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Distribution: Further dissemination onlyas directed by Federal AviationAdministration, Office of SupersonicTransport Development, Washington, DC,20553, Sep 1966, or higher DoD authority.
AUTHORITYFAA ltr, 10 Oct 1972; FAA ltr, 10 Oct 1972
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THIS DOCM T CONTAINS INFORMTION AFFECTING TIlE NATIONAL DEFENSE OFTHE UNITED STATES WITHIN THE MEANING OF THE ESPIONAGE LAWS, TITLE 18,U.S.C. SECTIONS 793 AND 794. THE TRANSMISS1ON OR THE REVELATION OF,ITS CO ENTS IN ANY .MNNER TO AN UNAUTHORIZED PERSON IS PROHIBITED BYLAW.
NOTICE: When government or other drawings, specifications or otherdata are used for any purpose other than in connection with a defi-nitely related government procurement operation, the U. S. Governmentthereby incurs no responsibility, nor any obligation whatsoever; andthe fLct that the Government may have formulated, furnished, or in anyway supplied the said drawings, specifications, or other data is notto be regarded by implication or otherwise as in any manner licensingthe holder or any other person or corporation, or conveying any rightsor permission to manufacture, use or sell any patented invention thatmay in any way be related thereto.
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NOTICES
When Government drawings, specifications, or other data are usedfor any purpose other than in connection with a definitely relatedGovernment procurement operation, the United States Governmentthereby incurs no responsibility nor any obligation whatsoever;and the fact that the Government may have formulated, furnished,or in any way supplied the said drawings, specifications, or otherdata, is not to be regarded by implication or otherwise as in anymanner licensing the holder or any other person or corporation, orconveying any rights or permission to manufacture, use, or sellany patented invention that may in any way be related thereto.
All distribution of this document is controlled. In addition toset.urity requiremetits which apply to this document and must be met,it may be further distributed by the holder only with specificprior approval of:
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The distribution of this report is limited because it containstechnology identifiable with items excluded from export by theDepartment of State (U. S. Export Control Act of 1949, as amended).
L] ~~CONFIDENTIAL SPEBR16
IEENIN -R PO POAL
L ~~ FOR -PHASE TM OF~ THEIUPERSONIC TRANSPORT DIEVELOI'MENT PROGRAM
TECHNICAL/ ENGINE.
EMTRD.D
INSTALLATION AND INLET eo 0mSYSTEM COMPATIBILITY 4 Q 2 U U LjL~
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IPratt &Whitney Alrcraft lpoOFrAUuFLORIDA REW RCH AND DEVELOPMENT CENTER
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Pratt & Whitney Aircraft Mw WT PT1 AM EUU94.PWA FP 66-100 --
Volume IIIReport D
CONTENTS',
SECTION *-/PAGE
I INSTALLATION COMPATIBI!.f-- -..-. I_ IA. Scope and Obj..lv. ---------------------- -- -------- DI..!B. Desscription of Cocmpil JTF17 Englns/Airframe
In~terface Bound* I___
C. Ia lstaliati nfiguration and Supperting Design Detaffs.-D I -38D. Soc diary Airflow Systemn Definition and Requirements ---D 1-42
,AC"Eigine Mockup Plan ----- 1i.44
/ F. Engine Installation Handbook .DI -47
II INLET SYSTEM COMPATIBILITY3 Di.A. Oblectives and Introduction il- 1 IB. Airframe/Engine Cornmon- - . -D1C. Engine Inlet ,ty Tedt Plan -D - --- - 12D. Ana Simulation --- --_ .. 11 -12
I Istortlon Prediction Development - -- D11 .37
IIIl ( EXHAUST SYSTEM COMPATIMIlTY --- D III- IA. Introduction ___D _____ DII - j1
B.1 Installation Effects _ _-- D III - I
I Prat & wV~ AircreftPWA FP 66-100
40 Volumne III
0) d0 M~ni i 4snt Trn i t
fit is recognized that airplane ~-. peration mlay exceed the
S321continut; uperiating lim its.- .. Notl to) ExeeCnrretivo action rou.It he Einitiated irnmediatelv.) I\
I 24 ~future enlgine-_airfrasiccoordination.
12 -Operating Envelope F
1 0 - ~ imit for Hot Dayz Operation 0 nly. Not
0 ~ to Ezceed Ma-ch 2.7.
1-100 0 100 200 300 400 500 600 700ENGINIE INLET TOTAL TEMPERATURE Tt 2 - *
Figure 3. Engine operating Envelope -Lockheed FD 17609i
6 NDO Permissible Rev~rer00Thnt Operating io
5000 N
p ~~~4000'__
1; J0o20.4 0.6 0.8 1.0I MACH NUMBER. M
Figure 4. JTFI-7 Reverser Thrust Operation -Boeing ED 16332
IBB 1-3
Pratt &WhftneyAircrftPRA FP 66-100V o l u m e 1 1 5
- -
r t
PNrnuea m- Rvvra- Thrus 0- ratin Re o
120
MACH NUMBER,
Figure 5. JTF17 Reverser Thrust Opera! ion -FD 16331 iLockheed BI
Time percentages established were applied to the engine life objectivento obtain the requirementb given below. Points indicated for various limit-ing conditions refer to figures 2 and 3.
For major cases, 50,000-hour life;
1. 30.000 mission cyc1aa,
2. 50,000 engine acceleration cycles - idle to SLTO
3. 21,000 hours at cruise (Point A)
4. 1000 hours at SLTO (M - 0-0.3)I5. 1500 hours at maximum compressor inlet temperature
and pressure (Point Q)
6. 750 hours at envelope limit (Line GCAII)
7. 50 hours at maximum transient point (Line DEF)
8. 4250 hours at hot day cruise (Point B)9. 1000 hours at maximum reverse thrust
Io dsi_6 20,UOJ-nour disk ifie:
1. 12,000 mission cycles
2. 20,000 engine acceleration cycles - idle to SLTO
3. 8500 hours at cruise (Point A)f
4. 400 hours at SLTO (M4 w 0-0.3)
5. 600 hours at maximum compressor inlet-temperatureand pressure (Point C)
B1-4
CONFIER AL - rta hn"ArrfPWA FP 66-100
Volume IIIleakage connect points that permit optimum engine maintenance And foolproofassembly. The tubes are designed for easy maintenance and inexpensivefabrication and with adequate support to keep resonant frequencies wellabove engine operating frequencies.
Specific objectives and requirements are:
1. Each tube is designed for its specific application, i.e.,function and environment, with a usable life that is consistentwith 50,000 hours aircraft life and with 10,000 hours engineTBO without mawjer repair.
2. Stresses associated with differential expansion and contractionof various parts of the engine during steady-state and traqsientoperation must be kept within acceptable limiLs through properrouting and support arrangements.
1 3. Tuh-s are routed to ensure a nominal clearance envelope of0.500 inch between engine components and other tubes. Aclearance of 1.000 inch nominal is maintained adjacent toairframe structure.
4. Fuel and oil supply lines must function satisfactorily inambient temperatures from -65OF to 700 F and fluid tem-perature to 400'P maximum.
5. Air lines, including breather and signal lines,and fluidlines that are drained during part of the flight envelopemust withstand ambient temperatures of 1050OF maximum.
3. Design Approach
a. Detailed Description
(1) Tubing
Experience gained from the design of the high Mach number and highH performance J58 engine is being directly applied to the JTF17 engine.
Development engine time in excess of 22,.000 hours has been accumulatedon tubing of the type used on the JTF17. Tubing systems for the JTF17engine are designed using refined tube computer programs which weredeveloped in conjunction with the J58 engine project. Two malor tubecomputer programs are utilized. These are: (1) the tube stress program,and (2) the tube clearance program.
The procedure ot-tlincd ir, figure 3 demonstrates the approach usedto design engine tubing systems. Each tubing systei, is analyzed todetermine optimum tube sixe, wall thickness and tentative routes.
Pump inlet line sizes are designed with a fluid velocity ofi to 16 ft/sec to prevent cavitation or erosion. Tube sizes for theouct heater fuel sysem _ire derived from a maximum allowable fill timerequirement with resu>Lng velocities and pressure drops in the systemcompatible with a pressure requirement at the nozzles. Other systems are
BII 1-3
CoIMT IAL
4
Prst hwy AircrCOfDENTIALPWA FP 66-100Volume III
designed to permit minimum tube sizen within allowable pressure loss iirequirements and minimum expansion loop length.
3. rlmnr Mockur Routing L
4. Plumbing Cordinatea vir Layout5. Stress and Frequnc, Program H6. Interference Program Check7. Mockup Fit Check8. Aiveptable Design Layout
I!Figure 3. P&WA Tube Design Process FD 17001
BiI !!Tube wall thicknesses that have been established for the JTF17 engine
are as follows: 1! 1tj
Tube --ime Maximum Operating Basic Wall(Outside Dia, in.) Pressure, psi Th.',ckness. in. {fUp to 2.500 Under 1000 0.035
Up to 0.438 Over 1000 0.035
0.500 to 0.688 Over 1000 0.049
0.750 to 1.500 Over 1000 0.065
Tubing routes are checked on engine mockups te provide a visual 3-dimensional aid for final route selection. Removal of components, ease ofengine maintenance with respect to inspection parts, access panels andports are considered during this phase of tubing design.
Tube OD wal1 thicknesses, coordinates, end points, bracket locationsand final thermal calculations are input into a tubing computer programwhich computes stresses along the tube length, relative thermal movementsat sliding bracket locations, and resonant frequencies between supportpoints. The tube is redesigned and recomputeZ, in an iterative procesh,until the design requirements of stressen and frequencies are met. Themaximum combined hoop, tension and bending design stress is limited to22,000 psi, providing ample stress margin for tolerances and fatigue stressesfor the material used. Tubing resonant frequencies are kept above 180 cyclesper second, which is 20% above high rotor frequency.
B, ,-4
CONFIDENIAL Pratt &Whitney AircraftPWA FP 66-100
Volume IIIleakage connect points that permit optimum engine maintenance and foolproofassembly. The tubes are designed for easy maintenance and inexpensivefabrication and with adequate support to keep resonant frequencies wellabove engine operating frequencies.
Specific objectives and requirements are:
1. Each tube is designed for its specific application, i.e.,
function and environment, with a usable life that is consistentwith 50,000 hours aircraft life and with 10,000 hours engineTBO without major repair.
2. Stresses associated with differential expansion and contractionof various parts of the engine during steady-state and transientoperation must be kept within acceptable limits through properrouting and support arrangements.
3. Tubes are routed to ensure a nominal clearance envelope of0.500 inch between engine components and other tubes. Aclearance of 1.000 inch nominal is maintained adjacent toairframe structure.
4. Fuel and oil supply lines must function satisfactorily inambient temperatures from -650F to 700°F and fluid tem-perature to 400OF maximum.
5. Air lines, including breather and signal lines,and fluidlines that are drained during part of the flight envelopemust withstand ambient temperatures of 1050OF maximum.
3. Design Approach
a. Detailed Description
(1) Tubing
E::i:ernce gained f-om the design of the high Mach number and highperformance J58 engine is being directly applied to the JTF17 engine.Development engine time in excess of 22,000 hours has been accumulatedon tubing of the type used on the JTFI7. Tubing systems for the JTF17engine are designed using refined tube computer programs which weredeveloped in conjunction with the J58 engine project. Two major tubecomputer programs are utilized. These are: (1) the tube stress program,and (2) the tube clearance program.
The procedure outlined in figure 3 demonstrates the approach usedt casign engine tubing systems. Each tubing systei is analyzed todet6,rmino optimum tube size, wall thickness and tentative routes.
lu jp inlet line sizes are designed with a fluid velocity of1 f ) if) ft/soe to prevent cavitation or erosion. Tube sizes for the
Ler h i- system are derived from a maximum allowable fill timewith resulting velocities and pressure drops in the system
&i ,it a pressure requirement at the nozzles. O-ther systems are
I11 1-3
N 01 F1 E NTIAL
CONHiDEHTIAL&rf W~hftney Aircraft M H =
PWA F? 66-100
Volume IIIdesigned to permit minimum tube sizes within allowable pressure lossrequirements and minimum expansion loop length.
2. Preliminary Routing on Developed Engine View 1
3. Preliminary Mockup Routing4. Plumbing Coordinates via Layout5. Stress and Frequency Program6. Interference Program Check7. Mockup Fit Check8. Acceptable Design Layout
[1Figure 3. P&WA Tube Design Process FD 17001
BII
Tube wall thicknesses that have been established for the JTF17 engineare as follows:
Tube Size Maximum Operating Basic Wall
(Outside Dia, in.) Pressure, psi Thickness, in.
Up to 2.500 Under 1000 0.035
Up to 0.438 over 1000 0.035
0.500 to 0.688 Over 1000 0.049
0.750 to 1.500 Over 1000 0.065
Tubing routes are checked on engine mockups to provide a visual 3-dimensicnai aid for final route selection. Removal of components, ease ofengine maintenance with respect to inspection parts, access panels and
ports are considered during this phase of tubing design.
Tube OD wall thicknesses, coordinates, end points, bracket locationsand final thermal calculations are input into a tubing computer programwhich computes stresses along the tube length, relative thermal movementsat :Ii'1ing bracket locations, and resonant frequencies between supportpoinurm. The tube is redesigned and recomputed, in an iterative process,until the design requirements of stresses and frequencies are met. Themaxitm combined hoop, tension and bending design stress is limited to22,000 psi, providing ample stress margin for toleran-es and fatigue stressesfor Vo naterial used. Tubing resonant frequencies are kept above 180 cycles
Sco , which is 207 above high rotor frequency.
BI 1-4
CONFIDENTIAL
CONFIDNTIALPrat &Whitny AircraftPWA FP 66-100
1.0 < Volume IIICOJ H
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II IA-7
CONIDETIA
PIQI FP 66-100Volume III LOCkHEED COMPETITNE DATA
Not.. Al Load. Are umiIndlividual Point.i:Y ., Velocit- rad'.CeY. Accolration. -l*cc/
pitkh VO-ity 05
fPh Arlerotlo,* R.i v.'ocl"ty az
R ll AcceleratioT A* 0. SPm~ Louittdinal Drag or Thrust 2Tr Ms Reverse Thrust 11b Fore A- t .0 (T Man Frmard Thrus.t (b) i
Eine Torq.. is Abo . . \ idw rFi.,d=.i-OplLongitudinal L Side Si 0 I rlad I
It I L S "Id !
sufbegnic 1Airigad a0. 5 PmI
LFVert,%,I
Emerecy Manewvrs E... mru.1t z "°J{
Foryaudp Alt 01'r'r. Aft G I
Sido G = J Engine Failure (M 3)2 R. 102,.000 lb (Alt)
4' 0 0. 10
Figure 8. JTF17 Maneuver Load Diagram- FD 16285Lockheed Installation [IA
Rotor integrity is assured by application of the following rotortiebolt or axial retention criteria:
I. -ho moment i=poscd by loftn u- ;G of the blades of any &iistage shall not stress the tiebolts more than the 0.2% yieldstrength of the material.
2. The strain energy produced from loss of 10% of the biades ofany one stage must be within the strain energy absorptioncapability of the tiebolts.
3. In addition, the tiebolts must prevent flange separation during j1normal operation at the extreme conditions of maneuver andflight envelope.
Energy absorptive techniques are used to prevent gross rotor fail--Iures. Generally, this technique is to utilize energy by deformationor rubbing friction on less critical or smaller mass parts to prevent Ithe sudden release of large masses of energy. This criteria is appliedto rotor-stator axial sp' ing-
The minimum parts life requirements shown below are established onthe basis of the general engine parts life requirements described inSection I (Introduction to Report B). I
Fan disks 20,000 hours IBlade lock rings 20,000 hoursFan rotor blades 10,000 hoursRotor tiabolts 10,000 hours
RIIA-8
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Pratt 8 Whitney AircraftPWA FP 6-100
Volume III
Prior to the use of fans, the criteria employed to ensure containmentequated the potential energy absorption capacity of the case to the
V kinetic energy of the blades released. IWile this method yielded generally
accevtable results for high compressors and turbines (figures 25 and 26),tests revealed that the energy absorption was not sufficient Ior the largerdiameter fans. The blade impact was determined to produce intense localcompressive stresses that caused a portion of the case to be sheared beforethe energy could be absorbed.
A technique has been developed to provide a type of containmentanalysis that employs the shape of the failed blade and the dvnamic shearstrength of a localized section of the engine case. This method has pro-
, vided consistent results in tests using actual blades and engine cases.From many tests, a factor has evolved that relates the actual blade velocityat impact to the total velocity ef the same weight blade required to shearthe case. Experience defines this factor. (Sc. figure 27.)
In the JTFl7, shear criteria has been used to size the fan case. The
shear criteria was also used for the compressor and turbine sections aswell as the energy criteria, Table 8 qhows the actual containment factors.shear and energy, provided in. the engine.
Containment Capbility of casea is judged by a "'containment factor,"
that in ampiricslly obtained fromn test.
This containment factor (CF) relates case energy absorptionpotential (PE) to blade kinetic energy (KE)
CP= (Of all cam"e surrounding ntage)
t(E (Ali lud. Instago)Few ~~~: -c.' - i).;6•5hrow:Mt blAdD CT =012(vch A! turbine nrage)
0.400 -- 2OXComlrao~E7
(Unakrouded Blades)
CA
A Not Covitaind
~ ~ -~ -. ~ 0-Contained -
- .M " I n i u ---2I- ,-
Figure 25. ContainmenL Experience FD 16232
BI.-
BIIA-27. 'j
Pratt &Whftney Aircraft4PWA FP 66-100Volume III H
0.22 I i a" t __|_ !___ -
°-_ I ,, _
0 T !n rINI1with Mad..
ro.0e -- 1
ii ~ NotContained
E-0.14 1 1 Dw¢u f[
S. - ---- -- ---------N -R oanmuId Inimum
- t.l tfJ
.00
0.013
t I 0O
z0
ilEl 013"
Figure 27. Contain~ent Experience ase nFD 16234hIA
Te : N.-. --Cfm rw4urqd but HIM&. 1.0 n
$W Bonm*
-A -28 0.6
_____ ____ _____ ____ 0.4
j ~' ECASE2; MATPERIA
K K ~ K K IFAIUREM
Figure 27. Contain~eiL. Experience Based on FD 16233Shear Penetration Testing 11A
BIIA-28
U Preftt8 Whitney PAutnrftPWA FP 66-100
Volumc IlI
Table 8. JTF17 Blade Containment
Stage Shear Criteria Energy Criteria(See figure 27.) (See figure 25.)
Factor Allowable Factor Allowable
Min Min
First (Fan) 0.68 0.68 11.3 8.6K Second (Fan) 0.68 0.68 12.6 8.6
Third (Compressor) 0.914 0.68 0.25 0.25Fourth (Compressor) 1.204 0.68 0.263 0.25Fifth (Compressor) 0.957 0.68 0.322 0.25
V Sixth (Compressor) 0.934 0.68 0.264 0.25Seventh (Compressor) 1,315 0.68 0.379 0.25Eighth (Compressor) 1.333 0.68 0.475 0.25First (Turbine) 0.890 0.68 0.230 0.250Second (Turbine) 0.755 0.68 0.120 0.120Third (Turbine) 0.702 0.68 0.120 0.120
L A containment failure involves an impulsive load of infinitely shortduration that results .n an extremely high strain rate (approximately100,000,000 times greater than standard tensile test strain rates).High strain research indicates that the true energy absorption and shearcapability of a material varies with strain rate, generally increasingsubstantially with increased strain rate.
Material dynamic shear factors have been obtained by Pratt & hlitneyAircraft through ballistic testing of materials at velocities related toblade impact and subsequently verified by whirl it testing with actualF, blades and cases. Figure 28 shows the relationship of dynamic shearfactor to st-ati" tM,611e for a variety ?w ' tral'ald.
AMS &54
[j00 Valuee Obtained from Ballistic Teto
( AMS 5624 Shear Ca bility = D, X Tensile Sength
AMS 5536B IS3 AJ48 55&9 Rc2O - S552-
0 AIMS 4910 (Ti.I PWAim (r-i
2 _____~ Ag KIL,(Ti)
[ ' 1~PWA i030 (Wasp)I AMlS 6604Rc42
60 80 100 120 140 I(40 184) 200STATIC TENSILE - psi (Thousand)
Figure 28. Dynamic Shear Factor vs Static FD 1.6279Tensile Strength 11A
BILA-29
Prett & Whfnmy AWrcrftPWA FP 66-100Volum III l(6) Intermediate Came II
The intermediate case, which is the section between the fan and thecompressor. is the support for the No. I and No. 2 bearings and also the fimajor support for the gao generator.
The intermediate case is designed to carry the radial and thrust loadsof the No. 1 and No. 2 bearings, thrust load, shear torsional load of the L
ges generator, and the main engine mareuver loads. In addition to theseloads, the intermediate case must withstand the unbalance forces causedby blade loss and supplement the 2nd-stage fan OD liner in providing bladecontainment.
The basic structure of the intermediate case is a titanium weldment Iutilizing AMS 4966 (A-110) forgings and AMS 4910 (A-110) sheet, Buttwelded joints will be used on all attaLhments by welding to integral,machined, contoured 6tandups in the wall forgings. (See figure 29.)This construction has the decided advantage of having all welds loaded insimple tension or compression. Areas with combined bending stresses willoccur in parent material away from the welds and heat-affected zone,
i s
Du" ag a joint-- iI
w, u -___ n nIt I
Butt Wakh,-
Butt Weldod
Figure 29. Intermediate Case Basic Structure FD 17666
The struts are fabricated from a combination of sheet stock andforgings and are butt welded to airfoil contoured platforms that are anintegral part of the wall rings 11
Vane to shrouzd cracking experienced in service on the JT3D inlet casehas led to the standup foot butt weld design. Cases such ae the JT3D I
BIIA-30 II
COFDEN L Pr & W W AirraftPWA FP 66-100
Volume III
(g) LCF Design Approach
Initial studies considered using engine 3rd stage (compressor lotstage) air a. a source for disk cooling. The temperature level was idealfor rapid disk metal temperature response but the pressure level was too
low for the required flow. Subsequent studies led to the selection of4th stage air. This source, in conjunction with antivortex tubes tocompensate for radial inward flow pressure drop, resulted in acceptablebore air temperature and pressure levels.
1:The next step was to design a flow system in which the bore coolingair would Lontact all disks and yet be separated from higher temperaturecompressor discharge air. This was accomplished by Incorporporating abore tube that returns the cooling air to the front of the high compressorwhere it is discharged through the intermediate case to the fan cavity.
The final cooling air flow path, which resulted in a disk thermalenvironment that satisfied LCF requirements, is shown in figure 63.
4...--- Antivortex Tuo_ Tbe
., jr:.\ o
Figure 63. Compressor Disk Cooling System FD 16417
To evaluate disk transient thermal response, a detailed analysis wasperformed on the most severe conditions of a typical aircraft missioncycle. This analysis was performed by creating a mathematical model ofthe entire rotor and programming all pertinent parameters on a digitalcomputer. The compter progr-m, in *upeac, simulaced the total in-
flight rotor environment.
To assure that optimum weight had been realized, the disks were firstanalyzed on a stress-limited basis. This means that the disk configura-tion was determined first from the standpoint of a noncyclic stresscriterion such as yield, burst, and creep. This disk configuration was
then adjusted to account for the specific number of dynamic and thermalcycles.
BIIA-75
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Prua b & W t e irrtPhi FP 66-100Vol -e III
Several supporting studies were conducted in conjunction with the main reffort to obtain the "-&sic cooling air flow path. One of these studiesinvolved optimization of the bore cooling air flow rate. Results of thisstudy are shown in figure 64. The curve shows that increasing the flowbeyond 0.6% of gas generator flow offered little advantage in decreas-ing disk temperature gradient.
400__ ---- __ - -i
300---------~.-I
~200
100 - -__ - - -t__ -- I
0.2 04 5 1 1.2 1.4 1.6 iGAS GENERATOR FLOW - %
Figure 64. Disk Temperature vs Gas Flow FD 17416IIA
The substantiation of the need for a design using cold-bore cooingair is illustrated in figure 65. This curve showed the relative weight F;required to obtain cyclic life with a hot-bore and cold-bore configuration.The hot bore configuration is that used in the initial experimental engine.
32U -1- 64
~24 [ .48*
wei-ht. I.-Cre-n tu 0,.i A
C, 12,000 Cycle. I ~I
~16 -1 32C.
LCF~iao Stress8 - Limited i.-,.- 16
-400 -300 -200 -100 0 100 200 300 II
COLD BORE AT t OF HOT BORE AT- OF
Figure 65. Cold Bore vs Hot Bore LCF Life FD 16359and Weight Trends I IA
BIIA-76
CONMA AL(This Page is Unclassified)
U
Pratt & Whitney AircraftPWA T 66-100
Volume IIIFuel is metered to the duct heater as a function of power lever position
and Tt2 as shown on figure 3. Power lever translates a 3D cam and Tt2rotates the cam. the output of which is the desired duct heater fuel flowburner pressure ratio. This ratio is multiplied by burner pressure result-ing in a signal proportional to fuel flow being generated.
The duct heater incorporates two zones of fuel injection. Within theunitized control, each zone is provided with a fuel shutoff valve and amanifold rapid fill system. This latter system reduces by a significantamount the time required for augmenter transients by providing a high rateof fuel flow from the gas generator boost pump dur q the fill period. Eachzone is also provided with separate fuel pressu- ;nals for operating thefuel manifold dump valves.
When the power lever is advanced beyond maximum nonaugmentation flatto the minimum duct augmentation flat, a seg acing valve -: the unitizedfuel control initiates the following events: manifold dumpvalve close&, (2) the Zone I rapid-fill valve opens, e a Zone I shut-oft valve opens, (4) the duct exhaust nozzle resets parL_ y open, and(5) the duct igniters are energized. Fuel is delivered to he Zone I fuelmanifold at a high flow rate until a pressure signal indi, es the manifoldis full. The rapid-fill valve closes, the igniters are tt ' off, andthe duct exhaust nozzle reset is removed.
Further power lever advancement increases duct fuel-air ratio and ductnozzle area on a coordinated schedile to hold the total engine airflowconstant.
If the power lever is moved to the Zone Ii range, the Zone II fuelmanifold dum[ valve is closed, the Zone II shutoff valve is opened, andthe Zone II rapid-fill vaive is opened to fill the Zone II fuel manifold.A constant fuel-air ratio t hild d .r.lr......... ^, tt.r Zu,,e Y,!" rapid-ft!! transient.
Pressure increasing in the Zone II manifold provides a signal. re3ulting inclosing of the rapid-fill valve and simultaneous routing of meLered fuelto the Zone 1I manifold. Total duct fuel flow is divided between Zone Iand Zone II by the fuel nozzle flow characteristics. Zone II fuel ignitesspontaneously when the fuel enters the burner. Continued power lever ad-vancement causes increased duct heater fuel flow, increased engine thrust,and increased duct nozzle area to maintain constant engine airflow.Maximum duct augmentation is scheduled by power lever position. Fuel flowfor quick filling of both the Zone I and Zone II fuel manifolds is suppliedfrom interstage pressure of the gas generator fuel pump. The duct heaterfuel flow schedule is shown in figure 8.
The total corrected engine airflow is controlled as a function of Tt2to the schedule defined ft the englue specification. The airflow control[is achieved by actuating the variable duct exhaust nozzle. In the cruiserange the nominal airflow schedule may be manually adjusted by the flightcrew between maximum and minimum limits to obtain optimum inlet per-formance. The total corrected airflow schedule, the maximum and minimumlimits, and the nominal schedule coordinated with Boeing are shown infigure 11 of paragraph B of this section, while those coordinated withLockheed are shown ii figure 10 of paragraph B of this section.
rBIII.D-7
__ _ __._ __
PIsUAWVhtn~y Airrft IPlA FP 66-100Volu. III
TIMi
Zone I Plus~~Zono I_ Zone It
POWER LEVER ANGLE
Figure 8. Duct eater Fuel Flow Schedule FD 16453BIll
Total engine airflow is the sum of gas generator airflow and duct air-flow. Gas generator airflow is determined by sensing high rotor speed and.engiue inlet temperature. Knowing this airflow permits determining the .duct airflow required to obtain the desired total. engine airflow. There-fore, desired duct airflow parameter will be scheduled as a function ofhigh rotor speed and engine inlet temperature, as shown on figure 4.
The duct correcten airtiow is measured using the duct reasure ratio LBparameter, which is the difference betweez fan discharge t ttl pressureand fan discharge static pressure divided by fan discharge total pressure,(Pt3-Ps3)/Pt3' This same parameter is utilized in superaonic aircraft air|Iinduction controls. The unitized control wi1 schedule the duct pressureratio necessary to obtain t.e dewired duct airflow. The actual ductpressure ratio will be determined by the control and compared with theIischeduled pressure ratio. The difference between the pressure ratiosinitiates corrective action through a proportional plus integral serv anda power boost servo to reposition thi duct exhaust nozzle as required in aclosed loop basis to obtain the desired duct airflow. .
c. Separately Mounted Components.F
The Litized fuel and area control system includes the followingseparately mounted components:
1. Turbopump controller, which regulates air supplied to the ductheater fuel pump by modulating a butterfly valve located in thepump air inlet supply. fr
2. Two engine inlet temperature sensors which sense temperature with K- i -
a gas-filled tube. The resultant g" pressure is transduced intoa fluid pressure and in turn sensed oy the control for use as agine -1inlet temperature bias.
BXIID-8
Prett & Whitney AircraftPWA P' 66-100
Volume IIIto close and opens the respective fuel manifold dump valves. Coolingfuel flow will also be reestablished through the duct heater tuel svs-tern components. In this e, ertt, the power lever must be retarded to thenonaugmented flat power lever position, or less, and then moved to anaugmentation position before duct heater operation can be reinitiated.
[ In the event an inflight engine shutdown is performed, the enginemay be restarted by placing ignition selector switches on the "on"
position, the power lever at idle position and moving the shutoff leverto the "on" position. The gas generator ignition selector switches shouldbe placed in the "off" position after engine restart is atccomplished.Normal engine control by power lever modlation may then 'ie rr,umed.
Table 1. Sample Cruise Thrust Setting Tables
Subsonic: 36,150 Ft. M = 0.9
Gross Wt. OAT - OF -70 -52EPR/EGT 2.26/1370 2.26/1450
480,000 N21NI 94.7/97.2 96.5/100.0I WFT 10600 10900AJD/DH 3.6/Not Lit 3.6/Not LitRAT/TAS -7/516 +14/528
EPR/EGT 1.91/Si05 1.91/11.80360,000 N2/Nl 89.6/92.7 91.5/94.7
WFT 7800 8000
AJD/Dh 4.3/Not Lit 4.3/Not LitRAT/ TAS - %5'P'_ +14/528
Supersonic: 65,000 Ft. H - 2.7
Gross Wt. or -34
EPR/EGT 0.743/1425 0.701/1425500,000 N2/NI 100.0/86.2 100.4/86.5
WFT 26000 27200AJD/DH 6.4/Lit 7.0/Lit
HRAT/TAS 494/1548 537/1534
EPR/EGT 0.743/1425 0.701/1425370,000 N2/Nl 100.0/86.2 100.4/86.5
WFT 191.00 20200AJD/DH 5.6/Lit 6.2/LitRAT/TAS 494/1548 537/1584
Symbol Description UnitsOAT Outside Air Temperature OF
EPR Engine Pressure Ratio (Pt7/Pt2) NoneEGT Turbine Discharge Total Temperature OF
N21N I High Rotor RPM/Low Rotor RPM 7.WFT Total Fuel Flow PPHAJD Duct Nozzle Jet Area Ft2
DH Duct Heater Lit/Not LitRAT Rai.A Air Temperature (TC2) OF
TAS True Airspeed Knots
BIIIB-19
'9-- - - - - -: :-- - - - -- _ : - -# . --
T - 7 - - - --- _: : -- ---- -
Pratt &Whitney AircraftSpecification No. 2710
26.0 TURBINE C®LING SYSTEM - The turbine cooling system is a
self-contained, fixed orifice and self-regulating system which does
not require an airframe supplied input.Ii
26.1 Turbine Temperature Measurement System - The engine shall
be equipped with thermocouples for use in conjunction with the airframe
temperature indicating system. The thermocouples shall permit consistent
measurement of exhaust gas temper ure. The system design shall be
such that it is possible to service check individual temperature probes
for continuity.
27.0 CUSTOMER REOUIRDINTS
27.1 Drive Power Extraction - The maximum allowable continuous
horsepower extraction and overload horsepower extraction at the power
takeoff pad for all operating conditions as a function of high pressure
compressor rotor speed is as specified in the form of torques and speed
ratios on the Installation Drawing.
27.1.1 D~ainic Loai ing - The dynamic loading limt of the drive
pad is specified on the Installatiou Drawing.
27.1.2 Power Takeoff Shaft Speed - Lower takeoff shftr speed ratio
is specified on Installation Drawing.
27.1.3 Shear Section - The shear section requirements are specified
on the Installation Drawing.
27.1.4 MiuntinR Pad and Power Takeoff PTP) Shaft Loads - The
mounting pad and P0 shaft load specifications are shown on the Installation
Drawing.
27.1.5 Hvdraulic Pump Drive Pads - Hydraulic pump pads shall be
provided as shown on the Installation Drawing.
27.2 Compressor Bleed Air
27.2.1 Quality - The air at the engine bleed ports shall not contain
quantities of engine generated noxious, toxic or irritating substances
above the maximum threshold limit values of the substances shown below:
Substancas Parts per Million (Volume)
Carbon dioxide 5000.0
Carbon monoxide 50.0
Carbon tetrachloride 50.0Decaborane 0.0UDiboramne 0.1
Pentaborane 0.01
Ethyl alcohol (ethanol) 1000.0
225
I- - - . .--- --- -- '-.' .----
Pratt & Vil ,itney AircraftSpecification No. 2710
Substances (continued) Parts per Million (Volume) |I
Fluori 0.1Fuels, aviation 250.0Hydrogen peroxide 1.0Methyl alcohol (methanol) 200.0Methyl bromide 20.0Monoch Iorobromomethane 40.0Nitrogen dioxide 5.0Oil breakdown products (aldehydes, acrolein, etc.) 1.0Ozone 0.1Unsym-dimethyl hydrazine 0.5
The air shall contain a total of not more than 5 milligrams per cubic jmeter of submicron particles.
Dirt or other foreign particle concentration in the bleed air afterexpansion to atmospheric pressure shall not vxceed that of the air atthe engine inlet on a per unit volume basis. If a demonstration isrequired P&WA will demonstrate on a normally functioning engine on aPFIWA plant test stand that the above requirinents are met within theaccuracy of the testing technique available to P&WA at the time of thedemonstration. However, it must be recognized that there may be occasionalinstances in service operation when the bleed air is contaminated.
27.2.1.1 Seals and Oil Lines - Accessory seals, bearing seals,and oil lines shall be designed so that a single failure (except forengine bearing failure) can not result in bleed air contamination. P&WA !shall submit a failure analysis to the airframe manufacturer to demonstratehow the design meets this requirement. Ii
27.2.2 Quantia - The engine shall provide for high pressure compressorair extraction, for aircraft use, as indicate2d:
a. Within the operating envelcpe of the engine, high I!compressor air will be available in quantities notto exceed 5% of gas generator airflow from idle toa .iust corres-ponding to a turin-Le exhaust gastemperature 800F (44.40C) less than that correspondingto maximum cruise.
b. Within the operating envelope of the engine, high compressorair will be available in quantities not to exceed 3% ofgas generator airflow from a thrust corresponding to aturbine exhaust gas temperature 80"F (44.4C) less thanthat corresponding to maximum cruise, to takeoff thrust. ,
The ntuaber, location and connection flange details of the bleed portsare defined on the Installation Drawing. Changes of compressor airextraction must not exceed 1% per second. For h1gh pressure compressoraig extraction, within the limits specified, the pressure and temperatureat the engine bleed ports shall be as provided by curve No. S-84. sheet1. The effect; upon eiuglne performance when bleeding air from the engineshall be as provided by curve No. S-84, sheet 1.
26
U
Pratt &Whitney AircraftSpecification No, 2710
1: 26.0 TURBINE COOLING SYSTE1- The turbine coolu,,p svstem is a
self-contained, fixed orifice eid self-regulating system which does
Snot require an airframe supplied input,
26.1 Turbine Temv rature Measurement kytm - The engine shall
be equipped with thermocouples for use in conjunction with the airframetemperature indicating system. The thermocouples shall permit consistentmeasurement of exhaust gas temperature. The system design shall besuch that it i possible to service check individual temperature probes
for continuity.
27.0 CUSTOMER REQUIREMENTS
1 27.1 Drive Power Extraction - The maximum allowable continuoushorsepower extraction and overload horsepower extraction at the powertakeoff pad for all operating conditions as a function of high pressurecompressor rotor speed is as specified in the form of torques and speedratios on the Installation Drawing.
27.1.1 Dvnamic Loading - The dynamic loading limit of the driveIpad is specified on the Installation Drawing.27.1.2 Power Takeoff Shaft Speed - Power takeoff shaft speed ratio
is specified on the Installation Drawing.
27.1.3 Shear Section - The shear section requirements are specifiedon the Installation Drawing.
27.1.4 Mounting Pad and Power Takeoff (PTO) Shaft Loads - Thej, mounting pad and PTO shaft load specifications are shown on the InstallationDrawing.
27.1.5 Hydraulic Pump Drive Pads - Hydraulic pump pads shall beprovided as shown on the Installation Drawing.
27.2 Compressor Bleed Air
27.2.1 Q ti - The air at the engine bleed ports shall not containquantities of engine generated noxious, toxic or irritating substances
iabove the maximum threshold limit values of the substances shown below:
Substances Parts per Million (Volume)
Carbon dioxide 5000.0Carbon monoxide 50.0Carbon tetrachloride 50.0Decaborane 0.05Diborane 0.1Pentaborane 0.01Ethyl alcohol (ethaaol) 1000.0
I25
[I|
I
Pratt &Whkrt y AircraftSpecification No. 2710
Substances (continued) Parts per Million (Volume)
Fluorine 0.1Fuels, aviation 250.0Hydrogen peroxide 1.0Methyl alcohol (methanol) 200.0Methyl bromide 20.0Monochlorobromomethane 40.0Nitrogen dioxide 5.0Oil breakdown products (aldehydes, acrolein, etc.) 1.0Ozone 0.1 IUnsym-dimethyl hydrazlne 0.5
The air shall contain a total of not more than 5 milligrams per cubic Imeter of submicron particles.
Dirt or other foreign particle concentration in the bleed air afterexpansion to atmospheric pressure shill not exceed that of the air at iithe engine inlet on a per unit volume basis. If a demonstration isrequired P&WA will demonstrate on a normally functioning engine on a IiP&WA plant test stand that the above requirements are met within the5!accuracy of the testing technique available to P&WA at the time of thedemonstration. However, it must be recognized that there may be occasionalinstances in service operationi when the bleed air is contaminated.
27.2.1.1 Sealz and Oil Lines -- Accessory seals, bearing seals,and oil lines shall be designed so that a single failure (except forengine bearing failure) can not result in bleed air contamination. P&WA Ishall submit a failure analysis to the airframe manufacturer to demonstratehow the design meets this requirement.
27.2,2 _Qantiy - The engine shall provide for high pressure compressorair extraction, for aircraft use, as indicated:
a. Within the operating envelope of the engine, high icomoressor air will bi AVAXahlP in nif--tt-_pq not
- - - - - -
to exceed 5% of gas generator airflow from idle to fa thrust corresponding to a turbine exhaust gas[Itemperature 807F (44.4C) less than that correspondingto maximum cruise.
b. Within the operating envelope of the engine, high compressor Hair will be available in quantities not to exceed 3% ofgas generator airflow from a thrust corresponding to aturbine exhaust gas temperature 80°F (44.40 C) less thanthat corresponding to maxiam cruise, to takeoff thrust.
The number, locarion md connection flange details of the bleed portsare defined on the Installation Drawing. Changes of compressor airextraction must not exceed 1% per second. For high pressure compressorair extraction, within the limits specified. the pressure and temperatureat the engine bleed ports shall be as provided by curve No. S-84, sheet j1. The effects upon engine performance when bleeding air from the engine I
shall be as provided by curve No. S-84, sheet 1.
26 [I
CONETAL Praft & Wftney AmsfVOIUUI
JIIFigr.! 4. Florida Research and Development FE 13329
Center Mlanufacturing and Office
Bu ilding
A I IMFigure 5. Eat Hartford Plant Z 24989
5
CONENTIAL
I
NPF 66-100 1Volume 1
The estimated cost of the JTFI7 development program proposed to meet ra
the SST objectives for engine life and reliability is $325 million fromthe end of Phase Il-C through engine certification in December 1971; acontinued development program is planned after certification to improvethe engine in service. The estimated costs for the various phases ofthe SST program are summarized below:
Assumed Time Frame Program Phase Estimated Cost
July 1965 through Phase II-C (demonstrator) $ 50 MillionDecember 1966 IiJanuary 1967 through Phase II (FTS development, 290 MillionSeptember 15, 1970 20 prototype engines, and p
support for 100 hours ofaircraft flying)
September 15, 1969 Phase IV (Engine certifi- 252 Millionthrough May 15, 1974 cation, plus continued Ii
development and flighttest support through air-craft certification) I
May 15, 1974 through Phase V (Continued engine 180 MillionMay 1979 development)
The estimated unit engine cost established in accordance with theFAA's instructions is $1.21 million, for the engines to be delivered duringPhase V.
The Mach 3+ flight time obtained with J58-powered aircraft each daynow exceeds the total Mah 3.0 flight time obtained by all other aircraftto date. Av this supersonic experience continues to accumulate in theyears before the flight of the first supersonic transport, the FloridaResearch and Development Center engineering team will continue to apply to ; -the JTF17 the hard lessons learned from Lhc Jbh program. In oirnerciat
servic' , othhr Pratt 6 Whitney Aircraft engines have demonstrated a main-tenance cost per Lb thrust per hour one-half that of competitive engines,and unmatched rate of TBO growth, and the lowest premature-removal rate in Ithe industry. The same development philosophy that made this record possiblewill be applied to the SST propulsion task. This extensive and continuinghigh Mach number experience, combined with extensive and continuing com- Lmercial turbine engine experiete, provides a singular understanding of theSST propulsion problems; an understanding which in the final analysis willresult in the most economical engine.
6
I
LiRDM A Prat awrietney RlrcrmftPWA FP 66-100
Volume I
5. Engine le.ght
The weights of the JTFl7A-21 production engines are d fin'd in theEngine Model Spjcifications as 9910 pounds and 9860 pounds for the Buingand Lockheed models, respectively. Prototype engine weights are T' higher.Pratt & Whitney Aircraft has developed a system of weight prediction andcontrol that was applied to the JTF17 engine design during Phase 11. Eachcomponent part was carefully estimated and controlled. Thle accuracy of.
this estimating system is confirmed by the fact that tlh actual weightof the first test engine proved to be 50 pounds less than the weightestimated for the engine. The extensive weight control program developedduring Phase 1I will be continued through Phase III.
6. Noise Attenuation
Phase II-C tests and analyses indicate that the followit.g FAA noiseattenuation objectives can be met:
1. 1500 feet from centerline of runway 116 PNdb2. 3 statute miles from start of takeoff roll 105 PNdb3. 1 statute mile from runway on approach 109 PNdb
Current levels of unsuppressed and suppressed engine noise for Con-dition I are shown in figure 15. The potential for suppression devicesto reduce the level of the predominant jet noise is also indicated.
Figure 16 shows noise levels for Condition 2 at the 3 mile point forthe thrust levels the airframe manufacturers anticipate. The suppressiondevices, current and potential, include acoustic treatment of the ductheater diffuser to absorb fan generated noise, and reverser-suppressorattenuation to reduce the jet noise. A fitrther reduction in total eL.ginenoise at thrust cutback after takeoff may be obtained through the use ofduct heating beyond the thrust cutback point. By means of this procedure,the required thrust level will be attained at a lower fan rotor speed.Engine noise reductions of 5 PNdb may be obtained using this technique.
.amTwo EginesCurrent Commercia Aircraft:1. To Enines98- 105 PNdL
2. Sideline Noise at 15430 ,t 1, 140- Extra Ground Attenuation- FAA S Objective; 116 PNdb
130 Max AugmentedUnsuppressed Thrust
o Current Suppression
0 40 50 60I INSTALLED THRUST - lb (Thousand)
Figure 15. Predicted Turbofan Airport Noise ED 16948
Levels 223__ _
hi. 1.1...0-i -2 :• . .. ..I -.
i
Pratt & Whitney Aircraft O DENTIAL1A FP 66-100
Volume I VFigure 17 shows predicted noise level for Conidition 3, the I mile
approach condition, where acoustic damping and opLimtin blade-vant, spacingreduce fan noise.
130--------- Current Comimercial Atnrrft. 125-1V7 PNdb
FAA SBTObjoctiw IMWPNdb
z 1 ,p__io Z r U Ilisuppresped
Z90 " Potential Suppresin
s o_10 12 14 16 18 20 22 24 26
NET THRUST - lb (Thousand) '1Figure 16. Predicted Turbofan Noiee Levels FD 16949
at Thrust Cutback After Takeoff i
1
9 10 n1 12 13 14 15 16 17 18 19 j INET TMUST - lb (Thousand) [
Figure 17. Predicted Turbofan AWise Levels D 16950 [f
at Approach Conditions
24
110m
Pre & V tney AlrcraftPWA F7 66-100
Volume III
SECTION IINSTALLATION COMPATIBILITY
A. SCOPE AND OBJLCTIVES
This section describes the design approaches and design requirementsIi used to define all interface boundaries between the Boeing B-2707 airplaneand the Lockheed L-2000 airplane and the JTFI7 engine. By virtue of thePhase II-C coordination effort, the section further delineites the highdegree of airplane/engine compatibility achieved in the areas of enginemounts, accessory drives, external plumbing, air bleed systems, secondaryairflow systems, instrumentation, and engine mockups.
Continuing coordination with the airframe manufacturer, airlines, andthe FAA during Phases III, IV, and V will ensure that engine compatibilityis maintained as the engine is developed and the airplane design Is refinedduring the prototype and flight test prcgrams. The plan and procedures torachieving this continued compatibility are described in Volume V, Report C.Configuration Management.
A set of the engine installation drawings, which are actually part ofthe engine model specifications, has been incorporated in this sectionfor ready reference. As such, the drawings represent the engine coordinatedconfiguration as approved by each of the airframe manufacturers.
B. DESCRIPTION OF COMPLETE JTF17 ENGINE/AIRFRAME INTERFACE BOUNDARY
All engine/airframe interface boundaries, including dimensions andload limitations for the interface, are shown on the Installation Draw-ing, figure I for the Boeing B-2707 and figure 2 for the Lockheed L-2000.
jThe symbols (X, AD, AE, etc.) which correspond to locating symbolsused on the Installation Drawing for cross-reference are listed alphabeti-cally in the nomenclature columns of Sheet I of the Installation Drawingand after each paragraph heading in this section and provide the location ofeach item by coordinate zones.
The protoype euine- iR ctrrenri[Iy identiral. to the prodiction engineLat the engine/airframe interface boundaries. During Phase III, required
changes to the prototype engine will be coordinated with the airframemanufacturer and the FAA as shown in the configuration management plan.The interface& are deecribed below.
1. Mountinf and Installation Attachment Systems
LI a. Engine Inlet (X)
A circular flange is located 5.30 inches ahead of the engine frontmount ring to mate with the airframe inlet duct. Connection and matingto the a4rframe inlet for Lockheed is by a V-band clamp to the engineflange. The engine inlet case also provides for a Lockheed-supplied inlet
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Volume IIIduct strut extension or flow splitter to extend into and be attached to
the engine inlet case. The provisions were requested to be compatiblewith the Lockheed airplane inlet system and consistent with Lockheed'sexperience. Engine development testing is required to confirm the meritsof such a configuration and that such a system will net adversely affectcompressor blade stress levels.
The connection to the Boeing inlet duct is accomplished with four pairof 2 bolt attaching points. While this design provides the maintenancefeature specifically requested by Boeing, it will require a stiffer inletcase front flange to minimize leakage problems.
Matching the airframe inlet centerbody with the engine inlet casefront bearing hub is required to provide a smooth airflow path to theengine compressor, although no mechanical connection is required.
b. Front Mount (AD)
The front mount ring incorporates two mounting points located in theupper left and right quadrants of the engice. The systems are describedin greater detail in paragraph C of this section. The design loads andmethod of loading are shown in the Loads Drawing, figure 3 for the BoeingB-2707 and figure 4 for the Lockheed L-2000. The engine mount structure,the materials selected, and the design practices used have been derivedfrom years of experience with many successful commercial engine programs and
Pratt & Whitney Aircraft's knowledge of high Mach number environment obtainedthrough the J58 program.
The rear mount ring incorporates two mounting lugs located in theupper right and left quadrants of the engine. The design loads and methodof loading are described in the Loads Drawing. The details of the systemsare described further in paragraph C of this section.
d. Reverser-Suppressor Shroud
The forward face on the nacelle mating flange of the reverser-suppres-sor consists of a belt flange for attachment of the airframe nacelle cowlseal. It is located (behind the rear mount ring) approximately 18 inchesfor Boeing and 10 inches for Lockheed. The specific contiguration of theflange was coordinated with the airframe manufacturers. The mating dimen-sional requiremeime and load limitations for the flange are given on theInstallation Drawing. The specific requirements and description of thereverser-suppressor as requested by the airframe manufacturers are describedfurther in paragraph C of this section.
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LOADS BASED ON THE FOLLOWING CONDITIONS:
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APPLIED WITH LOD CENTER LOCAfED AT ARADIAL VISTANCE FROM ENGINE OF 32-504) 1.050 AND RESULTANT 0, AR2 THL ALLOWABLE LIMIT LOADS FOR REAR MOUNT ATTACSMN .
UNIFORSILY DISTRIkJ1 ED ON S,fl-CES AS SHOWN ALL VALU S GIVEN BELOW, ITESE LOADS REPRESENT TAND RES41.TANT R3IWuED WITH LOOS DISTRIBUTED TO 4 BOLTS. LOADS AT EACH LOCATIk AND I) NOT OCCUR SIMRTANE
2. LOAD 1v ACTING ON FRONT MOUNT MAY BE APPLIO WITH LOAD CENTER LINRIf 7a,O0 LaS 21 A,11 (-G ].ITi I, ')
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THE ALLOWABLE LIMJT LOADS TOR EACH FRONT MOUNT ATTACHMENT MUST SATISFYALL VALUES GIVEN BELOW, IHL. LOADS REPRIFSF'NT THE MAXIMUML 7 AT L..AOLCATION ABC, DO NOT O.CUB SIWMIT14EIJSLY
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Prat & Whitney AircraftPWA FP 66-100
Volume III
tU:LGHT REAR MIOL*JT DATA'EITIC REAR VIEW OF SEAR NI0111
DIECION I* In DO'T AS CHWN,IEIIETIVE IN OPPOSITE DIRECTsON
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MPESENT TIC M&AWM LOADS AT LA41 LOCATION AND DO NOT OCCIAR PRATTI&tnITNEV 411 RAFTtiT WATAEOUStY COJRtAED.
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DI-23 DI
Pratt & Whi y AircraftPWA FP 66-100
i 2. Accessory Drive System Volume II
* a. Power Takeoffs (A) (B)
The power takeoff (PrO) gearbox and drive pid conaecLion provided for
driving afrFrame accessories and for starting the engine Is located at thetop of the engine. The drive pad connectin can be oriented in the direc-tion required to match the location of the airframe supplied drive shaftsas coordinated to meet the airframe manufacturer requirements. The Boe-ing connection !s on the top of the gearbox. The PTO for Lockheed can betiground installed for either a left- or right-side connection.
For Lockheed only, an additional power takeoff-environmental controlsystem drive pad (B) is located on the lower right side of the engine fordriving the airplane environmental contrcl system air compressor.
Drive attachnent requirements and horsepower extraction, torque, andmoment limitations are tabulated on the Installation Drawing.
The J58 high Mach number experience for power takeoff gearbox design,
gear trains, and lubrication has been used extensively in the design of the
pover takeoff gearbox drives and accessory drives.
b. Accessory Drives (R)(S)
For Boeing only, two additional accessory drive pads are provided onthe engine power takeoff gearbox for direct mounting of two airplanu systemhydraulic pumps. The location, configuration, and horsepower extractionrequirements for the design of the gearbox were coordinated with Boeing.Prive pad requirements; spline definition; and horsepower, torque, andmoment lmi.tat-ons are taulLEJ on !.le Tnntaiiation Drawing.
c. Power Takeoff Decoupler Actuation (BV)
An electrically-actuated power takeoff drive shaft decoupler is pro-
vided to permit in-flight or ground decoupling of the airplane remote gear-box drive system. The systeta permits decoupling in the event of an air-plane accessory failure so that the engine can continue in operation.Only ground recoupling with a static engine is possible. The electricalconnection description is located in figure 5. This provision is currentlyshown only on the Lockheed engine.
3. Fuel System
a. Fuel Inlet (VI)
A common fuel inlet to the main and duct heater fuel pumps is providedon the upper right side of the engine for Boeing and on the left ride forLockheed. The Boeing Inlet flange connection is of the bolted flange typewith a conical metal seal. Lockheed requested and coordinated a specialMarman Clam,) Conoseal type flange for their requirements with sealing
configuration to be established.
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Pratt &Whtney AircraftPWA FP 66-100
i Volume III
The additional inlet lines for a common fuel inlet connection art pro-vided as optional equipment as requested by and coordinated with the air-
i frame manufacturers.
b. Fuel Return to Airframe (AF)
Fuel is returned to the airplane from the engine foel contrt,l durengine windinilling operation or when the fuel inlet teMPCe7at1Lre is highenough so that the added oil and accessory cooling load of the fuel system
rezults in reaching the fu-l system temperature limit. Details of thefuel control system are given in Volume Ill, Report B. Provisions aremade to permit a minimum fuel flow in the return line at all operating
conditions to cool the fuel return line and prevent the fo#rnation of coke.
c. Fuel Pump Outlet Vent (FV2)
For fuel pump venting, a threaded connection with seal provisions islocated at the fuel pump discharge boss. Venting of the fuel pump will
expedite the pump recovery following component replacemewt or in theevent the airplane fuel tank to engine line inadvertently becomes airlocked.
d. Fuel Drains
Fuel drains are located at the low point of cacti of the following fuelsystem components to allow fuel drainage overboard for safcty and cleanli-ness:
1. Combustion chamber - gas generator2. Fuel manifold drain valves - gas generator, duct heater Zone Ij
All drain points have threaded connocrtion with seal nroviion Andhave been sized for engine shutdown fuel drain flows. For Lockheed these
fuel drain locations (FDI, FD4, FDIS, FDI8A) are shown on the installation
drawing and the expected drain quantities have been coordinated. ForBoeing the fuel drains are connected to an engine supplied drain tank.
e. Engine Fuel Power Control (AJ) IThe engine fuel control and fuel system concept and functions are
described in Volume III, Report B. Only the interface requirements areconsidered in this section. The rapid removal unitized control conceptincorporates a universal type coupling between the control shaft and an
engine mounted drive. For the Lockheed installation, the fuel control~~thrust settlngs ate accomplished by an airframe-supplied electrical drive.The drive is located aft of the unitized fuel control on the left side of Ithe engine and is connected to the control through a splined shaft andpulley system. Fuel control settings and allowable installation loads
for the drive attachment to the shaft are shown on the Installation Draw-ing. ':
D'-29
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Pratt &W ney AiraI!fPWA K? 66-l100Volume .11 Ii'
A different power control system was coordinated for the Boeinginstallsation. Boeing required a nechanically actuated fuel control rack-and-pinion shaft with a maximum of 5 inches of stroke travel. The ettach-ment is adjacent to the unitized fuel control on the upper left side ofthe engine. The attachment requirements and load limitations are on theInstallation Drawing.
f. Engine Fuel Shutoff Control (AK)
A detail description of the configuration and functions of the fuelshutoff control are presented in Volume 1i, Report B. I
The fuel control shutoff &,4ten for the Lockheed installation consistsof an airframe-supplied elect:1- c,1 drive system mounted directly on thefront face of the fuel control.
An engine mounted shutoff control spitned shaft located on the forwardside of the unitized fuel control has been provided for Boeing. Mechanicalactuation has been coordinated with Boeing for the fuel shutoff control.
4. Oil System
a. Oil breather (LB2)
The breather pressurizing valve mounted on the engine main gearboxprovides pressurizing and venting of the engine oil system. Pressuriza-tion of a fixed absolute and delta pressure is controlled by an aneroidvalve system and a relief valve, For Lockheed, overboard breather vent-ing from the deoiler is accomplished through the breather pressurizingvalve by connecting to a four-stud flange with a metal seal. Allowableloads are shown on the lstallation Drawing. For Boeing, overboardbreather venting is accomplished by an engine supplied vent pipe to the
reverser-suppressor cowl.b. Oil Tank Inlet (L2)
The ol tank fill port is located on the upper portion of the oil tankon the right side of the eng ,e. A dipstick is attached to the cap as a 1means for determining the oil level.
A separate remote oil filling system connection with seal provisions ifis provided on the aft side of the oil tank as requested by Boeing.
c. Oil Tank Overflow3
For Boeing, oii tank overflow connections for the manual and remote
fill (LD2) systems have been provided aL the bottom of the tank. Themanual fll drsainys is routed to an tngine supplied drain tank,while the remote fill drain connection is made by a two-bolt flange withan "0" ring. For Lockheed, an oil tank overflow connection for the manualfill (LD) system has been provided at the bottom of the tank as a two- Ibolt flu ge with an "0" ring seal. Use of the overflow provisions withoverboard drains will ensure that engine oil does not drip on the engineor nacelle.
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Pratt & Whftney AircraftPWA FP 66-100
d. Oil Drains (LDI) Volume III
A manual drain valve is provided at the lowest point of the oil tankto permit oil tank draining. A threaded connection is provided on thevalve to permit the attachment of an overboard drain line.
Threaded plugs are located at low points of each of the following com-ponents to provide means for draining engine oil for maintenance purposes:
1. MaLn gearbox2. Oil pump strainer3. Oil pump gearbox
4. No. I and 2 bearing and seal compartment sump.
All connector sizes and locations have been coordinated with the air-frame manufacturers and are shown on the Installation Drawing.
e. 1,al Drains
Overboard seal drain provisions at the seal cavities of each of thefollowing components are provided to facilitate seal drainage overboardfor engine safety, maintenance, and cleanliness:
1. Unitized fuel control drive2. Main fuel pump accessory drive3. Hydraulic pump and hydraulic pump accessory drive pad4. Duct heater fuel pump5. Duct heater pump controller6. Aerodynamic brale actuator7. Secondary al valve actuators (Boeing only)8. Environmental control system power takeoff (.ockheed or-"y9. Tachometer gczni:Lor accessory drive
10. P.T.O. accessory drive11. Hydraulic pump PTO pads (Boeing only).
The attachment requirements for Lockheed have been coordinated and areshown on the Installation Drawing. For Boeing these drains ar- connectedto an engine supplied drain tank.
f. No. 1, 2, and 3 Bearing Seal Vents (CG) (CH) (CJ) (BA)
Venting of the bearing compartment labyranth back-up seals to ambientis required as an added measure of protection to ensure thac engine oilleakage cannot get into the air passages and contaminate the bleed aii2and to ensure that surrounding hot air does not enter the ccmpart et andauto-ignite or coke the oil. 'or Boeing this is achieved by providing anengine supplied overboard vent line from the separate vents to the reverser-suppresscr cowl. For Lockheed, circular bolt flange overboard connections(CG, CH) are provided for the No. I and 2 bearing seal vents on each sideof the top centerline of the engine and a threaded connection (CJ) is pro-vided on the bottum. A common bolt flange is provided for the No. 3 bear-ing seal vent (BA) on the left side of the top centerline. The No. 4bearing compar..ment vents through the exhaust nozzle; no airframe connec-tion is required.
DI -3 1.
Prtt & Whiney AircraftPWA FP 66-100Volume III
Allowable loads for these connections are shown on the InstallationDrawing for Lockheed only.
5. Electrical Systems
a. Ignition Exciter (M) ITwo ignition system exciter boxes located on the lower left and lower
right sides of the engine are provided as voltage boosters for the engineelectrical igniters. Engine-supplied individual leads connect the boxesto the main and duct heater combustion chamber igniters. This systemprovides a redundant dual ignition system for both primary engine and ductheater igniters. The ignition system power requirements and the exciterbox and igniter locations were coordinated with the airframe manufacturers. 11Several changes in the location of the system were required to satisfythe maintainability requirements. Electrical connectors are located oneach exciter. Descriptions of the connectors can be found on the Electrical LiInstallation Diagram, figure 5 for the Lockheed L-2000 and figure 6 for the
IBoeing B-2707.Lb. Control System Remote Adjustment Input (BS)
Means to electrically adjust the primary fuel flow as an engine pres-sure ratio adjustment and the fan nozzle area as an airflow adjustmentare included on the engine. An electrical connector is located on thefuel control for the remote adjustment signals. The above system andoptional variations to provide manual or automatic remote adjustment are Idescribed in detail in Volume III, Report B, Section III. An electzical
&connector description is located on the Electrical Installation Diagram.
The automatic remote adjustment system will bp nvallable a optIonal epfllnm for ther an tngine-mounred or airplane-mounted electronic
computer. In either case, additional electrical connections, which arenot shown on the Installation Drawing, are required. The remote adjust-ment system is further described in Volume II, Report B, Section III.&
6. Air Bleed and Vent Systems Ia. High Pressure Air Bleed Ports (U)(AM)
Air bleed ports for cabin or air pressurization systems are located Ion the engine diffuser case in positions specified by the airframe manu-
* 4facturers. i itA bolt flange type connection, located on the engine cane 4,,a$t above
the right horizontal centerline was coordinated with Lockheed to provideI bleed air for the aircraft cabin air-pressurization compressor (environ-I mental control system).
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The Boeing installation requires an additional bleed air duct to mani-fold the engine bleed pads in the upper quadrants into a single largecircular bolt flange at the top of the engine. The flange, as in the case
of the power takeoff drive pad, faceq up to directly accommodate the air-craft-mounted air conditioning system. A separate, smaller connection(AW) is located on top of the engine to provide high pressure bleed aMrfor che airplane anti-icing system. In all cases, the coordinated pro-visions meet the bleed pressure load, installation vibration load, and
environmental design requirements.
I b. Bleed Pilot Valve Ambient Vent (BD)
The engine starting bleed system requires an ambient pressure sense
to properly schedule the start bleed system. This is accomplished by ascreen protected port on the bleed pilot valve located on the left side
of engine. No airframe attachment is required.
c. Duct Heater Fuel Pump Exhaust (CE)
A duct is provided on the aft side of the duct heater turbo fuel pumpfor exhausting the pump turbine, which is driven by engine bleed air.
For the Boeing installation, a discharge duct to the secondary air systembulkhead will be provided. For the Lockheed installation, a short dif-fusing duct is attached to the pump exhaust.
d. Secondary Air Flow Control
The installed engine secondary airflow systems differ markedly for
the Boeing and Lockheed concepts.
Lockheed requires no secondary airflow control on the engine sincecortrol 12 nr--nplshcd by t'h airfrj e i" L system and is essentially
I a flow-through system. Boeing has a requirement for an unpressurizednacelle. Six secondary air flow ducts which Are routed froin the engineinlet to a bulkhead at the forward end of the reverser-suppressor are
provided for Boeing ins* llation. All six ducts have a check valvelocated at the aft bulki.ead to prevent backflow during reverse thrust andairplane take-off. Four of the six ducts contain hydraulically operatedvalves to control the secondary air flow requirements for reverser-suppressor performance and cooling flow during cruise. Valve positionindicator switches are provided. The electrical connectox descriptionsfor the switches are located on the Electrical Installation Diagram.
e. Aerodynamic Brake Position Valve (CX)
The aerodynamic brake position valve requires an ambient pressurevent. This is accomplished by a screen protected part and is shown on
the Installation Drawing.
7. Instrumentation and Instrumentation Provisions
a. Fuel Flowmeter Provisions (AW) (AX)
I =
DI-35
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Prst &aWhnyArcraft IPWA PP 66-100Volume IIl IProvisions for fuel flowmaters to measure main engine and duct heater
fuel flows are included in both the main fuel and the duct heater fuel
supply lines as shown on the Installation Drawings.
b. Main Fuel Pump Inlet Pressure (FP-l)
Instrumentation provisions for sensing fuel p imp inlet pressure are jfprovided on the engine. Definitions of the connections are shown on theInstallation Drawing.
C. Fuel Inlet Temperature (FT1)
Instrumentation provisions for sensing fuel inlet tamperature areincluded on the engine. Definitions of the connections are shown on the 0
Installation Drawing.
d. Fuel Filter Inlet and Outlet Pressures (FPlO) (FP11)
Instrumentation provisions for sensing fuel pump filter inlet pressureand outlet pressure, or a delta pressure, are included an the engine. jDefinitions of the connections are showni on the Installation Drawing.
e. Oil Level Sensing (LL-l)
Oil level or oil quantity indicating instrumentation provisions are iincluded on the engine. The oil tank flange connection and the internalsensor envelope are depicted on the Installation Drawing.
f. Oil Pressure and Temperature (LP1) (LT1)
Instrumentation provisions with a threaded connection having seal riprovisions are included in the main oil pressur line from L- c oier toonin- for snalm oii -rtsaura. An identical connection for instruren-tation provisions for sensing oil temperatuve is also provided. Pro-visions for direct engine mounting of a pressure transmitter will beincorporated if required. The above provisions are shown on the Installa-tion Drawing. fi
g. Oil Pressure Transmitter Vent (LV3)
A threaded connection with seal is provided on the oil pump gearboxfor venting the engine oil system pressure transmitter.
h. Oil Filter Inlet and Outlet Pressure (LP4) (LP5)
Instrumentation provisions for sensing oil pressure to and from thefilter; or a drlta pcessure, are include on the engine and are shown onthe Installation Drawing.
i. Turbine Exit Temperature (TT7) (ITT7) (Average and Individual) :An averaging and individual exhaust gas temperature measurement syste-
is provided on the engine. The airframe electrical connection is locatedabove the right horizontal engine centerline fer Lockheed and on the lower
-D--36
Praft & Wh*ney AircraftPA FP 66-100
Volume III
right side for Boeing. These locations were coordinated to provide con-nections for both the aircraft turbine exit temperature cockpit inJicating
fsystem and for maintenance checking of the nine individual probes. A
description of the system is provided om the Electrical InstallsadonDiagram.
J. Turbine Exit Pressure (PT7)
Turbine exit or exhaust gas pressure instrumentation is provided on
the engine. The coordinated airframe threaded connection for cockpitindication is located on the right side below the horizontal centerlinefor Lockheed and above the horizontal centerline for Boeing.
f k. Duct Heater Nozzle Position Feedback (AL)
A linear variable differential transformer to indicate duct heaternozzle position is provided on the engine. The electrical connection islocated on the right side of the engine ahead of the rear mount ring. Adescription of the connector is provided on the Electrical Installation
IDiagram.1. Reverser Position Indicator (BC)
[An indication of reverser ejector nozzle clamshell position is
obtained from a limit switch located on the right side of the engineahead of the rear mount ring. A description of the electrical connectiin[is provided on the Electrical Installation Diagram.
1m. Low Rotor Speed Transducer (BT)
transduepr in provided on thc eajlu, Lu ildicale the Jow rotor specdby sensing 2nd-stage fan blade rotation. The location anid the descriptionJof the connection are provided on the Electrical Installation Diagram.
n. Aerodynamic Brake Position (AS)
An aerodynamic brake position indication is supplied by a switch actu-
ated by the aeredynamic brake actuator arm. This provides an indicationof the inlet guide vanes in the start and cruise position. An indicationfor brake-on position is not considered necessary since rotor speed isaffected significantly when the brake is on. A description of the connector
is shown on the Electrical Installation Diagram.
0. Vibration Pickup (2.-
Vibration pickup mounting brackets are provided on the engine caves
as depicted on the Installation Drawing.
p. In-flight Power-Setting Systems Provision
The use of engine pressure ratio as the basic power-setting parameterhas proved successful on both turbojet and turbofan engines with fixedjet nozzle areas. By adding the measurement of total fuel flow to en-
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Pratt& Whiny AicrftPWA FI 66-100Volume III
gine pressure ratio, slightly be.ter thrust setting accuracy is achievedfor a duct-burning turbofan engine. The use of turbine exit total pres-sure provisions, inlet total pressures, and total engine tuel flow mess- iurements provides the required engine parameters for an in-flight powersetting system. Such a modified engine pressure ratio system satisfies 91
the requirement for a simple accurate method of setting engine thrust.The details of such a system are described in Volume II1, Report A.
C- INSTALLATION CONFIGURATION AND SUPPORTING DESIGN DETAILS
1. Installation Configuration and Design for Airframe CompatibilitvF
The engine mount points are primarily ed to effectively reactthe engine thrust and weight to the airp e ,ing or other primary struc-ture. This necessitates locating the w it points in the upper quadrants.One of the front mount pointB reacts t total thrust load and a proportion-ate amount of the vertical and side It er front mount car-ries a proportionate amount of the vert. rear mount attach-ment points provide for airframe-supplied links t fow for axial thermalgrowth of the engine. The rear mounts also carry th- greater part of theengine vertical and side load. Further information out the engine mountsystem design may be found in Volume III, Report B, rtion II, paragraph 11.
The resulting engine-coordinated mounting points and the requiredairframe mount structure necessitates positioning the major engine con-trols, components, and accessory drives to the lower side areas of theengine for the Lockheed installation and on either side of the horizontal Hcenterline for the Boeing installation. For the Boeing airplane, the com-ponent positioning must be compatiable with the Boeing requirement forsecondary airflow ducts. The secondary airflow is required for reverser- [_suppressor cooling and ejector nozzle performance. The dted ssystc Pet-
mits designing c nonpreuLiveti naccl. Lir Lockheed engine configurationanci component arrangement are greatly affected by the requirement to mount 11an engine-driven airframe environmental control system air compressor andheat exchanger on the bottom of the engine.
The power takeoff gearbox and drive pad and the locations differ be-tIitween the Boeing and Lockheed installations. The Lockheed power takeoffis positioned at the top of engine to drive the airframe remote gearbox-mounted accessories. This gearbox has the added capability of being en-gine-mounted for either # right or left side airplane installation. Anadditional power takeoff-environmental control system drive pad is pro-vided on the lower right side of the engine for the bottom-mounted environ-mental control oystem air compressor. For the Boeing airplane, the en-1gine power takeoff gearbox and drive pad incorporate two additional hy-draulic pump acca@ory drive pada for the airframe inlet control system.The gearbox is located at the top of the engine with the power takeoffpad facing upward.
The high pressure air bleed ports and manifold are located near thetop of the engine for Boeing to facilitate direct routing of the air ductand vents into the airplane. The Lockheed installation requires only asingle bleed sit port located in the lower right aide of the engine to
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Pratt Whitney AlrcmftPWA FP 66-100
Volume IIIfacilitate direct routing of the air du.:t to the bottom-mounted enginedriven air compreasor. All the heariog compartment labyrinth seal vents
i- are located iT the upper quandrants of the engine for ease of installation.
A common fuel inlet and manifold to the main and duct heater fuel pumpsis provided near the top of the engine to facilitate connection to the
airplane fuel manifold as requested bv the airframe manufacturers. The
locations differ slightly as a result of airplane differences.
The enIne reverser-suppressor consists of the following major ex-ternal components and assemblies: Reverser-suppressor to nscelle shroud.reverser doors, tertiary air blow-in doors, main strtcturc, and aerodvnam.-icelly-actuated variable nozzle flaps. The nozzle is canted down 5 tie-grees for Lockheed and 6 degrees 30 minutes for Boeing. The canting isrequired for airplane performance and installation compatibility. The
forward face of the reverser-suppressor provides a flange for bolted at-I tachment to the nacelle outer contour air seal at the engine-to-airframe
boundary. Thrust reversing is provided by two hydraulically actuatedsemicircular reversing clamshelit. Backflow o the reversed exhaust
gases during reversing operations is prevented by a series of air balancedflapper doors which close because of exchaust gas pressurt. The doors arelocated around the engine nozzle near the reverser-suppressor front face.
Secondary airflow for reverser-suppressor perforuinsce and cooling is ob-tained from the airfrme air inlet system. This so~ondary-air flowsalong the outside of the engine and through the flap,,er doors to the eject-or nozzle for the Lockheed installation. For Boeing, the secondary air
[flows through six ducts,
The thrust reverser doors are configured and positioned to providethe required reversing gas flow targeting that has been coordinated with
the airframe manufacturers. The door locations are Ai function of the de-sign of the airplane wing and engine installation.
2. Installation Configuration and Design for Maintrinability and Servicing
Borescope inspection ports have been spaced circumferentially around
the engine diffuser case clear of the engine components and plumbing topermit inspection of the main burner, fuel nozzles, nozzle guide vanes,and burner swirl guides. Thc Coiihpi essor horeseope inspcctlon iotts arepositioned along the bottom of the engine in a clear area that allows view-
ing of all compressor blade stages. The turbine inspection borescope portsare positioned at two locations on the rear mount ring case to allow viewingall turbine blade stages. The turbine 3rd-stage can be inspected fromthe exhaust end of the engine. Provisions to rotate the high compressorduring borescope inspection have been incorporated into the main gearbox.All of these functions can be accomplished with the engine installed, if
i permitted by the nacelle design,
Components of the engine have been packaged or positioned to provide
access to all inspection and servicing points on the engine. Specifically,
the packaging allows access to the eight large access panels located cir-
cumferentially around the engine. The access panels permit inspection ofthe main burner fuel supply p.u.bang, rornresso bleed valve plumbing and
i compressor air bleed plumbing, replac .ent of the primary gas generator
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Pratt &Whtney AlrcraftP4A FP 66-100Volume III
fuel n, szles. and removal of the compresso., bleed valves. The engine hasbeen designed so major plumbing is not routed over these panels. Thereverser-suppressor design incorporates access nnje for inspectionof hydraulic actuitors, plumbing, and reverser mechanism and for minoradjustments and parts replacement. Adjustmrits of the reversing andtertiary blow-in door system are provided to allow complete interchange-ability of parts, thus not requiring replacement with matched parts orsets of parts. The clamshell hubs and bearings are easily removed andinspected after removal of the outside access panel and inside cover-plate. The complete reverser-suppressor assembly and nozzle can be in-stalled or removed krom the airplane without requiring other disassemblyor rigging adjustment.
The engine oil system has been designed to utilize threaded type 1chip detector plugs. These are accessible at strategic locations to Iprovide early warning of failure or excessive wear of bearing or powertrain areas of the engine. 5
I :
All oil scavenge pumps are provided with inlet screens .-,r pro-tection against foreign objects in the oil sumps.
Lines and tubing on the outside of the engine are provided with suf-ficient mechanical connectors to permit easy removal of components and Jhot section inspection. All liues that cross Flange "G", which is the ]separation flange for a hot section inspection, have an additional con- Inection permitting removal of short lines.
Fuel and exhaust nozzle area control management functions have been [packaged on the left side of the engine in a single unitized controlwhich can be removed by removing nine bolts and disconnecting the elec-trical and pneumatic connections. All fluid lines are connected to thepedestal base behind the uuitized control so that disassembly and as-sembly of these lines is not required. The fuel controls linkages aredesigned so that removal of the fuel control will not disturb the air-frame fuel control actuation system. j
Fluid filters which are provided in the components of the fuel andhydra-lic sytc at auLiUIkcG 60 rtmovai ran be accot-pished easilywithout the prior removal of other engine components.
All components weighing 45 pounds or more are provided with liftingeyes or other lifting provisions to facilitate ground handling for fieldinstallation or replacement.
Space for ground handling at the front and rear mount rings has been Iicoordinated with the airframe contractors and is specified on the Instal-lation Drawing. Ground handling fixtures for transportation or engine in-stallation and removal from aircraft may be attached to the engine in '>these areas as desired by the airframe manufacturers.
Further )etailed information about engine configuration to facilitate
Lervlr-ng and mamntenancc my be fouind it Volume IV, Report F, Section I; ji
Volume IV, Report D, Section III; Volume V, Report C; and Volume III,Report B, Section II. Further maintenance and inspection informationmay be found in Volume I1, Report B, Section II, paragraph H. i
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" CODENTIAL Pratt & Whiey AlrcmftPWA FP 66-100
a Volume III
3. Installation Configuration and Design foj. Safety Considerations
Engine-mounted components, lines, and fittings have heen positionedto be within the envelope defined by the front and rear mount rings asfar as possible with the constraints of the airplane requirements andphysical space. This provides engine structural protection to the fluid-I filled components and plumbing. Similarly, all fluid-filled components,except for the overboard drain valves, have been positioned away fromthe bottom of the engine for improved installation safety. For theBoeing Installation a fuel drainage collector tank is provided on thebottom of the engine. Overboard drains from fuel components of theengine will be plumbed to this ceittral collection tank and thus keep
lU the airport ramp areas free of fue!. While the airplane is parked, noU fuel wili drain overboard. The collection tank will be emptied by
aspiration through an internal ejector during engine operation.
I All fluid-carrying lines have been designed for maximum safety bythe use of lines with integral ferrules at all connections. Integralferrules are incorporated by forging and subsequent machining of thetubing ends to fotrn the fetiule, thus eliminating the need for brazedconnector ends. All line connectiona are threaded connectors using astandard 37-degrec coiie end and a nickel seal. Use of this type ofconnection is based upon a record of excellent reliability in flight atMach 3+ with the J58 engine. All engine and enginc-to-airframe inter-face boundary connections through 1.25-inch diameter are of this type.
Above 1.25-inch diameter size, connections are bolted flanges using theHaskell design "K" seal. The K-seal has nroved superior for sealing dur-ing thermal shock test.o and from 158 experience in these sizes. All en-gine connections are the types described, except where other types arespecifically requested by the airframe manufacturers. N1 brazed con-nections are used because of the lack of reliable inspection methods.All lines and brackets have been designed for thermal deflection, in-ternal pressures, manufacturing misalignment, and vibratory and maneuverloads. Tube and bracket stresses resulting from thesc loads are main-tained nt a safe margin below the material yield strength at all oper- mating temperatures to achieve the hpst iow cycic fatiue lifet Linehracket spans ae checked to assure that the line and bracket natural
frequencies are outside the engine rotor speed range. Details of thedesign of lines, fittings, and brackets are given in Volume III, L-port B.
All engine and engine-to-airframe interface boundaries for lines andelectrical connections that are in close proximity to each other havebeen sized or designed to prevent improper connection to the engine oiairfrare.
The reverser-suppressor reversing actuation system has been designedso that if a hydraulic system failure should occur, the aerodynamic forces
t acting on the clamshell segments will automatically move thcm to the sub-sonic cruise position so that no reversing will result and forward thrustis available The aerodynamically-actuated tertiary blow-in doors andthe re,,zosr- doors have been provided with an interconnecting lockoutmachenism so That reversing cannot take place while the tertiary blow-it doors are in the closed (supersonic) position.
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R~an £Whftny Aircraft CFIOE TfALPWA "-' 66- 100Volume III
As optional equipment, an automatic restart system switch is avail-able. This engine-mounLed switch senses a drop in compressor dischargepressure and automatically closes the ignition circuit for main burnerrelight. The system is not currently shown on the Installation Drawingbut is shown in Volume III, Report B, Section III.
All filter housings in the main fuel and oil supply systems havepressure sensing provisions at the inlet and outlet of the filter to !
permit measurement of filter AP as an indication of filter element clog-ging.
Provisions for chip detectors in engine oil system are covered above.
D. SECONDARY AIRFLO4 SYSTEM DEFINITION AND REQUIREMENTS 'i1. Description
The purpose of the secondary air system is to provide airflow forreverser-suppressor performance and for cooling of the reverser-sup-pressor shroud and ejeCLor trailing edge flaps. in addition to thesefunctions, the system is used to bypass air from the airplane inlet,commensurate with engine and aixframe supersonic inlet requirements.
The Lockheed and Boeing airplanes use distinctly different sec- I.ondary airflow systems as a result of different inlet and nacelle systemconcepts. The differences have little effect on the basic engine, buta significant effect on the installation configuration.
The Lockheed airframe system is shown schematically in figure 7.Secondary air enters the engine nacelle cavity through airframe-suppliedinle control valves in the plane of the engine air inlet. This alsoprovides a means for shutting off the air and isolating the engine aresin the unlikely event of a nacelle fire. The system utilizes the fullnacelle annulus as a flow path, resulting in a oragnirzed naccllc car-ity aa therefore sealed from amblent static pressure. An engine flappervalve system is incorporated into the secondary air flow path at a re-verser-suppressor bulkhead just behind the engine rear mount. The put-pose of the valve system is to isolate the upstream environment fromhot gases during reversing.
The Boeing airframe utilizes a ducted secondary airflow systemshown schematically in figure 8. The system incorporates six ducts ex-tending from an annular manifold aft of the engine front mount ring tothe bulkhead at the forward end of the reverser-suppressor. The air- iflow is bled from the boundary layer in the engine gas path and flows inthe ducts to an engine nacelle cavity just forward of the reverser-sup-pressor. Check valves are I.ncorporated in e4ch of the ducts at the aftbulkhead to prevent the reverse flow of hot gases during roveree thrustoperation. Four of the duc'ts incorporate hkdraulically-actuated valvesto throttle secondary flow as required for pf" formance and cooling. Twoof these valves are scheduled by the engine control system to close atabout Mach 2.0 to mAtch the inlet and engine performance requirements.At cruise the option of manually closing the other two valves is avail-
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[I CONFBENTAL P~t W fn ArnfPWA FP 66-100
Volume III
i able if the engine is being operated at low cruise powters. Since thesecondary flow is contained in ducts, the nacelle annulus will be main-E tained at or near altitude ambient pressure.
S 1--' ine '-, -' iE ine
.I.- Engine Cay,-- H[Prenaurized NcleCavity
.....l F Floating--'Nacae Case Flapper Valvti Fow t Reverser-Suppressor ---
Figure 7. Lockheed Secondary Air Flow System FD 17035DI
Engine
I! I U " "
U Nacelle
B lo sCavity
,
Airflow Duct ActuatedControl Valve
Free Floating Flapper Valve
Flow to Rmvrser-S..- ---c,
Figure 8. Boeing Secondary Air Flow System FD 17034
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Pratt SWhtnsy Abvraft CONADLPW& 1? 66-100Volume III
2. Design Objectives
Within the framework of system design and airframe coordination,the design objectives of the secondary airtlow systems were: j
1. To provide secondary air to the aft section of the nacelle cavitycompatible with the ejector nozzle performance and reverser-sup-pressor cooling requirements.1
2. To bypass air from the engine inlet at a flow rate compatible withengine/airframe supersonic inlet requirements.
3. Design Requirements I
A minimum of 2% corrected flow is required at all operating conditionswhen the reverser-suppressor tertiary doors are closed. II
The airframe requirements relative to the airflow and inlet compat-ibility are being coordinated on a continuing basis with both airframemanufacturers.
4. Design Criteria 1Consideration was given to the following factors:
1. Minimize all resistance to flow in the nacelle cavity and in l11the bypass ducts
2. Design the secondary airflow inlet geometry to have a minimumeffect on engine inlet requirements II
5. Design Approach
Balanced areondery flow requitements for the engine and the airplane Liinlet systems were determined by the same gereral method used for the J58engine YF-12A and SR-71 bypass system flow requirements. Flow parameter
characteristics of each system were determined using component blockage,net effective flow areas, and compressibility relationships. After theflow characteristics of the system had been established, the balanced flow -
was determined using the airframe-supplied inlet pressure and the ejectorinlet pressure for various flight conditions. Li
E. ENGINE HOCKUP PLAN U1. Scope and Objectives
This plan describes the methods and planning used to fabricate, pro-vide, and maintain up-to-date full~acale engine mockups throughout t1heJTFI7 program. The program provides the airframe manufacturer and En-gineering Design with the moat useful and complete mockups possible andwith a clear conception of the engine external configuration. These mock-ups are primarily used to ensure engine installation compatibility andwill aid in achieving practical solutions of general configuration de-tails covering installation, accessibility, serviceability, and maintain-jiability throughout the design and development of the engine, The planfurther states the accomplishments achieved during Phase I-C in pro-,viding full-scale engine mockups to both airframe manufacturers.
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Pratt & Whitry AircraftPWA FP 66-100
Volume IIIIi 2. Organization
The engine mockup parts list is established under the direction of theManager, JTFi7 Development. A Development Project Engineer is responsiblefor coordination between the Product Support Group, who contact the air-frame manufacturer, and the Design Installation Group and for ensuring thatthe mockup configuration is in accordance with the airframe manufacturer'srequirements as well as the engine design requirements. These will bereviewed by the Chief of Configuration Managements. The Product Support,Installation and Field Engineering Groups are responsible for the coordina-tion of installation requirements between the airframe manufacturer andProject Engineering. The JTFl7 Engine Design Group is responsible for the
engine design and drawings necesaary to manufacture and maintain theIengine mockup.
The Prodtct Support Grcup is responsible for the construction andmaintenance of the mockup engines, including the processing of materialnecessary for updating, used at the airframe manufacturers facilities.
3. Description
IEngine mockups which are discussed in the following paragraphs canbe described by the following class definitions:
1. Class I Envelope Engine Mockup - This class of mockup representsthe overall size and approximate contour of a new engine modeland is intended only to facilitate early installation studies.
2. Class II Preliminary Engine Mockup - This class of mockup isconsidered satisfactory for preliminary design work and isprovided early in the design program, prior to the establish-ment of the complete engine design. The parts are fabricatedo la,- .ad ns.,tonis imte to verification ot theprincipal dimensions. Construction materials are usually fiber-
glass with metal flanges for engine cases, wood with metal in-serts for accessories, and steel tubing for plumbing. Thistype engine mockup can be updated and maintained to the latestdesign configuration.
. 3. Class III Installation Design Engine - This class of mockup isintended to be used as an installation fixture for final installa-tion design work and is an exact representation of the externalfT details of the released engine design. All parts are manufacturedto released engineering drawings and are inspected for conformanceto the manufacturing tolerances specified on the drawings. Construc-tion materials are metal and fiberglass. -
4. R-qirement-:
Experience has shown that full-scale engine mockups, because of theirthree-dimensional nature, play a very important role in the initial estab-lishment of engine component clearances, plumbing routing, and for sub-sequent checks of the engine in the airframe.
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PlA FP 66-100
Volume IllDuring the Phase IIC program, one full-scale engine mockup was de-f
livered to each airframe manufacturer; each mockup represented the JTFI7specifically designed for the respective aircraft. At the same time, twoother engine mockups have been maintained and used as in-house engineering Idesign mockups. Each incorporates the major features of the basis engine Iand the differing airframe configuration.
During Phase 1I, a Class 11 engine mockup will] be maintained at IPratt & Whitney Aircraft and continuously updated to incorporate the latestdesign changes. Concurrently. a duplicate set of parts will be fabricatedand forwarded to the airframe manufacturer for incorporation into and up-dating of their Class II engine mockup. h
The Boeing Company has requested one and possibly three additionalClass II engine mockups of the Boeing JTF17 engine configuration for !delivery early In Phase III. Approximately two years later, floeing has arequirement for two Class III engine taockups representing a firm con-figuration for the prototype engine. Lockheed has requested an updatedI;Class II engine mockup in 1967 and two Class I1l engine mockups for deliveryin 1968 representing a firm configuration for the iTFI7 prototype engine.
The updated Class II and the Class Ill Pratt & Whitney Aircraft mock-ups will be used as design tools for coordineting all proposed changeswith the airframe manufacturer and for visualizing the Installation Draw-ing changes. These engine mockups will be manufactured as rtquired anddelivered in accordance with the schedule,
Throughout the SST engine program, emphasis will be placed on main-talc ability and accessibility features of the engine, the engineering [Imockups will aid in identifying such areas. Some of the features whichwill be emphasized are:
1. Quick disconnect features for components "2. Clearly d caLilien service chek points such as filters, chip
detectors, pressure and temperature check points, and oil tankservicing points.
3. Inspection provisions such as borescope, access plugs, andaccess covers.
5. Mockup Changes and Review
Should the airframe manufacturer's review of the Class III engine Ifmockup Indicate a need for configurational changes, field survey lay-outs will be prepared for si-frame manufacturer review and approval.This procedure is described in detail in Volume V, Report C. If thechange affects the engine mockup, parts will he made, checked out on Uthe Engineering Design engine mockup, and sent rn the zirframne manu-
.fnstalation on the engine mockup. When the revision issatisfactory, and the field survey approved, the engineering change will Iibe processed and incorporated into the Assembly Parts List Complete.
The engine mockuF maintained for Pratt & Whitney Aircraft Engineering IDesign will be available for inspection or Mockup Review after completion.Tentative suggested schedule dates are shown on figure 9. On notification
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I- Pratt & Whtney Rirct afPWA FP 66-100
Volume IIIby the FI, Pratt & Whitney Aircraf. will provide the facilities, services,personnel, and material for an engineering inape-tion of this JTF17 engine
mockup, tollowing the general procedures for Mockup Review in ANA Bulletin(406a.
6, srv'-sro--d and Facilities
Pratt & Whitney Aircraft has accumulated over 200 man-years of ex-perience in building jet engine mockups for both military and comnercialprograms.. The Installation Engineering Mockup Group and FRDC Engineer-ing Design have provided mockups in recent years for aircraft programssuch as the C-141, F-Il, B-52, 707, 727, DC8, SR-71, F-12, and the Jet-star. Photographs of typical mockups are provided in figures 10 through15. The experience of the Connecticut and FRDC personnel, the 11,300 sq.
LI ft of mockup design, manufacturing, and assembly facilities in Connect-icut and the 6,400 sq ft of mockup and assembly area at FRDC are avail-able for the SST program.
F. ENGINL INSTALIATION HANDBOOK
[ "The Installation Handbook will be provided as a ceparate enclosure.
1967 181*N APR U I JAN APR JULY OCT JAN APR JULY
... 4LDNGNOCKE
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BOEING B-270
1 1
i i , -j 4
P.WA(E NGINEERING "
I ~ ~EIGN MO)CKV I1 ~ i
[Figure 9. JTFl7A-21 Mockup Schedule Fl) 17641DI
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Prat & WhfneyAircraftPWr& F? 66-100V'iume III
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Figure 10. JTFl7A-21B Mockup, 314 Right FE 60737Front View DI Il['I
_
IIm
Figure 11. JTFI7A-2IL Mockup Delivered to FE 56829I
Tlamed / Right Front View
PWA Fl' 66-100Volume III
II
'Iigure 12. JTF17 Mockup, Typical Closeup FE 60414,Showing Main Fuel Control, Main DIGearbox, Main Fuel Pump, andEngine Hydraulic Pumn
Figure 13. JTIID-20 Mockup 0J58), Right Side FE 48692
IFigure 14. TF30P-l Installation Design Engine, H1-39671134 Left Front View Di
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Pratt & Wh-wy Aircraft CONINIWAIPWA FP 66-100 lU ftVolume III
F i u r 5 .J T D I n t a l ti n D e i g ng n eH 5 2 8
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- i Volume III
SECTION IIIINLET SYSTEM COMPATIBILITY
A. OBJECTIVES AND INTRODUCTION
IThe development of a compatible high performanct, engine/inlet systemis a primary requisite for initiation of supersonic transport flight
testing. In order to accomplish this development within the allowed timeEz schedule, P&WA and the airframe manufacturer will pursue, from the ear-liest possible date, a vigorous, coordinated program consisting of analyt-
ical studies, rig and model tests, and full-scale inlet and engine te~ts.As shown in the following sections of this report, analytical studiesrelating to determination of inlet and engine steady state requirements,dynamic compatibility, and distortion compatibility are already underway.
IP&WA has over eight years experience in development of engines designedto operate at flight speeds equal to or greater than those planned for theSST. For example, the J58 engine has accumulated more flight time at andabove Mach 2.7 than all other engines in the free world combined. Theknowledge acquired as a result of this extensive experience has been incor-
porated into the design and development plans for the JTF17.
Our experience has shown that it is necessary to integrate the engineand airframe inlet duct at the earliest possible date. Therefore, toavoid the possibility of delays in the flight program and the complete
engine-nacelle test program aL AEDC, initial compatibility tests will beaccomplished early in the engine development program on compressor cm-ponent rigs and development engines. In view of the rapid modificationsthat are desirable in the early development cycle (and long lead times onexperimental parts) the most efficient way to accomplish this testing isin the manufacturer-s test facility. Thus, a large portion of the inlet/engine compatibility testing will be run in the extensive P&WA sea level
and altitude engine and rig test facilities in preparation for the scheduledcomplete engine-nacelle tests at AEDC.
1 B. AIRFRAME/ENIGlNE LOMPATIBILITY AGREEMENT
The Engine Inlet Compatibility Test Program has been coordinated with[ the airframe manufacturer and included in the following documents:
1. Boeing:"Coordinated Inlet/Engine Test Plan" Comercial Supersonic Trans-port Program Report D6AlO007-2Engine/Airframe Technical Agreement Number D6AI0199-1 P&WA
2) Lockheed:L-2000 Airframe/Engine Compatibility Agreement Exhibit A
P&WA Interface Control DocumentLockheed PhA.e ITT PrVporal 11^", Se.tion 8AEDC Inlet/Engine Test Plan
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-iPraft £ Whimey AircraftPWA FP 66- 100Volume III
C. ENGINE INLET COMPATIBILITY TEST PLAN
The compatibility of the engine and inlet will be tested in logical,sequential steps leading to the prototype flight test.. Significant testmilestones are shown in Table I and the tests are more fully described itthe following paragraphs.
1. Program Integration of Fan/Compressor Rig Testing
Steady-state evaluation of simulated inlet total pressure distortionduring fan rig tests provides early indications of engine distortiontolerance and engine/inlet compatibility. Concurrent full-scale engine Itesting will incorporate the fan/compressor cnnfigurations developedduring this test phase that exhibit major improvements in distortionattenuation. Full use will be made of analytical simuiatioa rpchniquesto complement the testing phase, and improved techniques used as they
are validated. Also, test data will be used as it is obtained to updatethe analytical simulation program. The testing and analytic processesnecessary to attain distortion compatibility are intLer-related and neithercan proceed rapidly without the other.
Information obtained from distortion testing with the fan will beutilized to obtain representative duct burner and high compressor inletprofiles for additional component tests. The importance of fan/1compres-sor distortion testing Is emphasized for two reasons: first, this isthe earliest and most economical means of evaluating distortion compati-bility problems and second, this testing provides data which is essentialfor the engine test phase. Ii
Table 1. Engine Inlet Compatibility Testing Proposed Milestones
1. 0.6 Scale Fan Rig - Distortion generators will be usedto simulate patterns obtained from airframe manufacturer Feb. 1967testing of model inlet.
2. Rematched Engine - Testing with distortion generators will ,b d.e to dis..rrion attenuating itfluence ot
high compressor. The data will be compared to fan rig May 1967
distortion test data.3. 0.6 Scale Fan Rig - The fan rig will be tested at Willgoos
Laboratory with the subsonic diffuser portion of the air-frame inlet. These tents will evajuate the subsonic dif- July 1967fuser performance with static pressure gradients imposed Iby the fan.
4. 0.6 Scale Fan Rig.- Variable vane simulation of high com-pressor. Test with required back pressures to reproducecircumferential variation of flow determined from thereautched engine. corroboration of this data when com-pared to the rematched engine distortion data will allow July 1967
development of fan to continue on isolated test basis.Data will else be compared with anaiytical calculationsof distortion effects for corroboration.
5. 0.6 Scale Fan Rig - Willgoos Laboratory testing with the Fsubsonic diffuser portion of the airframe inlet. The Sept. 1967July testing will be repeated to evaluate any changesindicated by the earlier test series.
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IPratt A Wh t ey Alrcraft~PWA F" 66-100
Volume III
ITable I. Engine Inlet Compatibility Testing Propos2d Milestones~(Continued)
6. Full Scale Fan - Testfng to be done at FRDC altitudefacility with variable vane simulation of high compres- Jan. 1968
sor. Distortion generators will be used to simulate
patterns obtained from al-f .-- manufacturers testing,
7. JTF17 - Engine will 6e teste. with distortion generators
to simulate patterns obtained by airframe manufacti rs Nov. 1967
testing. Steady state and transient performance will be Nevaluated.
8. JTFI7 - Engine will be tested with the subsonic diffuser,
flight configuration bleeds, and their controls. Data
will be taken at cruise to check performance and previewdynamic compatibility before AEDC testing. Sea level
static testing will be done to evaluate performance and Aug. 1968[ distortion effects with a full inlet on the engine.
Starting schedules will be optimized, The inlet will be
operated in a near choked condition for fan noise atten-uation studies.
9, JTF17 - AEOC Test Facility - A ground test engine withthe supersonic inlet will be tested. Configuration will
be a complete engine nacelle system to simulate the free*stream inlet.
10. Flight Tests - Propulsion system performavce and engine/
inlet compatibility will be evaluated during the initial
* 100 hour flight test program. Flignt tests are scheduled
to begin as follows,
Boeing 1st Prototype December 1969
2nd Prototype February 1970
Lockheed Ist Prototype March 1970( I2nd Prototype June 1970
2. Inlet DiStertcn Cznzrctok TebLs
The early compatibility tests will consist of compressor component
tests with simulated steady state distortion generated by distortion
screens (figure 1). Since the distortion will vary both in magnitude
and pattern as a function of flight Mach number, several representative
distortion generating screens must be avatlable. This also means that
the simulation screen should be readily changed as a function of rotor
speed and thus development of a remotely variable distortion generator
will be undertaken. During these tests the presence of bleed ports and
obstrujctions in the vicinity of the compressor inlet will Le simulated
as they can have an sppreciabie effect on the velocity distr'bution.
Also important is the dupliration of the static prcqsure fie !. This
will be accomplished by constructing the inlet section of the test
facility to conform with the Sometry of the iniet diffuser, of partic-
ular importance is the section within approximately one diameter of the
I fan face inicluding simulation of bleed flows. Coupling the fan to the
Inlet diffuser will modify the inlet distortion, but this effect isI expected to be of second order importance.
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Pratt & Wh sy AWcvAPWA PP 66-100
Ill/ Exit Guide
biloFlw Paee P11 2nd-Stage vane
l_ a FlowDuct B*elmouth Rtr-
Distortion Screen
Airlowlet-Stage Rotor--"
Figure 1. Fan Compressor Distortion Screen FD 16680Test Rig DII
In supersonic aircraft, such as the YF-12, F-ill and Mirage IIIF,the surge tolerance is related more to the maximum instantaneous distor-tion produced by the inlet than it is to the time-averaged distortion.If the distortion pattern that exists just prior to an in-flight surgeIs recorded, it can be reproduced statically by distortion screens.When this is done, a test engine which is run with these screens willsurge at the same operating conditions that accompanied the in-flightsurge. I
It is imperative that the instrumentation used in the development ofinlet-engine compatibility possess high response capability. Experiencehas shown that inlet flow perturbations exhibiting time-variant character-istics that change in a period less than 0.01 seconds are not uncommon insupersonic propulsion systems. Such non-steady distortion finds itssource in flow perturbations generated both externally and internallywith respect to the inlet. Distortion generated externally to the inlet,and subsequently ingested, may be caused by gusts, landing gear, or otheraircraft wakes. Internally generated distorrinn may be cue bv flue- Oft,
....... I * sock pvniLion, boundary layer separation from the -Isubsonic diffuser walls, instability in the inlet control system, andother related causes,
Figure 2 shows a time history of inlet pressure stability for an I1engine stall that occurred during high altitude inlet-engine developmenttesting at Mach 2-plus. The inlet pressure fluctuation that generatedthis en3ine stall was delineated by high-response instrumentation. IIDuring earlier tests time averaged distortion measurements failed toindicate the cause of surge. When high-response instrumentation wasincorpora.ed, a sharp increase in distortion was observed to precedesurges resulting from distortion.
Inasmuch as the instantaneous extremes of Inlet distartion are ofsignificant intc:rest in the development of a compatible inlet-engineiisystem, the test inlet instrumentation should have response times on theorder of one millisecond.
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[Volume III
1DALTA
PERCENTCHAGE 40
4030
s inle 0- -440"&*I-201- -
2D0-Pt 2 10
APPROX 60 7.0 0105TIME PRIOR TO STALL - sec
Figure 2. Tine History Stability of Inlet FD 17643
Pressure Stability for an Engine DII
Stall
3. High Compressor Simulation Tests
The necessity to simulate the high compressor during fan distortioniestIng is supported by analysis and testing. This work has shown thatthe high compressor has a stabilizing influence on the performance of
the fan while operating in a distorted flow field. This beneficialinfluence of the compressor is the result of the characteristic of thefront stages of the compressor wherein the stability of the flow increasesas the corrected speed approaches 100%. Because the corrected speed of
the high compressor remains above 80% at all conditions between takeoffand cruise, the airflow at a given N2./jq is nearly constant. Thiseffect induces a stable flow at the back end of the fan which, in turn,reduces the effect of inlet distortion. The reason is that the fan rootvelocity and flow are maintained by the pumnip effect ,f the cmpesn rteVLt thogh the total pre~surv is low.
It should be noted that high compressor effects on fan flow do notalter the total pressure field at the fan face. Rather, the high cor-pressor exerts a substantial influence on the fan root static pressurezone whose flow is destined for the high compressor by depressing thestatic pressure field in those regions corresponding to low total pressurezones in the fan flow exiting into the high compressor. The net resultof approximating the static pressure field strength to that of the totalpressure field is the tendency to maintain a constant velocity throughou;.
the flow field, despite the presence of non-uniform total pressure gra-dients. A uniform velocity flow field is beneficial. to fan performance
in that dasign levels of blade loading are not exceeded because low-velocity-induced rotor blade stall is avoided. This characteristic isdiscussed in detail in Report A, Section IlIA - Fan/Compressor Performance.
A circumferentially-variable back pressure capability will be incor-poLated in the engine annulus portion of the fan rig discharge and will
simulate the above characteristi of the high compressor by diverting high
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PtA yr 66-100Volume III
energy air flow into those engine flow zones of low energy. The degree Iiand circumferential extent of back-pressuring will be initially predicted,and later modified as required when testing of the rematched engine rig isinitiated for quantitative determination of high compressor-fan inter-action effects. Figure 3 shows a representation of this apparatus. TheLndividual groups of flow controllers are independently positioned at theexit of each duct partition such that the proper circumferential distribu-tion of the engine flow Mach number is attained within each segment of theengine duct.
Me sp ,t.IInletla. Ductrod Uw$lo. ~
A D Duct ISDut Partitio. / A Y
Typical Plances spitte.
Basin. .et SEngine Ductview Catnlio,(5ee sketch 11)
agn D View A V&ariah %asn Allow lnvdoen~evt
egine Dtuct Part:,if ~itionnl___ c . a of Each Pafltltoned
,. 4. 0.Sasele the u.m 0o Which
Inital an ompoenttesing illbe compished Basin.lo aDuctal
dta Bait Fla to
tsngine m -
Figure 3. Inlet Test Rig Schematic n D 16779 jrDII
4. 0.6 Scale Fan Rig Tests IT
Initial fan component testing will be accomplished using a 0.6 sgcalefan rig with distortion screens that duplicate the total pressure inletdistortion of the airframe manufacturer's inlet. in order to attain a-.ire realistic and meaningful test, the fan rig will include circumferen-1)t:.ally variable, back-pressure apparatus described in the preceding para-graph. Without the flow controllers in the engine duct, distortion test- Ijitsg of the 0.6 scale fan rig is useful, but in matching the predictedannular distribution of cowptesaor air flow the stabilizing effects of thehigh compressor prs .. e are simulated. II
The majority of testing to refine the analytical simulation programused for the prediction of distortion effects will be done using thisIirig. Distortion screen testing intended to support and/or modify analytic i
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theories will not be limited to usage of airframe inlet patterns, butwill also incorporate distortion screen-generated perturbations aimed1at the clear delineation of specific effects and their related causes.This testing is necessary to supply the calibration constants that willbe progressively incorporated in the analytic prediction theory.
5. Rematched Engine Tests
Early in the program, distortion screen testing using a modifiedengine will be accomplished. This engine will be modified by means ofrematching to allow operation near its surge line at wheelspeeds corres-ponding to cruise operation. The method employed to induce fan surgewill consist of progressively back-pressuring the au:- ;entor duct, drivingthe fan into surge. Closure of the augmentor nozzle provide.- this back-pressure capability (A description of this engine is presented in VolumeIII, Report E, Section III).
The test objectives of this rematched engine are to determine thesignificance of high compressor interactions with the fan. These testF results will be correlated with 0.6 scale fan rig distortion data, bothwith and without the latter's engine duct flow controllers. Necessaryadjustments to the 0.6 scale fan rig flow controllers will be made onthe basis of rematched engine test results, allowing further usage of
the less expensive fan rig.
6. 0.6 Scale Boilerplate Subsonic Diffuser and Fan Rig Tests
An additional step in inlet-engine component compatibility testingmakes use of a "boilerplate" subsonic diffuser as supplied by the air-Iframe contractor in conjunction with the 0.6 scale fan rig. This unitwill be constructed such as to duplicate the inlet subsonic diffuser.A properly-positioned normal shock and preceding oblique shock will beproduced at the entrance to the subsonic diffuser. The influence oi thebleeds, and to a degree, the shock-boundary layer interactions will pro-vide a partial simulation of inlet dynamics. The 0.6 scale fan rig usedwith the boilerplate inlet will permit an evaluation of the subsonicdiffuser performance under more vpp , . tis. operating Coaidi ci tian A
were poasi'uie during inlet model tests.
7. Full-Scale Fan Rig Tests
Following the 0.6 scale rig program will be full-scale fan distortiontesting with engine duct flow controllers. Those fan configurations thatshow sufficient gains in distortion tolerance can be directly transferredto the concurrent complete engine program for further tests. As pre-viously noted, the duplication of high-compressor interactions with the
fan by use of a fan rig with controlled engine duct flow should provideaccurate simulation of the engine fan characteristics.
Periodic cross checks will be made between results of this rig and theengine tests.
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8. Complete Engine Tests (Ref. Vol. III, Report E)
As a logical extension of the rig and modified engine test work de-scribed in the preceding section, the development of a compatible inlet- |engine combination will be pursued through full-scale testing. Totalduplication of the operating environment can be accomplished only by Jiactual flight test. However, very early in the development cycle,rigorous experimental and analytical investigation can be conductedsuch as to include the great majority of in-flight variables. These testswill lead up to the full-scale simulated flight Mach number inlet engine Icompatibility test program that is planned for the AEDC facilities at IITullahoma, Tennessee and continue into the 100-hour prototype flight testprogram.
.he primary objectives of full-scale engine testing include:
Correlation of results with rig disrorLion testing.Evaluation of engine transient response in the presence ofdistortion.Evaluation of the effects of distortion on:
Steady-state engine performanceTurbine irniet temp.,ature profilesDuct hater operationAltitude relight
Checks on the validity of the analytical simulation programsand updating as required.Establishment of inlet and engine control systems compatibility.Correction of gross inlet-engine problems before proceedingwith the AEDC test plan,
a. Testing with Distortion Screens
The airframe manufacturer will supply to P&WA the inlet distortionprofiles obtained during the airframe inlet model test program. Theseprofiles will depict both the time averaged distortion and the maximuminstantaneous distortion recorded for cakh of tho N-h nunbcrs and aLti-Lde. i.vegt....d. Dltt rtion generaturs will be used to reproducethese patterns during engine test so that ultimately the engine fanwill accommodate the distortion pattern of the actual airframe inlet.
Experience in developing engines to operate in harmony with super- Iisonic inlets has demnstrazed that blockage screens installed one or twoengine diameters ahead of the engine face in a properly constructed shortduct can be used successfully to subject an engine or compressor rig tothe distorted pressure field produced by an actual aircraft inlet, If,in the actual inlet duct, any bypass bleed doors, bends, etc., existwithin the above-mentioned distance from the engine face, they will beduplicated in the test stand representation. i
It is not sufficient to conaider only the overall time-averagc iA- - -i
tortLon. As was indicated in the preceding section, the distortion patcernproduced by an inlet ahead of an engine can be expected to vary with time.Experience has shown that if the distortion pattern that exists just prior
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CONFIDNTIL Pi att a vtimwy AircraftPWA FP 66-100
Volume IIIto a surge is recorded, and reproduced by thc above mentioned distortionscreens, the test engine will surge at the same operating conditions
Fthat accompanied the surge.
As discussed in Report A, the fan and compressor development will bedirected toward attenuating the distortion profile. First stage turbinevane leading edge thermocouples will be installed to measure the effect ofdistortion on the turbine inlet temperature pattern.
The program will begin with testing in the P&WA sea level test stands.The initial tests will be designed to permit a comparison of the engine'stakeoff performance obtained with a conventional bellmouth to tnat whichresults from running with the distortion screens described above. Theengine/inlet characteristics will be evaluated at all engine conditionsto demonstrate the ability to tolerate the distortion levels associated
with landing approach, and after takeoff, with the thrust cut-back forcommunity noise abatement. A demonstration of the JTFI7's ability tofunction without performarce loss or surge during these critical phasesof operation will help ensure the development of an airplane that is both[safe and quiet.
Following the sea level test, the distortion program will be shiftedto the P&WA altitude test facilities. The program to be run in thisfacility will permit a detailed determination of the effects of inletdistortion on overall engine performance, required control schedules,duct heater efficiency, lighting limits, augmentor-blow-out limits,windmill relight characteristics, and turbinke inlet temperature patterns.
In the event the distortion levels from the aircraft inlet mod'. testsexceed the limits presented in the engine model specification, the testprogram wi'l be aimed toward an evaluation of the effects of this distor-tion. Modifications will be proposed either to improve the distortionpatterns supplied by the airframe mnufacturer or to accommodate theinlet distortion in the engine. These modifications will be evaluatedin a coordinated program with the airframe manufacturer. This wil nurCite development of a pripulfiun package that provides optimum performanceand st.ability in a minimum amount of time and with minimum cost.
The simulated altitude program will consist primarily of steady statecalibrations plus normal augmentor transientL. This will provide not onlya demonstration of satisfactory engine performance but will verify theJTF17 insensitivity to reasonable inlet distortion during augmentor lights,
S-throttle excursions with the augmentor lit and augmentor shuto-f transients.
b. Testing with Simulated Inlet
(1) Boilerplate Subsoniic Diffuser Tests
To more nearly simulate in-flight inlet dynamics, a boilerplatofacsimille of the subsonic diffuser portion of the inlet duct (schemat-i.ally illustrated in figure 4) will be installed ahead of the engine in
the P&WA altitude test facility. This facsimile will be designed toproduce a representation of the last oblique shock and the terminalnormal shock associated with the actual inlet. The duct representation
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will intclude the inlet boundary layer bleed and bypass systems _ad their Iiassociated controls. The program to be run with this configuration willbe primarily directed toward an assessment of the effects of inlet andengine dynamics on propulsion system compatibility. The test objectiveswill include an evaluation of the effects of terminal shock location,turbulence, and bypass flow variations.
Modified Centerbody for !1Subsonic Diffuser Testing atFROC Pert Bleed Extraction Overboard Bypass(,w nd (."an r.__ody) Bleed Extract ionI
AirframeNorm! Shock Fan
*Used to Attain the Correct Mach No. Ahead of the Normal Shock
Figure 4. Inlet Duct Subsonic Diffuser FD 17647 U1Schematic DII
Within the limits imposed by the facility, tests will be conducted Vto determine the inlet duct response to engine transients and engineresponse to the inlet. These will include normal duct heater lights andshutoffs. engine acceleration and deliberately provoked airflow excur- Iisions beyond the normal tolerance limits.
In addition to evaluating the effects of engine transients as describedabove, the program will also include an investigation of the effects ofa|
bypass bleed valve excursions, variations in boundary layer bleed extrac-tion, etc.
Artificially induced flow disturbances will be used to permit anevaluation of the sensitivity of the propulsion system to inlet boundarylayer separation.
(2) Boilerplate Full Inlet Duct Tests
A boilerplate representation of the full inlet duct will be testedfor compatibility on a JTF17 engine on a P&WA sea level test stand.These tests will be designed to investigate the effect of distortionon transient operating characteristics of the engine. This will be 91accomplished by subjecting the engine to a series of increasing ratesof acceleration and deceleration.
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A proper evaluation of the efects of distortion on engine transient
characteristics cannot be limited solely to gas generator transients but
will include augmenter transients and in investigation of the effects
of inlet pressure loss on starting characteristics. This phase of the
test pi:ogram will verify the engine's tolerance to distortion during
typici.1 duct heater light and shutoff transients. In addition to pro-
Sviding high efficiency and good stability during steady state operation,
the augmentor and control systems are specifically designed to avoid tan
surge during transients. A series of starting tests will also be con-ducted to define the optimum fuel control schedule with realistic inlet
duct loss and distortion,
An important part of the noise abatement plan is the use of a chokedor near-choked inlet duct to reduce forward propagated rotor noiseduring landing approach. A series of tests will be run in the P&WA
sea level test facility to ensure freedom from any problems during this
mode of operation.
9. Instrumentation
The instrumentation needed for inlet-engine compatibility is divided
into two categories, static instrumentation for quasi-steady engine
operation and dynamic instrumentation for time variant processes that
require a high degree of definition. Both types of instrumentation are
useful in the isolation of component interaction effects, these effects
being obtained by comparison of the component performance within the
system to individual test performance. In addition to the instrumenta-
tion required to determine overall engine performance, special instru-mentation will be installed to permit evaluation of the effects of dis-
tortion on turbine inlet temperature profiles, duct heater discharge
temperature and pressure profiles, duct heater efficiency, etc. Dynamic
instrumentation with response times of less than one millisecond will
allow accurate definition of transients during the control compatibility
and distortion testing.
10. Compatibility Tests in AEDC Facilities
The test program described above will lead up to compatibility
tests of the complete propulsion package that are scheduled to be run
in the AFDC propulsion wind tunnel in Tullahoma, Tennessee. The details
: of this program are covered in the engine-inlet compatibility test plans
that have been coordinated between Pratt & Whitney Aircraft and the two
airframe manufacturers. The Pratt & Whitney Aircraft test program is
designed to reduce the number of problems that may be encountered during
the complete propulsion package compatibility evaluation at AEDC. The
time spent in the 16 foot wind tunnel at Tullahoma may then be devoted
exclusively to evaluation of those facets of the propulsion system com-patibility investigation that can only be handled with re'istic external
flow.
ii. Flight Tests
The ultimate test of the engine/airframe compatibility will begin with
the 100-hour flight test of the prototype supersonic transport. The
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propulsion system performance will be monitored by instrumentation and 1
recording devices. Engine/inlet compatibility and performance and theoperation of inlet and engine control systems will be evaluated invarious steady state and transient flight conditions throughout theflight envelope. In addition, extreme and abnormal flight conditionswill be simulated in flight. Pratt & Whitney Aircraft will prepare apropulsion system flight test program in coordination with the airframemanufacturer and will provide full support for the flight tests through-out the program.
12. Cross-Check with Analytical Simulation I;Throughout the test program described above, a continuous cross- ,J
check between engine test results and the analytical simulations described 9!in the following section will be maintained. This will permit updatingof the various simulation system gains and time constants and will ensurethat the desired stability and performance requirements are met. Further- Imore, this effort will permit an incorporation of the effects of inletdistortion and representation of certain failure modcs into the simula-tions.
D. ANALYTICAL SIMULATION
1. Introduction
Engine/inlet compatibility over the entire range of operation isimportant to the design and development of a supersonic propulsion system.The interactions between the engine and inlet are so many and complexthat it would be hopeless to develop the system with a seaL-of-the-pants,cut-and-try approach. Computer methods have been developed that allow idynamic; simulation of inlet, engine and control systems. Use of these f#simulations permits the choice of proper control modes in the designstage rather than after flight test. Simulations continue to be valuablediagnostic tools throughout the development and flight test program,when measured component performance can be substituted for the originally
useful in simulating flight test problems and exploring various means ofsolution through control system changes and adjustments prior to actual..flight testing of such changes.
Pratt & Whitney Aircraft has made extensive use of dynamic simula- iition to design and imprcve its engines. For example, during the designand development of the RLIO, liquid hydrogen fueled rocket engine, dynamicsimulations of the system required approximately 13,700 computer hours.Solutions to pump stall and control instability problems were found thatwere then confirited by hardware tests, leading to a record of 48 enginesflight tested without a failure. Similarly, for the J58 high Mach numberturbojet engine, dynamic simulations of the engine/inlet system conducted I!in cooperation with the airframe manufacturer early in the program,involving 8430 cmputr hours, led to Inlet control and. engine controlmodifications before expensive hardware was committed. As a result,compatibility between the engine control system and the inlet controlhas presented no serious problems in the J58 installations.
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Following this precedent, dynamic simulations of the JTFl7 engine andcontrol systems with the airframe inlet have been prepared and studied forapproximately 2900 hours to aid the design and development of these systems.As rig and engine test data become available they will be used to verifythe simulation and contribute to further improvement of performance. DuringPhase III and the remainder of Phase II, gains and time constants of theengine and control system will be continuously updated as required. (Referto Report BIIIC and D and Report El and II for complete discussion ofengine and control test plans.)
2. Analytical Technique and Simulation
a. Technique
The objective of a dynamic simulation dictates the analytical technique..ngine/inlet compatibility simulation techniques range from the use of
several simple test signal. type simulations to IBM 360 computer simtula-tions. These simulations compliment each other. Simple simuldati ons providerapid parametric analysis of quasi steady state operation. Complexsimulations provide detailed analysis of large scale engine/inlet traoelnnts.
'The dynamic analysis of the JTF17 engine/inlet propulsion systtm requiresboth simple and complex simulations. Simple simulations have been orogrnrmledon digital and analog computels. A complex detailed simulation has beenprogrammed on the IBM 360 digital computer. These programs have been closelycoordinated with both airframe manufacLurers to ensure the integrity of boththe steady state and dynamic characteristics of the engine/inlet simu)A tlots.
As shown in Figure 5, typical engine operation at a given flight conditionconsists of a constant airflow regime and a non-constant airflow regime.Aircraft operation, except for descent, is experted to b ifn tHU duct heaterlit constant airflow regime. A complex highly detailed simulation is ttsedfor the study of transients involving large variations in airflow, flightcondition, and for failure analysis.
b. nimuiotjn
(1) Digital
Digital Computers are utilized for dynamic simulations by includingthe proper differential equations necessary for rotor inertia, mass storage,and control dynamics. Transients are calculated by computer cycling throughthe simulation for finite time increments. This time increment is selectedto be compatible with the simulation dynamic characteristics. To illustrate,a simple first order lag is programed as follows:
In Laplace Notation Output IInput ys + I
The calculation can be unstable if the time increment for calculation isnot compatible with the time conitant of the system.
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The desired result is shown in figure 7.
The above plot demonstrates a stabitt:, criterion necessary in digital I
programming of dynamic 61niut6ions. The time inczrement used in the calculationshculd be equal tco or less than tl)e smallest tir,, constant fi the system. Theexact solutiov is approachetd by decreasi.ag the size of the time increment.
irio - . C-..onstant Airflow S-hjuin- -- 1I Schwlag
LEIGEND-
S / 0 Airflow Idle/ Maximum Rating - Duct H3ater Not Lit
0 /Minimum - Duct Heater Lit0 Maximum Rating - Duct Heater LitU Vi
Idle Maximum Rating Maximum RatingDuct Heater Not Lit LVaet Heater Ut
POWER LEVER ANGLE, PLAj
Figure 5. Typical Engine Operation FD 16695
DII
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S - Laplace Operator..
+ Integrator
Input - Output I
Feedback
Figlure 6. Schematic Representation of a FD 1669iSimple First C dec Lag DII
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The Dsired Result is:
1r m Time Constant
ri Initial Slope
CLt '-- Exact Solution z Input (I-e- /Yr
TIME
Unstbl\
Possible Digital Soluti *ns T= 2r
Iz
--0.5T"
TIME
Fi&.re 7. Stability Criterion for Digital FD 16692Programming of Dynamic Simulations DII
(a) IBM 1620 Simulations
Simple inlet simulations received from both airframe manufacturers have
been programmed for checkout and study on the IBM 1620 computer. Fi-re 8Aiw a ilock diagram and a schematic of the Boeing Aircraft inlet simulation.Figure3 9 represents the duct dynamics and bypass control system of theLockheed Aircraft inlet simulation. The use of these simulations is dis-
cussed in paragraph 3, Analytical Studies. These simulations represent
[ early Qstimates ox the dynamics of the two inlets. The response and
dynamics of both inlets are being improved through continued airframedevelcpmex.t and test. Revisions to all computer simulations will beincorvorated as they become available.
(b) IBM 360 Simulations
(1) Engine/Inlet Dynamic Siuulations
A compiex, highly detailed engine and control simulation has beenprepared for study on a.. 11W 360 type digital computer. An interface
deck has been written to enable coupling of the engine/control deck withan inlet simulation. The program can operate with or without an inletsimulation. This engine/control simulation has been sent to the airframemanufacturers for their engine/inlet compatibilit. studies. The simulation
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is very versatile. All power lever transients from idle to maximum duct [heat are availalble, The simulation operates at all flight conditionswithin the flight envelope. Flight condition and/or ambient conditioncan be ,hanged as c function of time. Failure modes such as windmilling [and blowout are av,-ailable.
Engine/inlet transients are simulated by computer cycling beginningat the aircraft inlet and progressing through the engine control andengine component calculations to the exhaust system. Figure 10 is asimplified block diagram that depicts the flow of information that occursit. the engine/inlet system. Figures 11 through 14 show detailed block |diagrams of the engine control as progreimned for this simulation. Inletsimulations from the Boeing Company and Lockheed California Companyhave been programned for JTF17 engine/inlet compatibility studies with theJTF17 engine and control. (See figures 8 and 9.) These simulationsrepresent early estimates of the two inlets and ate dynamically identicalto the slmulations discussed above. j
11C,_. __
------. l-...- - -- [- - -. IAi
Figure 8. Boeing Simplified Ilet Block FD 16700Diagram and Schematic DII
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Figure 16. Lockheed Ine. utiynri..nlet FD) L6701EneControl se DII
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Figure 11. Gas Generator Control Function FD 14758A 1.Diagram DII
!,
Fiur 2. Exas ozeAe o-r ucin F 67
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:Figure 12. Exhaust Nozzle Area Control Function FD 16777 1-Diagramii
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-u I
I Figure 13. Augmentation Fuel Flow Control FD 16748Function Diagram DII
imp
Figur'e S cnedi.le Servo and Anci!tary Valves FD 14770 ..
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(2) Stability Analysis
An TIM 360 calculation routin, is used for stability estimates. Toanalyze system stability a mathematical model cf the dynamic propertiesof the system Is constructed in the form of a block diagraw. Each blockin the diagram contains a mathematical expression (ir Laplace notatio,)which repres-ents the dynamic characteristics of a system component. The blocks are connected by lines representing the flow of interact Iou betweevcomponents. The blocks are combined in a manner similar 'to combination ctseriec and parallel electric components until the system ;s represented byone mathematical xpression in one block with oe input and one o putd.I.This single block transforms the input into the output and is called thesystem transfer function. Thte transfer function facilitates the frequencyresponse analysis of system stability. Using the transfer functicn, the Vsystem output in response to a sinusoldal input can be determined. Themagnitude and phase oi the output wave relative to the input wave arecalculated for a wide range of input frequencies. This output/inputrelationship is compared with established, mathematically derivedcriteria for divergence and limit cycle instability to determine if asystem will be stable.
The IBM 360 calculation routine performs the following functions:(1) makes the detailed algebraic comptations of reducing the blockdiagram to transfer function form, (2) calculates the magnitude andphase of the output relative to the input, for a given frequency input,(3) automatically plots frequency response results in any of severalformats, (4) and in addition calculates parameters of interest includinggin and phase margins and error coefficients.
(2) Analog Simulations I
The analog computer is used to study inlet/engine compatibility anddynamic response when simple rimulations can be used (cruise mode ofoperation is an cxanple). The inherent desirable characteristic of ananalog comput.er is that it simulates the Dtimary syste! in a c..taiu...raaihIet um opposed to the discrete time interval simulation of a digitalcomputer. This provides excellent response characteristics. As anexample, coesidcr the first order lag described abeve for programmingon a digital computer. The analog ,-omputer gives the exact solution asshowz in figure 7.
Simple engine/control simulations have been generated and have beenSnt to 0..: airframe manufacturers for thema engine/inlet compatibilitystudles. Figare 15 illstrates a simple engine/control simulation I;suitable for pewet lever transients from maximum rated duct heater not
"'A lit to maximum thrust duct heater lit at all inlet started flightconditions. Figure 16 illustrates a simlla, dyvo;ic simulation for partpower engine/Inlez compatibility studies at the supersonic cruise flightcondition. Figure 17 Is a schematic of a simple engir.e/inlet system usedfor aalog studies. These simulations permit rapid acueasment of currentand proposed engine/inlet operation.
A..-
DII-20
GOE ENTlALF
I CWONFDETAL Praft & Whitney AircraftPWA FP 66-100[ Volume III
[ Figure 15. Engine/Control Simulation for Power FD 15581Lever Transients DIJ
[t
I~~ Ain ~ ~
Lm
L[t -.
, 7
Fjg~,jre 16. -foe/otil ~ 'r Part FD 1c"82I.Power Engine/Inlet Compatibility DII
I DII-21
CONFID[NTAL
Pratt & WNhftney R~crf wCuwwiufu-ui
PWA FP 66-100Volume III
X I f
Z~i .< .
Figure 17. Engine/Inlet System for Analog FD 1b703Studies DII
3. Analytical Studies
a. Response and Stability I|a.
Tb excellent engine control frequency response characteristics andscheduling tolerances that exist over the entire JTF17 operation envelopefurther improve engine/inlet compatibility. Using the extensive experiencegained by the J58 and other supersonic engine control programs, the JTFI?control, Report B-ITl, provides excellent scheduling tolerances andstability r- ...... th cccd. Lop veiilier control, i!gh frequencyresponse characteristics are provided with open loop basic control.
This combination of basic control and vernier control, used byearlier supersonic engine controls and now further improved by theJTF17 control, decreases the level of engine induced disturbances uponthe inlet.
Figure 18 illustrates the JTF17 control concept used to combine theadvantages of both basic control and vernier control. JTF17 turbinetemperature control is similar to the J58 control. The best features of [the J58 and other supersonic engine control systems are incorporated inthe JTF17 control. Figure 19 is a simple analytical model of the JTF17control concept. A relatively slow highly accurate closed loop vernier |control is used to adjusta high response reliaule open loop basic control.Exceptional control accuracy, response, and reliability are the result.
Engine response and stability is further improved by JTFI7 control [gain schedules that are similar to J58 gain schedules. These schedulesprovide overall gain compensation. Gain compensation maintains thesame overall duct aiea control sensitivity. JTFI7 control gain schedules
DII-22
CONDENMAL i
CONFIDENTIAL Pra Whftney 1rcmftPWA FP 66-100
Volume II1
achieve this objective by varying control gain inversely with duct heatergain. Gain compensation iml'roves ergine/inlet compatibility by maintainingJTFI7 response and stability WfiihouL .Jevintion from nominal design levels.
' ITJYIT -d Ju TI.. T- C- j., -.
Figure 18. JTF17 Control Concept FD 16704DII
BLOCK D1AGMiA
i k~____ _Verni er Gat LOI
B..ic C~ntroII
a
EQUIVALENT CIRCUIT "r.A10SFER FJNCTION
Input Output T: t ;
.Al.r.:
S L LaPta* OpOrCAoK T :u 3.0IL Ve ier C0.,trol G. T
T - G C.o rI C Ti C. -tt
Figure 19. JTFI7 Control Concept Analytical FD 16705Model DII
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Pra & V fney Alrcfft CONFDENTALPWA FP 66-100Volume. Ill
b. )ostream Transient Feedback Effects
Studies have been made to minimize downstream transicnt effects on
the tan nurge margin and inlet shock position. Protecion is providedagainst adverse engine/inlet interaction due to duct hearer light or
rapid power lever movement.
One method of minimizing these effects is to minimize the fuel/airratio required for ignition of the duct heater. During Phase Il-C the
minimum duct heater fuel/air ratio has been reduced from 0.008 to 0.002.This reduction came about when the duct heater configuration was changedfrom an aerodynamic flame holder type to a ram induction type. The raminduction configuration is characterized by very soft, smooth ignitionat fuel/air ratios of 0.002 ol less. Figures 20 and 21 are traces of
data taken from oscillograph records of Zone I ignition at a 0.002 fuel/airratio setting in the full-scale annular duct heater rig at conditions
of pressure, temperature, and airflow corresponding to sea level staticand cruise respectively- The pressure pip for the cruise light was g
3.6% of the absolute pressure level and for the sea level light was 3%of the absolute level. These pressure increases were measured at the
duct hearer station. These pressure pulses will attenuate before reachingthe fan discharge station. (See Engine Performance, Duct Heater, Sec- I:
tions I1-A., Ill-B, Ill-C, and III-D for a complete description of theduct hater development.) Figures 22 and 23 show calculations from thedigital simulation for duct heater light at sea level static and cruisewith light-off fuel/air ratio of 0.002. Figures 24 and 25 are simulation
traces of a 0.008 fuel/air ratio light-off with a ram induction ductheater at sea level static and at cruise. As shown by the traces, a
reduction of light-off fuel/air ratio reduces tne effect on surge marginand the airflow error signal. This low ignition fel/air ratio alsosignificantly reduces the thrust discontinuity at light-off.
3n 1.2 psi
_______ pal____ BurrPra _r__,___ [.__Burner Outlet Temperature 270F !
Spark Ind -ator0 0.6 1.0 1.5 2_0 2.5 3.0
TIME-sc
Figure 20. Sea Level Takeoff F/A 0.002 Full FD 16696
Scale Annular Duct Heater Rig DI
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CONFIDEN TAL
( CONFIDENIAL Pratt Whitney .,rcrattPWA FP 66-100
Volume 111
255pi Burner Preitre.
' -
Figure 21. Full Scale Annular Duct Heater FD 16697Rig Crlse Ignition (65K) DI
FAN SURGEMARGIN -T -
ORRECTEI3TOTAL ENGINE~ 1021
AIRFLOW 10
IL
104 -......THRUST. -9 102,
100
CONTROL POW ER 85 _ _ _ _ _ __ _ _ _ _ _ _
LEVER ANGLE -deg
AIRC OWT POWER riLEVER ANGLE-deg
0 11o 2.6 3.0TIME set
Figure 22. Maximum Nonaugmented to Minimn FD 16706Augmentation Currert Duct Fuel/Air DIIRatio
G II- 25" GCONFIDENTIAL
-. 7
&WhftneyRrcat CON HDETIALPWA FP 66-100Volume III
IILOCrWID COW W ll
3,7 'INLET SHOCK .POSITION -f
CORRECTED TOT AL1ENG!NE AIRFLOW |,2
TH1RUSTl
FAN SURGE 40 -MARGIN .
CONTROL POWERLEVER ANGLE - deg
82IAIRCRAFT POWER
LEVER ANGLE - deg
01.0 2.0 3.0 T
Figure 23. JTF17 Cruise Maximum Nonaugmented FD 1671,&to Minimum Augmentation Current Duct DII
Fuel/Air Ratio
CORRECTED 102 -
TOTAL ENGINE 10 (-
AIRFLOW, Waft 2 / t 2 - IFAN SURGE 20.0MARGIN -% 10.0
120I11,5
110 aTHRUST - % 105 -
100
CONTROL POWER sLEVER ANGLE - d- 75-
AIRCRAFT POWER 75LEVER ANGLE - deg 70
0 1.0 2.0 3.0TIM -sec
Figure 24. JTF17 Sea Level Static Phase I1-B FD 16713
Minimum Duct Heater Lightoff DII
(F/A = 0,008)
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CONFIDENTIAL Pratt & Whitney AircraftPWA FP b6-100
i Volume III
LOCKHI[ED COMFPTITIVI ibilA
INLT S10c:K "35POSITION - fauj
+ (1t)l2(1i2( HFFI) ~TOTAIl I NuIN h ii)--A JI? V!.,, W q+[
I~W isL (2 --
I . ++ '
201 V
( h d~8 71.|
LEVER AN F
- deg 72I. 2,11 3)
TIME - se'
Figure 25. JTF17 Cruise Phase I-B Mi-imum Duct FT) 16715
Heater lightoff (F/A 0.008) DII
Another method of minim zing the pressure ratio and airflow variation
Is; to correlate duct nozzle area and fuel flow. Ideally, or with perfectcorrelation, there would be no effect on fan pre. sure rqtin or airflow &.e-o power ChanE ill ti, dc,,t h -t g lude. Figure 20 illustrates two
correlation schemes, the iF30 and the JTF17. In the TF30 scheme, onlyone zone of the five separate burning zones is shown. Power lever angleschedules gross duct nozzle area which then schedules duct fuel flow withthe correlation cam. A proportional plus integral controller makes
vernier adjustments to the system. The correlation cam scheme for theJTFI7 is an improvement over these earlier configurations. The nozzlearea feedback scheduling of fuel has been eliminated. Duct nozzle areaand fuel flow are scheduled directly with power lever angle and nozzle
area is adjusted by a proporLional plus integral controller to eliminateany airflow errors. Since perfect correlation is not possible with any
I real ;ystem, further steps are necessary to insure protection.I The JTF17 control definition provides nonlinear control. characteristics
that maintain inlet margin and fan surge margin equal to or oreeter thansteady state margins. Figure 27 illustrates typical nonlinear gain and
response circuits. Scheduling; increased inlet bypass door area andincreased dus-t nozzle area increases inlet unstart margin and fan surgemargin. figures 28 and 29 show the desired inlet and engine control
action during transients, If a limit cycle of either duct fuel or area
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I CONFIDENTIAL
Pratt& gW ,Itney ,rv-", t CONFIDETI~ALPWA FP 66-100Volume III
occurs, the fan surge margin is increased due to nonlinear duct hearergain and response characte-istics. A nonlinear gain i. used with the ductcorrected airflow error signal such that increases in duct airflow occurrapidly and decreases in duct airflow occur slowly. A similar nonlineargain is used with the duct area qlide valve. A nonlinear response Is ucedwith the duct hearer fuel cirruit such that increases in duct fuel occurslowly and oieceases In duct fuel occur rapidly. The control furtherimrrr,,es engine and inlet transient margins by rese,ting duct correc-edairflow upon initiation of duct heating. Figure 30 illustrates Lhe desiredeffect of ducc nozzle area reset during power !ever angle modulation inithe duct heating mode. TransinL surge margins are maintained greater
than steady state surge margins. Figure 31 is a digital simulation traceof a power lever angle transient from maximum nonaugmented to maximum ductheat at cruise with neither inlet nor duct: area reset. Figure 32 shows r.esame transient with duct area reset. Inlet unstart margin and fan surgemargin are maintained at safe margins during the trausient.
F~llJTFI7 Scheme
L_~
TF30 Scheme
I- .
Figure 26. Duct Heater Control FD 16698
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CONFIOENTIAL
I COWFONJA Pratt 9 Whitney (AIrcrafPWA P 66- 100
Vo.une II
KOI,.II Aa GATh KfIIwIyIrAt PJZ5++uI
J ,~
CI~a-k valve
Figui-e 21. Typical Nonlinear Circuits FD 15547
DII
1.=-
WITHOUT NONLINEAR WITH NONLINEAR
RESPONSE AND GAIN RESPONSE AND GAINAcceleratii
re i /////"
1.4 De leratio , nn
Unsar U litart/Beglegio
INLET BYPASS DOOR AREA INLET BYPASS DOOR AREA
Figure 28. Inlet Co'trol Action ID 15545B
DII
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Pratt& Whitney Alrcraft CONFIDENTIALPWA FP 66-100Volume III
WITI !OUT NONLIN EAR WITH NONLINEAR;ESPONSE AND GAIN JONS E AN[ GAIN
Fan. Surg, Ara/ Fan urt e Anre
' A c' , cr t io n
A c e l r a t io n E -a n d '
r L lDeceeration
DUCT JET AREA,AJD DUCT JET ABA,AdD _
Figure 29, Duct Control Action FD 155A6
.. .. . Heat 1uc7...-
;Il x'll aKted P -
= 14 . 17 1 1 iI
.._...L
- la
12./ Nk
t ~ _ a Aucteaerio E-
Figure 30. Analog Simulation Trace
FD 16750
DII
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CONFIDENTIAL,- Pratt 8 Whitney AircraftPWA FP 66-100
Volume III
x ;
rQ CU
,x -
[J4
0 .j,.0
'.
j in
\ (
DI -3! CONFIDENTIAL
. . . t • •
I
Pratt & Whitney Aircraft CONFIDENTIALPWA FP 66-100
Volume III
Studies by the Lockheed and Boeing Companies show that inlet reset
is not necessary and none is provided. A disadvantage inherent withinlet reset is the increase in distortion that accompanies supercritical
inlet operation. If the inlet can be designed to operate safely withoutreset, this gould, of course, eliminate the transient increase in dis=tortion due to reset.
Figure 33 is a trace obtained from an early lockheed inlet simulationwith a severe ramp in engile airflow demand. The inlet simulation did
not disgorge the shock and handled the airflow transient very well,Figure 34 is a s~mulation trace of a power lever transient from maximum
rated to idle at cruise. Safe inlet unstart margins are maintained forthis large transient in airflow.
100 y
~ 90
80L02 zC 40-
20
0 -
4 0 K"-- .
0 :3: 1312 V=________________
C6 1.4
Q C
0 0.4 0.8 1.2 1.6 2.0TIME sec
Figure 33. Inlet Response to Change in Corrected F 16745
Flow -
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CONFIDENTIAL
CONFIDENTIAL Pratt & Whitney AircraftPWA FP 66-100
Volume III
LOCKHE|W COMPITiiMVI DiAINLET ,SIbOCK [,4 ,,
POSITION- ft 3.2
2 .$
CORRECTED O'A. ,,nENGINE AIRFLOW.
W5 ;#2/t2 4
80
60THRITS'" %1 40
20
-20
AIII(RAIFh JOWI'At 7lAIKII ANGI.I" - rg
1 1 2J A.0 I i ' . (0 7.
TIME -
Figure 34. JTF17 Cruise Maxi mum Nouaugrmnted YD 167]8to idle DI I
By means of duct fuel --area correlation and nonlinear response and v ain,
the ITF17 control achieves tolerance to acheduling and dynamic : ismatch.
Further tolerance is achieved by coot rollinlg the slew velocity of the
power lever angle input servo for the duct hootin g mDOdc. Tls' h n cno-vja'i' efecr o1 slowiny, down ail power lever transients above maximum
rated. The effect can be 3een by comparing power lever angle, PLA, tocontrol servo position power lever angle, PLAL. The sle,w rate limit in
the control tolerates a larg;e amount of mismatch in the correlation cam
and circuit dynamics. Figure 35 illustrates this tolerance on a naximumrated-maximum duct heat-maximum rated transient. Three analog compuersimulation traces are shown: a base case, a +10% duct fuel flow error
case, and a case with duct nozzle area rate decreased to a fourth of
the base case value. The system tolerated both cases of large mismatchand maintained safe margins.
c. Pressure, Temperature, Airflow Transient Effects
The engine/inlet system is relatively insensitive to large ai-flow
disturbances. Figure 34 shows data, calculated with the digital simulation
of a power lever angle transient from maximum rated to idle at cruise.
This is a normal transient, that is, one that should occur at least once
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Pratt &Wthltney Aircraft CONFIDENTIALPWA FF 66-100Volume III
on every flight. The system compensated for the large decrease in en .-_
airflow without disgorging the shock from the inlet. Figure 36 shows thetolerance of the system to another abnormal airflow transient, In-flightshutdown at cruise. The engine/inlet system retained the shock for thi
severe airflow transient.
I .
24
!AJ :1.2INIET .SHOCK :3j)
6
24
C TE,,I )I'AL &T l llEN(IINF. 0-.
W8 ,9t2/8t2 - 7OL "-
o601
40HTHRUST 20
-2 0 - " -
-40-
POWERI "VEf(aGL:. d0 .0 20 3-.11 "Co 50 611 7.0 NIP 9.0 I1
Figure 36. JTF17 Cruisc Maximmn riagmen ed FD 16719Fuel Shutoff Valve Turned Off - DIIEngine Windmills
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CONFIDENTIAL
CONFDENTIAL Pratt &Whitney AircraftPWA FP 66-l00
Volume Ill,
The most severe pressure transient the engine will feel occurs if the
inlet shock disgorges. The dynamics of the unstarted inlet are not avail-able but an estimate of the total pressure variation of the engine face
* during unstart and restart has been made. Figure 37 shows the effects ofthis pressure perturbation on the engine. The sequence of events is a.follows: (1) the inlet total pressure is ramped down and decreases thecorrected airflow, (2) the duct nozzle area opens to correct the negativeairflow error, (3) duct nozzle area modulates to correct the airflow error.At the beginning of this transient the engine feels a momentary high over-temperature and decrease of surge margin on the fan and high compressor.This is a normal occurrence on a high Mach number engine diring an inletunstart. Past experience on the J58 engine has shown that this momentarytransient has n(, adverse effect on the integrity of the engine. The
transient occurs too quickly for the hardware to feel significant over-temperature; see plot of the turbine leading edge temperature, figure 37.A possible source of trouble is the minimum fuel flow limit. If this limit
is set too high, serious engine over-temperature might occur with the low
unstarted recovery. The JTFl7 fuel flow limit offers safe operation withan unstarted inlet at cruise altitudes at Mach 2.7.
( SOING COMPIYIIIVI GAlA
RAM RuF'VEtY~1 tarn 60
10
TUFF3INE VANE -iirr.LEADING FI)E 2TEMPERATIEINCREASE - F
DU(}T NOZZLE (; A,.-.4 51
Ii
TEMI'ERATI!,E -
21h1'
110 ________
CORIECTE D'TOTAL ENGINE 74
.30
TH1RUST 40fi
Ir 20to fr) .0 .2- i 3 - 5., -"
TIME -r
Figure 37. JTFlI Cruise Inl(t Unstart and FD 16729
Restart DIl
DJ T- 3'?
CONFIDENTIAL
Pratt &Whitney Aircraft CONFIDENTIALPWA YP 66-100Volume III
4. Phase il - Analytical Simulation Test Plan
A closely coordinated effort will be maintained with the airframe
manufacturer to make maximum use of all available data affecting thecompatibility of the engine and inlet with respect to both steady stateand transient operation. Analog and digital simulations of both inletand engine systems will be employed. Revised inlet steady state anddynamic characteristics as obtained from the airframe manufacturer willbe included in the computer program as provided.
The analog and digital simulations will be updated a- additionalperformance characteristics become available. Modifications to thesesimulations will he made to effect satisfactory solutions. Off designvariations studies will be conducted to further refine engine/inletcomparibility estimates. These revised and updated simulations willprovide the nece.w-lury information for the more specific and detailedstudies requirzd by the development program.
Transient eitine/inlet operation studies will continue with increasedemphasis toward physical control hardware. Engine and rig test data willbe used to insure the validity of the simulations. This will be a closelycoordinated effort by both the engine and the airframe manufacturers.As new inlet model test data is obtained, the inlet simulation will berevised to include significant changes in dynamics or concept. Enginetest program data will be used t, check the validity of the enginesimulation gains and time constants. See Report E-I and TI for a completediscussion of the engine test plans. The envine control test plan,(see Peport B-Ii-C), will provide data to check the control simulation. IIRefer to table 1 for proposed milestone time chart of inlet/enginecompatibility testing. The computer simulation can then be used concurrentlyduring the test programs to ensure that desired stability and performancerequirements are met. This will begin a continuing cycle of test,refinement of simulation, analytical studies and further tests until aconfiguration suitable for the AEDC compatibility 1r-tst is: cv.lvcd. Thf. Uprogram will eveilate phenomena such as the effect of variable gains,
dcadband, random noise, interaction between control components for theeffects on the overall system.
The dynamic simulation is being enlarged to include inlet flow distortioneffecte and certain failure mode effects on the system. An imposed inletflow distortion will always tend to degrade the flow stability and performance 9charact-.ritIcs of fans and multi-stage axial flow compressors. Methodsfor exact analytical solutions for the processes by which a distorted flowaffects the stability and performance characteristics are being studied. IUn,:il these methods are perfected, the effects of flow distortion determined
from rig test results will be programmed into a simulatlon. This simulatiolcan be used to determine safe system operating lilits, performance trade Iifactors, and general tolerance to distortion over a broad range of system 3
operating condit ons.
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.... .- ... <-£ II
CU IOEWI Pratt 8, Whitney AilrcratPWA FP 66-100
j Volume III
E. DISTORTION PREDICTION DEVELOPMENT
1. Empirical Distortion Factors
Empirical distortion correlation factors have been previously usedto predict inlet distortion tolerance. Their popularity is due to theirrelative simplicity rather than to their complete effectiveness. Sincethey are founded upon test data they have the inherent capability ofproducing reasonable predictions of distortion tolerance for conditionsthat are close to the correlated data. hcwever, when differences in thedistortion pattern occur or changes in the compre.sor (sometimes quitesmall) are made, these correlations are usually inadequate. As an example,a partia. listing of the effectiveness of one typical distortion factoris shown in table 3. Figure 39 presents a number of presently used simi-lar statistical distortion factors, As an example of the unreliabilityof presently used statistical correlation methods to predict distortioneffects, figure 38 presents the application of these distortion factorsto the data shown on the performance maps of two very similar compressorswith both undistorted and distorted inlets.
11 Compressor A UNDIS'FORTE D1 0 . . ... .1. K d =0.0
o 2. Kd = 0.010 -.... a. Rd = 0.04. W%,; = 0.0
,' 90 - -. _ --. , 1 i -I. 5. R .=0.0ZI6. KdJt0
U~80 _4 ' DISfOfl'I'ED1. Rd = 0.022. Kd = 1a8.O
701i 3. Kd=143.0z4. Kd' 0.03
U 60 1 ] 7'' 5. Rd = 0.064
T'1,, 6. Kd =2.26COnrTI;EnCIV-1, AIRFLW -%
Compressor B0 100 UNDIST_-ORTED)P J 1. Kd = 00
2. Kd = U0Rd 3. Kd .0
|) -4. Kd = 0.05. DF -fl,80- 6. Kd = 0.0
0 - DIS'ORT, I)x 1. Kd = (.02
6 - 2. Kd = 13u.0SYA. 3. Kd = 155.0
'50 1 I , - 4. Kd = 0.03Vi 70 80 go I00 5. Kd = 0.06HCOtRRECTED AIRFLOW - % . Kd =- 1.87 , -
Figue 38. Effect oj a Specific Distortion FD 16682On Two Compressors DII
Although the difference between these two multistage compressors wasonly an alteration of the inlet guide vane camber, a s-ibstantial Increasein the ttrge line deterioration due to distortion Is evident for the "A"configuration. Other changes within the compressor or to the flow paLh,adding stages to the front of the compressor, changing hub-tip ratio, etc.,would simliarly negate the confidence in a completely empirically deriveddistortion factor.
CONFIOFNTIA[
Pratt &Whitney Aircraft CONFIDENTALPWA FP 66-100Volume III
Table 3. Pratt & Whitney Aircraft Experience with Kd
Engine Type Aircraft Experience
J57 Jet F4D Kd correlated wellJ57 Jet FlIl Kd correlated wellJ57 Jet FlOO Kd correlated well
J57 Jet F102 Kd correlated well -
J75 Jet P105 Kd correlated wellJ60 Jet Jet Star Kd not able to handle the
vorticlty encounteredTF33 Faii-Jet B52-11 Kd did not correlate well.
Correlation was improved byweighting the fan portion of
the flow less heavily.JT8D Fan-Jet 727 Kd did not correlate well
TF30 Fan-Jet Fill Kd did not correlate well
K l t2 avg - t2 min avgd P
Pt2 avg
M 0.2 F j~/ 6(2) Kd Pt2 avg t2 min av& .0 W 110[2 avg J ][~~6design]J
=4 )NI
(4) Kd = 2 aptvPt2 min avg - (0 + (0 +) l-2
() K t 2 avg 2
- ,,S - i I I}
L2 avg j L J'
IIIFiue 9 t a tsi Ditrto FActr 52
DIIIDII-I38
COFIDENAL
CONFIDENTIAL Pratt Whitney AircraftPWA PP 66-100
Volume III
2. Considerationis for Analysis of Distorted Flow
Two factors make synthesis of an analytical prediction system dif-
ficult. Most oLvous is that axial symmetry can no longer he assumed.This requires that the conservation and momentum equations be satisfied
with an additional degree of freedom and recognizes that circumferentialas well as radial gradients will exist in all variables, such as flow
rate, pressure, etc. Less obvious is the fact that the rotor flow cannotbe made steady by selection of a frame of reference rotating with therotor. This is apparent when one considers the changing angle of attack,Mach number, etc. as "seen" by a rotor as it passes through the regionsof differing inlet distortion. These two considerations; the (a) :on-axisymmetric and (b) non-steady, nature of the flow within the compressorundergoing inlet distortion, are the reasons which have prevented therelatively rigorous analysis used for non-distorted flow from beingextended to include distortion. Lack of such a model has consequentlyforced reliance upon the frequently inadequate techniques of empiricalcorrelation. Several ways of improving these analytical tools are beinginvestigated at P&WA. The most promising of these methods will bedescribed in the paragraph 6, "Analytical Prediction of Distorted Com-
pressor Performance.
3. Undistorted Inlet Compressor Information
One very useful tool in the analysis and prediction of distorted
compressor performance is a library of undistorted compressor performancedata, collated for efficient access. In this area, P&WA has conductedover 100,000 compressor cascade tests, the pertinent data for which arecontained in digital computur "memory" for automatic inclusion in digitalcomputations of the flow through a multistage axial flow compressor.
* Similar information from both P&WA single stage and multiste comprez.orsis availblue for comparative purposes.
4. Analytical PredicLtcn of Distorted Compressor Performance
Since the existing analytical compressor design calculations are vallafor axisymmetric flows they ace completely capable of accounting for theeffects of the radial component of the distortion. By then approximatingthe combined circumferential and radial distortions generally encounteredwith a number of sectors, within which only radial distortion is present(figure 40) the present methods may be readily extended. The number ofsectors used In this approximation is dependent upon the severity of the
circumferential distortion and the accuracy desired. Inherent in this
portion of the procedure is the assumption that the blading, In pn.ing
from sector-to-sector, can instantaneously adapt to the new environment,Within these assumptions, a complete definition of the Internal, non-axisymmetric, flow field is made to satisfy a given set of boundary con-
ditions. From the results of these computations the lift that any blade
element will be required to generate in order to remain In equilibriumwith the flow field is given (figure 41). Since the circulation (lift)
about the blade cannot instantaneously be altered to maintain this equi-
librium to the non-steady conditions, it becomes necessary to determine
the extent of the departure from apparent equilibrium that is actually
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CONFIDENTIAL
Pr-.tt&Whftn y Aircraft CONFIDENTIAL IPWA FP 66-100Volume III
experienced. To accomplish this, the cyclic lift distribution is approxi-mated with discrete incirements. At the occurrence of each of these incre-ments a vortex is shed by the lifting surface and "washed" downstream withthe flew as depicted in figure 42. At any instant in time the lift on theblade eli.*ient is determined through integration of the effect of all oftheae vortexes on the circulation about the lifting line (or blade). Inthis fashion, the transient response (figure 43) of the various bladeelements is calculated and then included to satisfy the required boundaryconditions as an iterative procedure, with the stage matching considera-tions already mentioned.
9=160
Figure 40. Ex-su lllb r licmeeta ode D 166.113
D -4I/ tat~~__-o ean HigiCrneey-iltGieI
1 0Ua *60
CIRUM F NTIAL LOATOFigure 40. Ex..... Cicr L n 16683
DIII 40
C a N", RD [ H1'1i
ICONFIDENTIAL Pratt Whitney AircratPWA FP 66-100
Volume III
Discrete Increments 01-, -Dumete
j Constant
1 00180* 3W~
CIRCUMFERENTIAL LOCATION
0 Flow
-Trail of Shed Vortexes Due toDiscrete Changes in Inlot Air Direction
Figure 42. Transient Response Prediction FD 16685Method DI I
~-Constant fliAn!'ter
I Apparent Equilibrium
S Predicted A1
Ii - Response
180* 30ii CIRCUMFERENTIAL LOCATION
Figure 43. Predicted Rotor Blade Respoise ED 16686DIIDI-4
I CONFIDEIAL
Pratt & Whitney Aircraft CONFIDEIALPWA FP 66-100Volume III
It may bt noted in ligure 43 Lhat the response will be out of phase
with the equilibrium loading and that the respo so to a seve-e but localregion of distortion, seen as AB, may be less than that to a relativelyless severe but- more extensive region such as AA. This effect togetherwith those attributed to rotor speed and flow velocity are automaticallyincluded since they determine tLhe position of the discrete shed vortexesrelative to the blade. Their importance is evi.denced by the fact Chatmany empirical prediction methods utilize weignting factors which are func-tions of the circumferential extent of the low pressure region, the rotorspeed and flow rate.
Having completed the iteration, a complete description of the condi-
tions occurring instantaneously throughout the compressor is available.It is now possible to evaluate these quasi-steady conditions using theextensive criteria already mentioned for evaluation of a non-distortedflow since the non-axisymmetric non-steady aspects of the flow are accountedfor.
It is recognized that successful completion of such an ambitious under-
taking is largely dependent upon a firm understanding of the effect of thedistortion on the internal aerodynamics. The interstage data referred to pr
previously will provide this necessary input as will similar data now beingobtained on current P&WA engine compressors, both of the "fan" and jet types,The superiority of this "Blade Element Response" method is illustrated I(figure 44) by comparing its comprehensiveness with that of six empiricalcorrelations which have found widespread use. 'Types of Distortion Factors"identified by numbers one through six on figure 44 correspond to similarlynumbered KD factors shown on figure 38. Ef
TYPE OF DISTORTION FACTORS RESPONSE
EFFECT 2 3 4 5 6 ANALYSISrelative t .2 to .F12
CiunaEfAL EXTZNT: (I.), (4r+) 0. -* * * *MDIAL DiSr Tli'l; Dihtctlioo 1. Wfighd -I
ocordlo to --naI i- ---tl.- ,
DiSiotIiol Tuo Ljy: Atcoante for . etc,locatim and tel'tive level of engine lice
distortion at expressed In racs ci oF2. ----- - - -
velcity, incideaca angle, etc. i
LTAE mf hI ING: opecttlc1
pout/oPe--
.linJ lirA c'etns.a co. icce dat. dietortion0iTflTION ArMl% ATION: Deoree ofr1.occi'n itt distort in lanievel ri - - - - - - - - StIc of rtor" bldefe(). A epenJ. -l of tenish D! WOADING: atanorml."clon of blade::Iura .1 r14 tv to -, xl io -.ur. -. -- - - - - - - -- - --
rie irecatlo.
APPLICATIOH TO NEW MISICH: PrldlctIonpyuta. rlinm ptenE*t aele etchin to - - ----- -------*cceandata ditoct . in.COPUSOR - LXIT I11.W TILR#: nenl .setanr l perturbat o la cet of di lt.o-n - --
on ,"Lbled e apnnaaie that mpis. ic. t.Braid proplsian ytae-.
Figure 44. Aerodynamic Distortion Considerations FD 16687DII L
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CONRfDENTIAL
CONUIDENUIIL Pratt & Whitney lircraftPWA Fl 66-100
Volume III
5. DistortEon Limits Based on Blade Element Characteristics
As discussed in the preceding paragraphs, the techniques used in the
prediction of blade element response to distortion were used as a guideto establish criteria to define distortion limits for inlet-engine compati-bility, Use of this general approach will be coo-dinated with the air-frame icianufacturers and would take the place of the presently used methodwhich treats inlet-engine distortion compatibility on the basis of corre-
lation of (Pt2 Max - PL2 Min)(Pt2 Avg)-1 data. Although the method pre-sented herein is an initial step toward a more sophisticated technique,it does have immediate merit in that it is based on more fundamental fac-tors than is the above referenced method of defining a distortion limit.
formance is that a more realistic distortion limit, dependent upon the
distribution and extent of the distortion is attained. It is recognizedthat absolute levels of both performance limits and surge limits will beset as a result of Phase III testing and analysis.
The blade element performance shown in figure 45 represents that of
a typical first stage blade. This blade characteristic is independentof engine configuration and is equally representative of turbojet firststage blades as of turbo-fan blades. It can be seen that the rotor bladeroot by virtue of its lower pressure Mise reserve, would be less tolerant
of low pressure regions of distortion than are the mild-span and tip. Con-versely, high pressure distortion regions at the root, leading to higher
than average flow, a characteristic of some early SST inlet distortion
data, are favorably handled with relatively small pressure ratio oaseCs.A distortion limit based on blade element prformance would accounu forthis. Response of a fixed point on the blading as it passes through thedistortion also needs to be considered, the response increasing as a
function of dwell time in a given low or high pressure zone. The cir-cumferential extent of pressure zones characterized as low or high with
respect to an annular average, as well as the variations in total pressure1U within the particular zone ,nlndr oonsideratiCn, W4lldctralr, Lie d-e
of rotor btace response to the inlet velocity of that zone. Rotor speed
effects must also be considered, as the performance of the blade elementsvaries significantly with both angular velocity and inlet mass flow.
6. Blade Element Distortion Limit Analysis Proceaure
Airframe inlet distortion is of an unsteady nature, in many instancesvarying in e time period during which the rotor makes considerably lessthan one revolution. However, in many cases the response of the rotating
blades is sufficient to almost fully react to a time variant itlet dis-tortion as if it were a steady process.
Therefore, in order to provide adequate definition of the airframe
inlet distortion, total pressvre pickups should be located at the Inlet-
engine interface on 10 to 12 circumferentially equa-distant riakes, eachrake containing 5 to 6 pickups which are placed at centers of equal annu-lar areas. The time required for a given blade to pass from one ofthese rake stations to another is approximately one millisecond, so thatthe total pressure instrumentation should have a -esponse time on tle
order of 1/500 to 1/1000 sec. The total of these equal annular areas
DII-43
CONFIDENTIAL
Pratt Whitney Aircraft CONFIDENTIALPWA FP 66-100Volume III
compriseS the inlet annulus. It should be recopn!zed that completedefinition of the inlet flow field excee(ds practical limitations. llowever,instrumentation possessing the characteristics noted above serves reason-ably well, dividing the engine-inlet interface into 5 or 6 annulil ofequal areas, the circunferential roessure distribution of each annulusbeing provided every 300 or 360, depeodent upon the number of rakes thatare utilized. Specific requirements for the instrumuntatlon used todefine distortion must be established mutually by the engine and air-frame manufacturers. Should instrumentation having a slower response beused for the purpose of defining engine face distortion, then compatibledistortion limits will be negotiated in an effort to cimpensate for theinaccuracies introduced by poorer response pickups.
Region of Stall
/-Low Inlet0 Tip Pressure Capability
Midapan
iii,I-A-t - it i
Low Inlet
Pressure Capability
Ii:0 Loz:=tion of Operational Point
FLOW CHARACTERISTIC
..iure 45 .Blade Element Performance FD 16688
DII
I] I
DII-44
CONFiDENTIaL
I I l l l l l l i lI I l
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I CONFIDENTIAL Pratt & Whitney PlrcraftPWA Fl' 66-100
Volume 111
The terms used in the "Blade Element" distortion limits method are
listed below:
Pt2 avg = Average area weighted enigine inlet total pres.ure
Pt2 max = Maximum engine inlet total pressure measured usingrecommended instrumentation.
Pt2 min = Minimum engine inlet total pressure measured byrecw,ended instrumentation.
Pt2 rad = Average area Weighted inlet total pressure for givenannulus
NI = Angular velocity of fan rotor, rpmet2 T t2 avg °R/518.7C Circumferential distortion factor that accounts for
zones of distortion in which blade response is complete.This factor may be considered to be ropresentative ofa square wave distortion pattern of long period.
01 = That angle, at a constant radius, subteoding the zonein which the total pressure deviation from average
reaches a limiting value; a maximum in the case ofhigh pressure, or a minimum for the case of low pres-
sure (see figure 46) dog.
62 = That angle, at a constant radius, subtended by thebeginning of a low or high pressure zone and thepressure level that is 90% of the difference betweenmax or min Pt2 and Pt2 rad occurring at !11, deg. (Seefigure 46)
03 - That angle, at a constant radius, subtended by the
terminus of 61 and the intercept of the response pres-sure level Pt2 rad - Z (Pt2 rad - Pt2 min or t2 max)with the pressure zone boundary following 61, deg.(See figure 46)
84 The angle, at a constant radius, between the occurrence
of the minimum (or maximum) pressure and a value of 90%
of the difference between Pt2 rad and that pressure,leaving a low or high pressure zone. (See fiv,,rp 46)
I,,C \
______ t2 rad
/1t rd '4Pt 2 rad"l'tu rain
t2 red m nut
Figure 46, Constant Radius Circumferential FD 17803InleL Distort .. DII
DII-45
) CONFIDENTIAL
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Pratt & hltney Ai1rcraft CONFIDENTIALPWA F2 66-100Volume 1II
These angles are used to define the significant circumferetial extentof a distortion pattern and to indicate its wave form as illustrated below.It may be noted by referring to figure 46 that when -1 = I
= 0 and64 = 03 the distortion pattern in question has a square wave form. Con-
versely, when 01 = 02$0 and 04 is very small, the circumferential patternunder investigation is sharply cusped.
Z = That factor which depicts the degree to whic-h a bladeelements respond to the pressure perturbation (Pt2 rad -
Pt2 rin or Pt2 max) as it rotates through the arc 'I.
An inlet distortion pattern, representing the total pres;ure field atan instant of time as measured with the previously described instrumenta-tion, is analyzed in two parts to determine its effect on the engine:
1) Radial and, 2) circumferential.
The radial component of the distortion is examined by dotermining the
area average total pressure for each concentric annulus (Pt2 red) and com-paring the resultant radial distribution (normalized by Ptt avg) with thelimits specified ii) figure 47 and figure 48. Operation with radial distortiwithin the band bounded by tile curves of figure 47 aill produce a-i negligibleeffect on engine performance. Operation within the band bounded by the curvof figure 48 will not so adversely affect engine operation as to precipitate
engine stall or flameout,
A radial total pressure gradient or distortion is the normal inletcondition anticipated for all fan/compressor stages. This occurs becauseeach stage provides the succeeding stage with a non-uniform profile.Since this condition is usual. and anticipated, it is included in the de-
sign procedures. The first stage design must include the radial distor-tion provided by the inlet and differs from the design of succeedingstages only in that the inlet distortion is usually poorly defined duringthe early phases of engine development. As compatibility testing of the
inlet provides a better dlefinition of the expected radial distortion itmay become necessary to modify the design of the fan stages to be toler-ant to a different radial nattern than .a ,rgia pl niakno for. qb1,,
lc ibillity is a part or the anticipated development program during whichcompatibility is achieved by early modification of both inlet and engine.
For this reason the curves depicted in figure 47 and fi ;ure 48 must be conlsidered as schematics, representative of trends and concepts but subject t
significant alteration during compatibility development.
The circumferential component of the distortion requires more complexlimitations because it is necessary to account for not only the radialdifferences in blade attenuation but also the transient response of a
blade element as it rotates through the non-uniform flow field. Thelimiting levels of circumferential distortion for each annulus are expres:
by the ratio of the maximum and the minimum total pressure within theannulus to the mean annulus total pressure. These limiting values are
expressed in the following equationg:
DI1-46 IiCONFIDENTIAL
CONFIDENTIAL Pa tt ftne AicraftPWA Fn 66-100
T Volume III
Equations Defining Allowable Circumferential Distortion
Pt2 max = 4- CxRxMxY
Ft2 rad
I Pt2 min I - CxRxMxY
Pt2 rad
The quantities C, R, M and Y are factors relating to the circumfer-ential "shape" of the pressure profile and the transient response of theblading and will have values defined by theoretical studies and compati--bility testing during Phase III. These tests will also provide theIboundary condition data needed for the previously described "BladeElement Response" method ot analytically prcdlct.ng distortion effects.I
-! -00 7W Negligible Effect on Performance
0 1 --~80 - _
4) High Pressure to be Supplied
X _
I t
'-4 -. ,12
-40
I Lwre~ure ______
0 0 20 40 60 80 100RADIAL LOCATION - % of AnnulusiU
-FiurL 4?. Etimuated Limiting aluec of Radial FD 17820SCorsponent oL Inlet Totl' Pessure DI_
U Distortion - NeglgLle Effect on-Per forirance
< .
!Z
DII 47I CONFW ,,AL
-mo- ---
Prat a Whlney Arcrft C-RDO ALPWA M 66-100Volume III
0Will Not Cause Stall or Flameout
"- k2 to be Supplied
HihPressure
p.4
U zDW 0
O r 20
E... 0 . .. .|
1W Low Prealture+*60 ...
naiie Diameter Outside Diametcr--.-
0 20 40 60 80 100 a,
RADIAL LOCATION - % of Annulus
Figure 48. Estimated Limiti'g Value of Radial FD 17821 - 1Component of Inlet Total Pressure DTTDistortion - will Not Cause StalorFleout
The maximum circumferential distortion allowable when the completeresponse is att. ned by the blading is accounted for by the quantity C(figure 49 and iigure 50) dependent upon the rotor corrected speed and thespanwise location of the annulus under consideration. Since complete response Iiis not the general case, the modifiers R, 14, and Y are provided to account
for less than complete response due to the circumferential distortiongradient and the dwell or residence time of the blade elements withinthe high or low pressure zones. These modifiers are set equal to unityfor the case of complete response and increase accordingly with de'-easinglevls of response.
The ron-onns to the pressure deviation (min or max pressure) fromthe average (Pt2 tad) is estimited by the factor Z of figure 51a a func- 1tion of the angle 01 between Pt2 rad And the minimum (or maximum) pres-sure and the angle, 02, required to attain 90% of the difference between
Pt2 min or Pt2 max and Pt2 rad" Because of its bivariate dependence, thequantity Z accounts for the element response to the minimum (or mnaximum)pressure based on the rapidity (0I) with which it is reached and the
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CONHOIENTIAL Pratt &Whitney AircraftPWA F? 66-100
Volume III
"shape" o; the leading pressure gradient (02 and 1). The resnonse, ?,is used to define tie portion of the low (or high) pressure region ofprimary interest through the angle, 63, terminating (figure 46) when thepressure returns to Pt2 rad - Z(Pt 2 rad -' t2 min or Pt? ma× )
The response, R, (figure 51b) is then determined by the total angulardisplacement "l + 03) of the significant portion of the low (or high)pressure zone. The modifier, M, given in figure 51c, is a function of Zand 03. This factor acsounts for both the initial response (Z) and thedwell, or residertce time, of the blade within the significant portionof the low (or high) pressure zone. Thle dwell time (F) becomes lesssignificant as the response to the leading portion of the zone (Z)approaches 1.0 denoting complete response.
Finally, the factor Y, figur,, 51Id, is deteris tnied from the ratio ot04 to 03 - This factor approximately accounts for projile shape wi thitthe dwell region.I
The maximum and minimuml pres'ures allowahlc in ea-h annulus are cal-culated through substitution of the above determined factors in the
"Equations Defining Allowable Circumferential Distortion." Operation with-in these limits when the value of C is determine:d by the curve of figure 49will not effect cgne performance. If the curve of igoer 50 is used todefine the quantity .C, the resultant a lowu ble circumferential distortionwill not cause surge or Ilameout, A linear interpolatiot shall be useud todetermine the value of the quantity C between the 11) and OD.
Separation of the distortion into radial and circumferential compon-ents provide a logical distinction because of their steady and non-steadycharacter, respectively. However, their combined effect on the fan/con-presser aerodynamics is, In the ffn~l analysis, the significant aspect.For this reason, the foregoing discussion is of a general nature andtrade-uffs may be made between radial and circumferential allowances to
provide adaptation to the distortion pattern oritor-,c by z-.pcfic
.,,AL. s. in the .ac of tfi iadial distortion allowancc, the circum-
ferential limits may be changed by modification of the ian/compressorif this measure is indicated during compatibility testing. Since modi-fications to increase the engine distortion tolerance sometimes lead toa performance decrement, carefully coordinated analysis between airframeand engine manufacturer is mandatory.
While the above improved method of defining di!tortion limits Includesfactors which are considerably more pertinent than does the former(Pt max - Pt inin ) definition it must be recognized that no sirmple method,
Pt avg
such as described, can provide a universally applicable system. Conse-quently, Special cases invulving unusual distortion pati,:r.g may bocon.idered individually by more rigorous and detailed analysis by coot-dination with the hirirame manufacturer.
UD-49! COHNiCENTAL
Pratt &Whitney PAt--craft CONFIEIALVolume III
oo r'-
(0 C- U
45-
C... 4 4I- 14il anv Ii.tun oC
-C C ~-4 J- J 4bj~~~- - EI~LUf~ o* .
Iti-44 C::
4
4~16 01 I
444J
CIO w 0 c a
IN 44NAIa-iw N14. '-48.111f 'H I'JD A
44 .V1t 50
44 A li
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CONFIENAL Pratt &%Whtney AircraftPWA FP 66-100
Volume III
1.0 92=0
I2 1 (a)
2.0
~(b) (c.R M B ou c
I.O1.o .0 0 9 1
01+93 Z
Y (d)
1.0L0 104
Fignte 5i. Sinpe an_ meeponse Factors to bc Used FD 17804in Int rmixing "Blade Element" DIIDistortion Limits
CN
CONFIDENTIALz
CONFDENTIAL V-ratt&Whtnyn'rcraftPWA FP 66-100
Volume III
SECTION IllEXHAUST SYSTEM COMPATIBILI TY
A. INTRODUCTION
The design of the engine exhaust system and its installation in theairframe must be analyzed and tested to optimize the performance and in-sure the compatibility of the airframe/engine combination. 'he i nlet andexhaust system compatibility ere both important to the performance of asupersonic transport. Pratt & Whitney Aircraft and the airframe manufac-turer will execute a coordinated program of analytical studies, designand testing to develop a compatible, high-performance exhaust system. Al-ready in Phase II, extensive wo-k has been done toward this objective.The results of this work, which have been fully coordinated with the air-frame contractors, are discussed in the following paragraphs. A morecomplete explanation, including such Interactions as the routing of sec-ondary air and the effect of tertiary air doors-open operation, is givenjin Volume ilI, Report A, Section Ill E.
B. INSTALLATION EFlFECTSIThe selection of the exhaust nozzle configuration for the JTF17 en.-
gine was based upon trade-off studies involving three different nozzleconfigurations. Although not the only criterion, the sensitivity of thenozzle to installation effects was an overriding factor in the evaluationof candidate nozzles. Such trade-off studies rely heavily on experienceof service aircraft for evaluation criteria, and the initial evaluationwas made utilizing the excellent service experience of the blow-in-doorejector in the YFI2 aircraft. When data became available from the lessfavorable F-1ll installation, these trade-off studies were repeated tossLre data from the less favorabl, installation received proper cqnsid-eration.
Three configurations, the plug nozzle, the long vnrlable flap nozzicVand thc blow-i-door ejctor were inluded in the study. Detail results
of these studies can be made available upon request; however, in summary Ithe plug nozzle was rated very low because of its relatively high sensi-tivity to performance loss from flow field variations, and its unsuit-ability to incorporation of sound suppressor devices. The blow-in-doorejector and long variable flap nozzle were rated about equal for periormi-ance loss due to installation effects during the initial evaluation andagain during the reevaluation. The final choice of the blow-in-doorejector was then made on the basis of its overall performance, as a resultof better sealing, iess complexity and weight, and better noise attenuationcharacteristic,. A complete description of the ejector amid its designfeatures is presented in Volume III, Report B Engine Design, Section II.
To achieve the performance and a'-"qht advantages of tlhe blow-fn-doornozrle, czre must e- taken in its inLk&ta,1-n to account for variationsof the aircraft local flow field. Mont,, crts, summarized in figure 1,have shown that tertiary dor-open nozzC )L'rformance is dependent uponproviding the required quantity of tertiary flow regardless of the amountof door 'blockage. Performance is directly related to tertiary flow quan-tity as indicated in figure 1, and adverse local flow fields are compen-sated by increasing tertiary door area. An exhaust system placed tel-
bJil-I
CONFIENTIAL(This Page 1a Unclassified)
Pratt&Whtney Aircraft CONFID.NTIALPWA FP 66-100Volume III
stive to the aircraft such that it operates in a uniform flow field withminimum deviation from free stream conditions will exhibit little or noinstallation effect. Both airframe manufacturers are aware of this, endhave provided wing-mounted installations that should produce minimum dev!ations of the nozzle local flow fields from free-stream conditons,
0.96- All Trertiary
XFlight Mach Number ,9 Doors Open-
i 01E_ [ All' Tertiary Doors0 Z Half-Closed-
All Upper Tertiarh tU 0.88 -Doors Cl"ed
All l owr ertiaryDoors Open
0.84 1 Te total correcte-d flow ratio is jefined asirJ. the product of the ratio (A the secondary
< (Ait Tertiary ajid tertiary airflow to engine gas flow and
0.80 Doors Clsed the square root of the compressor inlettoial temperature divided by the engine
total temperature. W t/T i/ r" I I I 'otit) Corrects Flo to =- "" + T _
0.7 I I I o I Wen ine 'Ttengine
0 0.02 0.04 0.06 0.08 0.10 0.12 0.14 0,16 0.18 0.20 0.22TOTAL CORRECTED FLOW RATIO
Figure 1. Variation of Exhaust System FD 17124Performance with Total Corrected DIIIFlow Ratio (M = 0.9)
The JTF17 exhaust eyf-tcm 1z bcit,6 designed to coipetiblae for varn-
otioan in local tlow coaditions when they occur, thereby minimizing in-
stallation effects. The local flow field approaching the exhaust noz- ';zle may vary from isolated test conditions due to the effect of thewing and adjacaitt bodieq, overboard bleed flows upatream of the nozzle.and differences in model and full-scale boundary layers. The differencebetween free-stream Mach number and the local Mach number around thenozzle can affect nozzle performance by changing the tertiary air in-let conditions and the external pressure drags. High local Mach num-berg, and consequently low local static pressure fields, may also causethe pressure-actuated tertiary doors to close prematurely and therebyproduct overexpansion losses in the nozzle. The local pressure fieldCan also affect the position of the trailing edge flaps which can pro- liIduce overexpansion losses in the nozzle if the flaps float to a largerexit area. A variation of boundary layer height between the installedand isolated test conditions will also effect the r.rtiary flow and in-ternsl performance. The isolated transonic wind tunnel investigationsof the JTF17 exhaust system are conducted in a free stream flow fieldwith 4-inci' diameter models. The models are mounted on a shaft ap-proximately 6 feet long that protrudes from a streamlined strut located t'just ahead of the test section. Calculations indicate that the relativeboundary layer heights between the isolated scale model tests and the
actual aircraft (approxLmately 11 inches full scale) are aimilar. In
DIII -2
CONFIDENTIAL ___ I
CONFIDENTIAL Pratt a Whitney AircraftPWA FP 66-100
Volume III
the absence of large quantities of overboard bleed and large variations
in boundary layers, installation effects arc primarily a function of the
local Mach number. produccd by the aircraft. Experience has shown thatI differences in nozzle performance occur in the transonic flight regionwhere the tertiary doors are open and the trailing edge flaps are closed.
At cruise Mach numUers, nozzle performance is not affected by the air-
craft fiow field.
1. Previous Experience
J- Installation effects have been thoroughly investigated in the Pratt
& Whitney Aircraft J58 engine installation in the YF12 aircraft. Wing-
nacelle model test programs and full scale flight investigations have
been conducted in conjunction with the: airframe manufacturer. Resultsof these programs show that installation effects are predictable and
that the exhaust system can be desigied to minimize these effects
in the transonic flight range, as shown in figure 2.
J58 experience has showni that an approximation of installation ef-
fects for a wing mounted nacelle can be made by testing isolated scalemodels at the local Mach number of the aircraft flow field. For ex-
ample, a simulation of the conditions at a flight Mach number of 1.05
could be made by testing at approximately Mach 1.2. This type of test
would simulate both the lower pressure air entering the tertiary doorsas well as the increment in boattail pressure drag caused by the air-craft flow field. This test would not, however, show the effect of lowpressures caused by the boattall on aircraft surfaces. A similar in-vestigation has been carried out for the JTF17 exhaust system instal-
I lation in the L: Wiv'ed aircraft, as discussed in the next paragraph.
c 0.06p~- -i---
Tertiary Doors Opet:
Tertiary Doors Closed
0.02
0
L -0.02 ---_ J--- __-
[" 0 Flight Data -4 Full Size Tertiary floor
-0.06, 2 - - ,hau.t Syste,_---
0.8 1.0 1.9 1.4 1.6 1.8 2.0 2.2 2,4 2.6 2.8 3.0
FLIGHT 'fACH NUMBER M
Figure 2. Performuance Comparislon of Flight aud FD 17100jScale Model Exhaust Systems D111
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Pratt & Vhtney Aircraft CONFIDENTIAL.-WA IP 66-100-Folume III
2. JTFI7 Exhaust System Installation Effects
Chordwise static prcssure distribations have been measured for theinboard and outboard nacelles on tliw Lockheed wing. Average local Machnumbers in the vicinity of the tertiary doors have been calculated fromthese data and e comparison is shown in the lollowing table:
Freestream Mach Number 0.9 1.3Average Local Mach Number 0.92 1.31
Figures 3 and 4 indicate the estimated performance increments as a func-tion of local Mach numbers for the JTF17 nozzle during transonic ac-celeration and subsonic cruise respectively. No appreciable instal-
lation effect is estimated from either figure with the maximum Cfpins tof 0.004 occurring during subsonic cruise operation. During t
operation at acceleration conditions, exhaust system throat areas andpressure ratios are larger than for subsonic part power and installation
effects will be less than at part power conditions.
I',I-I__________ M ach W)~ E ngineu 01 v ratin ~jond it ions
H i- Clu ist C.nl rt~
- t' IIt inst
*- 1 _ I, "
. "r
L _ _ __ I , __
IIi
>e L.
-.- 141-_
F1 .2ILJ' .4 .6 .8 Mj) 1.2 1.4 1I.6
~~~~EXAS YTM 'TEINAL LOCAL MACH NUMBER, MI,
Figure 3. Estimated Jil] Lxhaust System FD 17125Installation Effects, Lockheed DiliInstallatiof,, Subsonic CruiseConfiguration
D11-,4
CONFIlHIAL |
CONFDENTIAL Pratt & Whitney AircraftPWA FP 66-100
Volume 11
Mach 1-3 Engine 0perating Conditions
r.,. Z oo 10 -- 7 1:. o.0 1 .)
-0.02
I Istllaio'fTanoni Acetra I
'.; rnsnc ccl; nConfiguration 1
1 0. - -- 0. iII I--
3. JTF17 Installation Tests
During Phase II--C the L-2000 propulsion installati-on details were
supplied by Lockheed, and the first of a series of tests with a JTF17I exho ist system wind tunrel model representing the Lockheed installation
t were conducted in the United Aircraft Research Labo~ratories main windtunnel Lo investigate installation eftects on exhaust nozzle perform-
Sance. Tho cnrmple ^+tcc t_, &t:Lius: io De co t e n" ' Lo'I= . '--,.= inse !I! wilt in-~veatigate the effects of the basic aircraft installation, angle of at-
tack, exhaust system trailing edge flaps, eleven deflection angle, andexhaust system axial location.
~The initial tests established the effects of the L-2000 installation
~a Ii
on the performance of the JTF17 exhaust system during subsonic part power~operation and investigated the wing and exhaust system local flow field
pressure and ,Mach npunber distribu~tions. ::
- The 1/20th scale installation model consists of a metric flowingIl nacelle witli the JTF17 exhaust system and a non-metric half wirg of
i cropped-delta plan form with an open channel dummy nacelle as shown in
figure 5. The metric ilowing nacelle consists of a 4-inch outsidei diameter shaft w4 t ertCnd5 I feet downstream of the trailing edge of
astreamlined fairing. The upstream end of the shaft to attached to athree-flow force balance and the exhaust system is attached to the down-
" stream end.
The non-metric half wing consists of a cropped-delta plan form,with a modifted double wiedge section and A 10 degree tit. dro(, .. A halfwing model is representative of the installation sincl- a splitter plateI
Dl1 I
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Pratt & Whitney Rircraft CCIIDENTIAL1PWA FP 66-100Volume III
at the wing root divides the flow field and the flowing nacelle is lo-cated several nacelle diameters away from the centerline. An open chan-nel dunmy nacelle simulates the inboard propulsion package. Replaceableelcvons are available to, simulate various trim positions but were nottested during this initial test program. The flowing nacelle shaft was Vset at a negative incidence of 2 degrees relative to the wing chord. Theinstalled tests were conducted at angles of attack of 2 degrees and 5 de- Igrecs. The wing was mounted on edge in the test section on a wing root Iplate and two support struts. The test instrumentation consisted ofthe following: U
1. A three-flow balance to record exhaust sys tea gross thrust-coefficients,
2. Three static pressure pip 2s located three feet from the upperwing surface, three feet Lrow the lower wing surface, adthree tnchcs_ ot, the wing tip.
3. Tunnel wall static pressure plates arc mounted on the upper
wing surface side of the tunnel and on the lower wing surfaceIside of the tunncl.
4. A tunnel spanning static and total pressure "T" rake.
Readings frm th~s rake were taken just upstream of the Iplane of the ejector te.-tiary doors.
5. The lower surface of the wing was instrumented with 35 staticpressure taps and the upper surface with 18 taps.
6. The exhaust nozzle had four static taps circumferenLiallyspaced just ahead of the tertiary doors. _
7. Two one-inch high boundary layer mice were mounted on the
nacelle just ahead of the tertiary doors above and belowthe wing.
I.
W- 7
Figure 5. Exhaust System Installation iodel FD 16905Tests - Lockheed Configuration DIll
DllI-o
CONFIDENTIAL(This Page is Unclassified)
CONFiDENsli Pratt Whitney AircraftPWA FP 66-100
4 Volume III
Preliminary results of these tests are shown in figure 6. The r'irves
show the variation of isolated and Installed nozzle performance with flight
Mach number for Mach 0.9 engine opaering conditions. The JTFI7 exhaust
system installed performance is approximately i/2K lower than isolated
model performance; however, the installed model performance meets the sub-sonic cruise performance goal. Angle of attack has little effect on the
installed performance. The excellent agreement between the theoreticaland test results indicates that installation effects for the JTF17 exhaust
system are predictable and the nozzle design can be modified when neces-
sary to minimize these effects.
The initial test program was limited to subsonic Mach iumber because
of wind tunnel limitations. Wind tunnel flow field investigations con-ducted with both Boeing and Lockheed wings (no exhaust nozzles) indicated
that satisfactory flow fields could be established for subsonic Mach
numbers only. However, a wind tunnel modification incorporating a per-
* forated wall test section to permit transonic Mach number testing isplanned early in Phase III.
Uo1.00 Subsonic Cruise Operation
G Gas Generator IZ .9 Pressure Ratio 1 am 2.
0.98 . .
Fz4 0.6 -0 -- I
0.94
!V 40 .90 JTF Isolated Exhaust -Model Test:IO Wing-Nacelle Teat,5.O-d4 Angle of Attack
0.88 LoiNck ele nt 2.0-deg An Is of Attack
0.86 - _- -- - - 1-o0.84- .06. .
0. 0.4 .5 06 .7 0.8 0.9 1.FLIGHT MACH NUMBER, M
Figure 6. JTF17 Exhaust System Wing - Nacelle FD 17127
Test, Lockheed Installation, Subsonic DIIICruise Operation
DIII-
CUNFIDENTiAL AIIII I~rI -A