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~A7OAOBA511 KAMAN AVIOSNE BURLINGTON MA F/6 21/5F URTHER EVALUATI'ON OF BLAST TESTS OF AN ENGINE INLET. CU)
MiAR 79 J R RUETENIK. R F SMILEY. M A TOMAYKO DNAO-78-C-0239
UNCLASSIFIED KATR-16 DNA-4994F N3ffffffffffff
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DNA 4994F
Oo FURTHER EVALUATION OF BLAST
o TESTS OF AN ENGINE INLET
:Kaman AviDyne
B3 Seqond Avenue
IBurlington, Massachusetts 01803
31 March 1979
Final Report for Period 1 April 1978-31 March 1979
CONTRACT No. DNA 001-78-C-0239
APPROVED FOR PUBLIC RELEASE;DISTRIBUTION UNLIMITED.
THIS WORK SPONSORED BY THE DEFENSE NUCLEAR AGENCY
UNDER RDT&E RMSS CODE B342078464 N99QAXAJ50401 H2590D.
DTICPrepared for ELECTE
Director SEP2 191M
DEFENSE NUCLEAR AGENCY U! Washington, D. C. 20305 B
808 4 137
Destroy this report when it is no longerneeded. Do not return to sender.
PLEASE NOTIFY THE DEFENSE NUCLEAR AGENCY,ATTN: TISI, WASHINGTON, D.C. 20305, IFYOUR ADDRESS IS INCORRECT, IF YOU WISH TOBE DELETED FROM THE DISTRIBUTION LIST, ORIF THE ADDRESSEE IS NO LONGER EMPLOYED BYYOUR ORGANIZATION.
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REPORT DOCUMENTATION PAGE BEFORE COMPLETING FORM
I. NMBE 12. GOVT AC.CESSION NO. 3 RECIP NT S CATALOG NUMBER
DNA494F. . --. PERIOD COVERE
-F1RTHER VALUATION OF BLAST TESTS OF AN" FinalRpar-for PeriodENGINE ADLET, ADDR 1 Apr 78-31 MarOCT-------------.... . .. . . A-E RFQBA
R G . R EPO ' N UMBER~
KAt -TR-165AREA W- GRANT NUMBER(s)
FRobert F. SiIe DNA,01-78-C-0239Michel .//omak o
-- ha om yk 1i r~~XIm/V mI*KIr AME AND ADDRESS 10 PROGRAM ELEMENT PROJECT, TASK
Kaman AviDyne AREA & WORK UNIT
83 Second Avenue SubtakbQAXA ]504-01Burlington, Massachusetts 01803 _ ___
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II. CONTROLLING OFFICE NAME AND ADDRESS 12- RIeT DATE
Director 31 Marc 479 _Defense Nuclear Agency .13 NUMBER OF PAGES
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(. /UNCLASSIFIED/ .ISa. DECLASSIFICATION 'DOWNGRADING
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16. DISTRIBUTION STATEMENT (of this Report)
Approved for public release; distribution unlimited.
17. DISTRIBUTION STATEMENT (of the abstract entered in Block 20, It different from Report)
1. SUPPLEMENTARY NOTES
This work sponsored by the Defense Nuclear Agency under RDT&E RMSS CodeB342078464 N99QAXAJ50401 H2590D.
19. KEY WORDS (Continue on reverse side if necessary and Identify by block number)
Aircraft Engine ShockBlast Experimental Test Shock TubeB-i Inlet SubsonicComputer Studies Pressure Wind Tunnel
20. ABSTRACT (Continue on reverse side If necessary and Identify by block number)
A test program was conducted to simulate blast wave intercepts with a scaledaircraft engine in subsonic flight. Initial results were reported in DNA4590F. This report presents additional test data and evaluated results.
Jumps in total pressure at the engine face produced by blast interaction withthe blastward and leeward inlets are evaluated with respect to effects offour test variables: inlet weight flow, Mach number, shock overpressure and
DD fJoA 1473 EDITION OF I NOV 65 IS OBSOLETE UNCLASSIFIED
SECURITY CLASSIFICATION OF THIS PAGE (When DeaEnired) 4
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UNCLASSIFIEDSECURITY CLASSIFICATION OF THIS PAGE(Whtin Data Entered)
0. ABSTRACT (Continued)
intercept angle. The total pressure rose more rapidly in the blast inlet thanthe leeward inlet, following blast intercept, and by a larger amount. In manytests the rise in the leeward inlet was followed by a rapid fall-off, indi-cating possible flow separation in the inlet walls.
The two-dimensional Blast Induced Distortion (BID) code provides good repre-sentation of effects essentially two-dimensional and inviscid, such as ramp,cowl and engine-face pressures of the blastward inlet.
Blast wave reflection from the engine is expected to be a potential cause ofdistortion through boundary-layer retardation. The effect on engine-facetotal pressure is examined with respect to the four test variables. Improvedcalculations of the effect of the engine-reflected blast wave on the boundarylayers in the inlets, made using the generalized NASA-Lewis BLAYER code,indicate that boundary-layer separation would result.
UNCLASSIFIEDSECURITY CLASSIFICATION OF THIS PAGE(17ten Date Entered)
Conversion factors for U.S. customaryto metric (SI) units of measurement.
To Convert From To Molt irls Rv
angstrom me'ter. 1n 0,t 1 00 X S I n1
atmosphere (normal) kln pas..ii (kl'a) 1.011 215 XC EC +2
ha kiln pa.,.1 (kPs i.0,10 '0 X EC 4.
barn mcer Wn) 1.000 000 C EC -."
British thermal unit (thermochemieal) Joule (J) 1.0Sf. 350 XC E +3
calorie (thermochemical bole, (J1) 4.184. 00n
cal (thermochemical)/cm2 met., joule/m, (MI/.') 4.18. 00n X E -2
curie gtig bcqocrel )Cq) .Ion 000 X r +i
degree 0-ngi0 radian (cad) 1.745 329 IC I -2
degree Fahrenheit degree kelvin (K) Wt f + 45q.67l/l.S
electron volt joule (J) 1.602 19 XC F -19
erg jool e (J) 1.000 000 X F. -7
ergt/.econnd watt (W) 1.0 no 0 OM x F -7
(not meter In) 1.0~ 00 " fton x -1
loot-pound-force joule (.3) i.lS5 ale
gallon (U.S. liqu.id) mterd We) 3.705, 412 X F -3
inch meter (.n) 2.%',n 000 X EC -2
jerk joule (J) Loon0 000o x E 49
JOUle/klOgR. (J/kg) (radiat ion dose
,absorbed) Cray ICV),* 1.000 000
kilotons teraJ.ules 4.1R)
hip (1000 lbf) ne..on WN 4..0 222 X E 43
hip/inch2 (kni) kilo ponil (hI~a) 6.894 757 XC E +3
ktap ncwton-occond/m2
(N-sis') 1.00n 000 x F +2
micron metr,* (oI.001 000 C IC -6
mil Pltr In)'. 2 n 00X F -$
mile (international) meter In I.1.09 34. X IC +3
ounce kilogram (hgl 2.814 952 X IC -2
pound-forte (lhf avoirdupoin) newton (N) 4.416A .2
pound-force inch newton-meter (Nm) 1.129 SO9 X E -1
pound-force/inch newton/meter (N/in) 1.7i1 26A X IC +2
pound-forccffont2 kiln pascal WI.) 4.788 026 X IC -2
pound-force/inch7
(psi) kilo Pascal (kPa) 6.894 757
pound-mann (Ibis avoirdupois) hilogra. (kig) 4.535 924 X IC -1
pound-mass-root2 (moment of Inertial kilogr3m-meter2
(khgm',) 4.2114 il X IC -2pound-mans/oot
1 ktilram/meter
(kg/n1) 1.601 A!, f, x F +1
red (radiat ion dose absorbed) Cray (rCy)* 1.000 000 C IC -?
roentgen rolooh/kilgream (C/k;) ?.579 760 XC F -4
shake s"eond (s) 1.000 000 IC IC -8
plus kilogram (kgt) 1.459 390 C IC +1
tort (wi Hg, 0* C) kilo Pa,ral NkP.) 1.311 22 X IC -1
*The becquorel (8q) in the Si unit of radinsettvitl; I Bq - I event/s.**The Cray (Cy) In the SI unit of absorbed radiation.
A mnts Complete listing of conversions may he found in "Metric Practice Guide E 380-74,American Society for Testi-ug and Materials.
PREFACE
This work was performed by the AviDyne Division of the Kaman
Sciences Corporation for the Defense Nuclear Agency under Contract
DNA-O0l-78-C-0239. CAPT J. Michael Rafferty of the DNA Shock Physics
Directorate served as technical monitor.
Dr. J. Ray Ruetenik of Kaman AviDyne was the project leader under
Dr. Norman P. Hobbs, Technical Director of KA. Mr. Robert F. Smiley
performed engineering functions.
Appreciation is expressed to CAPT Rafferty for his continuing
interest and support of this program.
ACCESSION for
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DISTIO)UII(AYROWIJ CMEDis? I - 4ii. /or 3PECI.
2
Lei= N
TABLE OF CONTENTS
Section Page
1 INTRODUCTION --------------------- 15
2 RESPONSE OF MEAN TOTAL PRESSURE AT ENGINE FACE - - - - 20
2-1 GENERAL FEATURES OF EFFECT ON MEAN TOTALPRESSURE --------------------- 20
2-2 MACH 0, 0 LB/S FULL-SCALE WEIGHT-FLOW ------ 232-3 MACH 0.55, 235 LB/S FULL-SCALE WEIGHT FLOW- - - 232-4 MACH 0.55, 350 LB/S FULL-SCALE WEIGHT FLOW- - - 242-5 MACH 0.70, 300 LB/S FULL-SCALE WEIGHT FLOW- - - 252-6 MACH 0.70, 350 LB/S FULL-SCALE WEIGHT FLOW- - - 262-7 MACH 0.70, 350 LB/S, ±50 YAW ----------- 282-8 MACH 0.85, 300 LB/S FULL-SCALE WEIGHT FLOW- - - 282-9 MACH 0.85, 350 LB/S FULL-SCALE WEIGHT FLOW- - - - 29
2-10 MACH 0.90, 350 LB/S FULL-SCALE WEIGHT FLOW- - - - 29
3 GENERAL FEATURES OF BLAST AND OPERATIONAL EFFECTSON ENGINE-FACE MEAN TOTAL-PRESSURE SIGNATURE ----- 56
3-1 EFFECT OF BLAST SHOCK OVERPRESSURE -------- 563-2 EFFECT OF BLAST INTERCEPT ANGLE --------- 58
3-3 EFFECT OF WEIGHT FLOW -------------- 583-4 EFFECT OF MACH NUMBER -------------- 59
4 SPECIFIC EFFECTS OF BLAST AND OPERATIONAL VARIABLES ON
ENGINE-FACE MEAN TOTAL PRESSURE ------------ 116
4-1 VARIATION IN FIRST PEAK OF MEAN TOTAL PRESSUREFOR BLASTWARD INLET --------------- 1164-1.1 Effect of Incident Total-Pressure
Increment --------------- 1164-1.2 Effect of Test Variables on Engine-Face
First-Peak Ratio, ------------ 1174-1.2.1 Shock Overpressure Effect- - - 1174-1.2.2 Shock Intercept Angle Effect - 1174-1.2.3 Weight Flow Effect ------ 1184-1.2.4 Mach Number Effect ------ 118
4-2 VARIATION IN SECOND PEAK OF MEAN TOTAL PRESSUREAT ENGINE FACE FOR BLASTWARD INLET- ------- 1184-2.1 Effect of Shock Overpressure ------ 1194-2.2 Effect of Intercept Angle ------- 1194-2.3 Effect of Weight Flow --------- 1204-2.4 Effect of Mach Number --------- 120
4-3 SUMMARY FOR BLASTWARD INLET ----------- 120
3
TABLE OF CONTENTS (CONCLUDED)
Section Page
4-4 FEATURES OF MEAN TOTAL PRESSURE FOR LEEWARDINLET -- --------------------- 121
4-5 FLOW SEPARATION WITHIN LEEWARD INLET -- ------ 1214-6 CONCLUSIONS- ------------------- 125
5 COMPARISON OF THEORY AND EXPERIMENT -- --------- 139
5-1 BLAST INPUT CONDITIONS -- ------------- 1395-2 INLET PRESSURES- ----------------- 1415-3 ENGINE FACE PRESSURES- -------------- 141
5-3.1 Blastward Pressures- ---------- 1425-3.2 Lpeward Pressures- ----------- 142
5-4 DISTORTION AT ENGINE FACE- ------------ 1435-4.1 IDC- ------------------ 1435-4.2 IDR- ------------------ 1455-4.3 IDL-------- ----- - - --- -- -- -- - - 1455-4.4 IDA- ------------------ 1455-4.5 IDT- ------------------ 146
5-5 CORRELATION SUMMARY- --------------- 146
6 EVALUATION OF LATE TIMlE LARGE DISTORTION VALUES- - -240
7 REFLECTED SHOCK-BOUNDARY LAYER INTERACTION- -- ---- 242
7-1 SHOCK-BOUNDARY LAYER CALCULATION -- -------- 243
8 CONCLUSIONS- -- -------------------- 250
REFERENCES -- --------------------- 252
APPENDIX -TRANSDUCER CALIBRATION EVALUATION- -- --- 253
4
LIST OF ILLUSTRATIONS
Figure Page
1.1 Sketch showing typical shock wave pattern and
transducer locations in the inlets ---------- 18
1.2 Engine face transducer locations ----------- 19
2.1 Typical engine-face mean total pressures. Run 8.
Mach 0.70, W2R=300 lb/s (nom), Aps=5.0 psi, =98 deg.- 30
2.2 Engine-face mean total pressures. Run 1. Mach 0,W2R=0 lb/s (nom), APs=2.7 psi (nom), 0=79 deg. 31
2.3 Engine-face mean total pressures. Run 2. Mach 0.55,
W2R=235 lb/s (nom), APs=4.7 psi (nom), f=91 deg. - - - 32
2.4 Engine-face mean total pressures. Mach 0.55,
W2R=350 lb/s (nom) ------------------ 33
2.5 Engine-face mean total pressures. Mach 0.70,
W2R=300 lb/s (nom) ------------------ 35
2.6 Engine-face mean total pressures. Mach 0.70,W2R=350 1b/s (nom) ------------------ 38
2.7 Engine-face mean total pressures. Mach 0.70,
W2R=350 lb/s (nom) ----------------- - 44
2.8 Engine-face mean total pressures. Mach 0.85,W2R=300 lb/s (nom) ------------------ 46
2.9 Engine-face mean total pressures. Mach 0.85,W2R-350 lb/s (nom) ------------------ 51
2.10 Engine-face mean total pressures. Mach 0.90,W2R-350 lb/s (nom) ------------------ 54
3.1 Effect of blast shock overpressure on engine-face meantotal pressure in blastward inlet at Mach 0.70.
Weight flow 350 lb/s (nom). Shock Tube 1 ------- 60
3.2 Effect of blast shock overpressure on engine-face meantotal pressure in blastward inlet at Mach 0.85.Weight flow 300 lb/s (nom). Shock Tube 1 ------- 61
3.3 Effect of blast shock overpressure on engine-face meantotal pressure in blastward inlet at Mach 0.70.Weight flow 350 lb/s (nom). Shock Tube 2 ------- 62
5
LIST OF ILLUSTRATIONS (CONTINUED)
Figure Page
3.4 Effect of blast shock overpressure on engine-face meantotal pressure in blastward inlet at Mach 0.85.Weight flow 350 lb/s (nom). Shock Tube 2 ------- 63
3.5 Effect of blast shock overpressure on engine-face meantotal pressure in blastward inlet at Mach 0.70.Weight flow 350 lb/s (nom). Shock Tube 3 ------- 64
3.6 Effect of blast shock overpressure on engine-face meantotal pressure in blastward inlet at Mach 0.85.Weight flow 300 lb/s (nom). Shock Tube 3 ------- 65
3.7 Effect of blast shock overpressure on engine-face meantotal pressure in blastward inlet at Mach 0.85.Weight flow 350 lb/s (nom). Shock Tube 3 ------- 66
3.8 Effect of blast shock overpressure on engine-face meantotal pressure in leeward inlet at Mach 0.70. Weightflow 350 lb/s (nom). Shock Tube 1 ---------- 67
3.9 Effect of blast shock overpressure on engine-face meantotal pressure in leeward inlet at Mach 0.85. Weightflow 300 lb/s (nom). Shock Tube 1 ---------- 68
3.10 Effect of blast shock overpressure on engine-face mean
total pressure in leeward inlet at Mach 0.70. Weightflow 350 lb/s (nom). Shock Tube 2 ---------- 69
3.11 Effect of blast shock overpressure on engine-face meantotal pressure in leeward inlet at Mach 0.85. Weightflow 350 lb/s (nom). Shock Tube 2 ---------- 70
3.12 Effect of blast shock overpressure on engine-face meantotal pressure in leeward inlet at Mach 0.70. Weightflow 350 lb/s (nom). Shock Tube 3 ---------- 71
3.13 Effect of blast shock overpressure on engine-face meantotal pressure in leeward inlet at Mach 0.85. Weightflow 300 lb/s (nom). Shock Tube 3 ---------- 72
3.14 Effect of blast shock overpressure on engine-face meantotal pressure in leeward inlet at Mach 0.85. Weightflow 350 lb/s (nom). Shock Tube 3 ---------- 73
3.15 Effect of blast intercept angle on engine-face meantotal pressure in blastward inlet at Mach 0.55.Weight flow 350 lb/s (nom). Blast shock overpressure
3.8 psi (nom) --------------------- 74
6
LIST OF ILLUSTRATIONS (CONTINUED)
Figure Page
3.16 Effect of blast intercept angle on engine-face meantotal pressure in blastward inlet at Mach 0.70.Weight flow 300 lb/s (nom). Blast shock overpressure
2,7 psi (nom) --------------------- 75
3.17 Effect of blast intercept angle on engine-face mean
total pressure in blastward inlet at Mach 0.70.Weight flow 350 lb/s (nom). Blast shock overpressure
2.9 psi (nom) --------------------- 76
3.18 Effect of blast intercept angle on engine-face meantotal pressure in blastward inlet at Mach 0.70.Weight flow 350 lb/s (nom). Blast shock overpressure
4.9 psi (nom) --------------------- 77
3.19 Effect of blast intercept angle on engine-face mean
total pressure in blastward inlet at Mach 0.85.Weight flow 300 lb/s (nom). Blast shock overpressure
4.4 psi (nom) --------------------- 78
3.20 Effect of blast intercept angle on engine-face meantotal pressure in blastward inlet at Mach 0.85Weight flow 300 lb/s (nom). Blast shock overpressure
4.9 psi (nom) --------------------- 79
3.21 Effect of blast intercept angle on engine-face meantotal pressure in blastward inlet at Mach 0.85.Weight flow 350 lb/s (nom). Blast shock overpressure
4.0 psi (nom) --------------------- 80
3.22 Effect of blast intercept angle on engine-face mean
total pressure in blastward inlet at Mach 0.90.Weight flow 350 lb/s (nom). Blast shock overpressure
3.7 psi (nom) --------------------- 81
3.23 Effect of blast intercept angle on engine-face mean
total pressure in leeward inlet at Mach 0.70. Weightflow 300 lb/s (nom). Blast shock overpressure
2.7 psi (nom) --------------------- 82
3.24 Effect of blast intercept angle on engine-face mean
total pressure in leeward inlet at Mach 0.70. Weightflow 350 lb/s (nom). Blast shock overpressure
2.9 psi (nom) ------ ------ ---------- 83
3.25 Effect of blast intercept angle on engine-face meantotal pressure in leeward inlet at Mach 0.70. Weight
flow 350 lb/s (nom). Blast shock overpressure
4.9 psi (nom) --------------------- 84
7
LIST OF ILLUSTRATIONS (CONTINUED)
Figure Page
3.26 Effect of blast intercept angle on engine-face meantotal pressure in leeward inlet at Mach 0.85. Weightflow 300 lb/s (nom). Blast shock overpressure4.9 psi (nom)- -------------------- 85
3.27 Effect of blast intercept angle on engine-face meantotal pressure in leeward inlet at Mach 0.85. Weightflow 300 lb/s (nom). Blast shock overpressure4.4 psi (nom)- -------------------- 86
3.28 Effect of blast intercept angle on engine-face meantotal pressure in leeward inlet at Mach 0.85. Weightflow 350 lb/s (nom). Blast shock overpressure4.0 psi (nom)- -------------------- 87
3.29 Effect of blast intercept angle on engine-face meantotal pressure in leeward inlet at Mach 0.90. Weightflow 350 lb/s (nom). Blast shock overpressure3.7 psi (nom) --------------------- -88
3.30 Effect of weight flow on engine-face mean totalpressure in blastward inlet at Mach 0.70. ShockTube 1. Shock overpressure 2.8 psi (nom) ------- 89
3.31 Effect of weight flow on engine-face mean totalpressure in blastward inlet at Mach 0.85. ShockTube 1. Shock overpressure 3.3 psi (nom) -0------
3.32 Effect of weight flow on engine-face mean totalpressure in blastward inlet at Mach 0.70. ShockTube 2. Shock overpressure 2.7 psi (nom)------ - 91
3.33 Effect of weight flow on engine-face mean totalpressure in blastward inlet at Mach 0.70. ShockTube 2. Shock overpressure 5.0 psi (nom) ------- 92
3.34 Effect of weight flow on engine-face mean totalpressure in blastward inlet at Mach 0.85. ShockTube 2. Shock overpressure 3.9 psi (nom)- ------ 93
3.35 Effect of weight flow on engine-face mean totalpressure in blastward inlet at Mach 0.70. ShockTube 3. Shock overpressure 3.0 psi (nom) ------- 94
3.36 Effect of weight flow on engine-face mean totalpressure in blastward inlet at Mach 0.70. ShockTube 3. Shock overpressure 4.3 psi (nom) ------- 95
8
LIST OF ILLUSTRATIONS (CONTINUED)
Figure Page
3.37 Effect of weight flow on engine-face mean totalpressure in blastward inlet at Mach 0.85. Shock
Tube 3. Shock overpressure 4.4 psi (nom) ------- 96
3.38 Effect of weight flow on engine-face mean totalpressure in leeward inlet at Mach 0.70. Shock Tube 1.
Shock overpressure 2.8 psi (nom) ----------- 97
3.39 Effect of weight flow on engine-face mean totalpressure in leeward inlet at Mach 0.85. Shock Tube 1.
Shock overpressure 3.3 psi (nom) ----------- 98
3.40 Effect of weight flow on engine-face mean totalpressure in leeward inlet at Mach 0.70. Shock Tube 2.
Shock overpressure 2.7 psi (nom) ----------- 99
3.41 Effect of weight flow on engine-face mean totalpressure in leeward inlet at Mach 0.70. Shock Tube 2.
Shock overpressure 5.0 psi (nom)- - ---------100
3.42 Effect of weight flow on engine-face mean totalpressure in leeward inlet at Mach 0.85. Shock Tube 2.
Shock overpressure 3.9 psi (nom) -i---------- 101
3.43 Effect of weight flow on engine-face mean totalpressure in leeward inlet at Mach 0.70. Shock Tube 3.
Shock overpressure 3.0 psi (nom) ----------- 102
3.44 Effect of weight flow on engine-face mean totalpressure in leeward inlet at Mach 0.70. Shock Tube 3.
Shock overpressure 4.3 psi (nom) ---------- - 103
3.45 Effect of weight flow on engine-face mean totalpressure in leeward inlet at Mach 0.85. Shock Tube 3.
Shock overpressure 4.4 psi (nom) ----------- 104
3.46 Effect of Mach number on engine-face mean total
pressure in blastward inlet. Weight flow 300 lb/s(nom). Shock Tube 1. Blast shock overpressure
2.8 psi (nom) -------- ------------- 105
3.47 Effect of Mach number on engine-face mean totalpressure in blastward inlet. Weight flow 350 lb/s(nom). Shock Tube 1. Blast shock overpressure
3.7 psi (nom) ---------------------- 106
9
LIST OF ILLUSTRATIONS (CONTINUED)
Figure Page
3.48 Effect of Mach number on engine-face mean totalpressure in blastward inlet. Weight flow 350 lb/s(nom). Shock Tube 2. Blast shock overpressure3.9 psi (nom) - --------------- 107
3.49 Effect of Mach number on engine-face mean totalpressure in blastward inlet. Weight flow 350 lb/s(nom). Shock Tube 2. Blast shock overpressure3.9 psi (nom) --------------------- 108
3.50 Effect of Mach number on engine-face mean totalpressure in blastward inlet. Weight flow 300 lb/s(nom). Shock Tube 3. Blast shock overpressure4.4 psi (nom) --------------------- 109
3.51 Effect of Mach number on engine-face mean totalpressure in blastward inlet. Weight flow 350 lb/s(nom). Shock Tube 3. Blast shock overpressure4.2 psi (nom)- - ----------------- - 110
3.52 Effect of Mach number on engine-face mean totalpressure in leeward inlet. Weight flow 300 lb/s(nom). Shock Tube 1. Blast shock overpressure
2.8 psi (nom)- -i- - - ---------------- 111
3.53 Effect of Mach number on engine-face mean totalpressure in leeward inlet. Weight flow 350 lb/s(nom). Shock Tube 1. Blast shock overpressure3.7 psi (nom) --------------------- 112
3.54 Effect of Mach number on engine-face mean totalpressure in leeward inlet. Weight flow 350 lb/s(nom). Shock Tube 2. Blast shock overpressure3.9 psi (nom)- ------------------- 113
3.55 Effect of Mach number on engine-face mean totalpressure in leeward inlet. Weight flow 300 lb/s(nom). Shock Tube 3. Blast shock overpressure4.4 psi (nom) --------------------- 114
3.56 Effect of Mach number on engine-face mean totalpressure in leeward inlet. Weight flow 350 lb/s(nom). Shock Tube 3. Blast shock overpressure4.2 psi (nom) --------------------- 115
4.1 Sketch illustrating first and second peaks of engine-face mean total pressure for blastward inlet ----- 127
10
LIST OF ILLUSTRATIONS (CONTINUED)
Figure Page
4.2 Increment in mean total pressure at blastward engine-face versus incident total pressure increment ----- 128
4.3 Engine-face peak ratio versus incident overpressure- 129
4.4 Engine-face peak ratio versus blast intercept angle- - 131
4.5 Engine-face peak ratio versus weight flow- ------ 133
4.6 Engine-face peak ratio versus Mach number ------- 136
4.7 Sketch illustrating features of engine-face mean total
pressure for leeward inlet -------------- 138
5.1 Blast input characteristics for Run 39 (Part 544)- - - 147
5.2 Blast input characteristics for Run 40 (Part 619)- - - 148
5.3 Blast input characteristics for Run 18 (Part 624)- - - 149
5.4 Comparison of theoretical and experimental time
histories of ramp pressures for Run 39 (Part 544)- - - 150
5.5 Comparison of theoretical and experimental time
histories of cowl pressures for Run 39 (Part 544)- - - 151
5.6 Comparison of theoretical and experimental timehistories of ramp pressures for Run 40 (Part 619)- - - 152
5.7 Comparison of theoretical and experimental time
histories of cowl pressures for Run 40 (Part 619)- - - 153
5.8 Comparison of theoretical and experimental timehistories of ramp pressures for Run 18 (Part 624)- - - 154
5.9 Comparison of theoretical and experimental timehistories of cowl pressures for Run 18 (Part 624)- - - 155
5.10 Comparison of theoretical and experimental timehistories of engine face total pressures for Run 39
(Part 544), blastward (outboard) inlet -------- 156
5.11 Comparison of theoretical and experimental timehistories of engine face total pressures for Run 39(Part 544), leeward (inboard) inlet ---------- 164
LIST OF ILLUSTRATIONS (CONTINUED)
Figure Page
5.12 Comparison of theoretical and experimental time
histories of engine face total pressures for Run 40(Part 619), blastward (outboard) inlet -------- 172
5.13 Comparison of theoretical and experimental time
histories of engine face total pressures for Run 40(Part 619), leeward (inboard) inlet ---------- 180
5.14 Comparison of theoretical and experimental time
histories of engine face total pressures for Run 18(Part 624), blastward (outboard) inlet -------- 188
5.15 Comparison of theoretical and experimental time
histories of engine face total pressures for Run 18(Part 624), leeward (inboard) inlet ---------- 196
5.16 Comparison of theoretical and experimental time
histories of engine face distortion for Run 39(Part 544), blastward (outboard) inlet -------- 204
5.17 Comparison of theoretical and experimental timehistories of engine face distortion for Run 39(Part 544), leeward (inboard) inlet ---------- 210
5.18 Comparison of theoretical and experimental timehistories of engine face distortion for Run 40
(Part 619), blastward (outboard) inlet -------- 216
5.19 Comparison of theoretical and experimental timehistories of engine face distortion for Run 40(Part 619), leeward (inboard) inlet ---------- 222
5.20 Comparison of theoretical and experimental time
histories of engine face distortion for Run 18(Part 624), blastward (outboard) inlet ------- 228
5.21 Comparison of theoretical and experimental timehistories of engine face distortion for Run 18(Part 624), leeward (inboard) inlet- --------- 234
7.1 Boundary layer separation and distortion from fan
reflected shock wave ----------------- 246
12
LIST OF ILLUSTRATIONS (CONCLUDED)
Figure Page
7.2 Predicted free-stream properties and form factor Hiof boundary layer on cowl of blastward inlet. Afterreflected wave enters inlet -------------- 247
7.3 Predicted free-stream properties and form factor Hiof boundary layer on ramp of blastward inlet. Afterreflected wave enters inlet -------------- 248
7.4 Predicted free-stream properties and form factor Hiof boundary layer on cowl of leeward inlet. Afterreflected wave enters inlet- ------------- 249
13
LIST OF TABLES
Table Page
1.1 16T wind tunnel test conditions ----------- 17
4.1 Post-test assessment of possible blast-inducedseparation within leeward inlet ----------- 122
5.1 Test conditions for correlation runs --------- 140
A.1 Ramp/cowl transducer evaluation ----------- 255
14
SECTION 1
INTRODUCTION
The problem of the interaction of a blast wave from a nuclear
explosion with an aircraft engine-inlet system is of importance for
military survivability/vulnerability evaluations.
An experimental study regarding this problem was recently
conducted at the Arnold Engineering Development Center (AEDC) (References 1
and 2). In this study blast waves produced by shock tubes impinged on
a 0.1-scale B-1 inlet pair mounted in the AEDC 16T transonic wind tunnel.
Tests were performed at tunnel (pre-blast) Mach numbers of 0, 0.55,
0.70, 0.85 and 0.90 for blast overpressures (scaled to 1 atm. ambient
pressure) from 2 to 6 psi for inlet flow rates representative of cruise
and maximum power conditions. These tests are described in detail in
Reference 1, which also presents a preliminary analysis of the test
data and a preliminary correlation of the test results with predictions
of the Blast Induced Distortion BID-2 computer code (Reference 3).
The present report is a continuation of the studies of
Reference 1 covering the following topics. Section 2 discusses the
blast response of the mean total pressure at the engine face. Section 3
discusses general features of blast and operational effects on engine-
face mean total pressure signature. Section 4 discusses specific
effects of blast and operational variables on engine-face mean total
pressure. Section 5 presents detailed comparisons of theoretical and
experimental inlet pressures and distortion parameters. Section 6
presents an evaluation of some large late-time inlet distortion data
discussed in Reference 1. Section 7 discusses engine reflected shock-
boundary layer interaction. Conclusions are presented in Section 8.
Since this report is closely related to Reference 1, it is
assumed herein, to limit repetition of material from that reference,
that the reader is familiar with Reference 1, particularly regarding
test geometry and test instrumentation. However, for the reader's
15
convenience, Table 1.1 and Figures 1.1 and 1.2 indicating test
conditions and pressure transducer locations in the engine inlets and at
the engine face are repeated here from Reference 1. (See Reference 1
for an explanation of the table and figures.)
16
TABLE 1. 1
16T WIND TUNNEL TEST CONDITIONS
Pirt Tube Shock Intercept YawPoInt Mach Flow Rate (lb/sec) Tube rressure Overpressure Angle Ang.
Run No.8 Inl IB Inlet No. (psia) (pSI) (deg) (deg)
1 501.01 0 0 0 2 69 2.7 79 0
2 615.03 .552 235 235 3 186 4.7 91
3 591.03 .550 351 348 1 157 3.7 764 589.03 .551 351 348 2 115 3.8 975 590.02 .549 351 349 3 124 4.0 94
6 602.02 .700 302 300 1 72 2.6 86
7 600.04 .70i 302 302 2 58 2.6 1068 573.04 .700 304 302 112 5.0 98
9 601.03 .701 302 300 3 69 3.0 10410 574.03 .701 303 300 132 4.4 97
11 621.03 .700 351 351 1 73 3.0 8412 519.02 .699 348 344 103 3.8 8813 527.02 .700 349 344 135 4.8 7814 626.02 .701 352 352 142 4.8 79
15 512.03 .700 351 344 2 59 2.8 10316 517.02 .700 349 344 | 85 3.8 10317 525.02 .701 348 344 113 5.0 10018 624.02 .700 350 350 139 '.2 99
19 513.03 .700 351 344 3 70 3.0 102
20 518.02 .700 348 343 102 4.2 9921 526.02 .700 349 344 133 4.8 9822 625.02 .701 350 350 155 5.6 92
23 570.03 .699 351 350 1 144 3.6 87 +5.024 568.04 .70(0 350 349 2 132 5.8 103 +5.025 509.03 .703 350 349 3 143 4.2 105 -5.0
26 559.02 .850 300 298 1 61 2.2 89 027 598.03 848 299 299 90 3.0 8528 584.03 .847 294 293 122 5.0 8229 608.04 .850 299 299 142 4.4 82
311 557.04 .850 300 298 2 55 - -31 596.05 .848 300 300 73 3.8 10832 582.03 .847 300 297 94 4.4 105
31 606.03 .849 300 299 120
34 558.03 .850 303 301 3 60 >2 -
35 597.03 .848 300 300 85 4.0 10436 583.03 .847 298 296 114 4.4 10237 607.03 .850 299 298 140 4.8 108
38 546.02 .847 348 347 1 121 3.6 84
39 544.04 .847 348 347 2 94 4.0 10740 619.02 .850 352 351 120 5.8 110
41 545.03 .847 348 347 3 113 4.4 10542 620.02 .850 351 351) + 139 5.6 100
43 553.03 .900 327 329 1 117 3.0 8644 550.02 .899 349 354 2 86 4.0 10745 551.01 .900 349 354 3 104 4.2 105
17
44i-4
00-- 7 0
H-
z 4L
cocE-44
00
0' C0
CYN 0'-.4 -4l
ac
ccno fn0 000
14 0E--
.1~
180
22.5 22.5 DYNAMIC AND TOTAL PRESSURE PROBE
2 8. 2605 28o1 180
, ,..0', 28 260, _Z .- 450 ,81h, 0 /?- 89 2 80A.- 183 160 1803 65]Wb2838 M, 2603 .1o9
1 12 38 183 ,,' 1603 -,,, 65"283T 26 26 161
2 63A G\ \ 2 6 10 ' . _6 5 / 16 ~ .37 6 1
,126 26236 2601 2801 2609 __ 163483 L501 1801\ l69~.
283 6- 2631 18330 632 I" 1V3Laos
283. / 2)'"1
2 611 2 16.7.,1 2621 ' / 4 D 2612 16271 AU2 / 1612 /266211 -. 261 61 16284 - 1826 1 1613 q
N30 26 2-28 28615 163029 262 0 L 28 1 = .o 1. 1 6 S 16 T D S 1615
93 121 q 18160 230 88 22 82 6 2211 \ 8~~~2b 28370 189, 8 80 ' 22 1622 1617\\ ~28l1 ' I 'u4
.623 268. DYNAMIC 6s 1623 ,3
,)"7 = 2.60
6 84 2*21"' 1624 82 16197 TEADY262,12257 PROBE 1625 1 25 1819k STATE
62 NUMBERS 12 TT
/LOOKING DOWNSTREAM PROBENUMBERS
INBOARD ENG~INE FACE RAKES OUTBOARDYj=0.777
* 2601 - 264.0 STEADY STATE y2z LO2S e 1601 - 164~0 STEADY STATE*@2801 - 284+0 DYNAMiIC Y3 1.393 .1801 - 184.0 DYNAMIC
.Y4: .683
Y6 : 2.149
Yy: 2.250
Figure 1.2. Engine face transducer locations. (Rockwell drawing)
19
L . ..... ..._.,.,.. . ...... ......, ... -.,,, - .... .. ..L.... ,b ;.d ,,-i
SECTION 2
RESPONSE OF MEAN TOTAL PRESSURE AT ENGINE FACE
The effect of the blast wave interaction upon the mean total
pressure at the engine face as measured in the tests is examined in
this section and the following two sections. In this section the test
records of the mean total pressures at the two engine faces are examined
qualitatively for each run on a run-by-run basis. In Section 3 the
records are grouped so as to show the effect of each test variable, one
at a time - snock overpressure, intercept angle, inlet flow rate and
free-stream Mach number - and the general effects of the variables are
examined qualitatively. In Section 4 the effect of these test variables
upon specific features of the records are examined on a quantitative
basis.
2-1 GENERAL FEATURES OF EFFECT ON MEAN TOTAL PRESSURE.
The blast wave first intercepts the cowl of the blastward
inlet, then diffracts around the cowl and into the inlet. It then
reflects from the splitter between the two inlets and diffracts around
the splitter and into the leeward inlet. At each engine face of the two
inlets a series of shock waves arrives as a result of the diffractions
and multiple reflections from the duct walls.
Typical records of the reduced mean total pressures at the
engine faces of the two inlets are shown in Figure 2.1. The quantities
R20 and R21 are the ratios of the instantaneous mean total pressures at
the engine faces of the outboard and inboard inlets, respectively, to
the preblast total pressure in the wind tunnel. The mean total pressures
for each inlet were obtained from averages of the 40 pitot probe measure-
ments at each engine face. Tube 2 was fired in this run, so R20 is the
record for the blastward inlet and R21 the record for the leeward inlet;
the same would be true for a Tube 1 firing and the opposite for a Tube 3
firing.
R20 in Figure 2.1 is constant at 0.99 until about 8.83 milli-
seconds (ms). It then jumps up to about 1.1 on shock arrival at the
engine face, followed by a steep ramp-up of about 0.33 ms duration to a
20
peak of about 1.4. The ramp is believed to consist generally of a
staircase of weak shocks produced by the multiple reflections of the
initial blast shock from the splitter and the walls of the inlet. R20
then drops off from 1.4 down to about 1.3 due to the weakening of the
upstream reflection by diffraction of the incident shock around the
splitter and other outside surfaces of the inlets.
At 9.7 ms R20 increases abruptly again due to reflection of
the shock wave from the downstream throttle formed by the control vanes.
This wave is an upstream facing shock followed by a compression wave.
The rate of rise of the engine-face total pressure varies with test
conditions and the particular characteristics of the initial shock and
compression wave within the inlet.
The strength of the upstream-facing shock/compression wave in
terms of pressure would be greater than might appear from records of
total pressure. This reflected wave slows down the flow passing through
it. The velocity reduction across the shock itself decreases the total
pressure, offsetting partially the increase in total pressure due to
pressure rise.
The reflection from a fan stage would be expected to produce
a stronger wave than occurs from the reflection from the throttle
(choked orifice). This is demonstrated by calculations for the upstream
reflection of a shock wave from a typical high-performance fan in
Reference 1, Section 8.5.
The concern regarding the upstream-facing shock/compression
wave is that it tends to distort the flow within the inlet passing
through it to an extent that stalling of the fan stages within the
engine might result. Calculations presented in Reference 1, Section 9,
demonstrated that the boundary layers on the ramp and cowl surfaces of
the blastward inlet, and possibly the cowl surface of the leeward inlet,
would separate for full-scale inlets (lOx model size) due to a repre-
sentative reflected wave (5-psi blast intercept at 90 degrees for
Mach 0.85 and 350 lb/s full-scale reduced weight flow).
21
Following the arrival at the engine face of the reflected wave
from the control-vane throttle, R20 at the engine face then decreases
slowly in this run until about 12 ms, at which time it falls off rapidly
because of the fall off of pressure within the portion of the blast wave
in the area of the inlet mouth. For firings from Tube 1, R20 often
increased at about 3 ms or so after shock arrival, due to arrival of the
cold driver gas from the shock tube.
At the engine face of the leeward inlet Figure 2.1 shows that
the inboard total pressure R21 was constant at 0.99 in the test until
9.4 ms. It then jumped to about 1.10 upon arrival of the blast shock
that had diffracted around the splitter. RTI then continued to rise
due to further blast flow around the splitter, then level off and
abruptly rose again at 10.1 ms due to the shock reflected from the
control vanes of the inboard inlet.
Beginning at about 10.4 ms R21 then fell off rapidly. This is
attributed to the large sideslip angle produced by the blast-induced
flow at the inlets, which will be discussed further. A large sideslip
angle would produce separation of the flow from the leeward surface of
the splitter or retard the flow sufficiently that it would be vulnerable
to separation by the interaction with the reflected shock from the
control vanes.
The R20 and R21 records are discussed below relative to the
effects of the tunnel Mach number, M, the inlet full-scale weight flow
at the engine face, W2R; the particular shock tube fired; and the (nominal)
overpressure of the incident blast shock, Aps.
One difference between firings from Tube 1 and those from
Tubes 2 and 3 is the blast intercept angle, *. Intercept angles from
Tube 1 ranged from 76 to 89 degrees from head-on; intercept angles from
Tubes 2 and 3 were greater, ranging from 91 to 110 degrees. The second
difference between the firings was that Tubes 1 and 2 fired from the
outboard side of the inlets and Tube 3 from the opposite side, over the
fuselage, so the blast wave from Tube 3 traveled over more of the
fuselage model.
22
2-2 MACH 0, 0 LB/S FULL-SCALE WEIGHT-FLOW.
The R20 and R21 records for Run 1 with the tunnel off and a
zero weight flow through the two inlets are shown in Figure 2.2. The
record for the blastward inlet R20 indicates that after the shock arrived
at the engine face (9.7 ms) R20 increased to about 1.26 and then fell
off with very little indication of any reflection from the control
vanes. RTO finally leveled off about 1 ms after shock arrival at about
one-half of the peak value. The abrupt rise in RTO, beginning 3.9 ms
after shock arrive (13.6 ms), is attributed to the arrival of the cold
driver gas from the shock tube, which ends the test.
The record for the leeward inlet indicates that R21 rose
stepwise at shock arrival (11.7 ms) to a level of about 1.08, continued
to climb to a maximum of about 1.14 and then decayed slowly. It is
speculated that the decay may be due to a gradual separation of the flow
from the leeward side of the splitter and ramp of the inboard inlet.
The dynamic pressure is small at Mach zero relative to the static over-
pressure so that separation would not produce as large an effect on R21
as in the example shown in Figure 2.1, where the Mach number is 0.70.
2-3 MACH 0.55, 235 LB/S FULL-SCALE WEIGHT FLOW.
Figure 2.3 shows the records for Run 2, which had the lowest
non-zero weight flow for the two inlets, 235 lb/s (full scale) each.
The tunnel Mach number was also the lowest non-zero value, 0.55, and the
incident shock overpressure was high, 4.7 psi (nominal). Tube 3 was
fired, so the inboard inlet (R21)was on the blastward side.
The record for R21 in Figure 2.3 has a high first peak, which
is attributed to the relatively high shock overpressure. The reflected
shock from the throttle control vanes, which arrived about 0.7 ms after
the first shock, appears to have been the strongest here of all of the
tests, in terms of the increment in R21. The weight flow was also low
in this test, and comparisons between tests indicated that low weight
flow increased the second rise.
23
The R20 record in Figure 2.3 shows the effect of a strong
initial shock for the leeward inlet, by the abrupt rise to a value of
1.13. It then continues to rise more slowly to 1.16, fall off, then
rise again. The latter rise is attributed to a weak reflection from the
control vanes that reaches the engine face about one millisecond after
shock arrival. The R20 record then exhibits a rapid decay which is
attributed to a large sideslip angle, produced by the blast wave, that
degrades the flow uniformity entering the leeward inlet. The combination
of a low Mach number, low mass flow and strong blast wave results in a
large sideslip angle.
This is an example where the RTO fall-off for the leeward
inlet commenced slightly after the arrival of the reflected shock at the
engine face. This would be expected to be the sequence of events if
the reflected shock caused shock-boundary layer separation within the
inlet. Further information would be needed to verify that separation
actually occurred, however it does appear clear from the record that
significant flow deterioration did take place within the leeward inlet.
This strong throttle-reflected shock in the blastward inlet
and the combination of a throttle-reflected shock in the leeward inlet
followed by a rapid decay could be detrimental to the operation of
engines attached to such an inlet. The throttle-reflected shocks,
simulating engine reflected shocks, would tend to separate the flow
within the inlets, resulting in a rapid decrease in total pressure and
possible stall of a gas turbine engine.
2-4 MACH 0.55, 350 LB/S FULL-SCALE WEIGHT FLOW.
The results of firings at Mach 0.55 and a full-scale reduced
weight flow of 350 lb/s, (Runs 3, 4 and 5) are shown in Figure 2.4.
The results for Run 5, Figure 4c, at 350 lb/s can be compared
with those for Run 2, Figure 2.3, at 235 lb/s. Both firings were from
Tube 3 with comparable shock overpressures, 4.0 and 4.7 psi, respectively.
The first peak for the blastward inlet is similar for the two weight
24
flows, but the second peak, from the throttle reflection, is much smaller
for the higher weight flow and it also is spread out more timewise.
This characteristic of weak throttle reflections is true also for the
other two firings at Mach 0.55 and 350 lb/s, Figures 2.4a and 2.4b.
For the leeward inlet at Mach 0.55, the first peak of Run 5,
Figure 2.4c, is similar to the first peak of the 300 lb/s firing,
Run 2, Figure 2.3. The throttle reflection for the higher weight flow,
Figure 2.4c, is weaker than for the lower weight flow, Figure 2.3, as it
was for the blastward inlet. The decay rate is not as steep either, as
at the lower weight flow. It is speculated that these effects may be due
to a lower blast-induced sideslip angle for Run 5 than for Run 2 because
of the higher weight flow and 4lightly lower shock overpressure.
2-5 MACH 0.70, 300 LB/S FULL-SCALE WEIGHT FLOW.
The records of the reduced total pressures RTO and RTI at the
engine face for tests at Mach 0.70 and a full-scale reduced weight flow
of 300 lb/s, Runs 6 to 10, are presented in Figure 2.5. For the firing
from Tube 1 (Run 6, Figure 2.5a) the shapes of the records are generally
similar to those for the example discussed in Figure 2.1. Again, the
second jump, due to the reflected shock from the throttle, which simulates
the engine, is substantial for both the blastward and leeward inlets.
The total pressures both remain high for about 3-1/2 ms after shock
arrival until cold gas appears to have arrived from the shock tube (R20)
or blast decay sets in (R21).
The weaker blast firings (2.6, 3.0 psi) from Tube 2 (Run 7)
and Tube 3 (Run 9) show a marked falloff in the blastward total pressures
after the throttle reflection arrives. This fall off is believed to be
due to decay in the blast wave arriving at the inlet. Therefore the
falloff for the leeward inlet is not attributed to sideslip angle in
this case. For the strong shocks (5.0 and 4.4 psi) from Tube 2, Run 8,
and Tube 3, Run 10, however, the blastward ratios held up well for 3 to
4 ms after shock arrival, so the falloff of the leeward ratios beginning
at about 1 ms after shock arrival is attributed to the large sideslip
angle produced by the blast wave (and not to decay in the blast properties
arriving at the inlet).
25
2-6 MACH 0.70, 350 LB/S FULL-SCALE WEIGHT FLOW.
The results for Mach 0.70, and a full-scale weight flow of
350 ib/s, Runs 11 to 22, are shown in Figure 2.6. For Tube 1 firings,
Figure 2.6a-d, the effect of weight flow on the interaction is found by
comparison with Run 6, Figure 2.5a, for the same Mach number but only
302 lb/s. For the blastward inlet the first peak in RTO for Run 11,
Figure 2.6a, which has a nearly equal shock overpressure (3.0 psi) to
Run 6 (2.6 psi), is similar to the peak for Run 6. The second peak in
comparison to the first, is much smaller for Run 11, than at the lower
weight flow of Run 6.
The total pressures for Tube 1 at 350 lb/s hold up in both the
blastward and leeward inlets for the full 3-ms nominal test period for
the higher overpressures of 3.8 to 4.8 psi (Runs 12-14). This means
that sideslip is not causing a problem to the leeward inlet. There is a
small falloff for 3.0 psi (Run 11).
There is a feature of the blast wave that should be noted,
for later discussion, from the results shown in Figure 2.6. It will be
discussed with respect to Run 12, Figure 2.6b. The blast shock front
arrives at the blastward inlet first, of course, which is why R20 rises
about 1/2 ms before R21 for Tube 1. The interior flow of the blast wave
would also arrive first at the blastward inlet. However, the rapid
falloff in the total pressure for the blastward inlet (RTO) at about
12-1/2 ms is preceded by the falloff for the leeward inlet (RTI) by a
time of about 1/2 to 1 ms. It also decays earlier at the leeward inlet
for the firings at higher shock overpressures, as is evident in Figures 2.6c-d.
Examination of measurements of the blast waves produced by this method
has shown that this earlier decay at the leeward inlet is a result of
the three-dimensional (vs. one-dimensional) character of the blast wave.
In this case a decaying portion of the blast wave reaches the leeward
inlet before a decaying portion reaches the blastward inlet.
26
The question then arises whether the decay in total pressure
observed sometimes for the leeward inlet, which is attributed in some
cases to blast-induced sideslip, might in fact be due to this decay in
the blast pressure instead. The question cannot be resolved with
certainty at this time. But, from an examination of trends with varying
blast strengths and from measurements with external claw probes, it
appears at present that the blast decay results in a definite signature,
such as appears in Run 12, Figure 2.6b.
Another feature of the blast signature is shown in Figure 2.6c
for Run 13. At about 13 ms, R20 rises at an increasing rate. This rise
is associated with the arrival of the cold driver gas from the shock
tube. It is also observed for the leeward inlet by the rise in R21
beginning at about the same time. This rise is followed by the pressure
decay of the blast wave cited above for the leeward inlet (R21), starting
at about 14-1/2 ms, and for the blastward inlet (R20), starting somewhere
between about 15 to 17 ms. The results for Run 14, Figure 2.6d, are
similar.
These records all indicate that there is a usually good 3 ms
or more of good test flow within the blast wave before these blast
anomalies affect the results for this model. The anomalies appear to
produce characteristic signatures in the records that make them rather
well defined.
Tube 2 firings at Mach 0.70 and 350 lb/s, Figures 2.6e-h, and
Tube 3 firings, Figures 2.6i-1, produce weak throttle reflections for
the blastward inlet until the blast shock overpressure approaches about
5 psi, Figures 2.6h and 2.6k-1. For the leeward inlet there is essentially
no throttle reflection. There is however a decay for the leeward inlet,
beginning about 1-2 ms after blast shock arrival, that is attributed to
blast-induced sideslip angle. The rate of the decay increases with
blast strength which follows the trend expected for an increasing angle
of sideslip, associated with a stronger blast wave.
27
2-7 MACH 0.70, 350 LB/S, ±50 YAW.
The results with yaw angles of ±5 degrees are presented in
Figure 2.7. In each test the model was yawed with the nose away from
the particular shock tube that was fired. This means that the sideslip
angle is increased by the blast wave.
The test conditions of Run 23, Figure 2.7a, compare otherwise
most nearly with those of Run 12, Figure 2.6b. The results for the
blastward inlet are rather similar. The leeward inlet shows a notice-
ably greater falloff beginning about 1 ms after blast arrival (or at
10.5 ms). This means that the 3.6-psi blast wave with the 5-deg yaw
angle evidently produces too much sideslip for the flow within the
leeward inlet to maintain total pressure.
The test conditions for Run 24 with Tube 2, Figure 2.7b,
compare most nearly to those for Run 18, Figure 2.7h, except for the
initial yaw angle. For the blastward inlet the throttle reflection is
weaker, but otherwise the results are similar. For the leeward inlet
the falloff due to yaw is much more marked than for Tube 1, Figure 2.7a,
beginning very shortly after shock arrival.
The results for Run 25 with a firing from Tube 3, Figure 2.7c,
are similar to those for Tube 2, Run 24. The test conditions are the
same as for Run 20, Figure 2.6j, except for the -5-deg yaw. The blast-
ward inlet results are about the same as for Tubes 1 and 2, Figures 2 .7a
and b. The leeward inlet has a rapid pressure felloff, as for Tube 2.
2-8 MACH 0.85, 300 LB/S FULL-SCALE WEIGHT-FLOW.
The total pressure results at the engine face for Mach 0.85
and a weight flow of 300 lb/s are shown in Figure 2.8. Firings with
Tube I are presented in Figures 2.8a-d. The jumps due to throttle
reflections (engine simulation) for the blastward inlet are stretched
out much more in time for the weaker shocks, below about 4 psi, than
they were at the lower Mach numbers. The reflected wave steepens up
as the blast strength is increased, so that at 4.4 psi for Run 29
(Figure 2.8d) it is quite steep. The jump to the second peak is also
large.
28
The total pressures in Figures 2.8a-d for the leeward inlet
with Tube 1 firings build up well and generally hold up for the blast
period of about 3 ms. The only particular falloff occurs with the
weaker shocks of 2.2 and 3.0 psi (Figures 2.8a and b) beginning about
1.5 ms after shock arrival, which is attributed to decay in the blast
wave, since it roughly follows the fall-off of the blastward total
pressure.
For Tube 2 and 3 firings at Mach 0.85, Figures 2.8e-j, the
results are also similar to the Mach-0.7 results, Figure 2.5, for the
same tubes. The rapid falloff in the total-pressure ratios beginning
about 1-1/2 ms after shock arrival is again attributed to the decay In
the blast wave at the inlet. There is a strong effect of the throttle
reflection for the blastward inlet, also, appearing in the steepening of
the second rise and increase in level of the second peak as the blast
strength is increased. The leeward records have typical first peaks,
but again only very weak effects of throttle reflections which are
masked by a rapid decay in total pressure. There is no clear indication
of separation occurring in the leeward inlet at Mach 0.85 as there was
for the higher shock overpressures at Mach 0.70 (Figures 2.5c and e).
2-9 MACH 0.85, 350 LB/S FULL-SCALE WEIGHT FLOW.
The results for Mach 0.85 and a full-scale reduced weight flow
of 350 lb/s are shown in Figure 2.9. For Tube 1, Figure 2.9a, the
blastward inlet for a 3.6-psi shock, Run 38, has a sharp initial peak
and a strong throttle reflection, similar to the results for Run 29,
Figure 2.8d, at 300 lb/s and a 4.4-psi shock overpressure. The results
for the leeward inlet are also similar to those with 300 lb/s.
For Tubes 2 and 3 at Mach 0.85, Figures 2.9b-e, the results
are similar to those for 300 lb/s, Figures 2.8e-1, including the rapid
decay attributed to the blast wave decay (not to sideslip).
2-10 MACH 0.90, 350 LB/S FULL-SCALE WEIGHT FLOW.
The results for Mach 0.90, Figure 2.10, are similar to the
results for Mach 0.85, Figure 2.9, for both the blastward and leeward
inlets.
29
PRRT/PT M P1 TUBE PLOT573-04 0.70 10.21 2 19
1.60 Fis peFal Relecti oto
1.40 -l aeRoEDec a ributed to1.20 Rm O- i
CC" I.00 S
0.80
0.60
1.60 Reflection frov control1.40 va e throttle
1.20 - a_____ to1.0Shock possibl ;'parationC' 1 .00 - -w-h *A-t
0.80
0.60
0.00 4.00 8.00 12-00 16.00 20.00
TIME IN MILLISECONDS
Figure 2.1. Typical engine-f ace mean total pressures. Run 8.Mach 0.70, W211-300 lb/s (nom), Aps5 5.O psi, 0-98 dog.
30
PRRT/PT M PT TUBE PLOT
501.01 0.00 7.10 2 19
1.60 I1.40
1.200CC1 I00
0.80
0.60
1.6011.40
-1.20
0.60d
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
Figure 2.2. Engine-face mean total pressures. Run 1.Mach 0, W2R-0 lb/s (nom), Aps-2.7 psi (nom), f-79 deg.
31
PRRT/PT M PT TUBE PLOT615.03 0.55 12-33 3 19
1.60
1.40
1.20 -____
(C* 1.00
0.80
0.60
1.60
1.40 -_ _ ____ _
1.20
1.-00
0.80
0.60
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
Figure 2.3. Engine-face mean total pressures. Run 2.Mach 0.55, W2R-235 lb/s (nom) Aps 4.7 psi (nom),*-91 deg.
32
PRRT/PT M PT TUBE PLOT591.03 0.55 12.36 1 19
1 .60
1 .40
1.20
C' 1.00
0.80
0.601 .60
1 .40
1 .20
C* 1.00 -
0.80
0.60
(a) Run 3, Ap=3.7 psi (nom), 0=76 deg.
PRRT/PT M PT TUBE PLOT589.03 0.55 12.33 2 19
1.60
1 .40
1.20
C\J ___0_
0.80
1.60
1.40
1.20
C' 1.00
0.80-
0.600.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
(b) Run 4, Aps3.8 psi (nom), 0=97 deg.
Figure 2.4. Engine-face mean total pressures.
Mach 0.55, W2R-350 lb/s (nom).
33
PART/PT M PT TUBE PLOT590.02 0.55 12.30 3 19
1.60
1.40
1.20I00n, 1.00 -,_____0.60
0.60
1.60
1.40
1.20
r_ 1.004 008 0
0.00 4.00 .00 12.00 16.00 20.00TIME IN MILLISECONDS
(c) Run 5, Apsi4.0 psi (nom), *i94 deg.
Figure 2.4. Concluded.
34
PRRT/PT M PT TUBE PLOT602.02 0.70 10.19 1 19
1.60
1.40
1.20
C*4 1.00
0.80
0.601.60
1.40- _ _ _ __ _ _ _
1.20
1 .00
0.80
0.60
(a) Run 6, Aps=2.6 psi (nom), 4=86 deg.
PRRT/PT M PT TUBE PLOT600.04 0.70 10.18 2 19
1 .60
1.40 _ _ _ _ _ _ _ _
1 .20
1" .00
0.80
0.601.60
1 .40
1 .20
C= 1.00
0.80
0.600.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
(b) Run 7, Ap 2.6 psi (nom), $'.106 deg.
Figure 2.5. Engine-face mean total pressures.Mach 0.70, W2R=300 lb/s (nom).
35
PRRT/PT M P1 TUBE PLOT573.04 0.70 10.21 2 19
1.60
1.40 - _ _
1.20
C' I .00)
0.80 _ _ _ _ _ _ _ _ _ _ _ _ _ _
O0.601.60-_ _ _ _ _ _ _ _
1.40-_ _ _ -- - - - _ _
1.20 -_ _ _
1.00 -_ _ _
0.80
0.60
(c) Run 8, Aps=5.0 psi (nom), 0=98 deg.
PRRT/PT M PT TUBE PLOT601.03 0.70 10.18 3 19
t .60
1.40
1.20
c1.000.80
0.601 .60
1 .40
1 .20
4.00 8.0
TIME IN MILLISECONDS
(d) Run 9, Ap=3.0 psi (nom), 0-104 deg.
Figure 2.5. Continued.
36
PRRT/PT M PT TUBE PLOT574.03 0.70 10.19 3 19
1.60-
1.40
1.20
(C~ 1.00
0.80
0.60
1.60
1.40___________I--
1.20
01.00 ~
0.80
0.60 ___ _
0.00 4.00 8.00 12-00 16.00 2600TIME IN MILLISECONDS
(e) Run 10, Aps."4.4 psi (noma), -97' deg.
Figure.2.5. concluded.
37
PRRT/PT M PT TUBE PLOT.621.03 0.70 10.23 1 19
1.60
1.40
1 .20A
C' 1.00
0.80 -
0.601.60
1.40
1 .20
0.60__ _ _ _ __ _ _ _
(a) Run 11, Ap=3.0 psi (nom), 4=84 deg.
PRRT/PT M PT TUBE PLOT519.02 0.70 10.20 1 19
1 .60
1.40 A _
1.200-0
0.80 - _ _ _ _ _ _ _ _ _ _ _ _ _ _ _
0.60 -
1 .60
1.40
1.20 -_ _ _ _
CC' 1.00.-
0.80
0.600.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
(b) Run 12, Apin3.8 psi (nom), *-88 deg.
Figure 2.6. Engine-face mean total pressures.Mach 0.70, W2R-350 lb/s (nom).
38
PPRT/PT M FT TUBE PLOT527.02 0.70 10.17 1 19
1.60 -__
1 .40
1.20
0 1.00
0.80
0.601.60
1.40
1.20 -_
o 1.00 -
0.60
0.60 - _ _ _ _ _ _ _ _
(c) Run 13, Aps=4.8 psi (nom), =78 deg.
PRRT/PT M PT TUBE PLOT626.02 0.70 10.22 1 19
1.60
1.40
1 .20 V
C 1.00
0.80
0.601.60
1.40
1.200-4
1.00
0.80
0.600.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
(d) Run 14, 6psf 4 .8 psi (nom), fi78 deg.
Figure 2.6. Continued.
39
PRRT/PT M PT TUBE PLOT512.03 0.70 10.18 2 19
1.60
1.40
1.20
C', 1.00
0.80
0.601 .60
1 .40
1.20
c' 1 .00 - __ ___
0.80
0.60
(e) Run 15, Ap,2.8 psi (nom), *=103 deg.
PPRT/PT M PT TUBE PLOT
517.02 0.70 10.19 2 19
1 .60
1 .40
1.20-
C* 1.00
0.80
0.601 .60
1.40
1.20 - - _ _ _ _
I-100 -
0.80
0.600.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
(f) Run 16, Ap -3.8 psi (nom), 0-103 deg.
Figure 2.6. Continued.
40
PART/PT M PT TUBE PLOT525.02 0.70 10.20 2 19
1.60
1.40
1.20
1.00
0.800.60
1.60
1.40 -_
1.20
0.80 ,,,,,.,0E . -
(g) Run 17, Aps=5.0 psi (nom), 4=100 deg.
PRRT/PT M PT TUBE PLOT624.02 0.70 10.22 2 19
1.60
1.40 A
1.20
~1.20 cr 1.00
0.80
0.60
Fiue .. otiud
1.40
1.20.. ,,,cl x
1.00 .. . ._ ,- _ ,
0.80
0.60
TIME IN MILLISECONDS
(h) Run 18, Apsw5.2 psi (nom), €=99 deg.
Figure 2.6. Continued.41
PRRT/PT M PT TUBE PLOT513.03 0.70 10-17 3 19
1.60
1 .40
1 .200-0
0.80
0.601.60
1.40 - _ _ _ _ _ _ _ _ _ _ _ _ _ _ _
1 .20
S1.00 ____
0.80 -_ _ _ _ _ _ _ _ _ _ _ _ _ _ _
0.60 - ____
(i) Run 19, lAp =3.0 psi (nom), *-102 deg.
PRRT/PT M PT TUBE PLOT518.02 0.70 10.20 3 19
1 .60
1.40
1.20
t 1.00 n
0.60
t .60 ____________ ____________ ____________ __________ ____________
1.40
1.20
0.80
0.600.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
(J) Run 20, Apo=4.2 psi (nom), 0-99 deg.
Figure 2.6. Continued.
42
PART/PT M PT TUBE PLOT
526.02 0.70 10.20 3 19
1.60-
1.40 -_ _ _ _- _ _ _ _
1.20 - _ _ _ _ _ _ _ _ _ _
0:I-100
0.80
0.601.60 - _ _ _ _ _ _ _
1 .40
1.2011.-0 0
0.80 - _ _ _ _ _ _ _ _ _ _ _
0.60 _
(k) Run 21, Ap5=4.8 psi (nom), *=98 deg.
PRRT/PT M PT TUBE PLOT625.02 0.70 10.21 3 19
1.60
1.40-- _ _ __ _ _ _ _ _
1.200:
0.601.60
1.40
1 .20
C~ 1.00
0.80
0.600.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
(t) Run 22, Ap in5.6 psi (nom), *-92 deg.
Figure 2.6. Concluded.
43
PRRT/PT M PT TUBE PLOT570.03 0.70 10.22 1 19
1.60
1.40
1.20
0.80
0.601 .60
1 .40
1 .20
C' 1.00 -- ~-_ _ _ _
0.60 -(a) Run 23, yaw angle =+5.0 deg, Ap-=3.6 psi (nom), 0=.87 deg.
PRRT/PT M PT TUBE PLOT
568.04 0.70 10.21 2 19
1 .60
1.40 -__ ____
1.20
S1.00 -
0.80 ____
0.601 .60
1.40
1 .20
C" 1.00
0.80 -_ _ _ _ _ _ _ _ _ _ _
0.600.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
(b) Run 24, yaw angle =+5.0 deg, Ap W5.8 psi (nom), 0-103 deg.
Figure 2.7. Engine-face mean total pressures.Mach 0. 70, W2R-350 lb/ (nom) .
44
PART/PT M PT TUBE PLOT569.03 0.70 10.20 3 19
1 .60
1.40
1.20C:)
, 1.00 -_
0.80
0.60
1.60
1 .40
1.20
S1.00 _ _ _ - -
0.80
0.60
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
(c) Run 25, yaw angle = -5.0 deg, Aps-4.2 psi (nom), 0-105 deg.
Figure 2.7. Concluded.
45
PRRT/PT. M PT TUBE PLOT559.02 0.85 11.78 1 19
1.60
1.40
1.20100
1 .60
1.40 -____
1 .20
I-100
0.80
0.60
(a) Run 26, Ap=2.2 psi (nom), 0=89 deg.
PPRT/PT M PT TUBE PLOT598.03 0.85 11.78 1 19
1.60
1.40
CJ1.0
0.80
0.601.60r 1.40
S1.0
0.600.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
(b) Run 27, Ap5'3.0 psi (nom), 0-.85 deg.
Figure 2.8. Engine-face mean total pressures.
Mach 0.85, W2R-300 lb/s (nom).
46
PRRT/PT M PT TUBE PLOT584.03 0.85 11.78 1 19
1.60
1.40
0 1.200-0
0.80
0.601.60
1 .40
1.20cc~ 1.00-
0.80
(c) Run 28, Ap,=5.O psi (nom), 0=82 deg.
PRRT/PT M PT TUBE PLOT608.04 0.85 11.77 1 19
1 .60
1 .40
1 .20 _____0C'4 I00
0.80
0.60 -
1.60
1.40____ _
1.20
C~ .0 -o
0.80 I I0.60 I
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
(d) Run 29, Ap-.4.4 psi (nom), 0-82 deg.
Figure 2.8. (Continued.)
47
PRRT/PT M PT TUBE PLOT596.05 0.85 11.78 2 19
1.60 !
1.40
1.20
1 .410
1 .21 t.00
0.80
0.60t .60
1.40
1.20
0
n,- 1.00
0.80
0.60
(e) Run 31, Aps=3.8 psi (nom), 0=108 deg.
PR::RT/P T M PT TUBE PLOT582.03 0.85 1178 2 19
1.60
1.40
1.20
N 1.00 -
0.80
0.601.60
1.40
1.20
0.80
0 .60
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
(f) Run 32, APsW4.4 psi (nom), 0-105 deg.
Figure 2.8. Continued.
48
.. -i i ii i i lm i i i ii i i . . . . . . . . . l4i i " . . . . . . . .] i
PRRT/PT M PT TUBE PLOT558.03 0.85 11.81 3 19
1 .60
1 .40
1 .20
I-100 -
0.80
0.601 .60
1.40 _ _ _ _ _ _ _ _ _ _ _
1 .20
C~ 1.00~ _0.60
(g) Run 34, Aps=2 + psi (nom).
PARI/PT M PT TUBE PLOT
597.03 0.85 11.78 3 19
1 .60
1.20------------
Of 1.00
0.601.60 -
1.40i
1.20
Of 1.00
0.80 -- -
0.600.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECO NOS
1h) Run, 35, Aps 4.0 psi (nom), 0-104 deg.
Figure 2.8. Continued.
49
PRRT/PT M PT TUBE PLOT583.03 0.85 11.77 3 19
1.60- _ _ _ _ .__ _ _ _
1.40 --- 4----1.20
1.00
0.80 -- - -- t- ---- --.
0.60 -
1.60 -
.40 -- ~-- ---- I------- --- 4- 1.20 ----- I- ----- ~--
0.60
(i) Run 36, Ap =4.4 psi (nom), =1O2 deg.
PRRT/PT M PT TUBE PLOT607.03 0.85 11.82 3 19
1 .60
1.40
1.200:C' 1.00
0.80
0.60 __ _ _ _ _ _ __ _ _ _ _ _
1 .60
1.40 _ _ _ __ _ _ _ _ _ _ _ _ _ _I
1.20
I-100
0.80
0.60
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
(J) Run 37. Ap=4.8 psi (nom), =108 deg.
Figure 2.8. Concluded.
50
PRRT/PI M PT TUBE PLOT546.02 0.85 11.76 1 19
1.60
1.40 - ~ - ~ I-1.20
I-100 -- _-
0.60
1.60
1.40H z i -- _ - _
-1.20 tof 1.00 _ _ _ _ _ _ _ _ _
0.80- - - - - - - - - - - _ -
0.60
(a) Run 38, Ap,=3.6 psi (nom), c =84 deg.
PRRT/PT M PT TUBE PLOT
544.04 0.85 11.76 2 19
1 .60
1.40
1.20
1. I00
0.80
1.60
1.40
1.204_
0.60
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
(b) Run 39, L6p =4.0 psi (nom), 0=107 deg.
Figure 2.9. Engine-face mean total pressures.
Mach 0.85, W2R-350 lb/s (nom).
51
PRRT/PT M PT TUBE PLOT619.02 0.85 11.80 2 19
1 .60 _ _ _ _ _
1.40
1.20
0,j I0
0' .0 ---
0.60 _____
1 .60
1.40
1.20
S1.00
0.80
0.60
(c) Run 40, Ap=5.8 psi (nom), 4=110 deg.
PRRI/PT M PT TUBE PLOT545.03 0.85 11-78 3 19
1.60
1.40
1.20 - _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _
W" 1.00 C -___
0.601.60
1.40
- 1.20 11 .00 . ~
0.80
0.0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
(d) Run 41, Aps4.4 psi (nom), 0-105 deg.
Figure 2.9. continued.
52
PRRT/PT M PT TUBE PLOT620.02 0.85 11.75 3 19
1.60
1.40
1.200-0
0.80
0.60
1.60
1 .40
1 .20
I-100 -- ze
0.80
0.60
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
(e) Run 42, Ap,=5.6 psi (nom), 0=100 deg.
Figure 2.9. Concluded.
53
PRRT/PT M PT TUBE PLOT1553.03 0.90 12.41 1 19
1.60-
1 .40
1.201
_ 1 .000.80 ___ {j__ ___-
0.60 -_ _ _ __ _ _ _ _ _ _ _ _
1.60 - _ _
1.20 ~~1 _ I - __
0.80- -j-__0.60 h~ N
(a) Run 43, Ap,=3.0 psi (nom), 0=86 deg.4
PART/PT M P[ TUBE PLOT550.02 0.90 12.42 2 19
1.60 -_ _ _ _ _ _ _ _ _ _ _ _
1.40__ - 1CD 1.20 -- - - - -- - - - _ _ _
O 1.00 --
0.80 _ _ _ _ _ _ _ _ _ _ _ _ _ _ _- _ _ _ _
0.60 -_ _ _ _ _____
0.60 _ _ _ _ I
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
(b) Run 44, Aps=4.0 psi (nom), =107 deg.
Figure 2.10. Engine-face mean total pressures.Mach 0.90, W2R=350 lb/s (nom).
54
PRRT/PT M PT TUBE PLOT551.01 0.90 12.41 3 19
1.60 - _ _ _ _ _ _ _ __ _ _ _
1.40 - _ _ _ _ _ _ _ _ _ _ _ _
1.20-- _ _ _ _ _ _ _ _ _ _ _ _ _
0-o
0.80 -_ _ __ _ _ _ _ '
0.60
1.60
1.40 -- - -_ _
- 1.20 ~-
0.80 - 1~0.60
0.00 4.00 -8.00 12.00 16.00 20.00TIME IN MILLISECONDS I
(c) Run 45 Aps=4.2 psi (nom), 0=105 deg.
Figure 2.10. Concluded.
55
SECTION 3
GENERAL FEATURES OF BLAST AND OPERATIONAL EFFECTS
ON ENGINE-FACE MEAN TOTAL-PRESSURE SIGNATURE
The effect of the blast parameters and operational parameters
on the mean total pressure at the engine face is examined in this
section. The blast parameters are the shock overpressure of the inci-
dent shock, Aps, and its intercept angle, , to the inlet. The opera-
tional parameters are the inlet full-scale reduced weight flow, W2R,
and wind tunnel Mach number, M.
The records of the reduced mean total pressure at the engine
face are grouped in Figures 3.1 to 3.56 according to test conditions
(Aps' ,, W2R, M) to show the effect of each variable separately. The
results are presented separately for the blastward and leeward inlets.
3-1 EFFECT OF BLAST SHOCK OVERPRESSURE.
The effect of the blast shock overpressure, Aps, on the RTI
and RTO records is shown in Figures 3.1 to 3.7 for the blastward inlet
and Figures 3.8 to 3.14 for the leeward inlet.
For the blastward inlet, the effect of increasing the shock
overpressure is to increase the magnitude of the first peak and to make
the shape of the peak a sharper spike as for example Figure 3.1. The
magnitude of the first peak will be examined quantitatively in Section 4.
The second peak, which is produced by the reflection from the control
vane throttle simulating the engine, steepens up in rise rate with
greater shock overpressure. The height of the second peak, referenced
to its initial point of rise, essentially does not vary with shock
overpressure.
The variation of the engine-face total pressure following the
second peak is generally independent of the shock overpressure until
the test is terminated by the arrival of the cold driver gas from the
shock tube. The cold gas appears to arrive in Figure 3.1b at about
56
11.7 ms (3.4 ms after shock arrival) and in Figure 3.1c at about 13.2 ms
(3.2 ms after shock arrival). The most obvious exception to this
independence of shock overpressure is shown in Figure 3.5 at Mach 0.70,
350 lb/s full-scale weight flow for firings from Shock Tube 3. This
decay in RTI following the second peak for the two lowest overpressuretests, Runs 19 and 20, is attributed to a decay in the blast overpressure.
The mean engine-face total pressure records for the leeward
inlet, Figures 3.8 to 3.14, have a characteristically different signature
from those for the blastward inlet. These records for the leeward inlet
show that the initial shock jump is followed generally by a continuing
rise. For example, in Figure 3.8a the total pressure following shock
arrival rises to a peak in-about 1 1/2 ms. The first part of the rise,
which reaches a plateau in about 1/4 ms, is attributed to post diffraction
effects of the blast wave around the splitter. The climb from the
plateau to the peak is attributed to reflection effects from the choked
throttle, that simulates the engine.
The magnitude of the initial (shock) jump in reduced total
pressure for the leeward inlet, Figures 3.8-3.14 increases with increasing
shock overpressure. For the seven firings from Shock Tube 1, Figures 3.8-3.9,
the shape of the pressure records for the 3-ms test period is pretty
much the same. The wave reflected from the throttle always has the
relatively iow rise characteristic for the leeward inlet.
The leeward inlet records for firings from Shock Tube 2,
Figu'es 3.10-3.11, are similar to those for Shock Tube 1 except that the
mean total pressure falls off significantly after about a millisecond,
depending upon the shock overpressure. This fall off is attributed to
the development of large flow nonuniformity and possible separation
within the inlet due to the large blast-induced sideslip angles. For
Shock Tube 3, Figures 3.12-3.14, the mean total pressure also falls off
rapidly for shock overpressures of 4 psi or more, but in these cases the
blast overpressure at the inlet exhibits a similar decay, so the
possibility of reparation being present cannot be assessed.
57
3-2 EFFECT OF BLAST INTERCEPT ANGLE.
The effect of the blast shock intercept angle, 0, on the RTI
and RTO records is presented in Figures 3.15-3.29. Firings from Shock
Tube 1 had the smallest intercept angles (most head-on). Firings from
Shock Tubes 2 and 3 produced intercepts with about the same angles
(from head-on), except that Shock Tube 3 fired from the opposite side of
the aircraft model from the other two.
For the blastward inlet, Figures 3.15-3.22, the first peak in
the reduced mean total pressure has a sharper spike for the more head-on
firings (Shock Tube 1). The second peak, from the reflection from the
throttle (simulated engine), in all cases appears to be larger for more
head-on interception. From firings from Shock Tubes 2 and 3, a signifi-
cant fall-off in mean total pressure is sometimes evident after about
one millisecond (Figures 3.16, 3.17, 3.19, 3.21, 3.22), but in all cases
the blast overpressure also decays similarly, so the decay cannot be
attributed simply to the intercept angle.
For the leeward inlet, Figures 3.23-3.29, the initial shock
jump (the first vertical rise) is generally greater for the smaller
intercept angle (Tube 1). Not much can be said about the effect of the
intercept angle for later times here because of blast decay at the inlet
for Tubes 2 and 3 in these runs, with the exception of Runs 17 and 21,
Figure 3.25. In the latter cases (Mach 0.70, 350 lb/s, 4.8-5.0 psi.)
there is a very marked effect of the intercept angle on the mean total
pressure, which is attributed to separation at the large blast-induced
angle of sideslip that results.
3-3 EFFECT OF WEIGHT FLOW.
The effect of the full-scale weight flow, W2R, on the RTI and
RTO records is shown in Figures 3.30 to 3.37 for the blastward inlet and
in Figures 3.38 to 3.45 for the leeward inlet.
For the blastward inlet the higher weight flow results some-
times in a sharper spiked first peak and nearly always in a weaker
second rise, from the throttle reflection. The magnitude of the first
58
peak does not appear to correlate well with flow rate, but it generally
decays further before the apparent arrival of the throttle-reflected
wave.
For the leeward inlet, Figures 3.38-3.45, the initial shock
jump is frequently somewhat greater for the higher weight flow. The
peak of the reflected wave, in the few cases where it can be observed,
because of the absence of the effect of rapid blast decay at the inlet,
Figures 3.38, 3.39, it is essentially independent of the flow rate. For
one case, Figure 3.41 (Mach 0.70, Shock Tube 2, Aps=5.0 psi), the flow
rather clearly appears to show the effects of separation within the
inlet. The case with the greater flow rate (Run 17) has a less evident
and weaker throttle reflection and a fall off that starts sooner.
3-4 EFFECT OF MACH NUMBER.
The effect of Mach number, M, on the RTI and RTO records is
presented in Figures 3.46 to 3.51 for the blastward inlet and Figures 3.52
to 3.56 for the leeward inlet.
For the blas'ward inlet, there is no significant effect of
Mach number on the first peak, either during the rise or decay portion.
Regarding the later period, there were only two conditions where there
were enough runs not having a large decay in the external blast wave,
Figures 3.46 and 3.47. There is no clear effect of Mach number on the
record after the first peak. In particular, there is no regular effect
on the second peak, from the throttle reflection.
For the leeward inlet, there is no apparent effect of Mach
number on the first peak either. Again, only two conditions had non-
decaying blast waves, enabling examination of the later record,
Figures 3.52 and 3.53, and neither of these show any discernable effect
of Mach number.
59
1.60
1.40
1.20
0.80
0.60
a. Run 11, Ap = 3.0 psi
1.60 -
1.40 _____1___ _
1.201
C' 1 .00 _____
0.60 -I
0.60
b. Run 12, Ap = 3.8 psi
1.60
1.40 -
1.2
1 .00 *
0.80
0.600.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
C. Run 13, xps . 4.8 psi
Figure 3.1. Effect of blast shock overpressure on engine-facemean total pressure in blastward inlet at Mach 0.70.Weight flow 350 lb/s (nom). Shock Tube 1.
60
1.60 -___ ________
1.40 -___ ___
1.20
0.80
0.60 -
a. Run 26, Ap = 2.2 psi
1.60
1.20
1~ _ _ _ _
0.80
0.60
b. Run 27, Aps = 3.0 psi
1.60
1 .20 TS1 .00
0.80
0.60
C. Run 29, Ap = 4.4 psi
1.60
1.40
1.20C0C,,jS1.00
0.80 -
0.600.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
d. Run 28, Aps 5.0 psi
Figure 3.2. Effect of blast shock overpressure on engine-facemean total pressure in blastward inlet at Mach 0.85.Weight flow 300 lb/s (nom). Shock Tube 1.
61
1.60
1 .40
1 .20C
t~I .00
0.80
0.60 ___ _
a. Run 15, Ap = 2.8 psi
1 .60
1.40
1.20 - _ _
C
1 1.00 -
0.80
0.60
b. Run 16, Ap = 3.8 psi
1.60
1.40- _________ _
0.80
0.60
C. Run 17, Ap = 5.0 psi
1.60
1.40A
1.200
CC" 4.00
TIME INMLIEO S
d. Run 18, Aps . 5.2 psi
Figure 3.3. Effect of blast shock overpressure on engine-facemean total pressure in blastward inlet at Mach 0.70.Weight flow 350 lb/s (nomn). Shock Tube 2.
62
1.60
1.40 _ _ _ _ _ __ _ _ _ _ _ _
C' 1.00 - -_ _ __ _
0.80__
0.60
a. Run 39, Aps 4.0 psi
1.60 1_ _ _ _
1.40 -I. _ _ _ _
1.20
0
1.00 . ~0.80 -
0.60
0.00 4.00 8.00 12.00 16.00 20.00TIMlE IN MILLISECONDS
b. Run 40, Aps = 5.8 psi
Figure 3.4. Effect of blast shock overpressure on engine-facemean total pressure in blastward inlet at Mach 0.85.Weight flow 350 lb/s (non).. Shock Tube 2.
63
1.60
1 .40
1 .20
r_ 0 .80
0.60
a . Run 19, Aps 3.0 psi
1 .60 _____
1.40- _______ _
1.20 _______ _
I .00 -
0.60
b. Run 20, Lp = 4.2 psi
1.60 ____
1.40 j _1.201
CJ1 .00 Liz
a: 0.80 ____ _ -
0.60
C. Run 21, Ap = 4.8 psi
1 .60
1.40
1 .20
C"J 1 .000.80
0.60
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
d . Run 22, Aps 5.6 psi
Figure 3.5. Effect of blast shock overpressure on engine-facemean total pressure in blastward inlet at Mach 0.70.Weight flow 350 lb/s (nom). Shock Tube 3.
64
1 .60
1 .40 _____
- 1.20
S1.00
0 .60 - __ __________
a. Run 34, Ap S> 2.0 psi
1 .60
1.40
C' 1.00
0.80 - - -
0.0b. Run 35, LAp 4.0 psi
1.60
1.20 -. __
- I0
0 .60I
C. Run 36, Aps 4.4 psi
1.60
1.40
1.20 ____
I .00
0.80-_____ ___
0.60
0.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
d. Run 37, Aps - 4.8 psi
Figure 3.6. Effect of blast shock overpressure on engine-facemean total pressure in blastward inlet at Mach 0.85.Weight flow 300 lb/s (nom). Shock Tube 3.
65
1 .60
1 .40
1.00
0.80
0.60
a. Run 41, Aps 4.4 psi
1.60
1 .40
1.20
1 1.000.80
0.60
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONOS
b. Run 42, Aps 5.6 psi
Figure 3.7. Effect of blast shock overpressure on engine-facemean total pressure in blastuard inlet at Mach 0.85.Weight flow 350 lb/s (em). Shock Tube 3.
66
1.60
1.40
- 1.20Cl 1.00 - __ _ _ _
0.60 I _ _ _ _ _ _ _ _ __ _ _ _
a. Run 11, Ap = 3.0 psi
1.60 1_ _ _ _ _ _ _ _ _ _ _ _
1 .40 _______ __________
1.20 ______ I_________C\j _ _ _ _ _W 1.00
0.80
0.0b. Run 12, As 38psi
1 .60
1.20
I .00
0.60 d ____ ~ _____
0.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
C. Run 13, Aps - 4.8 psi
Figure 3.8. Effect of blast shock overpressure on engine-facemean total pressure in leeward inlet at Mach 0.70.Weight flow 350 lb/s (nom). Shock Tube 1.
67
1.60
1.40
1.201.00 .. . .... - _ . . -.-
0.80
0.60
a. Run 26, APs 2.2 psi
1.601.40
1.20
o 1.00
0.800.60
b. Run 27, APs 3.0 psi
1.60
1.40
1.209-4
(%% _ _._ _ -_ _ _
S1.000.80
0.60
c. Run 29, Aps = 4.4 psi
1.60
1.40
1.00
0.80
0.6010.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
d. Run 28, Aps = 5.0 psi
Figuze 3.9. Effect of blast shock overpressure on engine-facemean total pressure in leeward inlet at Mach 0.85.Weight flow 300 lb/s (nom). Shock Tube 1.
69
1.60
1.40
1.20
I.100
0680
0 .60a . Run 15, Ap = 2. 8 psi
1 .60
1 .40
1 .20
b. Run 16, Ap, 3.8 psi
1 .60
1.40
~1.20-- _
0.80
0.60
C. Run 17, Ap 5.0 psi
1.60 j '
1.40
1 .20
1.00
0.80
0.60
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
d. Run 18, Aps = 5.2 psi
Figure 3.10. Effect of blast shock overpressure on engine-facemean total pressure In leeward inlet at Mach 0.70.Weight flow 350 lb/s (nomn). Shock Tube 2.
69
1.60 __ _ _ _ __ _ _ _
1.20
0.80 - _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _
0 .60
a. Run 39, Apr= 4.0 psi
1.60
1 .40
1.20 _ _ _ _ _ _ _ _
c I1.00 - _ _ _ _ _
0.80 - _ _ _ _ _ _ _ _ _ _ _ _ _ _
0.60
0.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
b. Run 40,' Aps = 5.8 psi
Figure 3.11. Effect of blast shock overpressure on engine-facemean total pressure in leeward inlet at Mach 0.85.Weight flow 350 lb/s (nom). Shock Tube 2.
70
1 .60
1.4011.20 - ____ ___ ____
0.80
0.60
a. Run 19, Ap = 3.0 psi
1.60 - ____
1 .40
1.20
1 .00
0.80
0.60-___ _
b. Run 20, Ap =4.2 psi
1.60
1.40- ___ ___ _
1 .20
C' 1 .00 -___________
0.80
0.60
C. Run 21, Ap =4.8 psiS
1 .60
1.40
1.20
I 1 .00
0.80
0.60 iL0.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
d. Run 22, Aps = 5.6 psi
Figure 3.12. Effect of blast shock overpressure on engine-facemean total pressure in leeward inlet at Mach 0.70.Weight flow 350 lb/s (nom). Shock Tube 3.
71
1.601.401.20C)
, 1-00
0.80
0.60
a. Run 34, Aps > 2.0 psi
! ,601.0 -_________. _______________1.40 I1.20 --- ---.. _ -
0.80......... .......................... .......... .............
0.60
b. Run 35, Ap = 4.0 psi
1 .60
1.40 -
1.20 -
1.00
0.80
0.60
c. Run 36, Ap = 4.4 psiS
1.60
1.40
1.20C)Co: 1.00
0.80
0.600.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
d. Run 37, Aps = 4.8 psi
Figure 3.13. Effect of blast shock overpressure on engine-facemean total pressure in leeward inlet at Mach 0.85.Weight flow 300 lb/s (nom). Shock Tube 3.
72
F ....60 t~~~~~ -' _____ ___
1.60
1.40
1.20
c 1.000.80
0.60
a. Run 41, Ap s = 4.4 psi
1.60oS
1.401.20
0.00
0.60I
0.00 4.00 8.00 1200 1600 20.0TIME IN MILLISECONDS
b. Run 42, Aps = 5.6 psi
Figure 3.14. Effect of blast shock overpressure on engine-facemean total pressure In leeward inlet at Mach 0.85.Weight flow 350 ib/s (nom). Shock Tube 3.
73
1 .60
1.401
1.20S1.00 ____
0.80
0.60
a. Run 3, Shock Tube 1, 4)=76 deg., Ap = 3.7 psi.
1.60 -
1.40 - _ _____ ____
1.20 -_ _
0.60
b. Run 4, Shock Tube 2, 4) 97 deg., Apr 3.8 psi.
1.60
t.40
t.20
C' 1.0= -
0.80
0.600.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
c. Run 5, Shock Tube 3, 4) - -94 deg., Ap = 4.0 psi.
Figure 3.15. Effect of blast intercept angle on engine-facemean total pressure in bJlastward inlet at Mach 0.55.Weight flow 350 lb/s (nom). Blast shock over-pressure 3.8 psi (nom).
74
1.60
1.40
1.200-0cJ 1 .00
0.80 " _
0.60
a. Run 6, Shock Tube 1, 4 86 deg., Aps 2.6 psi.
1.60
1.40
1.20
S1.000.80
0.60 ,
b. Run 7, Shock Tube 2, 4 = 106 deg., APs = 2.6 psi.
1.60
1.40
1.20
:: 1.00 ,
0.60
0.60
0.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
c. Run 9, Shock Tube 3, 4 = -104 deg., Aps = 3.0 psi.
Figure 3.16. Effect of blast intercept angle on engine-facemean total pressure in blastward inlet at Mach 0.70.Weight flow 300 lb/s (nom). Blast shock over-pressure 2.7 psi (nom).
75
1.60
1.40
1.20
Ct 1.00 I
0.60 -
a. Run 11, Shock Tube 1, =84 deg., tAps' 3.0 psi.
1.60
1 .40
1.20
CC I .00
0.80
0.60
b. Run 15, Shock Tube 2, 4)=103 deg., Ap = 2.8 psi.
1.60
1 .40
1.20
S1.00-
0.60
0.00 4.00 8.00 12.00 18.00 20.00
TIME IN MILLISECONDS
C. Run 19, Shock Tube 3, 4)=-102 deg., Aps 3.0 psi.
Figure 3.17. Effect of blast intercept angle on engine-facemean total pressure in blastward inlet at Mach 0.70.Weight flow 350 lb/s (nom). Blast shock over-pressure 2.9 psi (nain).
76
1.60
1.40-
1.201000c,,In 1- 0I _ , _ .. , _ _\
0.80
0.60
a. Run 13, Shock Tube 1, ( 78 deg., Aps = 4.8 psi.
1.60
1.40
1.20
c - 1 .00 ,_
0.80
0.60
b. Run 17, Shock Tube 2, = 100 deg., Aps = 5.0 psi.
1.60
1 .40
1.20
W~ 1.00-
0.80____
0.60
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
c. Run 21, Shock Tube 3, (P = -98 deg., Aps = 4.8 psi.
Figure 3.18. Effect of blast intercept angle on engine-facemean total pressure in blastward inlet at Mach 0.70.Weight flow 350 lb/s (nom). Blast shock over-pressure 4.9 psi (nor).
77
1.60
1.40
1.20CC" 1.00
0.80
0.60
a. Run 29, Shock Tube 1, 4 = 82 deg., Aps 4.4 psi.
1.60
1.40
C3 1.20 - _ _ _
0-0
0.80
0.60
b. Run 32, Shock Tube 2, 0 = 105 deg., Ap s 4.4 psi.
1 .60
1.40ure 4.)
1.20
cc 1.00
0.80 ----
0 .60
0.00 4.00 8.00 12.00 16-00 20.00TIME IN MILLISECONDS
c. Run 36, Shock Tube 3, 0 -102 deg., Ap = 4.4 psi.
Figure 3.19. Effect of blast intercept angle on engine-facemean total pressure in blastward inlet at Mach 0.85.Weight flow 300 lb/s (nom)j. Blast shock over-pressure 4.4 psi (nom).
78
1.60
1 .40
C3 1.20
I-100
0.80-___ _
0.60
a. Run 28, ShockTube 1, 82 deg., Ap = 5.0 psi.
1 .60
1.40
1.20
0.80-- _ __ " _- ___"
0.60
0.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
b. Run 37, Shock Tube 3, = 08 deg., Aps 4.8 psi.
:~;r' '.C. ff.~ct of blast intercept angle on engine-facemean total pressure in blastward inlet at Mach 0.85.'-"#ixht flow 300 lb/s (nom). Blast shock over-
;. ir 4.9 psi (nom).
1.60
1.40 - !1.20
CD
ctf 1.00
0.80 . .. ....
0.60
a. Run 38, Shock Tube 1, = 84 deg., Aps = 3.6 psi.
1.60
1.40 -
1.20C)
1.00 - --- _ "- -
0.80 10.60= 1
b. Run 39, Shock Tube 2, $ = 107 deg., Aps 4.0 psi.
Sf
1.60
1.40
1.20
0.80
0.60 -_"
0.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
c. Run 41, Shock Tube 3, * f -105 deg., Aps = 4.4 psi.
Figure 3.21. Effect of blast intercept angle on engine-facemean total pressure in blastward inlet at Mach 0.85.Weight flow 350 lb/s (nom). Blast shock over-pressure 4.0 psi (nom).
80
1 .60
1.40 - f1.20 .
S 1.00 - ---- -. __ _ --- _ _
0.60 V
a. Run 43, Shock Tube 1, 4=86 deg., Ap = 3.0 psi.
1 .60
t.40- _ _
CD 1.20 _ _ _
0.80 - - -- - _ _
0.60 _ _ _ _ _ _ _ _ _ _ _ _ _
b. Run 44, Shock Tube 2, 0 107 deg., Aps 4.0 psi.
1 .60 -- _ _ _ _ -1.40 - _ _ _ _ _ _ _ _
- 1.20 -_ _ _
1 .00 _ _ _ _--
0.80 K 4-
0.60
0.00 4.00 8.00 12.00 16.00 20.00TIMlE IN MILLISECONDS
c. Run 45, Shock Tube 3, =-104 deg., Aps = 4.2 psi.
Figure 3.22. Effect of blast intercept angle on engine-facemean total pressure in blastward inlet at Mach 0.90.Weight flow 350 lb/s (nom). Blast shock over-pressure 3.7 psi (nom).
81
1.60
1.40
1.20
1.00 -
0.80---
0.60
a. Run 6, Shock Tube 1, 0 = 86 deg., Aps 2.6 psi.
1.60
1.40
1.20
1.00
0.80
0.60
b. Run 7, Shock Tube 2, 4) 106 deg., APs 2.6 psi.
1.60
1.40 -
1.20
S1.00
0.80
0.60
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
c. Run 9, Shock Tube 3, 4 104 deg., Aps 3.0 psi.
Figure 3.23. Effect of blast intercept angle on engine-facemean total pressure in leeward inlet at Mach 0.70.
Weight flow 300 lb/s (nom). Blast shock over-pressure 2.7 psi (nom).
82
1.60
1 .40
1.20 -
S1.000.80- ____ ___ _
0.60
a. Run 11, Shock Tube 1, 4)=84 deg., Ap = 3.0 psi.
1.60- _______ ___ _
1.40 -____ ____ ___ _
1 .20
S1.00 -____
0.80
0.60
b. Run 15, Shock Tube 2, 4) 103 deg., Ap = 2.8 psi.
1.60
1.20 - ____ ____ ____
C)N 1 .00
0.80
0.60 - ____ ____ ____
0.00 4.00 8.00 12.00 16.00 20.00
TIMlE IN MILLISECONDS
C. Run 19, Shock Tube 3, 4)=-102 deg., Ap, 3.0 psi.
Figure 3.24. Effect of blast intercept angle on engine-facemean total pressure in leeward inlet at Mach 0.70.Weight flow 350 lb/s (nom). Blast shock over-pressure 2.9 psi (nom).
83
1 .60
1:40 _ _ -____- _
0.60 --
0.60
a. Run 13, Shock Tube 1, =78 deg., Aps 4.8 psi.
1.60 ___ _
1.40
1 .20
0.80
0.60
b. Run 17, Shock Tube 2, 0 100 deg., *Ap 5.0 psi.
1 .60
ix00
0.04.00 8.00 12 .00 16.00 20.00
TIME IN MILLISECONDS
C. Run 21, Shock Tube 3, *=-98 deg., Aps 4.8 psi.
Figure 3.25. Effect of blast intercept angle on engine-facemean total pressure in leeward inlet at Mach 0.70.Weight flow 350 lb/s (nom). Blast shock over-pressure 4.9 psi (nom).
84
1.60
1.40
1.20 -_ _ _____
04 S1.00
0.80
0.60
a. Run 28, Shock Tube 1, 4-82 deg., Lp - 5.0 psi.
1 .60
1 .40
1.20 -_______ ___- ________
CD1.00
0.80 -___ _
0.600.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
b. Run 37, Shock Tube 3, *=-108 deg., Aps = 4.8 psi.
Figure 3.26. Effect of blast intercept angle on engine-facemean total pressure in leeward inlet at Mach 0.85.Weight flow 300 lb/s (nom). Blast shock over-pressure 4.9 psi (nom).
85
1.60
1.40
1.20c, 1.00 --
I=
0.80 - -i~I
0.60
a. Run 29, Shock Tube 1, € f 82 deg., Aps = 4.4 psi.
1.60
1.40
1 .20
1 .00-
0.80 "
0.60
b. Run 32, Shock Tube 2, * = 105 de., Aps = 4.4 psi.
1.60! .40 . . . .. .
1.20ot 1.00
0.80
0.60
0.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
c. Run 36, Shock Tube 3, 4 - -102 deg., Aps = 4.4 psi.
Figure 3.27. Effect of blast intercept angle on engine-facemean total pressure in leeward inlet at Mach 0.85.Weight flow 300 lb/s (nom). Blast shock over-pressure 4.4 psi (nom).
86
1.60
1.40 _ __ 1_tf 1.00 -- - .
0.800.60
a. Run 38, Shock Tube 1, ¢ = 84 deg., Aps = 3.6 psi.
1.60 -_ _ _ _
1 .40
1.20] .1 ,000.60
b. Run 39, Shock Tube 2, 0 = 107 deg., Aps = 4.0 psi.
1.60
1.40
S1.20 - *I_________ I0 .80 - ' _-_
0.80
0.60
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
c. Run 41, Shock Tube 3, 0 = -105 deg., Aps = 4.4 psi.
Figure 3.28. Effect of blast intercept angle on engine-facemean total pressure in leeward inlet at Mach 0.85.Weight flow 350 lb/s (nom). Blast shock over-pressure 4.0 psi (nom).
87
1.60 i_'
1 .40 i
1.20 -- -I-c1.00 -i
0:80 -____ __ ___ n_
a. Run 43, Shock Tube 1, 0 86 deg., Ap 3.0 psi.
1.60
1 .40- - - - - -_ _ _ _ _
1.20
0.60
b. Run 44, Shock Tube 2, f 107 deg., Aps = 4.0 psi.
1.60
1.40 -
1.20
0.80
0.6C
0.00 4.00 8 00 12.00 16.00 20.00
TIME IN MILLISECONDS
c. Run 45, Shock Tube 3, * - -105 deg., Aps - 4.2 psi.
Figure 3.29. Effect of blast intercept angle on engine-facemean total pressure in leeward inlet at Mach 0.90.Weight flow 350 lb/s (nom). Blast shock over-pressure 3.7 psi (nom).
88
1.60
1.40
1.20100
0.60
a. Run 6, W2R-FS =302 ib/s, Ap = 2.6 psi.
1.60
1.40
1.20C0(" 1 .00 .
0.80
0.60
0.00 4'.00 8.)0 12.00 16.00 20.00
TIME IN MILLISECONDS
b. Run 11, W2R-FS = 351 ib/s, Aps = 3.0 psi.
Figure 3.30. Effect of weight flow on engine-face mean totalpressure in blastward inlet at Mach 0.70. Shock Tube 1.Shock overpressure 2.8 psi (nom).
89
1.60
1.40
1.20
t 1.00
0.80
0.60
a. Run 27, W2R-FS = 299 lb/s, Aps 3.0 psi.
1 .60
1.40
1.20 -- _°* I-0c 1.00 . . ......
0.80
0.60
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
b. Run 38, W2R-FS = 348 lb/s, Ap. = 3.6 psi.
Figure 3.31. Effect of weight flow on engine-face mean totalpressure in blastward inlet at Mach 0.85. Shock Tube 1.Shock overpressure 3.3 psi (nom).
90
1 .60
1.40_____
C3 1.20
" 1.00
0.80
0.60
a. Run 7, W2R-FS =302 ib/s, Aps 2.6 psi.
1 .60
1.40
1.20
S1.00 *
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
b. Run 15, W2R-FS = 351 ib/s, As= 2.8 psi.
Figure 3.32. Effect of weight flow on engine-face mean totalpressure in blastward inlet at M4ach 0.70. Shock Tube 2.Shock overpressure 2.7 psi (niom).
91
1 .60
1 .40
1.20C
CJ 1.00 ~
0.80
0.60
a. Run 8, W2R-FS =304 ib/s, Ap = 5.0 psi.
1.60 -
1.40
1 .20
0.80
0.60
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
b. Run 17, W2R-FS =348 ib/s, Aps = 5.0 psi.
Figure 3.33. Effect of weight flow on engine-face mean totalpressure in blastward inlet at Mach 0.70. Shock Tube 2.Shock overpressure 5.0 psi (nom).
92
7 A AB 511 KAMAN AVI TNE ULINGTON MA
F G 21/5FURTHER EVALUATI ON OF BLAST TESTS OF AN ENGINE INLET.(U)
_L MAR 79 J R RUETENIK. R F SM ILEY, M A TOMAYKO DNAOO-78-C-0239UNCLASSIFIED KA-TR-165 DNA-4994F ML
inuuuuuuuI-EEIIIIIEEEEE-EEEEEIIIEEEE-I.E.E.E..'
1.60
1 .40
1.20C0
I .00
0.80
0.60
a. Run 31, W2R-FS =300 ib/s, Ai = 3.8 psi.
1:60 Z j_1.20
1.00 =-
0.80
0.60- ____ ___ _
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
b. Run 39, W2R-FS = 348 ib/s, Aps = 4.0 psi.
Figure 3.34. Effect of weight flow on engine-face mean totalpressure in blastward inlet at Mach 0.85. Shock Tube 2.Shock overpressure 3.9 psi (nom).
93
1.60
1.40
1.20
I- 100
0.8010.60
a. Run 9, W2R-FS =300 ib/s, Aps 3.0 psi.
1 .60
1.40
1.20
C1. 1.00
0.80
0.60
0.00 4.00 8.00 12.00 16.00 20.00
TIME IN MI.LLISECONDS
b. Run 19, W2R-FS -344 Ib/s, Lips - 3.0 psi.
Figure 3.35. Effect of weight flow on engine-face mean totalpressure in blastward inlet at Mach 0.70. Shock Tube 3.Shock overpressure 3.0 psi (nom).
94
1.60
1.40
1.20
1.00
0.80
0.60
a. Run 10, W2R-FS = 300 ib/s, Aps = 4.4 psi.
1.60
1.40
1.20
1 1.00
0.80
0.60
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
b. Run 20, W2R-FS f 343 ib/s, Aps = 4.2 psi.
Figure 3.36, Effect of weight flow on engine-face mean totalpressure in blastward inlet at Mach 0.70. Shock Tube 3.Shock overpressure 4.3 psi (nom).
95
1.60
1.40
1 .00 _______
0.60
1.60
1.40 - - ___ _
1 .20 _____________I
W 1 .00 -- __ ___
0.80 -0.60
0.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
b. Run 41, W2R-FS - 347 ib/s, Ap8 44pi
Figure 3.37. Effect of weight flow on engine-face mean totalpressure in blastward inlet at Mach 0.85. Shock Tube 3.Shock overpressure 4.4 psi (nom).
96
1 .60
1.40
1.20
I-.00 - _ _
0.80
0.60
a. Run 6, W2R-FS =300 lb/sb A1)P 2.6 psi.
t.60r 1.401.20 ____
c~1.00 - ~ ____
0.60
0.60
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
b . Run 11, W2R-FS =351 ib/s, Aps = 3.0 psi.
Figure 3.38. Effect of weight flow on engine-face mean totalpressure in leeward inlet at Mach 0.70. Shock Tube 1.Shock overpressure 2.8 psi (nom).
97
1.601.40
1.20
I-100
0.80
0.60
a. Run 27, W2R-FS = 299 ib/s, Aps 3.0 psi.
1.40 _ ___
1.60
1 .20 --- 1 -- -
S0.80 ' ... " ::.... '
0.60
0.00 4.00 8.00 12.00 16.00 20.00TIMF IN MILLISECONDS
b. Run 38, W2R-FS = 347 ib/s, Aps = 3.6 psi.
Figure 3.39. Effect of weight flow on engine-face mean totalpressure in leeward inlet at Mach 0.85. Shock Tube 1.Shock overpressure 3.3 psi (nom).
98
1 .60
1.40
1.20-'-4
S1.00
0.80
0.60 ____ ____ ____ ___ _
a. Run 7, W2R-FS =302 ib/s, Ap = 2.6 psi.
1 .20
Cr_ 1.00
0.80
0.60 -_____
0.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
b. Run 15, W2R-FS =344 ib/s, Al,5 = 2.8 psi..
Figure 3.40. Effect of weight flow on engine-face mean totalpressure in leeward inlet at Mach 0.70. Shock Tube 2.Shock overpressure 2.7 psi (nom).
99
1 .60
1.40 -___
1.20 __ _
S1 .00 - _ _ _
0.80
0.60
a. Run 8, W2R-FS =302 ib/s, Aps 5.0 psi.
1.60
1.40 -___ ____
N 1 00 _j
0.80
0.60
I0.00 4.00 8.a0 12-00 16.00 20.00TIME IN MILLISECONDS
b. Run 17, W2R-FS - 344 lb/s, Aps = 5.0 psi.
Figure 3.41. Effect of weight flow on engine-face mean totalpressure in leeward inlet at Mach 0.70. Shock Tube 2.Shock overpressure 5.0 psi (nom).
100
1.60
1.40____ _
1.20 _ _ _ __ _
I-100-
0.80
0.60
a. Run 31, W2R-FS 300 ib/s, Ap = 3.8 psi.
1.60
1 .40 Z_
16.0 20.00
TIME Z*N MILISECNDb. Run 39, W2R-FS 347 ib/s, Aps = 4.0 psi.
Figure 3.42. Effect of weight flow on engine-face mean totalpresur inleeward inlet at Mach 0.85. Shock Tube 2.
Sokoverpressure 3.9 psi (nom).
101
1 .60
1.40
1.200-0
0.80
0.60
a. Run 9, W2R-FS =302 ib/s, Ap = 3.0 psi.
1.60 -_ _ ____ _
t.40
1.20
1 .00 ~
0.80
0.60
0.00 4'.00 8*.00 12.00 16.00 20.00TIME IN MILLISECONDS
b. Run 19, W2R-FS =351 ib/s, Aps 3.0 psi.
Figure 3.43. Effect of weight flow on engine-face mean totalpressure in leeward inlet at Mach 0.70. Shock Tube 3.Shock overpressure 3.0 psi (nom).
102
1.60
1.40
1.2
I-100
0.80 'I11 ____
0.60
a. Run 10, W2R-FS =303 ib/s, Ap = 4.4 psi.
1.60
1.40 _______ _
1.20
0.60 - ____________________
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
b. Run 20, W2R-FS = 348 ib/s, Aps = 4.2 psi.
Figure 3.44. Effect of weight flow on engine-face mean totalpressure in leeward inlet at Mach 0.70. Shock Tube 3.Shock overpressure 4.3 psi (nom).
103
1.60
1.40 ___ ___ ___-
S1.0 --
0.60
a. Run 36, W2R-FS =298 ib/s, Ap = 4.4 psi.
1.60
1.40
1.20- -_ _- _ _ _ _
100
0.80 _______________________-
0.60
0.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
b. Run 41, W2R-FS = 348 ib/s, Ap = 4.4 psi.
Figure 3.45. Effect of weight flow on engine-face mean totalpressure in leeward inlet at Mach 0.85. Shock Tube 3.Shock overpressure 4.4 psi (nom).
104
1.60
1.40
1.20
I-1 00 ~
0.80
0.60
a. Run 6, Mach 0.70, Ap = 2.6 psi.
1.60
1.40 " '
1.20 A
CI 1.00
0.80
0.60
0.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
b. Run 27, Mach 0.85, Aps 3.0 psi.
Figure 3.46. Effect of Mach number on engine-face mean totalpr ssure in blastward inlet Weight flow 300 lb/s(nom). Shock Tube 1. Blast shock overpressure2.8 psi (nom).
105
ism
1.60
1.40
C3 1.20Vck 1.00 -_ _ _ _ _ _ _ _
0.60
a. Run 3, Mach 0.515, Ap, 3.7 psi.
1 .60
1.40- _____ _
1.20
1 1.00
0.80 -___
0.60
b. Run 12, Mach 0.70, Aps 3.8 psi.
1 .60
1.40
1.20
0.60
0.00 4.00 8. *00 12.00 16.00 20.00TIME IN MILLISECONDS
C. Run 38, Mach 0.85, Ap = 3.6 psi.
Figure 3.47. Effect of Mach number on engine-face mean totalpressure in blastward Inlet. Weight flow 350 lb/s(nom). Shock Tube 1. Blast shock overpressure3.7 psi (nom).
106
1.60
1.40
1.20
1 1.00
0.80
0.60
a. Run 4, Mach 0.55, Aps = 3.8 psi.
1.60
1.40
1 .0
0.80
0.60
b. Run 16, Mach 0.70, AP = 3.8 psi.
1.60
1.401-I ___
1.20 r
.00-
0.80 _
0.60
c. Run 39, Mach 0.85, LPs = 4.0 psi.
1.60
1.40 - - -- __ _
1.20
Ck 1 00 .. . . . . .. -
0.60 -___'__
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDSd. Run 44, Ilach 0.90, Ap 4.0 psi.
Figure 3.48. Effect of Mach number on engfne-face mean totalpressure in blastward inlet. Weight flow 350 lb/s
(nom). Shock Tube 2. Blast shock overpressure
3.9 psi (nom).
107
1.60
1.40
1.20 A(D
o 1.00
0.80
0.60
a. Run 16, Mach 0.70, Aps = 3.8 psi.
1.60 _
C)1.20 /%
S1.00 -
0.80 __ _ __ _ __ _ __ _ __ _
0.600.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECON5
b. Run 39, Mach 0.85, Aps f 4.0 psi.
Figure 3.49. Effect of Mach number on engine-face mean totalpressure in blastward inlet. Weight flow 350 lb/s(nom). Shock Tube 2. Blast shock overpressure3.9 psi (nom).
108
1.60
1.00
0.80
0.60
a . Run 10, Mach 0.70, Ap = 4.4 psi.
1.60
1.40 -
1.20
CC" 1.00 I
0.80 j- -
0.60
0.00 4.00 8.010 1.2.00 16.00 20.00
TIME IN MILLISECONDS
b. Run 36, Mach 0.85, Aps 4.4 psi.
Figure 3.50. Effect of Mach number on engine-face mean totalpressure in blastward inlet. Weight flow 300 lb/s(nom). Shock Tube 3. Blast shock overpressure4.4 psi (nom).
109
1.60
t1.40
1 .20 A-~ 1.00100_1_1 _1
0.'80
0.60
a. Run 5, Mach 0.55, Ap - 4.0 psi.
1.60
1.40
1.20
0.80- ____ ___ ___ _
0.60
b. Run 20, Mach 0.70, Ap = 4.2 psi.
1 .60
1 .40
1.20
0.00 4.00 6.00 12-00 16.00 20.00TIME IN MILLISECONDS
C. Run 41, Mach 0.85, Aps 4.4 psi.
Figure 3.51.. Effect of Mach number on engine-face mean totalpressure in blastward inlet. Weight flow 350 lb/s(nom). Shock Tube 3. Blast shack overpressure4.2 psi (nom).
110
1.60
1.40
1.20
S1.00
0.80
0.60
a. Run 6, Mach 0.70, Aps = 2.6 psi.
1.60
1.40
-. 1.20"
-1.00
0.80
0.60
0.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
b. Run 27, Mach 0.85, Aps 3.0 psi.
Figure 3.52. Effect of Mach number on engine-face mean totalpressure in leeward inlet. Weight flow 300 lb/s(nom). Shock Tube 1. Blast shock overpressure
2.8 psi (nom).
- 111
1.60
1.40
1.20
0.80
0.60
a. Run 3, Mach 0.55, Aps 3.7 psi.
1.60
1.40
1.20
(C" 1.00 -_ _ _ _ _-
0.80 I
0.60 - ____ 1
b. Run 12, Mach 0.70, Ps = 3.8 psi.
1.60
1.40 I1.20
0.60
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
c. Run 38, Mach 0.85, Aps - 3.6 psi.
Figure 3.53. Effect of Mach number on engine-face mean totalpressure in leeward inlet. Weight flow 350 lb/s(nom). Shock Tube 1. Blast shock overpressure3.7 psi (nom).
112
1 .60
1.40
1.20-40cu 1.00 ..... . .. .. ..
0.80
0.60
a. Run 16, Mach 0.70, Aps = 3.8 psi.
1.60
1.40
1.20t"'l 1.00 - ]
0.80
0.60 _.
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
b. Run 39, Mach 0.85, Aps = 4.0 psi.
Figure 3.54. Effect of Mach number on engine-face mean totalpressure in leeward inlet. Weight flow 350 lb/s(nom). Shock Tube 2. Blast shock overpressure
3.9 psi (nom).
113
1.60-
1.401
1.20 IC',1.0o
0.80-- 1iij ____ _
0.60
a. Run 10, Mach 0.70, LAp 4.4 psi.
1.60 __
S1.20 ___
0.80 - . I-
0.60
0.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
b. Run 36, Mach 0.85, Apr= 4.4 psi.
Figure 3.55. Effect of Mach number on engine-face mean totalpressure in leeward inlet. Weight flow 300 lb/s(nom). Shock Tube 3. Blast shock overpressure4.4 psi (nom).
114
1.60
1.40 I____ _
1 .20
0.8
0.60
a. Run 5, Mach 0.55, bps= 4.0 psi.
1 .50
1.40 - _ _ ___ ____ __
1 .20
W 1.00
0.60 ___ _
0.60
b. Run 20, Mach 0.70, tAp 4.2 psi.
1.60 -____ .____
0.801 ____I
0.60 - ____ ______________
0.00 4.00 6.00 12.00 16.00 20.00TIME IN MILLISECONDS
C. Run 41, Mach 0.85, bps= 4.4 psi.
Figure 3.56. Effect of Mach number on engine-face mean totalpressure in leeward inlet. Weight flow 350 lb/s(nom). Shock Tube 3. Blast shock overpressure4.2 psi (nom).
115
SECTION 4
SPECIFIC EFFECTS OF BLAST AND OPERATIONAL
VARIABLES ON ENGINE-FACE MEAN TOTAL PRESSURE
In Section 3, the general features of the reduced mean total
pressure at the engine face, RTI and RTO, are examined as a function of
the test variables. In this section factors affecting the first and
second peaks for the blastward inlet, for which the first peak is the
larger, and separation within the leeward inlet are examined.
4-1 VARIATION IN FIRST PEAK OF MEAN TOTAL PRESSURE FOR BLASTWARD
INLET.
The first peak in the mean total pressure increment at the
engine face is generally during the blast period the highest. Therefore
its magnitude is an evident measure of the effect of the wave in the
inlet. The factors affecting this magnitude will be analyzed in this
section from the test results and BID code calculations.
4-1.1 Effect of Incident Total-Pressure Increment.
In Reference 2.1, Section VII, it is shown that the change in
the mean total pressure at the engine face for the blastward inlet has
about the same ratio to the incident total-pressure change (i.e.,
across the blast shock) for intercept angles of 90, 105-and 135 degrees.
This comparison was based on results of BID code calculations. In this
section the relationship will be examined for the test data.
The jump in mean total pressure at the engine face, from the
preintercept value to the first peak after shock arrival, Ap' int
2Figure 4.1, was measured for each run. It is plotted in Figure 4.2
against the jump in total pressure across the blast shock Ptos , in the
wind tunnel. The straight line through the median of these data
(50 percent on each side of the line) is given by
P2= 1.41 Apt (4.1)t2
t0
Ninety percent of the data points fall within the +11 percent of this
straight line.
116
This result means that the jump in the mean total pressure at
the engine face to the first peak was roughly 1.41 times ar large as the
jump in the total pressure across the blast shock in the free stream.
Because the spread was essentially only ±11 percent, it is concluded
that the jump at the engine face was primarily a function of the jump at
the shock.
There are two conclusions from this result. First, the blast-
ward inlet acted somewhat as an amplifier of the total pressure increase
produced by the blast shock, by a factor of approximately 1.41. Second,
the jump in total pressure across the blast shock is the principal
factor causing the jump in mean total pressure at the engine face.
Perhaps the 11-percent accuracy of prediction (90 percentile)
may be sufficient for most purposes. But an examination will be made in
the remainder of this section to identify the factors involved in this
remaining spread.
4-1.2 Effect of Test Variables on Engine-Face First-Peak Ratio.
In this subsection, the data will be examined to determine the
Ap I
dependence of __2 , called hereafter "the engine-face peak ratio,"AP
t
0
upon the test variables.
4-1.2.1 Shock Overpressure Effect.
The engine-face peak ratio is plotted in Figure 4.3 as a
function of shock overpressure for Mach 0.70 and 0.85 and full-scale
reduced weight flows of 300 and 350 lb/s. There is no apparent con-
sistent trend with shock overpressure.
4-1.2.2 Shock Intercept Angle Effect.
The variation of the engine-face peak ratio, with intercept
angle is presented in Figure 4.4 for full-scale reduced weight flows of
300 and 330-350 lb/s. Over the limited range of intercepts tested,
there is no clear variation with intercept angle at either mass flow rate.
117
Resulcs of calculations made with the BID code are plotted for
comparison in Figure 4.4b. Results for angles of 135 degrees and
greater are not plotted because the reflected wave from the throttle
returned to the engine face before the first peak was attained. The BID
results indicate the peak ratio is essentially constant over the region
of the test data. The BID value is at the low end of the test data,
which is attributed to some dissipative effects in the calculations.
The BID code results indicate that the peak ratio would be less than
the approximate value of 1.41 from these tests.
4-1.2.3 Weight Flow Effect.
The engine-face peak ratio is plotted in Figure 4.5 as a
function of the full-scale reduced weight flow, W2R.
The effect of the weight flow on the peak ratio appears to
depend upon the strength of the incident shock, based on the data for
Mach 0.70, Figure 4.5a. For shocks of 2.2 to 3.0 psi, the peak ratio
decreases with weight flow (comparative data are available only for
Mach 0.70). For stronger shocks, of 4.0 to 5.0 psi, the ratio at
Mach 0.55 decreases with weight flow, at Mach 0.70 it is constant or
increases and at Mach 0.85 it increases markedly with weight flow.
4-1.2.4 Mach Number Effect.
The engine-face peak ratio is presented in Figure 4.6 as a
fun, ion of Mach number.
There appears to be some effect of Mach number on the peak
ratio, but it is a weak effect, and the dependence appears to be a
function of the weight flow, W2R. At 300 lb/s, the trend is weak, but
it slightly decreases with Mach number. At 350 lb/s, the peak ratio
increases with Mach number.
4-2 VARIATION IN SECOND PEAK OF MEAN TOTAL PRESSURE AT ENGINE
. FACE FOR BLASTWARD INLET.
The wave entering the inlet, that produces the first pressure
peak In the reduced mean total pressure at the engine face, RTI and
RTO , Figure 4.., partially reflects from the engine (and also passes
118
through). A principal concern about the reflected wave is the possibility
of its producing separation of the flow within the inlet, by means of
the adverse pressure gradient it presents. The separation in turn could
produce unacceptable distortion at the engine face, resulting in stall
or surge and engine flameout.
The characteristics of the second peak at the engine face are
a function of the blast and inlet variables, as shown in Section 3.
These characteristics include the steepness of the rise to the second
peak, the magnitude of the second peak and the time separation between
the first and second peaks. The factors causing separation of the flow
under these transient conditions are not well understood. Therefore no
attempt will be made here to relate quantitatively the characteristics
of the second peak with the blast and inlet variables, as was done for
the first peak in Section 4.1. Instead only the qualitative characteristics
of the second peak will be reviewed here from the data presented in
Section 3.
4-2.1 Effect of Shock Overpressure.
The shape of the rise in RTI and RTO to the second peak is
found to be a function of the shock overpressure. As the overpressure
is increased, the rise steepens from a compression wave to a quite
distinct shock wave (instantaneous pressure rise). This steepening is
expected to have a significant effect on the boundary-layer separation.
This steepening of the rise to the second peak is in contrast
to the observed rise time to the first peak. The latter was essentially
independent of the shock overpressure.
The magnitude of the rise, from the initial point of rise to
the second peak, is found to be essentially independent of the shock
overpressure.
4-2.2 Effect of Intercept Angle.
The rise to the second peak is steeper to some degree for
lower intercept angles (more head-on). In general the magnitude of the
jump to the second peak also increases with decreasing intercept angle.
This is to be expected from the dependence of the first peak on the
119
intercept angle: as the angle is decreased, the jump in incident total
pressure increment increases, so the magnitude of the first peak is
greater. This stronger wave then results in a stronger reflection from
the throttle (simulated engine effect).
4-2.3 Effect of Weight Flow.
The magnitude of the rise in RTI and RTO to the second peak is
essentially always greater for the lower weight flows. This resulted
for most of the Mach numbers tested and for all three shock tubes.
This result may have importance for aircraft in flight under
high thrust conditions, such as during base escape, where the weight
flow would be high. The weaker reflected wave might result in reducing
the possibility for separating the flow within the inlet and distorting
the flow at the engine face.
4-2.4 Effect of Mach Number.
There was no observable effect of the Mach number on the
strength of the reflected shock wave, as measured by the rise to the
second peak of RTI and RTO.
4-3 SUMMARY FOR BLASTWARD INLET.
The magnitude of the jump in mean total pressure at the engine
face,AP' in these tests was primarily proportional to the jump in
total pressure across the incident shock, Aptos, within +11 percent
(90 percentile). (The remaining effect was a function principally of
the engine-face corrected weight flow and slightly a function of Mach
number.) Therefore the jump would increase with increasing shock over-
pressure and decreasing intercept angle.
The first peak of the mean total pressure at the engine face
was more peaked (sharper, spike) for intercepts at smaller angles,
more head-on.
The reflected shock from the engine-simulation throttle pro-
duced a second peak in the mean total pressure at the engine face. This
reflected wave is of concern because of possible separation of the flow
within the inlet that might result from its presence.
120
The pressure rise to the second peak steepened with increasing
shock overpressures and, to some degree, with lower intercept angles.
The magnitude of the jump at the engine face definitely increased with
lower intercept angles. Lower weight flows also produced greater rises
to the second peak, but Mach number had no apparent effect.
The BID code correlates with the trend of the variation of the
jump to the initial peak in total pressure as a function of shock
intercept angle.
4-4 FEATURES OF MEAN TOTAL PRESSURE FOR LEEWARD INLET.
The profile of the engine-face mean total pressure for the
leeward inlet varied considerably between runs. However there were
definite characteristic features of most of the profiles which are
illustrated in Figure 4.7.
The initial shock jump for the leeward inlet, as for the
blastward inlet, is one-half or less, generally, of the jump to the
peak. This initial jump is produced by the first shock that comes down
the inlet. That shock is followed by other shocks, produced by shock
reflections and diffractions, which produce the further pressure rise.
The total pressure in the records then generally levels off
until reflections of these waves return from the throttle (simulated
engine). For the remainder of the test period after the second peak
the total piessure either remains essentially steady or it falls off.
The fall-off is attributed in some cases to decay of the blast wave at
the mouth of the inlet; in other cases it is attributed to possible
separation of the flow from the inlet walls. Of all these effects,
separation is believed to be the most detrimental to engine operation,
so attention is focused here on identifying the test conditions leading
to this possible separation.
4-5 FLOW SEPARATION WITHIN LEEWARD INLET.
The assessment of blast-induced separation is tabulated for
each run in Table 4.1. The test conditions are tabulated in columns
two through five; the assessment of separation is listed in the last
three columns.
121 -; -"
TABLE 4.1
POST-TEST ASSESSMENT OF POSSIBLE BLAST-INDUCED SEPARATIONWITHIN LEEWARD INLET
Assessment ofNominal Blast-InducedFull-Scale Nominal Separation in InletReduced Shock
Nominal Weight Flow Tube Overpressure Masked ByRun Mach No. (lb/s) No (psi) None Likely Blast Decay
1 0 0 2 2.7 X
2 0.55 235 3 4.7 x
3 0.55 350 1 3.7 x4 0.55 350 1 3.8 x5 0.55 350 1 4.0 x
6 0.70 300 1 2.6 x7 0.70 300 2 2.6 x8 0.70 300 2 5.0 ?9 0.70 300 3 3.0 x
10 0.70 300 3 4.4 x
11 0.70 350 1 3.0 x12 0.70 350 1 3.8 X13 0.70 350 1 4.8 x14 0.70 350 1 4.8 X
15 0.70 350 2 2.8 x16 0.70 350 2 3.8 x17 0.70 350 2 5.0 x18 0.70 350 2 5.2 x
19 0.70 350 3 3.0 x20 0.70 350 3 4.2 x21 0.70 350 3 4.8 x22 0.70 350 3 5.6 X
23 0.70 350 1 3.6 ?24 0.70 350 2 5.8 x25 0.70 350 3 4.2 x
26 0.85 300 1 2.2 x27 0.85 300 1 3.0 X28 0.85 300 1 5.0 x29 0.85 300 1 4.4 X
122
TABLE 4.1 (Continued)
Assessment ofNominal Blast-InducedFull-Scale Nominal Separation in InletReduced Shock
Nominal Weight Flow Tube Overpressure Masked ByRun Mach No. (lb/s) No (psi) None Likely Blast Decay
30 0.85 300 2 -
31 0.85 300 2 3.8 X32 0.85 300 2 4.8 x33 0.85 300 2 -
34 0.85 300 3 >2 x35 0.85 300 3 4.0 X
36 0.85 300 3 4.4 x37 0.85 300 3 4.8 X
38 0.85 350 1. 3.6 X39 0.85 350 2 4.0 x40 0.85 350 2 5.8 x41 0.85 350 3 4.4 X42 0.85 350 3 5.6 x
43 0.90 350 1 3.0 ?
44 0.90 350 2 4.0 X45 0.90 350 -3 4.2 x
123
Separation is assessed in Table 4.1 as either: (a) none,
(b) likely or (c) masked by blast decay. "None" means that there is no
apparent sign of separation. "Likely" means that the mean total pressure
drops as qualitatively expected if separation were present within the
inlet, so it is deduced that separation was a likely cause. (This
possibility has been examined in Section 7 for one run by boundary layer
calculations.) "Masked by blast decay" indicates that separation is not
excluded, but that it cannot be verified by visual examination of the
records because of decay that was present in the blast at the inlet
mouth.
At Mach 0 and 0.55 separation appears to have occurred for all
but one case. That case was for the weakest shock of the tests at 350 lb/s
full-scale reduced weight flow.
At Mach 0.70, for the inlet unyawed, separation evidently dia
not occur for Tube 1 in any test. Separation is believed to have
occurred for Tubes 2 and 3. This would indicate that the intercept
angle is an important factor. Whether shock overpressure is also an
important factor cannot be assessed satisfactorily from these runs
because the low overpressure firings from Tubes 2 and 3 also had blast
decay present at the inlet. For the remaining tuns, Tubes 2 and I had
higher overpressures, except for one firing which was equal (4.8 psi).
For all firings with the inlet yawed, separation was either
likely or, in the case of Run 23, possible.
At Mach 0.85 there was no separation evident for Tube 1,
and blast decay masked the results for all firings from Tubes 2 and 3.
At Mach 0.90 it is possible that separation occurred for Tube 1, but
separation was masked by the present of blast decay for Tubes 2 and 3.
Blast decay at the leeward inlet was a result of the limita-
tions in the size of the blast wave that was produced by the shock
tubes. in spite of the large size of the tubes (22.6-in. ID), the
fraction of the volume of the blast wave that was satisfactory for blast
simulation was relatively small by the time the blast wave reached the
inlets. At the higher Mach numbers the wave was masked downstream
enough that the wave was unsatisfactory at the leeward inlet after a
millisecond or so. Great care was taken to locate the shock tubes to
124
meet the greatest range of test conditions within the restrictions of
the wind tunnel structure, but the blast wave at the leeward inlet could
not be maintained without sacrificing blast simulation at the blastward
inlet. The limitation in the size of blast waves that can be produced
in a wind tunnel does present a restriction to wind tunnel testing of
blast effects.
4-6 CONCLUSIONS.
The jump in mean total pressure at the engine face of the
blastward inlet following blast intercept is found to be nearly pro-
portional (+1 percent) to the jump in total pressure across the blast
shock for the range of blast conditions tested (Aps=2.2-5.8 psi,
9=76-110 deg, W2RFS=0-354 lb/s, M=0-0.90). Therefore it primarily
increases with increasing shock overpressure and decreasing intercept
angle.
The reflection of the blast wave upstream from the engine
(simulated by a choked throttle) is expected to be a principal factor to
distortion, particularly for the blastward inlet, because of shock-
boundary layer interaction, possibly resulting in separation of the flow
within the inlet from the walls. The reflection produces a second peak
in total pressure at the engine face for the blastward inlet. The rise
in total pressure approaching the second peak steepens with increasing
shock overpressure, tending to result in formation of an upstream-facing
shock within the inlet at a point nearer to the engine end of the inlet,
where the reflection enters. The magnitude of the rise to the second
peak increases for lower intercept angles (more head-on) and lower
weight flows (low engine thrust conditions). It is essentially unaffected
by Mach number.
The mean total pressure at the engine face increased more
slowly with time for the leeward inlet, following blast intercept, and
by a smaller amount.
125
A rapid fall-off in total pressure from one millisecond or
more after shock arrival that was observed in many cases for the leeward
inlet is attributed to possible separation within the inlet. It occurred
for most of the tests at Mach 0 and 0.55. At Mach 0.70 it occurred
essentially only for firings from Shock Tubes 2 and 3, which produced
higher angle intercepts. At Mach 0.85 and 0.9 it could not be assessed
because of the limited size of the blast wave.
126
First peak
Second peak
AppPt 2 it 2
Time
Pt2
RT = RT = RTI for inboard inletPt0 = RTO for outboard inlet
Apt 2
ART =
Pt 0
Figure 4,1. Sketch illustrating first and second peaks of
engine-face mean total pressure for blastward inlet.
127
0.6
Ap' AP' 1. 41 Ap 0t2 t t -
2_ 2 s
Pt0 05 -0
0 000
0
0.4 0
CD0
0
0.2
0.1
0 _0 0.1 0.2 0.3 0.4 0.5
Pt 0
Figure 4.2. Inczement in mean total pressure at blastvard engine-faceversus incident total pressure increment.
128
AP'
Ap~
s W2R-FS
2.0Tube 300 350
1 A2 0 U3 0 *
1.8
1.6
1.4 00
1.2
1.0 I0 1 2 3 4 5 6
Ap 0 (psi)S
a. Mach No. 0.70
Figure 4.3. Engine-face peak ratio versus incident overpressure.
129
2.2r
Ap't 2 I
Ip Tube W2R-FSs 2.oQ 300 r350
2 0 U
3 0 01.8 L....... _ _
1.6
@00
1.4
1.2
0 12 3 456
Ap 0 (psi)a
b. Mach No. 0.85
Figure 4.3. Concluded.
130
4)4
'a 00
0-C3
00
.14
0w 00 Q
0
9: 41-4
0 C
4 44
1313
.0 C-4 crr -jc0 -0
1N 00-:0 ~ ~ 0 nn 000 % 41
0 0 0; 0
C14
I C,
+0
0
-01-4
c0 c00
<0 0
Y) 0
0 0
.
414
0.3
2.2-
2.0-
t 2
6p t
S 1.8-
1.6-
Tube_2
1.4-
1.2
1.0 I
200 250 300 350 400 450W2R-FS (ib/'s)
a. M = 0.70, Ap = 2.2-3.0 psi.
Figure 4.5. Engine-face peak ratio versus weight flow.
133
2.2
2.0
t2
S
1.8
1.6
Tube 2 M 0.7
1.4 - Tube 3
1.2 Tube 3 M =0.55
1.c III200 250 300 350 400 450
W2R-FS (ib/s)
b. M - 0.55-0.70, Aps . 0-5.0 psi.
Figure 4.5. Continued.
134
2.2
Tube
2.0
t 2 2Apt
0s 1.8
1.6
Tube1/2
1.4 - 3~-- y
0
1.2
1.0 1
200 250 300 350 400 450
W2R-FS (ib/s)
C. m= 0.85, ts = 3.8-4.4 psi.
Figure 4.5. Concluded.
135
2.2
Tube Ap
1 2.2-3.0
2.0 0 2 3.8-5.0
0 3 4.0-4.8
1.82
Ap tP0
S
1.6
Tube 1
1.4
2 -
1.2
1.0 i I I I
0.5 0.6 0.7 0.8 0.9 1.0Mo0
a. W2R-FS = 300 lb/s.
Figure 4.6. Engine-face peak ratio versus Mach number.
136
Tube 6PS0
A 1 3.7-4.2
2.0 -. 2 3.8-4.0
S 3 4.0-4.4
t2
Ap t
1.6-Tue1
1.4
1.2
1.0vII0 .5 .6 .7 .8 .9 1.0
1110
b. W2R-FS =350 lb/s.
Figure 4.6. Concluded.
137
Throttlereflection
Typicalwithout
Further Peak decay
Shock rise
jump
• L
4 Plateau
pN
Typical decayattributed to either
Pre-intercept separation withinlevel inlet or decay of
blast at inlet
I
Figure 4.7. Sketch illustrating features of engine-face meantotal pressure for leeward inlet.
138
SECTION 5
COMPARISON OF THEORY AND EXPERIMENT
Concurrent with the performance of the inlet blast tests of
Reference i, Kaman AviDyne developed a theoretical two-dimensional
computer code, designated BID (blast-induced-distortion), for predicting
the transient flow field produced in a inlet by a blast wave striking
the inlet at an arbitrary angle of incidence (Reference 3). Preliminary
calculations of inlet pressure time histories were made with this code
for four conditions similar to those of the experimental tests and the
results of these calculations were compared with the test results in
Reference 1. These preliminary calculations were made under the assumption
that the blast wave could be represented as having a constant orientation
and strength throughout the blast event. The code results based on
these simplifications indicated that the major features of the transient
pressures observed in the inlet tests are well represented by the BID
code results. However. a more extensive comparative study of the test
results appeared to be desirable to more definitively establish the
limitations of the code predictions. To provide such a correlation, a
detailed analysis was made of the time histories of the blast wave
strength and orientation after striking the inlet for three blast test
runs (see Table 5.1) and BID calculations of inlet response were made
using these quantities as inputs. Results of the code runs and data
correlations are presented below.
5-1 BLAST INPUT CONDITIONS.
In order to perform calculations of blast response charact-
eristics by the BID code it is necessary to represent the blast pressure,
density and velocities incident on the inlets as a time dependent plane
wave. These characteristics were determined for the three runs considered
here by a detailed synthesis of the data obtained from three claw-static
probes located around the inlet (see Reference 1), with first order
corrections being made for probe-model interference. The resulting
blast input variations for the three runs are presented in Figures 5.1
139
TABLE 5.1
TEST CONDITIONS FOR CORRELATION RUNS
Run Number 39 40 18
Part Number 544 619 624
Pre-Blast Conditions:
Mach Number 0.85 0.85 0.70
Mass Flow (ib/s) 348 352 350
Blast Conditions:
Nom. Shock Overpressure (psi) 4.0 5.8 5.2
Intercept Angle (Deg) 107 110 99
140
- -- . C'°-
through 5.3. It should be noted that the blast characteristics are
defined for a long duration of over 12 ms for the first two cases
(Run 39 (Part 544) and Run 40 (Part 619)) but only for less than 4 ms
for the third case (Run 18 (Part 624)). In the last case the test
geometry resulted in large not-readily-analyzed non-linear flow dis-
turbances at later times, associated with arrival at the inlets of the
cold air jet from the shock tube which produced the blast wave.
5-2 INLET PRESSURES.
Theoretical calculations of inlet ramp and cowl pressure time
histories from the BID code (Reference 3) are compared with the corres-
ponding AEDC test data (Reference 2) in Figures 5.4 to 5.9 for Runs 39,
40 and 18. Ordinates in these figures are the ratio of inlet static
wall pressure tc pre-blast wind tunnel total pressure and the numbers to
the left of the ordinate scales represent the transducer identification
number, transducer locations being indicated in Figure 1.1.
It is seen that the BID code results are generally in good
agreement with all major features of the test data, particularly for
Runs 39 and 40. The only apparent conspicuous differences are questions
of the relative amplitudes of the calculated and experimental pressures
for a few transducers (e.g., transducer 2902 for Run 39 in Figure 5.5)
and there is strong evidence that most of these differences can be
attributed to errors in calibration factors used to reduce the test data,
as is discussed in the Appendix.
5-3 ENGINE FACE PRESSURES.
Considering next total pressures at the engine face location,
Figures 5.10 to 5.15 present comparisons of BID code predictions with
time histories of experimental total pressures for a variety of positions
at the engine face for both the blastward (outboard) and leeward (inboard)
,
In order to permit a reasonable quantitative comparison of theory andexperiment in this report, it was necessary to take into account the factthat the pre-blast (steady-state) pressures were slightly different forthe two cases. This difference was taken into account in pressure-timeplots by vertically shifting the theoretical curves so that the pre-blaststeady-state values were the same for theory and experiment.
141
inlets. Ordinate scales are the ratio of total pressure to pre-blast*
wind tunnel total pressure and the ordinate label designates the trars-
ducer identification number, transducer locations being indicated in
Figure 1.2.
5-3.1 Blastward Pressures.
BID predictions of engine face total pressures in the blastward
inlet are compared with the corresponding AEDC test data for Runs 39,
40 and 18 in Figures 5.10, 5.12 and 5.14, respectively.
Considering first Run 39, the BID pressure variations are seen
to follow well both qualitatively and quantitatively the major features
of the test data for all of the transducer locations. To be sure, the
code results do somewhat tinderestimate the initial rate of pressure
rise, the first pressure maximum, and some very rapid shock-like changes.
Considerably smaller cell sizes would have had to be used in the BID
calculations in order to permit resolution of such high frequency
variations.
5-3.2 Leeward Pressures.
BID predictions of engine face total pressures in the leeward
inlet are also generally in fairly good agreement with the test data
with respect to the maximum blast pressure level and the duration of
the principal pressure pulse (see Figures 5.11, 5.13, and 5.15). However,
it is evident that the BID pressures noticeably lag the experimental
pressures with a noticeably slower initial rise of the BID pressures
to their peak values. These two related differences may be attributed
to the fact that the BID code is a two-dimensional code which assumes
that the essentially side-on blast wave for these runs can reach the
leeward inlet only by the relatively diffuse process of diffraction
around the apex point of the inlet ramp (see Figure 1.1), whereas in
actuality the blast wave can also enter the leeward inlet more rapidly
by passing directly over or below the outboard inlet. In addition it
should be noted that the blast input characteristics used for the BID
calculations (Figures 5.1 to 5.3) may not be as applicable to leeward
See footnote on preceeding page.
142
inlet calculations as for blastward inlet calculations, since the
incident blast wave could be distorted significantly by fuselage inter-
ference effects before reaching the inboard inlet.
5-4 DISTORTION AT ENGINE FACE.
Engine face distortion time histories computed from the BID
code runs are compared with the corresponding AEDC test data in Figures 5.16
to 5.21 for the same runs discussed above. Test data and BID calculations
are shown on left and right hand sides of facing pages, respectively.
Since the BID code is a two dimensional code and the distortion
definitions normally used in B-1 studies (see Reference 1) are three
dimensional concepts involving pressures at 40 locations at the engine
face location, it was necessary to relate the BID cell pressures to the
pressures at the 40 locations. This was done for BID computations
simply by taking the BID pressure at each of the 40 engine face locations
to be equal to the pressure in the BID cell within which the corresponding
engine face transducers is located.
In Figures 5.16 to 5.21 the time of blast arrival at the blast-
ward inlet is indicated approximately by the start of the BID curves on
the right hand pages and the time of blast arrival at the engine face is
indicated by the start of the first ripples or ramplike rises of the
distortion parameters.
5-4.1 IDC.
Consider first the circumferential distortion parameters IDC1
through IDC5 for the individual engine face rings 1 to 5 in parts "a"
of Figures 5.16 to 5.21. The pre-blast values of the BID distortion
coefficients (before blast arrival at the face) are seen to be qualitatively
in good agreement with the corresponding test data for both blastward
and leeward inlets to the extent that the distortion increases signi-
ficantly in going from the inner ring (IDCl) to the outer ring (IDC5).
Quantitative agreement is only fair, with the experimental distortions
being generally larger than the BID computed distortions.
See Reference 1 for an explanation of distortion parameters and otherterminology used here.
143
After blast arrival both theory and experiment indicate
circumferential distortion increases generally lasting at least several
milliseconds. The details of distortion behavior and correlation are
somewhat different for the two inlets and for the different runs as
discussed below.
For the blastward inlet for all three runs (Figures 5.16a,
5.18a and 5.20a) the computed blast-induced circumferential distortions
are seen to be somewhat similar to the experimental distortions for the
outer engine rings (IDC4 and IDC5) but the experimental distortions are
larger than the computed distortions for the inner rings (IDCl and
IDC2). These larger experimental distortions can probably be attributed
partly to the basically three-dimensional effects of the bullet nose
hub of the model engine which is not taken into account in the two-
dimensional BID code.
For the leeward inlet for Run 39 there is somewhat better
correlation between theory and experiment for the inner engine rings
(Figure 5.17a). Here, after blast arrival, both calculated and experi-
mental distortions tend to rise to similar Levels for the inner rings
(IDCl and IDC2). However, the calculated distortions are too large for
the outer rings (IDC4 and IDC5).
For the leeward inlet for Run 40 about the same trends are
observed as mentioned above for Run 39 (see Figure 5.19a) except that
all the computed BID distortions rise to clearly too large values at
late times. This large predicted distortion might be attributed to the
fact that the BID code does not contain an adequate simulation of
viscosity effects for the circumstances of this run.
For the leeward inlet for Run 18 (Part 624) there appears to
be fairly good agreement between the experimental and calculated dis-
tortion trends for the relatively short time of the calculations.
Calculated and experimental time histories of the total cir-
cumferential distortion parameter IDC may be compared in parts "c" of
Figures 5.16 to 5.21. This parameter is a weighted combination of the
144
individual values of IDCl through IDC5 discussed above. Generally
the calculated pre-blast values of IDC are too low. The calculated
blast-induced values are too low for the blastward inlet for Runs 39
and 40 but are too high for the leeward inlet at late times for these
runs. For Run 18 calculated and experimental IDC values are similar
for both inlets.
5-4.2 IDR.
The engine face radial distortion parameters IDRl through
IDR5 and the total radial distortion parameter IDR are shown in parts
"b" and "c", respectively, of Figures 5.16 to 5.21. Generally the
experimental distortions are seen to appreciably exceed the BID-computed
distortions, particularly for the outer ring (IDC5). Some of the dis-
tortion variations are qualitatively similar, e.g., the rising distortion
variations for IDR4 and IDR5 in Figure 5.20b, but for other cases, e.g.,
Figures 5.17b and 5.19b, the experimental distortions indicate sub-
stantial blast induced distortion drops which are not predicted by
the BID code.
The differences observed here may be attributed primarily to
three-dimensional effects and viscosity effects not covered by the
BID code.
5-4.3 IDL.
The overall distortion parameter IDL is a weighted combination
of the TDC and IDR parameters (see Reference 1). The calculated pre-
blast values are generally substantially less than the experimental
values (see parts "c" of Figures 5.16 to 5.21). The calculated blast-
induced values are generally lower for the blastward inlet (Figures 5.16c,
5.18c and 5.20c) and they appear to become too high for late times for
the leeward inlet (Figures 5.17c and 5.19c).
5-4.4 IDA.
The calculated pre-blast average distortion parameter IDA is
generallv substantially less than the experimental value (see parts "c"
of Figures 5.16 to 5.21). The calculated blast-induced distortions
145
are also lower for the blastward inlet (Figures 5.16c, 5.18c and 5.20c)
but reach levels similar to the experimental values for the leeward
inlet (Figures 5.17c, 5.19c and 5.21c).
5-4.5 IDT.
The calculated pre-blast total distortion parameter IDT is
generally substantially less than the experimental value (see part "c"
of Figures 5.16 to 5.21). The calculated blast-induced distortions are
also less for most conditions except for the inboard inlet at late times
for Runs 39 and 40 (Figures 5.17c and 5.19c) where the late-time cal-
culated and experimental distortions are similar.
5-5 CORRELATION SUMMARY.
In summary, it may be concluded from the preceeding comparisons
that the BID code provides a good representation of those features of
the blast-induced inlet flow which can be reasonably represented by a
two-dimensional inviscid approach, particularly the inlet ramp and cowl
pressures and engine face pressures on the blastward inlet. For flow
characteristics which may be appreciably affected by three-dimensional
effects, such as the leeward inlet and engine face pressures, agreement
is not as good but still fair. For flow distortion characteristics,
which may be affected both by three-dimensional effects and by viscosity
effects (not included in BID), the agreement is less satisfactory. The
BID code generally underpredicts the pre-blast experimental distortions.
For the blast induced distortions, the code generally substantially
underpredicts distortions for the blastward inlet. For the leeward
inle', the code underpredicts IDR, and, for some cases, predicts too
high values of IDC and IDL at late times.
The comparisons would be expected to be generally improved by
extension of the BID code to three dimensions, particularly in regard
to distortion. This extension would require a significant increase in
the computer storage and computation time requitements. The rapid
strides in improvement of computer capacity and speed makes this
extension to three dimensions feasible now.
146
8
6V
0. 001
a) 44
950 -
44 850L
-H 40
40> CI
0
-14
6
w CO 4Q) -
*0 C/ 0
950
850 i
400
0 4 8 12 16 20
TIME (MSEC)
Figure 5.2. Blast input characteristics for Run 40 (Part 619).
148
CD____/ _
~7)
$.4 6oJ
~4
ht
0.004., O.01r
'- .. '
8 0-
Cd
0 .
" L,0 I I I
8 00 r
CU11 700 F L . ~ .
o 4~oO
0 4 8 12 16 20
TIME (MSEC)
Figure 5.3. Blast input characteristics for Run 18 (Part 624).
149
1 .25S _ _ _ _ _ _ _ _ _ _ _ _ _ _ _
a)
0.95L) 0.80
1.1
ca 0.95
2 0.65
- 0.95
1.2S
- 0.950.80
4 8 12 16 20TIME (MSECS)
Figure 5.4. Comparison of theoretical and experimental time historiesof ramp pressures for Run 39 (Part 544).
150
1.Z0__
1:2 -____
O.95
1.25
1.2s
LA 0.s
- 0.80
U. S0 11 .25 _ _ _ _ _ _ _ _ _
1. 1 a
m 0.80L
1 . 2S D _ _ _ _ _ _ __ _ _ _ _ _
1.151
1.25BID11d
1.95C'dtTLst:Dt
1.21
- 1.95 L0I .81L
1.25
1.10
1.801.65
1.51
- 1.15
- 1.95
0)13
6 12 16 21TIME CMSECS)
Figure 5.6. Comparison of theoretical and experimental time historiesof ramp pressures for Run 40 (Part 619).
152
1.*25
N 1.951.80
- 1.65
1.11
0.95
- 1.25
1.1
3.95
1.21
1.11
1.95
8 .65
1 .25
N 0.95
'46 12 is 21TlIME CMSECS)
Figure 5.7. Comparison of theoretical and experimental time histories
of cowl pressures for Run 40 (Part 619).
1.'41
1.25
L 1.2S
Uv)
(0.d'% - Cd
O.I) I
- 1.19
0.95
00
1.2S4
- 1.13
3.154
1 .25
1.1.
Cv 1.95
0.80
1. 10TetDa
Cvn 0.95
1.25S _ _ _ _ _ _ _ __ _ _ _ _ _ _ _ _
0.95
8 .65
1.50
1.10
0.95
0.65
1.10
N1 0.95
N 0.65
O . SI) _ _ _ _ _ _ _ _ __ _ _ _ _ _ _ _ _
'4 8 12 16 2nTIME (MSECS)
Figure 5.9. Comparison of theoretical and experimental time historiesof cowl pressures for Run 18 (Part 624).
155
1.11
BID Code'~-Test Data
1.33
1.23
1.1
1.33,
I 1 12 114 lie1 23TIME (MSECS)
Figure 5.l0a. Transducer 1807.
Figure 5.10. Comparison of theoretical and experimental time historiesof engine face total pressures for Run 39 (Part 544),blastward (outboard) inle2t.
156
1.413
BID Code
1331 A-~~ Test Data
1.93
@.Sol6 S 3 12 14 S 1 23
TIME CUSECS)
Figure 5.10b. Transducer 1810.
157
1.33 BID Code''--Test Data
1.93
C12 1.12
TIME (MSECS)
Figure 5.10c. Transducer 1812.
158
l.7-1
BID Code,-' Test Data
'.159
BID Code
1 33 -~-~--~--Test Data
1.23
OD
3.93
3.A3N6 3 12 14 Is is 23
TIME (MSECS)
Figure 5.10e. Transducer 1827.
160
1.1
BID CodeSTest Data
1.161
BID Code
1.33'-'-'--Test Data
1.262
BID Code*--- Test Data
1.21
1.1
1.33,
6 3 12 14 is Is 23TIME CMSECS)
Figure 5.10h. Transducer 1835.
163
BID Code
1.33 1~-' Test Data
1.2
6 312 1'# 1is 23TIME (MSECS)
Figure 5.11a. Transducer 2807.
Figure 5.11. Comparison of theoretical and experimental time historiesof engine face total pressures for Run 39 (Part 544),leeward (inboard) inlet.
164
BID CodeivA-''- Test Data
1.21
1.33
6 IN1 12 111 is 18 21TIME CMSECS)
Figure 5.11b. Transducer 2810.
165
BID Code
~--Test Data
1.21
1. 13
1.336
BID Code1.33 '~- Test Data
CD
Figure 5.11d. Transducer 2815.
167
BID CodeTest Data
Cq
TIME (MSECS)
Figure 5.lle. Transducer 2827.
168
1.38 BID Code---- Test Data
O. D
6 0 12 1'4 16 18 20TIME CMSECS)
Figure 5.11f. Transducer 2830.
169
- BID Code
1.38~ Test Data
3.9,
8.88D10 i 12 14 is 1 23
TIME (MSECS)
Figure 5.11g. Transducer 2832.
170
BID Code
f'---Test Data
1.20
U).1.
lGD
U. 8
6 8 10 12 14 16s 1 20TIME (MSECSJ
Figure 5.11h. Transducer 2835.
171
BID Code1.33 Test Data
1814is is3
1.172
BID Code
1 .38 ~" Test Data
3.90
*TIME CMSECS)
Figure 5.12b. Transducer 1810.
173
BID Code
' ' Test Data
1.31
CO
68 132I 14 is 18 23TIME (MSECS)
Figure 5.12c. Transducer 1812.
174
BID Code
' - Test Data
1.33
1.2
3.93
3.83 1is 1 12 14 is is 23
TIME (MSECS)
Figure 5.12d. Transducer 1815.
175
BID Code
STest Data
1.31
1.190OD
1.93
6 8 13 12 1'I 1s 18 20TIME (MSECS)
Figure 5.12e. Transducer 1827.
176
BID CodeTest Data
1.23
1.13 1
1.177
1.141
BID Code~--- Test Data
1.33
6 8 3 12 14 16 1 21TIME (MSECS)
Figure 5.12g. Transducer 1832.
178
1.41
1.33 BID CodeTest Data
in
1 1.13 2
TIME (MSECS)
Figure 5.12h. Transducer 1835.
179
1.413
BID Code
STest Data
1.21
1.13
6811 12 16i 113 23TIME (MSECS)'
Figure 5.13a. Transducer 2807.
Figure 5.13. Comparison of theoretical and experimental time historiesof engine face total pressures for Run 40 (Part 619),leeward (inboard) inlet.
180
1.413-- BID Code
'-Test Data
1.33
CD
1.33
ILI
S S 1 12 16 is 8i 21TIME (MSECS)
Figure 5.13b. Transducer 2810.
181
BID Code'- Test Data
Cq
1.33
TIME (MSECS)
Figure 5.13c. Transducer 2812.
182
1.13
BID Code
1.33 "'-~'-Test Data
1.O
3.183
BID Code' Test Data
1.31
1.21
( 1.13GoCq
1.91
1.61 I -6 1 1I 12 1' 16 l6 21
TIME (MSECS)
Figure 5.13e. Transducer 2827.
184
BID Code'-'"--Test Data
I1.3e3
1.213
c) 1.1,
1.383
3 6 8 13 12 16 Is i 21
TIME CMSECS)
Figure 5.13f. Transducer 21830.
185
BID CodeTest Data.
1.33.
0
Cq
TIME CMSECS)
C Figure 5.13g. Transducer 2832.
186
BID Code133 ~Test Data
CO
6 3 12 14 is is 23TIME CMSECSJ
Figure 5.13h. Transducer 2835.
187
BID Codelest Data
OD S
IL
0.888
AOAOBO 511 KAMAN AVIDTNE BURLINGTON MA F/B 21/5F URTHER EVALUATION OF BLAST TESTS OF AN ENGINE INLET U)
UA 79 J UTNK LT ATNYO DAO 8C03UNCLASSIFIED KA-TR-165 DNA-4994F NL
-mEhE3E-EhhEhEinnunnunuunnuu-mEEhhEEEE
-- BID Code~-Test Data
12 4 640 2
Fiur 5.4.TrndceU30
189
BID Code*-Test Data
19
BID Code-.- Test Data
1.40D
W" 1.0
6 8 10 12 1'4 118 20TIME CMSECS)
Figure 5.14d. Transducer 1815.
191
-- BID Code4$v-Test Data
N 1.20
1.0
I . 6oe6 is le 12 1'4 6 18 20
TIME (MSECS)
Figure 5.14e. Transducer 1827.
192
BID Codei w Test Data
roe 6
6 8 18 1 1'4 6 18 20TIME (MStCS)
Figure 5.14f. Transducer 1830.
193
BID CodeSTest Data
0
6 82 ID1Z 4 1 is1 23TIME CMSECS)
Figure 5.14g. Transducer 1832.
194
BID CodeNW"-Test Data
1. 3D
TIME (MSECS3
Figure 5.14h. Transducer 1835.
195
BID Code
1 .6 4-,-- Test Data
N*
1.20
-- 1. Ut
I.II I
6 6 13 12 14 16 16 20TIME (MSECS)
Figure 5.15a. Transducer 2807.
Figure 5.15. Comparison of theoretical and experimental time historiesof engine face total pressures for Run 18 (Part 624),leeward (inboard) inlet.
196
1 * GDBID Code-'- Test Data
CD
IAI
S 0 12 1'4 16 Is 20TIME (MSECS)
Figure 5.15b. Transducer 2810.
197
____BID Code
1 * 61 - Test Data
0
6 S ID ia 1'4 is 1S ZTIME (MSECS)
Figure 5.15c. Transducer 2812.
198
BID Code
I 1 GO' Test Data
O. S
S S 10 1ZI4 is I 20TIME (MSECS)
Figure 5.15d. Transducer 2815.
199
____BID Code
1 * 0 -w--~---.Test Data
032
048
6 8 tO 12 1'4 6 18 26TIME CMSECS)
Figure 5.15e. Transducer 2827.
200
BID Code1 . 0 t-~-~Test Data
1.40
0.601
BID CodeI. GO0~-v~ Test Data
Ct)
6 312 14 is is 20TIME (MSECS)
Figure 5.15g. Transducer 2832.
202
BID Code
* .~-w'~Test Data
0
ID 12 1'4 6 S 20TIME (MSECS)
Figure 5.15h. Transducer 2835.
203
OUTBORRO INLET
0.25
0.20
0.15
0.05 _,__ __,h_. ___ I1 __,. _
0.00- ____
0.25
0.20
c* 0.15
en 0.10 -_
0.05 f, I . ,, 1 , -
0.000.25
0.20
Mn 0.15
- 0.05 NJJAW #
0.000.25
0.20
L 0.15 -
en 0.10 -
- 0.
0.000.25
0.20 --
uo 0.15I
0.00'0.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
EXPERIMENTAL DATA
Figure 5.16a. Circumferential distortion.
Figure 5.16. Comparison of theoretical and experimental time historiesof engine face distortion for Run 39 (Part 544),blastward (outboard) inlet.
204
.25
.20
.15
(aso .1 0
.25
.20
N' .15
Uo .10
.25
.21
en .15
-4 o. Oo S
.25
.20
.15UC3
0.00
U. D
.25 -_ _ _ _ _ _ __ _ _ _ _ _ _ _
.25
.20
_ .15_
U .10
S .05
0.000 I I I
0 4 8 12 1
TIMECMSEC)
BID CODE
Figure 5.16a. Concluded.
205
OUTBORRD INLET0.100.08
- 0.06
0.04
0.02 I "
0.00
0.08
CJ 0.06
C2 0.04
0.02 I
0.00 *,Li0.100.08
mr 0.06
CD 0.04
0.02
0.00
0.06 .. . .
C2 0.04
0.000.10
0.08Lf) 0.06
23 0.04
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
EXPERIMENTAL DATA
Figure 5.16b. Radial distortion.
206
.02
10
.02
.10
.02
.08k.08k
.02k
o ~ .04
.02
0.00 4 a 12 16 21
TI ME(MSEC)
BID CODE
Figure 5.16b. Concluded.
207
OUTBORRD INLET
1 .50
0.60 nj, ", .1 P.
0.30 ., - , - - -
0.000.25
0.20------..j-
0.15Xi IL_- --
0.000.10
0.08 - - - - _ -
OI rk
0 .02 .. .. .. .. .. . . i ..
0.001 .00
0 .8 0 . . ... . . .. . +- .. . .. . .
r-0.60 - ~ --- --- ~ -
CD 0.40 ' . . . . . . .
0.20 ... .
0.00
1.000.80 --- _ _ _- _ __.
0.60
C:' 0.40
0.000.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDSEXPERIMENTAL DATA
Figure 5.16c. Overall distortion.
208
1.20
. 90
.30
.20
0 * I.0
as8
Q 20
.80
.20
.40
.20
0.0000 '4 8 12 16 20
TIME CMSEC)
BID CODE
Figure 5.16c. Concluded.
209
INBOARD INLET
0.25
0.20 10.15 -
C] 0.10 _ __ _I'
0.05 A
0.000.25
0.20
Nj 0.15 t
0.05 -V
0.00 _ _ _ _ _ _ I "' i'li m0.25
0.20 --Mr 0.15
-10
0.05 -1.----v-0.000.25
0.20 __-~ -- ~- _______
"* 0.15 - _
D- 0.10 1 70.05
0.000.25
0.20 { -
Lro 0.15
C: 0.10
0 .05
0.000.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
EXPERIMENTAL DATA
Figure 5.17a. Circumferential distortion.
Figure 5.17. Comparison of theoretical and experimental time historiesof engine face distortion for Run 39 (Part 544), leeward(inboard) inlet.
210
rJ
.25 ------
*20
atsa
.25__ _ _ _ _ _ _ _
i s
0.110
.25 _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _
.2a~cr) IF)
02
us
0
.2s
. 20
Cr)
a-- 12 1620
TIMECMSEC3
BID CODE
Figure 5.17a. Concluded.
211
INBORRO INLET
0.0
•- 0.06.04
___o_____ --._ ___-V __0.000.10
0.08
cJ 0.06 - _-_
r- 0.04
0.02
0.000.10
0.08 -- -- _
S0.04f I
0.000.10
0.08 -- __ __ _
0 .06 .. ... . .. . . . . .
C 0.04 .
-0.02 ~.A
0.10
0.1o8
0.08 Li
C030.04 (jJ "i~ii~ A~'0.02
0.000.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
EXPERIMENTAL DATA
Figure 5.17b. Radial distortion.
212
0.0I
.10
.08
.02k_, .06
cr_
.02 L
.n2
.08 _
C 08
r_
04 _
.02 _
. 0 12
.08k
.06_
o7 .]4 1_
F~. .102 __ _ __
.08
:U- .06 _
c .04
"" .02
0.00 I " I I,
0 4 8 12 16 20TIME(MSEC)
BID CODE
Figure 5.17b. Concluded.
213
" k i I I2
INBORRO INLET
1 .20 ____1l.~
0.90_: _ 0 .6 0 1
0.30 __I I0.000.25
0.20
0.152 o.'o T- _ ... _ "
0.05 A, WA0.00
0.10,o0.08
0 .0 4 . . . . .. .1I . .. .. ..
0.02 ...
0.001 .00
0.80 ___
0.60 - _ 4,C30.40- _ _
0.20
0.00
1 .00
0.80 _ - _
0.60 ...
o 0.40
0.20 - _ - - - _
0.00 I I0.00 4.00 8.00 12.00 16.00 20-00
TIME IN MILLISECONDS
EXPERIMENTAL DATA
Figure 5.17c. Overall distortion.
214
1. 2 0 L -_ _ _ _ _ _
-J.90
.25
. 20
.10
o ~ .Lf
.02
. 80
- .20
0.00 L
1.00 -_ _ _ _ __ _ _ _ _ _
. 80
.Go0
C2 40l
.20
0 .0 00 '4 8 12 1 G 20
TI MECMSEC)
BID CODE
Figure 5.17c. Concluded.
215
OUTBORRO INLET
0.25
0.20
--' 0.15
2 0.10 ?0.00 ki
____
0.25
0.20
0.15
- 0.05 VO
.00
0.00 o* , , V .,. , _____________.....__
0.25
0.20
') 0.15
-3 0.10 ____ ______0.150.000.25
0.20
0.15
2D 0.100.05 A A Ititrf' , rw0.00
0.20
u' 0.15
0.000.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDS
EXPERIMENTAL DATA
Figure 5.18a. Circumferential distortion.
Figure 5.18. Comparison of theoretical and experimental time historiesof engine face distortion for Run 40 (Part 619),blastward (outboard) inlet.
216
.28
.23
- .15
P . go
.21
o .10- .15
.25
.21
.15
o .13
.25
.28
.15
0.10
.20
LP .15U-0
3.333 4 612 16 2
TIMECMSEC)BID CODE
Figure 5.1.8a. Concl1uded.
217
OUTBOARD INLET
0.10
0.08
- 0.06
0.04
0.02 _,JILL ii0.000.10
0.08
CI 0.06
C3 0.04
0.02
0.00
0.10
0.08
' 0.06
0.04
0.020.00 I .... m... .
0.10
0.08
, 0.06o: 0.04 f .
0.02 r Til rjT--~ ~ . 4 '+'~iit .j
0.000.10
0.08{
0.06 I 11
- 0.04
0.00
TIME IN MILLISECONDSEXPERIMENTAL DATA
Figure 5.18b. Radial distortion.
218
.16
.06_
0: ....
o o
.82
.10
DDu t ,I I I
* D
.08.84
.02
- . .2
uOu DDIt I
" .02
uTuI I EI MSC.18
.18.0D2
D.D Be*
a1 __ __ _ _12__ __ _ __is__ __ __ _21___ _
TIMCusC
ID CODSEC
Figure 5.18b. Concluded.
219
OUTBORRO INLET1.50
1.20
0.90 -. __ __ _
C3 0.60 ___1 WZ Aki4 tJ0.30 f____0.000.25
0.20
0.15
0.00 ____
0.10____ _
0.08 - _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _
0.06
0.04 Ihhdj r0.02-[
0.00 -
1.00-
0.80
0.60
2: 0.40
0.20
0 .00
0.80
0.20
0.00 _____
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
EXPERIMENTAL DATA
Figure 5.18c. Overall distortion.
220
1.23
. 93
.30
.21
o .10
0.10
.138
.12
.63
- .20
1.63
.83
.23
3.033 4 IN 12 16 21
TIMECMSEC)
BID CODE
Figure 5.18c. Concluded.
221
INBO:RO INLET
0.25
0.20
0.15
2 010 -L] 4
0 .00 "" . . . • *
0.25
0.20
0.15 i i j).L , w ___,____2l 0.10
0.05
0 .00 - • "_ -0.25
0.20 --
') 0.15
00 _____ ,____ _ ______ _____
0.25
0.20 - _ _ _ _ _ _ _ _ _ _ _
0.15 Jul,
0.000.25
0.20
LI) 0.15
0.0120 000 000.,o .o,.o o oo
TIME IN MILLISECONDSEXPERIMENqTAL DATAFigure 5.19a. Circumferential distortion.
Figue 519.Comparison of theoretical and experimental time historiesof engine face distortion for Run 40 (Part 619),leeward (inboard) inlet.
222
.25
.21
2.18
L)o .11
- . go
.2S
.1
U .10
- . Be
3.2S
.21
.15
Po .15
.2S
.21
C- Io3 al
AuSL
1 4 1 12 to 23
TIMECMSEC)
BID CODE
Figure 5.19a. Concluded.
223
0.1 _____INBORO INLET
0.08
0.06D0.04
0.00
0.020.10
0.08
CJ 0.06C3 0.04
0.02-
0.00 lJ AM Li0.10
0.08
m0.06o0.04
0.02
0.00.10
0.08
** 0.06
S0.04 ____1
0.0 li t____ "l________
Ll 0.060.10__ I
o0.04 8iA
LO0.003 0.00 4.00 8.0 20 6000
0-224
o .s
.13
.36
.04
1.13 0
o B
.3
3.30
.36
- .12
LO .46 21,2
TI ME CMSEC)
BID CODE
Figure 5.19b. Concluded.
225
INBOARD INLET
!.50
1.20
0.90 "" ' j'
0.00
0.25
0.20
0.15
2, 0.10
0.05
0.000.10
0.08
0.06
2 0.04
0.021A
0.001.00
0.80
0.80
0.60
0.20
0.000.00 4.00 8.00 12.00 16.00 20.00
TIME IN MILLISECONDSEXPERIMENTAL DATA
Figure 5.19c. Overall distortion.
226
1.23
.90
C33
.25 __ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _
.20
p.-5
.13
ct* 06
o- .02
S. of
.91
- .21
O * 12t32
.227
OUTBORRO INLET
0.25
0.20
- 0.15_ _ _ I i2 0.10 )IT r !r !S
0.05 1:, i$[I } il I
0.00 _____rI
N0.25
0.20-
2R 0.10
0.205
M 0.15 I-20.10
0.00L__ _ _ _
0.25T
0.20 _ _ _ 1_ _ _ _ _ _
LI) 0.152: 0.101 W
0.05 20.00
0.00 402.0150 60 00TIEINMLLSCOD
EXEIMNA DATAFiue52aUicmeeta itrin
Figure gur 5.2.2omario ofrctheorential adiexperienta.iehitre
of engine face distortion for Run 18 (Part 624),blastvard (outboard) inlet.
228
.25
.20
.2S
.20 I a
co .15
_.LU _
O.* 25 _ _ _ __ _ _ _ __ _ _ ____ _ _ _ _ _ _ _ _ _
,25
.20k
C.1,
o .10
0 .005 _ _
.2U
0.20_ 0.00 I I I I
14 8 12 1620
TI ME MSEC)
BID CODE
Figure 5,20a, Concluded.
229
OUTBORRO INLET
0.10
0.08
-0.06
C20.04 ]A0.02
lt
0.040~0.10
0.08
('j 0.04
0.08
S0.060
20.10 if ~ -
0.02
m0.00
0.04
0.02 _______i. iI0.00
0.00 .0 80 20 60 00
0.230
S04
.08L
CI I,131
.02
0.00 I I I I
.08 _8
.04
.02
08 _
.02_
0].O0 L._.-.- .
U G.02 C
.I] _41
T I ME C MSE CBID CODE
Figure 5.20b. Concluded.
231
OUTBORO INLET
1.50
2 0.60141; I_ _
C30.30
0.000.251
0.20
0.15 4
0.0504 4 ~ I~1___0.10
0.08 - _ _ _ _ _ _ _ _
0.06
C, 0.0C4 iil1
0.02
0.001 .00 _____
0.80
CC0.60
0.00____ _
1.00__ _ _ _
0.80
0.60
CR 0.40__ _ -- if0.00 _____
0.00 4.00 8.00 12.00 16.00 20.00TIME IN MILLISECONDS
EXPERIMENTAL DATA
Figure 5.20c. Overall distortion.
232
1. Inv._ _ _ _
1.20k
.go
.30L
.25
.20f-
.Bs
,08
*06
.02
Ua a _ _
.810
.GD
<]i0
.20
a.80 a
.00 8 1212020
TIMECMSEC)
BID CODE
Figure 5.20c. Concluded.
233
0.25
0.20
- 0.15
2C3 0.100.0o .os I, ,,..i ll), . , ., ,. . .,,,
0.25
0.20
C\I0.15
2 0.10 .,
0.000.25
0.20
Cr 0.15(..)
2 0.10
0.05 j ''iF.t' I 1 *.O"
0.00 OW
0.25
0.20
V, 0.15 Jj
62 0.10 -t0.05 'ti.t '
.0.000.25
0.20
LID 0.15
0.05 4v 1u rv * r
0.000.00 4.00 8.00 12.00 16.00 20-00
TIME IN MILLISECONDSEXPERIMENTAL DATA
Figure 5.21a. Circumferential distortion.
Figure 5.21. Comparison of theoretical and experimental time histories
of engine face distortion for Run 18 (Part 624),leeward (inboard) inlet.
234
.25
.20
.2o L
n_ .l a
.25
.20V
uLF
0 .D L I
.25 _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _ _
Figur 5.1. ocldd
* 230
I-. .
*, . L * J -- '-I-
.1LV
i 0 0 .L 4
U 812 1G 2
BID CODE
Figure 5.21a. Concluded.
235
INBORRO INLET
0.10
0.08
O.0 'IC 0.04
0.02
0.10-
0.08
(Y) 0.06
c 0.04
0.02 -_"_
0.10
0.08
*r) 0.06
o 0.04
0.02 -4 lta
0.00 ..... LJ . . ..
0.10
0.08
0.06
0.00
0.02
0.000.10
0.08
0202
0.00 4.00 8.00 12.O0 16.00 20.0WTI!ME I N MI]LLIS ECONDS
EXPERIMENTAL DATA
- Figure 5.21b. Radial distortion.
: 236
* oe.10
o02
04O
2-0
LL .°04 f
00
0.00 I I i
. 02
c-- .F02
. 08 L
_7 .04- -02.02
0.00 I I ____
U4 8 12 lG 20TI ME CMSEU J
BI:D CODE
Figure 5.21b. Concluded.
237
INBORRO INLET
1 .50
1.20
0.90
23 0.60 ~I
0.30 '_____
_____
0.000.25
0.20
0.15-___ ____
____
0 0.10 -____ ________
0.05 '
0.10
0.080.06 1jJ
0.02
0.80 _ _ _
C: 0.40
0.20N-6
0.00____ -
0.0040080120160200
0.238
1.20
.90
.30
. 25 __ _ _ _ _ _ _ _ _ _ _ _ _ _ _
. 20
. 15Uo .10
0.0 as L
.08
0 .04
.02
.80
os
- .20k
a 0
08 12 is20
TIMECMSEC)
BID CODE
Figure 5.21k. Concluded.L - 2394
SECTION 6
EVALbATION OF LATE TIME LARGE DISTORTION VALUES
It was found in the inlet-blast tests of Reference 1 that
generally very few large distortion values were obtained during the
early-time, definitely blast-type, flow periods of the tests. However,
there were some observed large distortion values at times after the
fairly definite blast type flow duration of about 3.3 ms, which appeared
to deserve more detailed study.
As part of the present study, a detailed re-evaluation was
made of the 16T inlet blast test results to determine whether the large
distortion values observed at late times (after the nominal blast event)
on some firings could be attributed to inlet response behavior or to
other effects. An initial appraisal identified 26 firings with signi-
ficant late time distortions which appeared worthy of consideration.
For these runs it was found that the large observed IDL values could be
correlated with one or more of the following circumstances.
1. Some IDL peaks correlated closely to significant rises in
total pressure, indicating that these peaks were produced
by the effects of the cold jet from the shock tube
passing directly into the inlet.
2. Some IDL peaks corresponded closely to rapid changes in
input pressures measured by the claw probes, indicating
that the distortion values were caused primarily by input
variations rather than inlet response effects.
3. For some runs, there were significant differences in the
input pressures measured by the different claw and static
probes, indicating a considerable distortion of the flow
entering the inlet. Hence, for these runs the engine
face distortion could have resulted from the input
distortion rather than from inlet response effects.
240
4. For several runs, large apparent distortion peaks were
caused by the readings of one pressure transducer (1802),
whose readings were somewhat erratic and were inconsistent
with the measurements of two adjacent transducers (1801
and 1804).
5. For at least one run, large apparent distortion values
were attributed to erratic behavior of a group of 4
pressure transducers which were recorded on the same
oscillograph track. These transducers gave pressures
inconsistent with adjacent transducers.
6. In some cases transducers bottomed, producing false
indications of large distortion values.
In summary, no cases were found where large late time distortion
values (IDL > 1) were obtained which could be definitely attributed to
inlet response behavior. In all cases there was evidence to suggest
that the large distortion values could be attributed at least in part to
either cold gas jet impact on the inlet, large input distortions or
transducer malfunctions.
The above observations do not completely resolve the question
as to whether any of the large late time distortions observed in the
tests can be attributed to inlet blast response effects. It can only be
said that the factors pointed out above would appear to make any such
determination from the late time test data a difficult matter and any
results of such a determination could be suspect.
241
-I. .' ' . . - 7 . . - - : '- ... . ........ > ' ...,, .. . : -iF .... -- _
SECTION 7
REFLECTED SHOCK-BOUNDARY LAYER INTERACTION
The reflected wave from the engine, as it moves upstream
through the inlet, makes an adverse piessure gradient for the flow
within the inlet. The adverse pressure gradient produces some dis-
tortion of the flow. If the gradient is large enough, boundary-layer
separation takes place and large distortion of the flow would result, as
pointed out in Section 3 and 4 and in Reference 1.
A sketch illustrating the effect of the reflected waves on
the boundary layers in the two inlets is shown in Figure 7.1. The waves,
shown here as two shocks, produce an adverse pressure gradient within
the boundary layers on the walls of the inlets. The pressure gradients
produce more rapid thickening of the boundary layers and retardation of
the flow within the layers. If the gradients are sufficiently large,
separation of the flow from the walls would take place resulting in
large distortion.
There were indications that separation by the reflected shock
may have taken place in some of these tests. The problem is that distortion
at the engine face would have resulted after the end of the test period,
about 3 ms, so effect of separation could not be verified.
Calculations are presented in Reference 1 using the NASA
BLAYER code of Reference 4. As pointed out there, the BLAYER calculations
are limited by the assumption of a constant total pressure and total
temperature in the free stream, whereas the variations produced by blast
interaction can be large (Reference 1).
NASA has generalized the BLAYER code, subsequent to the work
published in Reference 4, to include variations of the total pressure
and total temperature in the free stream. A copy of the generalized
BLAYER code was provided to Kaman AviDyne (Reference 5), and it has been
employed for the calculations reported below.
242
7-1 SHOCK-BOUNDARY LAYER CALCULATION.
Calculations of the boundary layers on the inlet cowls and
ramps were made using pressure, temperature and velocity data from BID
code results for the free-stream conditions. BID results for properties
in the cells contiguous to the respective wall were employed. The BID
data used were the same as for the calculations reported in Reference 1:
BID Run 10/26/77, M=0.85, W2R=350 lb/s, 4=90 deg and Aps=5 psi. The
data selected were for a fixed time of 33.6 ms (full scale) after blast
arrival, when the reflected waves were well into both inlets.
The generalized BLAYER code assumes a steady-state boundary
layer, i.e. the conditions do not vary with time. Therefore the BID
data distributions for the free-stream properties were assumed to be
frozen, i.e. non-varying with time. In actuality the reflected waves
move upstream with time, so the properties are time varying. The
assumption of a steady state flow is believed to be conservative, because
the pressure gradients experienced by the boundary layers with the
moving wave would be greater than for the assumed frozen conditions.
The results of the calculations are expressed in terms of a
parameter called the boundary-layer factor Hi, essentially representing
the velocity distribution across the boundary layer (normal to the
wall). If the velocity distribution is more uniform, the parameter
approaches unity. A value of 1.2 to 1.4 is typical for a turbulent
boundary layer on a flat plate in the absence of a pressure gradient.
In an adverse pressure gradient (increasing pressure in the direction of
flow) the velocity decreases towards the wall so Hi increases. When the
velocity gradient normal to the wall goes to zero, the flow can separate
from the wall. Experience shows that separation may occur for Hi
values as low as 2.0 and definitely by 2.8.
The generalized-BLAYER calculations were carried out for three
cases: the boundary layers on the two cowls and on the blastward
splitter ramp. The boundary layers were assumed to be turbulent from
the leading edges.
243
Calculations were not made for the leeward splitter ramp,
because of undefined starting conditions. The data indicate a possible
separation bubble near the leading edge with probable reattachment. The
boundary-layer conditions are not believed to be well enough defined at
this point for meaningful boundary-layer calculations.
The results of these calculations are presented in Figures 7.2
to 7.4. The input data (PI' Pt' Tt) and output (Hi) are presented as
functions of the station along each inlet, measured from the leading
edge of the splitter ramps.
Within the blastward inlet the primary part of the reflected
wave at 33.6 ms extends between ramp Stations 12 and 17 (ft), as indicated
by the rise in total pressure. On the blastward cowl there is an adverse
(positive) pressure (static) gradient between Stations 8 and about 13
due to recovery within the inlet. The pressure-gradient decreases
somewhat between about Stations 12 and 13.6, where the pressure gradient
of the reflected wave is picked up. The form factor H. increases to a1
peak value of 1.57 at station 11.7, which would be well below the
minimum separation value of 2.0. H. then decreases to 1.38 at Station 14.81
where the adverse pressure gradient of the reflected wave causes it to
increase rapidly, reaching the separation value of 2.8 at Station 16.8.
At the ramp the adverse pressure gradient begins at about
Station 10. H. is fairly constant to about Station 16 where the gradient1
increases markedly. The separation value of 2.8 is reached at Station 18.3.
In the leeward inlet the reflected wave has not had a signi-
ficant impact on the flow at this time (33.6 ms). There is some effect
over about the downstream half of the inlet but the effect is to reduce
somewhat the large negative gradient in total pressure produced by the
incident blast wave. The static pressure gradient produced by the blast
wave is large, so Hi rises all along the cowl, and the boundary-layer
separation value of 2.8 is reached by Station 20.6. The pressure
essentially levels off beyond that point, so the determination of
separation is not as definite as for the leeward cowl.
244
It is concluded that separation caused by the reflected waves
would have occurred in the blastward inlet and possibly in the leeward
inlet. There are several factors that are believed to make the cal-
culation conservative (under-predict separation). First, a fan stage is
expected to reflect a stronger wave than the choked throttle that was
employed in the BID calculation. Seccnd, the BID code is dissipative,
so adverse pressure gradients are expected to be somewhat under-predicted.
Third, the generalized BLAYER code applies for a steady-state boundary
layer, and the unsteady effect is expected to increase the tendency
toward separation.
Tests are needed having blast durations that are sufficiently
long to observe the separation and resultant distortion at the engine
face. The test period for these test conditions should be roughly
doubled.
245
BLAS, WAVI
SHOCK
SHOCK DISTORTION
1INL ET FAN
Figure 7.1. Boundary layer separation and distortion from fan reflected shock wave.
246
3. DI
-
2. t1t
1.50.....
O S I O 15 20 35STATION (FT)
T tlTt
t
l . to
3.63 Pto
0
S 13 is 2 2sSTATION (rT)
Figure 7.2 - Predicted free-stream properties and form factorH of boundary layer on cowl of blastvard inlet.Ater reflected wave enters inlet.
247
3.10
I.SI
I-t
0. 0
0.88
..... .. ....... ....... ........ I .. . .AAA'D AA.6
I I is252
STATION (FT)
Fiur .3 Prdi tedfe-tempoete n omfco
3.838
3.33
<-
Ixc_2.UD
a
U2.28 2SsT,^TioN crT)
t
t 0
1.61
3 1 -t3 232
STATION (rT)
Figure 7.4 -Predicted free-stream properties and form factorH of boundary layer on cowl of leeward inlet.Ater reflected wave enters inlet.
249
Iwill
SECTION 8
CONCLUSIONS
From evaluating the results of blast-wave engine-inlet inter-
action tests with a B-1 type engine inlet, the following conclusions are
reached.
1. The blast interaction with the windward inlet produced a rise
in the mean total pressure at the engine face followed by a
decay. A second rise occurred due to reflection from the
choked control vanes simulating the engine. The magnitude of
the first peak is found to be nearly 1.41 times as large
(+ll percent) as the increment in total pressure across the
incident blast shock. This ratio varies some with inlet
weight flow and Mach number but it is unaffected by changes
in shock overpressure and intercept angle, over the range
tested (76 to 110 deg).
2. The reflection of the blast wave from the engine is expected
to be a potential cause of distortion through boundary-layer
retardation. The rise rate approaching the second peak in
engine face total pressure increased with shock overpressure.
The magnitude of the rise to the second peak increased with
lower intercept angles (more head-on) and lower weight flows
(low engine thrust conditions) and was essentially unaffected
by Mach number.
3. The mean total pressure at the engine face rose more slowly in
the leeward inlet than the blastward inlet, following blast
intercept, and by a smaller amount. In many of the tests the
rise in total pressure was followed within a millisecond or so
by a rapid fall-off, attributed to possible separation of the
flow from the walls of the inlet.
250
4. The two-dimensional BID code provides a good representation of
those features of the blast-induced inlet flow which can be
reasonably represented by a two-dimensional inviscid approach,
particularly the inlet ramp and cowl pressures and engine face
pressures for the blastward inlet. For flow characteristics
which may be appreciably affected by three-dimensional effects,
such as the leeward inlet and engine face pressures, agreement
is not as good but still fair. For flow distortion characteristics,
which may be affected both by three-dimensional effects and by
viscosity effects (not included in BID), agreement is less
satisfactory. To improve this situation, extension of the BID
code to the three-dimensional case appears feasible.
5. A study of large apparent distortion values observed at late
times during the AEDC tests, after the limited test period of
about 3 ms, indicated no cases where these large distortion
values could be definitely attributed to inlet response behavior.
The limited test period was too short to permit observation of
possible large late time distortion effects not masked by
extraneous factors. It is recommended that the test duration
for future tests be increased, by a factor of two or more.
6. Calculations of the effect of the engine-reflected blast wave
on the boundary layers in the inlets, made using the generalized
NASA-Lewis BLAYER code, agree with the previous results
(DNA 4590F) that indicated boundary-layer separation would
occur on the cowl and splitter of the blastward inlet and
possibly on the cowl of the leeward inlet. The blast shock
overpressure was 5 psi and the intercept angle 90 degrees.
251
REFERENCES
1. Ruetenik, J.R., and Smiley, R.F., Wind-Tunnel Shock-Tube Simulationand Evaluation of Blast Effects on an Engine Inlet, Kaman AviDyneReport TR-147, DNA 4590F, 1978.
2. AEDC Reports, Data Tabulations, Plots and Tapes on the DNA B-1Blast Effects Test, AEDC Project P41T-D2A, Test TF-419, 1976-1977.
3. Thompson, J.H., Tomayko, M.A., and Ruetenik, J.R., BID Code forComputing Aircraft Engine-Inlet Response to Blast Disturbances,Kaman AviDyne. Unpublished.
4. McNally, W.D., FORTRAN Program for Calculating Compressible Laminarand Turbulent Boundary Layers in Arbitrary Pressure Gradients,NASA Technical Note D-5681, May 1970.
5. McNally, W.D., NASA Lewis Research Center, Correspondence toDr. Ruetenik, 28 April 1978.
252
L i___ _ _
APPENDIX
TRANSDUCER CALIBRATION EVALUATION
The experimental ramp and cowl pressures measured inside the
subject model inlet during the tests of References 1 and 2 are of
interest both for assessing blast-induced inlet loadings and for pro-
viding a basis for evaluation of the BID code for predicting inlet
pressures and velocities. However, in using these data for such
purposes it is important for the data user to appreciate that some of
these data appear unreliable to some extent because of experimental
difficulties experienced during the tests. This appendix points out
briefly the principal problems encountered and indicates the degree
of reliability of different parts of the data.
The primary problem experienced with the ramp/cowl pressure
transducers was an inability to calibrate the transducers accurately
during the test period due primarily to unanticipated constrictions
in some of the tubes connecting the transducers to the calibration
pressure source. This problem could not be resolved in the very limited
time that was available for the model tests. The resulting data as
presented in References 1 and 2, therefore, had to be reduced on the
basis of sometimes nominal or questionable calibration results.
To clarify this calibration problem, KA examined the test
data for all ramp/cowl transducers with the aim of identifying questionable
data and providing correction factors if possible. Data calibration
errors were identified by such means as comparing transducer pressures
for the same transducer for similar runs, and/or by comparing transducer
pressures for adjacent transducers which should have about the same
pressure on the average. E.g., transducers 1970 and 1903 face each
other at the same axial location inside the outboard inlet, and normally
have quite similar late-time blast-response pressure time histories,
both according to many test runs and to the BID code. Consequently,
since transducer 1970 had no apparent calibration problem, any strong
253
difference between the indicated pressures for these two transducers
can be reasonably interpreted as an indication that the calibration
factor used for transducer 1903 is unreliable.
Using comparisons of this type, Table A.1 was prepared, which
indicates the KA estimate of the degree of reliability of the calibration
factors used to reduce the data presented in References 1 and 2. In
this table, questionable data are identified as either H or L, depending
on whether the pressure values presented in References I and 2 appeared
conclusively to be too high (H) or too low (L). Data for some other
runs also appears questionable to a lesser extent (not indicated in
the table), but the evidence for such cases is less conclusive.
Some attempts were made to obtain correction factors for some
of the questionable data, but it appeared that additional information
(not available for this study) would be required from AEDC to effectively
accomplish this purpose.
254
TABLE A.1
RAMP/COWL TRANSDUCER EVALUATION
AEDC Ramp/Cowl Transducer NumberPartNo. 1903 1905 1935 1950 1970 1990 2902
512 X X H X X H H513 X X H X X H H517 X H H X X H H518 X H H X X H H519 X X H X X H H525 X X H X X H H526 X H H X X H H527 X H H X X H H544 X X L H545 X X L H546 X L H550 X X L H551 X X L H553 X X L H558 X X H559 X X H568 X X H569 X H570 X H573 X H574 X H582 L H583 L H584 L X X H589 H H590 H H591 H H596 L X X X H597 L X X H598 L X X X H600 L X X H601 L X H602 L X X X H607 X H608 X X H615 X H619 X X H620 X H621 x H624 L X x625 L626 L X
*Transducers 1902 and 1980 appeared generally reliable; transducer 1904
provided no useful data.
X Designates no useful data obtained.H Indicates pressure values in Reference 2 appear to be too high.L Indicates pressure values in Reference 2 Pppear to be too low.
255
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Boeing Wichita Co. Kaman Sciences Corp.ATTN: R. Syring ATTN: D. Sachs
Calspan Corp. Los Alamos Technical Associates, Inc.ATTN: M. Dunn ATTN: P. Hughes
ATTN: C. SparlingUniversity of Dayton
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258 :
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