacing
05 0.6 0.7 0 8 0.9 1.0 OUTBOARD MASS-FLOW BfiiT^.im^M^jQ'Q
c. 1.00-lnlet-Diameter Spacing
«3 CO
fn a "» c z o
« m o
Fig. 13^9) Mutual Interference Effects of Nonstaggered, Dual-Inlet Orientation on Inlet Performance; 1,27 Spike, MM = 2.06, a = 0 deg, ß = 0 deg, X/R = 1.63
IO
(wV0/82) INB 0 180 lbm/sec
> K 111 > O U Ul <r in P 5 5 (0 £ OT CM Ui z E a. _i
eM ? ^mt o
i^m' »- /■■K K1
SB *" z o r> ^E) »- r^^EcL K P»-JW* £ Cfl y l/i
01 Q Q
*n 111 at
m 5 5 o ::'-S 8 ^"» ui c £ 2 i
Q -i
f ' P * o 0> If» •n rti a
0.92
0.88
0.84
0.2
0.1
UNSTABLE OPERATION c-
ISOLATED-DUAL
L •—O-CHXO
OUTBOARD STABILITY. LIMIT
UNSTABLE OPERATION
0 -I
*== 0=0=00*
a "a>
0.92
0.88
0.84
0.2
Q m
o CM
O
0.5 0.6 0.7 0.8 0.9 1.0
OUTBOARD MASS-FLOW RATIO, ( IT>2 / m , >0UTB'D
(w>/c9/82)OUTBlD
O 180 lb>m7»ec D 165 lbm/sec
ISOLATED-DUAL
INBOARD STABILITY LIMIT
0.5 0.6 0.7 0.8 0.9 1.0
INBOARD MASS-FLOW RATIO, (mg/m,) ^
Fig. 14 jfitf Mutual Interference Effect of Staggered Dual-Inlet Orientation on Inlet Performance; I, 27 Spike, MM ^ 2.06, a = 0 deg, ß = 0 deg, (X/R)|nb.d = 1.63, (X/R)0utb.d = 1.67
S a
ui
P c z o
■ r- > >w ni in o
> m o
o» • IO
KJ
o
1 (A *n Si o
.10
a i. IS
1.00
0.88
0.3
O o HI m S; z
in £> UJ o K a.
o
0.2
0.1 ^-o
"B—^ C -<j
0.4 0.8 1.2 1.6 2.0
INBOARD SPIKE POSITION, (X/R)INB'D
SPIKE
SPIKE
1.00
o 0.96 an i- 3 O
CM Z 0.92
0.88
0.3
a 0.2 *CB
O CM
° 0.1
9= *= N b \
INLET - AIRPLANE
1 \
> m o n
o ■
W m W
c
a m o r- >
m O
c Z O
07 eg Tfl
0.4 0.8 1.2 1.6 2.0
OUTBOARD SPIKE POSITION. (X/R)0UTB<D
a.jm Moo = 0.85, wVfJ/ö, = 260 lbm/sec
Fig. 15 (U) Effect of Spike Position on Inlet Performance; a - 5 deg, ß = 0 deg
O I|10 SPIKE
G 1115 SPIKE
I 00
0.88
0.3
m o
cr o l- u> a p id m
in £? uj o (t a.
o t-
02
0.1
0.4 0.8 1.2 1.6 2-0
INBOARD SPIKE POSITION, (X/R)INB-D
1.00
o 0.96 m 3 O
CM Z 0.92
0.88
03
p 0.2 CD I- 3 O
CM ° 0.1
0=
} ̂U
INLET-AIRPLANE
1
Ml —^^— ___________ - ______________
c/» 0>
s- c- Z ■ o • r > a» m ni m o
0.4 0.8 1.2 1.6 2.0
OUTBOARD SPIKE POSITION, <X/R)<xiTe'D
b. XÄ H» = 1.20, WVö7&2 = 260lbm/Sec
Fig. 15 Continued
o o
n
m o r- >
o •». c z CO
5 <0 CO
a
UJ > O U HI
" o OJ ID
m H « CM UJ ■ a. ■ a.
1.00
0.96
032
O I-
0.88
o
E o I- OT 5 o u m % z
a«? cc a.
■ _i < i- o
0.3
0.2
0.1
fb INLET-AIRPLANE
•Un
O 1120 SPIKE
1.00
Q 0.96 "OD t- 3 O
CM Z 032
0.88
0.3
INLET • AIRPLANE
I w In
o (M o
0.2
0.1
r
o m o r- - >
w CO
: T» ni O
c .35 o CO CO
m o
0.4 0.8 1.2 1.6 2.0
INBOARD SPIKE POSITION. (X/R)IN8-0
0.4 0.8 1.2 1.6 2.0
OUTBOARD SPIKE POSITION, (X/R)0uTB>D
e.jHTM«, 1.75. wVtVfi, = 220 lbra/»« Fig. 15 Continued
O I I24 SPIKE *C .-v-r»
0.92
m o E -.z o .r
r:B ' =n
ft
> o u 111
Of UJ 3 Z U) H « IM IE Ä
CL
< I- O H
0.88
0.84
0.80
0.3
* i ■I INLET-AIRPLANE
I- CC O I- (0 o o uj m T z
in Jf U u a: a.
P
0.Z
0.1
%
-ft-
0.4 0.8 12 1.6 2.0
INB0AR0 SPIKE POSITION, (VR)INB'D
0.92
a 0.88 o i- o
CVI
0 84
0.80
0.3
a m i- 3 O
CM o
0.2
0.1
INLET-AIRPLANE
r»
I Rv
s P" > CO 0»
m a
o p"
s CO
m a
0.4 0.8 1.2 I 6 2-0
OUTBOARD SPIKE POSITION, (X/R)0UTB-0
J-JWM« = 2.05, wVö7s, = 196lbm/..e Fig. 15 Continued
> m
o
CO to
s (HI O
w.
m o c z o
>- CE U > O u 111 ce
a 0; CO
5> H to CM Id z
a.
CO «0
ce P CO
a a Si "«n i z
CO JN UJ o ce a.
o
0.90
0.86
0.82
0.78
0.74
0.4
0.3
0.2
0.1
JL i
INLET-AIRPLANE
1 J
O 1127 SPIKE
0.90
1 >_ —
0.86
CO
3 0.82 o
eg
0.78
0.74
0.4
03
INLET-AIRPLANE
1 o
n
••J
K> In
a m o
CO I-
o CM a
0.2
0.1
1 , -
I-
22 9
(»9 01
9
0.4 06 1.2 1.6 2.0
INBOARD SPIKE POSITION, (X/R)INB-D
0.4 0.8 1.2 1.6 2.0
OUTBOARD SPIKE POSITION. (X/R)QuTB>D
•■& M,,«, = 2.20, W\Zö7S2 - 180 lbm/sec
Fig. 15 Concluded
AEDC-TR-67-252
DECLASSIFIED / UNCLASSIFIED
O INBOARD INLET, X/R=l.63 D OUTBOARD INLET, X/R=l.67
m >- UJ > 0.92 o u a: INLET-AIRPLANE UJ 01 3 (M 0.88 O— —O-Oh-ffm
-PR
ES
S
N
_i < 0.84
| I— O
z g H
o H CO
LJ a: cvi 3 o co CO UJ a: a.
r- O
0.2
0 0.01 0.02 0.03 0.04 SPIKE BLEED MASS-FLOW RATIO, (mst>/nri|)
Fig. 16 yt) Effect of Spike Bleed Flow on Inlet Performance; 1, 27 Spike, M«, = 2.20, a = +5 deg, ß = 0 deg,
WVÖ7S2 = 180 lbm/sec
DECLASSIFIED / UNCLASSIFIED 33
AEDC-TR-67-252
O D o A A
oeSsK'"NCLA8Sff,EO
Mao wv^/82 (X/RIJNB'D 0.85 1.20 1.75 2.05 2.20
260 260 220 196 180
0.60 0.60 1.60 1.63 163
(X/RIOUTBTD SPIKE
I,10
I ,0 j 20 I 24
1.67 ij27
0.60 0.60 1.63 1.67
DARK SYMBOL - INLET STABILITY LIMIT
1.00
yS 0.96 DC UJ > o o 111 ac
Q
§ zO.88 V>^ UIZ
OS *V 0.84
0.92
g 0.80
0.76
Ü—c ̂ r—TL
\, ^
c^| r \
C^ A
INLET-AIRPLANE
0.76
0.5
0.4 ae
to 5 u a 0.3
0.2 2^ tc a.
■
< 0.1 o
/
7 #p- «■,- "-»OB« /
/
A—=1 _A _ /*
r 2>— X*3-
0 4 8 12 16 ANGLE OF ATTACK, CL
0 4 8 12 16 ANGLE OF ATTACK, a
Fig. 17 Effect of Angle of Attack on Inlet Performance; ß = 0 deg
34
DEGLASSlF ütois-eo'UNCLASS,F,eD
AEDCTR-67-252
DECLASSIFIED / UNCLASSIFIED
Moo Wv^/82 <X/R)|NB'D (X/R)0UTB'D SPIKE
O 0.85 260 0.60 □ 1.20 260 0.60 O 1.75 220 1.60 A 2.20 180 1.63
?l 10 0.60 0.60 163 -,,7 1.67 I 27
Ii10 I 20
0.96
0.92
m
CM 0.88
0.84
0.80
0.76
•4048 ANGLE OF SIDESLIP, ß
8-4048 ANGLE OF SIDESLIP, ß
Fig. 18 Effect of Angle of Sideslip on Inlet Performance; a = 6.2 deg
DECLASSIFIED / UNCLASSIFIED
35
CE ÜJ > O o Ui cr
1.00
0.96
0.92
5/. V)
a. Q.
5 o
a
so 69
m
Z g CE
P </> Q
LU CC
CO CO LU tr a. i
_i
o
0.88
0.84
0.80
0.2
""*=*
V \
O INBOARD INLET
\
\ D OUTBOARD
INLET \
SINGLE ISOLATED INLET
AT a = 0 DEG
\ \ \
s \
\
(M 0. I
> m O n i H 73
\ dt ä
I
0.8
\
I
Fig.
1.4 1.6 1.8
FREE-STREAM MACH NUMBER, Mm
19 Optimum Inlet Performance; a = +5 «leg, ß = 0 lieg
2.2 2.4
UNCLASSIFIED AEDC-TR-67-252
TABLE I
ACCURACY OF MEASUREMENTS
Uncertainty
Free-stream Mach number, Ma Tunnel 16T ±0.006 Tunnel 16S +0.010
Free-stream total pressure, pt , psf ±3.000
Compressor-face total-pressure recovery, N2 ±0.004
Compressor-face total-pressure distortion, D2 ±0.008
Compressor-face mass-flow ratio, mg/m^ ±0.010
Spike bleed mass-flow ratio, ra^/mj ±0.001
Simulated full-scale corrected weight flow, W^/ög, lbm/sec ±2.000
Model total pressure, psf ±4. 000
Model static pressure, psf ±1. 500
Ratio of distance spike tip extends forward of cowl lip to cowl radius, X/R ±0. 003
Free-stream total temperature, T+ , °R ±5.000
Model angle of attack, a, deg ±0. 150
Model angle of sideslip, ß, deg ±0. 150
37
UNCLASSIFIED
UNCLASSIFIED Security Classification DECLASSIFIED/ UNCLASSIFIED
DOCUMENT CONTROL DATA -R&D (Security classification of fftfe, body of abstract and indexing annotation must be entered when thm overall report Is classified)
I. ORIGINATING ACTIVITY (Corporate author)
Arnold Engineering Development Center ARO, Inc., Operating Contractor Arnold Air Force Station, Tennessee
2a. REPORT SECURITY CLASSIFICATION
2b. GROUP
3 REPORT TITLE
WIND TUNNEL INVESTIGATION OF A l/8-SCALE AMSA AIRCRAFT-INLET MODEL AT TRANSONIC AND SUPERSONIC MACH NUMBERS (U)
4. DESCRIPTIVE NOTES (Type oi report and Inclusive dates) Final Report - August 7 through September 9, 1967
5- AUTHORIS) (First name, middle initial, laat name)
Lawrence L. GaTLigher, ARO, Inc
6. REPORT DATE
November 1967 7a. TOTAL NO. OF PAGES
46 76. NO. OF REFS
1 Ba. CONTRACT OR GRANT NO.
AF40(600)-1200 b. PROJECT NO. 139A
e.Program Element 6340683F
9a. ORIGINATOR'S REPORT NUMBERIS)
AEDC-TR-67-252
9b. OTHER REPORT NOISI (Any other number» that may bo aaalgned this report)
N/A 10. DISTRIBUTION STATEMENT^
foreWn rageig (A/sag), "*-- merit
II. SUPPL
Available in DDC
12. SPONSORING MILITARY ACTIVITY
Aeronautical Systems Division, Air Force Systems Command, Wright- Patterson AFB, Ohio
13. ABSTRACT
Test results are presented for a 1/8-scale inlet model of a proposed AMSA air induction system at transonic and supersonic Mach numbers. The effects of spike bleed pattern and inlet orientation are presented. Compressor-face total-pressure recovery and flow distortion data are presented as a function of spike position, compressor-face mass-flow ratio, spike boundary-layer bleed mass-flow ratio, angle of attack, angle of sideslip, and free-stream Mach number. (U) — — ■,
Distribution limited to U. S. Gov't. Agencies 'ltPolerfsQd only; Test and Evaluation; Nov^j||^^&er ^ei f | e«cK4
nf tiff 6ASZI
requests for this docujp^^WrT-ue referred to Commander, AeoM#Ȁical Systems Div., Attn: YHT^a^B^t-Patterson AFB, Ohio l^lffi^jg^TAB 73-2, 15 January, 1973-
peign o-e^ASD
DD ,Fr.,1473 UNCLASSIFIED DECLASSIFIED / UNCLASSIFIED Security Classification
UNCLASSIFIED Security Classification
14. KEY WORDS
LINK A LINK B LINK C
ROLE »T ROLE WT ROLE WT
ff . 4 air induction systems "* AMSA
wind tunnel tests spike bleed patterns air inlets pressure recovery flow distortion mass-flow ratio angle of attack effects angle of sideslip effects Mach number effects transonic flow
^^* ' l IV-SA J*" supersonic flow . W ̂ ■
,t Ay
,/^
'■•<t ***
^
)
0 1 ' " 11 H %
r l
-9
^"'-d^ ̂
u 7 i
if n syiSp
5, "
I-* *
*-<*i ' '> ■\t
*
UNCLASSIFIED -Security Classification
