Advanced turboprop testbed systems study. Volume 1: Testbed program objectives and priorities, drive system and aircraft design studies, evaluation and recommendations and wind tunnel test plans

DIAGONAL MEMBER AL. AL. 0.10 SQ INS SKIN AL. AL. 0.05 INS TYPICAL SECTION AA Figure 38. G II Aft Nacelle Structure B. L. 145 FAIRT.NG TRA IL INC EDGE NACELLE LEADING FRAME TYPICAL F'AIRING WING BOX Figure 39. Gil Aft Nacelle Structure B.L. 185 supports the drive system installation by means of diagonal members from the upper attachment points downward and aft to the rear spar and by skate angle and nacelle lower longeron extensions which are attached to the front spar. The nacelle structure is fabricated from aluminum-alloy skin, frames, and longeronsl stiffeners. The upper, semi-circular portion of the nacelle is removable to provide access to the jet pipe installation. The main attachment of the nacelle to the wing upper surface is by means of chordwise "skate" angles. A fairing is provided at the tail pipe to protect the upper surface of the flap from the jet efflux. Because the turbine section of the power unit has moved aft to a posi- tion above the primary structure of the wing, provision for blade containment is required in this area.

GIl Flutter Analysis Preliminary wing flutter analyses were performed for the GIl testbed con- figuration to determine the effects of the prop-fan power plant installation at the two candidate locations. The same mathematical model was used as for the KC-135A.

Flutter Analysis Results, Drive System at BL 145.0 - The results of the wing flutter analysis are summarized in Figure 40. The unmodified GIl wing was analyzed first to compare the results with the Grumman analysis and thereby validate the mathematical model. The flutter boundaries agreed within 2 per- cent, as indicated by the circle symbols on Figure 40, even though the Grumman mathematical model included flexible fuselage and empennage effects, which were not included in the Lockheed analysis. The flutter mode involved is a 7 to 10-Hz antisymmetric wing bending-torsion mode.

The addition of the prop-fan powerplants at BL 145 caused a 5-Hz symmetric flutter instability inside the testbed dive speed envelope, as indicated by the solid square symbol. When rotating prop-fan aerodynamic and gyroscopic couplings effects were added, the speed of this instability increased by about 23 mls (75 ft/sec) , but was still unsatisfactorily low, as shown by the open square symbol.

To increase the flutter speed to a satisfactory level, a substantial in- crease in the wing torsional stiffness inboard of BL 145 is required. The FTII KOT UNMon n'l 1m Gil (NO l'1I0P-FAN) .•. ----G UNMOI)U'lJm WING (11-10 II" AtiYM) - ... .-----,/A~ 60% G./ INGKEAtiE - Wti 0 to Wti 172 C:JtUHMAN ANALYSIS-.p 1000 FT 1000 IU

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0 mls I 400 500 600 700 200 )00 0 100 KNOTS EQUIVALENT AIRSPEED Figure 40. Gil Flutter Boundaries - Prop-Fan at W.S. 145 effect of a 60 percent increase in torsional stiffness is shown by the solid and open triangular symbols for the feathered and rotating prop-fan conditions, re- spectively. Although a somewhat smaller stiffness increase might be satisfac- tory, a more elaborate and comprehensive flutter analysis will be required to determine a precise figure.

Flutter Analysis Results, Drive system at BL 185.0 - Relocating the QEC to BL 185, and with the prop-fan plane one diameter ahead of the wing leading edge, increases the flutter speed above that with the powerpplant located at BL 145. The damping of the fundamental wing torsion modes (both symmetric and unsymmetric) is, however, unsatisfactorily low at airspeeds well wi thin the limi t-speed envelope. Attempts to stabilize the mode by increasing the wing torsional stiffness actually reduced the damping, so that it became obvious that no reasonable amount of wing stiffening would solve the problem.

Moving the prop-fan plane aft, however, 0.914 m <3.0 ft) improved the damping of these modes, which when combined with a 60 percent increase in wing stiffness out to BL 200, provided satisfactory damping wi thin the limit speed envelope. It should be noted that the damping is marginal and is sensitive to changes in altitude, power-plant mounting stiffness, prop-fan aerodynamic characterisitics, and other parameters not investigated.

GIl Wing Modification The GII wing structure consists of integrally stiffened upper and lower skin panels and front and rear spar structures, which together form the wing box beam structure. Increasing the torsional stiffness 60 percent, for either of the drive system locations investigated, requires the addition of doublers to the upper and lower surfaces of the wing and to the front and rear spars. Be- cause double curvature exists on the wing from BL 145 inboard, perfect matching of the doublers and skin is not possible and liquid shim would be applied to the faying surfaces.

Adding doublers to the forward face of the front spar and to the rear face of the rear spar requires removal of the leading and trailing edge structures.

No problems are anticipated with the front spar reinforcement, but the doubler applied to the rear spar presents a major undertaking because removal of the landing gear support is involved. Finally, modification to the spoiler system is necessary which would eliminate the ground spoiler for the inboard location or deactivate the inboard flight spoiler for the outboard drive system location.

GIl Operating Envelope The operating envelope for the GIl, Figure 41, was established by analyzing the 0.13 rad (7.5 deg) upset condition for 20 seconds to determine the dive speed. The points analyzed were those at altitudes of 9118 m <30,000 ft) and 10,668 m (35,000 ft), starting the upset at a Mach number of 0.8. The upset condition onset at 9118 m (30,000 ft) results in a Mach number increase to 0.89 at the end of 20 seconds and an end altitude of 8534 m (27,990 ft). Below this altitude the testbed aircraft speed is restricted to 172 mls (565 ft/sec) EAS in order to minimize weight penalties arising from wing torsinal stiffness in- creases.

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~~ 160 :50 '00 ~IT"'''L['~''IIIS'UD J I I! 'tooTS !OO ZOO 300 "-00 SOO EQUIV.IU:it AtRSnD) PROP FAN LOCATED AT WS 185 Figure 41. G II Operating Envelope Figure 42. G II Trim Capabi I ity - Prop-Fan at W.S. 185 GIl Prop-Fan Testbed Trim Capability

The outboard limit for locating the drive system is BL 185, 1] = 0.45 and is

dictated by the aircraft trim capability. The data of Figure 42 show that, with one prop-fan windmil1ing and 100 percent power on the other, the testbed air- craft can be trimmed for engine-out conditions at 77.7 m/s (275 ft/sec) EAS in free air with 0.087 rad (5.0 deg) angle-of-bank or at 90 m/s (295 ft/sec) EAS on the ground. The use of the T56-A-14 gearbox restricts the power input to 4101 kW (5500 shp) so that, when this constraint is applied to the data of Figure 42, the power setting of the prop-fans, at the conditions indicated, is limited to approximately 75 percent of takeoff power. At this power setting, the engine- out, free-air trim capability can be achieved at a speed of 66.3 ml s (218 ft/sec) EAS and at 77.0 m/s (253 ft/sec) EAS on the ground.

GIl Testbed Weight and Balance .

Weight data are presented for both drive system locations in Table X. The essential difference between the weights is due to the increased doubler weight for the BL 185 drive system location. The operating weight of the unmodified aircraft is 15,464 kg (34,020 lb), which increases to 21,508 kg (47,318 lb) and 21,622 kg (47,568 lb) for BL 145 and BL 185, respectively. The difference in fuel weight for the two configurations is about 113 kg (250 lb). Balance checks of the testbed configuration show that, for either of the drive system location~ the aircraft center-of-gravity can be maintained within the envelope for the existing aircraft at all weights. Placing the test equipment in the passenger compartment eliminates center-of-gravity problems.

TABLE X. Gil TESTBED WEIGHT AND BALANCE WS 185/FS 332* DRIVE SYSTEM LOCATION WS 145/FS 385.98' WEIGIIT WEIGHT WEIGHT COHPONf.NT % HAC ARM FS ARM FS % HAC Kg Kg LB L8 0 OPERATING WEIGHT-UNMODIFIED 39.3 15,464 (34,100) 462.0 39.3 15,464 (34,100) 462.0 XT701 PROP-FAN PACKAGES 3,907 ( 8,614) 342.2 3,907 ( 8,614) 395.1 OVERWING NACELLE STRUCTURE 233 ( 514) 424.9 233 ( 514) 441.8 WING DOUBLERS 544 ( 1,200) 408.0 657 ( 1,450) 410.9 TEST EQUIPMENT 1,360 ( 3,000) 538.0 1,360 ( 3,ODO) 530.0 0 ZERO FUEL WEIGHT 26.6 21,513 (41,428) 443.3 32.8 21,626 (41,678) 452.4 FUEL 6,837 (15,072) 418.5 6,723 (14,822) 418.3 0 RAMP GROSS WEIGHT 22.5 28,350 (62,500) 437.5 27.3 28,350 (62,500) 444.3 *PROP-PJ.ANE LOCATION GIl Testbed Performance The mission performance of the GIl twin prop-fan testbed is shown in Figure 43. At a ramp weight of 28,344 kg (62,OOO lb), the start cruise weight at 10,668 m <35,000 ft) is 27,317 kg (60,000 lb) and the end cruise weight is 22,109 kg (48,640 lb). Cruising at Mach 0.8 gives a test mission duration of 2.68 hours. The speed/altitude performance also shown in Figure 43, shows that a Mach number margin of 0.04 to 0.05 exists over the design conditions for the twin prop-fans operating at full power with the primary "Spey" propulsion slightly above idle power setting.

- 1000 m 1000FT TWIN PROP-FAN V FULL MAX POWER ON PROP-FANS MAIN ENGINES AT APPROXIMATELY IDLE PLACARD LIMIT .76 .80 .1:54 .88 .92 MACH NO.

SPEED/ALTITUUE BEGIN TEST AT 10,668 m (35,000 FT.)

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RUNWAY CONDITIONS ELEVATION - 700 m (2300 FT.)

o TEMPERATURE - 26.6 C .,&>()o 0 ~ TAKEOFF (80 F) LANDING NO WIND NO GRADIENT RAMP WEIGHT 28,350 kg (62,500 LB.)

START CRUISE WEIGHT 27,322 kg (60,235 LB.)

ENU CRUISE WEIGlIT 22,114 kg (48,752 LB.)

FUEL WEIGHT 6,837 kg (15,072 LB.)

TEST TrnE 2.68 HRS Figure 43. Gil Testbed Mission Performance GIl Prop-Fan Near-Field Noise Characteristics Free-field peak sound pressure levels and noise contours were generated for the GII fuselage for the flight conditions shown by Table XI. A peak noise

level of 147.7 dB is experienced at M = 0.8 with a tip speed of 249 m/s (817

2 2 ft/sec) and a disc loading of 301 kW/m <37.5 Shp/d ). The noise levels decrease as Mach number, tip speed, and disc loading decrease. Relative sound pressure levels estimated for conditions up to the tenth blade passage frequency harmonic for tip speeds of 183, 213 and 244 m/s (600, 700 and 800 ft/sec) are shown on Figures 44, 45 and 46, respectively. These data represent the explicit cruise conditions of Table XI and cannot be extrapolated for other conditions.

The noise contours on the fuselage are shown on Figure 47 for the XT701 and SR3, 10 bladed prop-fan drive system at the cruise conditions of Table XI. At these conditions, the sound pressure level of blade passage frequency harmonics on the noise contour may be determined by algebraically adding the data on Figures 44, 45 and 46 to the OASPL for the appropriate tip speed of Table XI.

The complete account of Task V-Conceptual Design of Testbed Systems is to be found in Appendix E of this report.

TABLE XI. FREE FIELD PEAK OVERALL SOUND PRESSURE LEVELS SR-3 CONFIGURATION ON GULF STREAM II TESTBED @ 10668 m (35,000 FT.) CRUISE ALTITUDE TIP SPEED kW/m2 OASPL (dB) (fps) (SHP/D ) mls CASE CRUISE M (600) 142.0 209 C!6.0) 183 1 0.8 (700) 146.7 (30.0) 217 0.8 241 ..

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5C ~ m m ~ c ~ ~ m ~ ~ ~ 1011.= Figure 47. Gil Fuselage Noise Contours WIND TUNNEL TEST PLAN The Wind Tunnel Test considerations structured around the Program Objec- tives and Priorities preferred methods of solution, shown on Table I, fall into three areas as follows: o Wind tunnel tests that demonstrate the operational readiness of the prop-fan and the drive system through proof testing procedures o Wind tunnel tests that validate and/or advance the fundamental state-of-the-art of prop-fan propulsion o Wind tunnel tests that validate the airworthiness and predicted performance levels of the selected testbed aircraft.

The first of these areas could be addressed by means of full scale tests in a low speed wind tunnel such as the NASA Ames 40 x 80. This tunnel is large

enough to take a "Gulfstream II" complete with two prop-fan drive systems

installed. In the case of the KC135A a reduced span wing with two drive systems installed and a mocked-up fuselage would be required to properly simulate the KC135A prop-fan testbed configuration. Lockheed-Georgia, however, does not recommend wind tunnel testing of the full scale drive system prior to actual flight tests for the following reasons: o Most of the available wind tunnels are not capable of simulating the prop-fan design flight environment in terms of dynamic pressure, Mach number and temperature.

o Most of the available wind-tunnel flow, solid-wall blockage limits are exceeded with the full size prop-fan testbed nacelle and wing section.

o Low speed testing does not directly address the design point of the prop-fan.

o Costs of wind tunnel testing would be high relative to the usefullness of the data obtained.

No wind tunnel testing is proposed for the second area - the development of fundamental data for the prop-fan concept - since the testbed aircraft is in- tended to augment test data in this area.

The third area of concern is testbed aircraft oriented and tests proposed in this area relate directly to the airworthiness and performance of the testbed vehicles.

RECOMMENDED STATIC AND WIND TUNNEL TEST PLAN The recommended test plan consists of a static test of the drive system and high and low speed wind tunnel tests of either testbed aircraft configuration to determine performance and/or flutter characteristics.

Drive System Static Test o Static test stand experimentation of the full scale testbed propulsion system would be conducted to demonstrate operational readiness. These tests will provide proof-of-operation of both the prop-fan and the drive system as well as some near-field acoustic environmental data for the nacelle and adjoining structure. The tests would be independent of the receiving airframe.

Wind Tunnel Test Plan The proposed wind tunnel test plan requires both high speed and low speed wind tunnel testing.

High Speed Wind Tunnel Test Plan o High speed tests of a semi-span 0.21 scale model of either the GAC GIl or the Boeing KC135A in the AEDC 16T tunnel are proposed. These tests would provide aerodynamic data for prop-fan blade classical and stall flutter characteristics, wing/nacelle flow field characteristics data and thrust/drag relationship of the prop-fan components.

o Test of a 0.13 scale full span dynamically simulated model of the GAC GIl in the NASA Langley 16 Ft TDT facility for the purpose of investi- gating testbed aircraft flutter characteristics.

Low Speed Wind Tunnel Test o A low speed wind tunnel (LSWT) test of a 0.10 scale full span model of the GAC GIl in the Lockheed-Georgia LSWT to verify handling qualities and stability and control of the GIl testbed aircraft.

WIND TUNNEL AND STATIC TEST PLAN SCHEDULE AND COSTS The schedule and costs for the recommended static and wind tunnel test plan are summarized on Table XII. The tests span a period of 3 years from the program go-ahead· and the costs in terms of manhours and Material and Direct Charges (M&DC) are: Testbed Aircraft Test Hanhours H&DC GAC GII $347,560 55,720 KC135A 45,380 $158,560 The low speed and flutter tests shown for the GAC GIl are felt to be un- necessary for the Boeing KC135A testbed aircraft as flight safety analysis has shown that no handling or stability and control problems exist with the addition of two prop-fan propulsion units. Furthermore, flutter testing of the KC135A is not necessary as analysis shows that no appreciable changes in flutter boundaries occur as the result of the prop-fan additions.

The detailed discussion of the Static and Wind Tunnel Test Plan is to be found in Appendix tlF': 83.

PROGRAM COSTS AND SCHEDULES Because of the proprietary nature of the Testbed Program Costs and the associated schedules, the detailed costs and schedules are presented in Volume II of this report. A summary of the salient features of the cost determination and of the schedule is, however, provided.

PROGRAM COST ASSUMPTIONS The program costs were developed in terms of manhours and Materials and Direct Charges for each of the conceptual twin prop-fan designs. Cost estimation methodology was based on the Lockheed-Georgia Company experience in the design and manufacture of a wide variety of aircraft. The consistency of the cost-prediction base was assured by assuming that Lockheed-Georgia would execute the total program and be supported by subcontract arrangements for those activities in direct support of Lockheed-Georgia.

The cost data also assumed that: o The aircraft for conversion to the testbed configuration would be GFE.

o The modified DDA XT701 drive systems would be GFE o The prop-fans and modified controls would be GFE PROGRAM ESTIMATED COST The cost to perform the entire program for either of the recommended 6 6 conceptual designs is in the range of $40 x 10 to $45 x 10 based on the value of the U.S. dollar in 1981.

PROGRAM SCHEDULE The program schedule, Table XII covers a period of 6-3/4 years from the initiation of the program to the completion of the documentation of the flight test results. The testbed program consists of a phased arrangement of seven technical tasks and an overall management task. These tasks are as follows: ADVANCED TURBOPROP TESTBED PROCRAM YEAR PROGRAM PHASE/MILESTONES/EVENTS I 3 4 2 5 6 7 8 10

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IIIII IIIII II II III II IIIII

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, iIIINTRACT ADVANCED TURBOPROP TESTBED SYSTEM .- fLUTTER ANALYSIS GIl ONLY. WS&C ANALYSIS TECHNICAL ANALYSES PIIASE I :::J YCONflGURAT JON DEFINED PRELIMINARY DES U;N PIIASt: I I TECHNICAL ANALYSES, PDR ~r .DDR DESIGN ,Ir DRIVE SYSTEM RECEIVEDI LOCKllEED ORGINGS,. REC,.

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FABR rCATION AND DRIVE SYSTEM INSTL.

Q.E.C. FAB AND ASSY COMPLETE I PIIASE III TECHNICAL ANALYSES AIRCRAFT MOD.

PIl~'---'IJDR _ A'OUSTIC MODS DETAIL DESIGN AIRCRAFT/MODS WING/NACELLE R ~TESIBED AIRCRAFT ~EC. tOCKllEED <j FABRICATION AND MODIFICATION PROP-Fji REC. LOCiHEED~ DRIVE SYSTEM INSTALLEIl ON TESTBED 10 PIIASF. IV NACELLE PREP. AND PROP-FAN INSTL II TEST STAND PREP.

DRIVE SYSTEM TESTS COMPLE~ DRIVE SYSTEM STATIC TEST

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fUNCTIONAL TESTS COMPLETE I PIIASE V 1l SYSTEM FUNCTIONAL TEST I , [I, TECHNICAL SUPPORT PIIASE V I I I ~ PI.ANNING INSTRUMENTATION DESIGN , GROUND TESTS C,MPLETE 16 GROUND TESTS AIRWORTHINESS TEST COMPLETE Fl.l1TTER AND AIRWORTHINESS TESTS ..

. I FIRST FlIGHT TEST • FI. H:IIT .TEST COMPI.,TE 18 TESTBFJ! FLIGHT TESTS 'Y AIlVANCED TURBOPROP TESTBEIJ PHASE VII 1'1 PROGRAM HANAC:E11ENT

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VI TABLE XII. ADVANCED TURBOPROP TESTBED SYSTEMS PROGRAM SCHEDULE o Phase I - Testbed Aircraft Definition and Development - This phase is essentially the Preliminary Design Phase in which the configuration development and geometric description of the aircraft nacelle lines and scantlings will take place. A structural analysis, including a rigorous flutter analysis, if necessary, to verify nacelle location and to si ze structural members for design, and weight and balance checks will be conducted. A detailed performance analysis to establish drag levels to predict the mission performance of the test- bed aircraft would also be conducted. Propulsion System Analysis to provide air induction system design data, installed drive system performance and subsystems installation design data will be carried out. Stability and control analysis to determine stability and control characteristics of the testbed aircraft with the prop-fan drive systems installed will be conducted. Finally, the Aircraft Modification Design Analysis to establish the preliminary design of the QEC, wing and aft nacelle structures and aircraft subsystems will be performed.

o Phase II - Drive System QEC Development Design and Fabrication - Thi s phase will produce an airworthy drive system for test-stand and test- bed application by performing analyses and design of the drive system QEC, which will include optimization of the air induction system, drive system performance predictions, flow-field analyses of the in- stallation, stability and control checks of the effects of the QEC design, and the detailed design of the QEC nacelle and mounting structures, which will include structural analyses of the QEC nacelle structure. At the conclusion of the analyses and design, the QEC parts will be fabricated and assembled, resulting in a flyable QEC drive system.

o Phase III - Design and Fabrication of Required Aircraft Modifications- This phase will take the selected airframe and convert it to the testbed configuration.

This will involve detailed analysis of all affected structures and systems including, not only the propulsion system and its related controls, but also the aircraft flight control system. This analysis will be conducted so that detailed design of the aircraft modification can be performed concurrently. Hardware will be fabricated as soon as design validation permits.

o Phase IV - Static Test of the QEC on the Engine Test Stand This phase consists of the static test of the flyable drive system designed and assembled in Phase II to establish operational readiness of the QEC and systems before assembly to the testbed aircraft in Phase III.

The test stand schedule will be arranged so that the engine tests will be concurrent with the Phase III work and will be completed to coincide with the aircraft modification for final installation on the testbed aircraft.

o Phase V - System functional test will be performed, following comple- tion of the Phase III assembly, to etablish the performance of all the systems associated with the operation of the testbed aircraft, prior to conducting airworthiness tests of the modified aircraft.

o Phase VI - The final phase of the testbed program will be the testbed aircraft flight test program which will consist of the airworthiness flight test, followed by the Advanced Turboprop Testbed Flight Test Program.

o Phase VII - Program Management - This phase covers the entire time span of the testbed program.

Details of the Task VI-Testbed Program Costs and Schedules is to be found in Volume II of this report.

SUMMARY OF STUDY RESULTS AND DISCUSSION The study results are summarized and discussed in the following text: o Program Objectives and Priorities necessary to enhance industry acceptance of the prop-fan for commercial aircraft and establish technology readiness have been identified and defined for a number of critical technology areas. A total of 30 objectives have been identified as crucial to the continued development of prop-fan propulsion. Furthermore examination of the objectives shows that the majority of the objectives are best approached through the medium of a test-bed aircraft.

o The survey of drive systems reveals the scarcity in the U.S.A. of power sections at power levels of 2984 kW (4000 shp) and above. This imposed constraints on the testbed aircraft system by limiting prop- fan size to that required by the highest power level available; that of the DDA XT701 at 6018 kW (8071 shp), which imposed restrictions on the testbed aircraft configuration. A prop-fan drive system can be assembled from existing hardware suitably modified. The drive system installation design approach is based on the Quick Engine Change (QEC) concept. Use of the Lockheed P-3C nacelle structural components re- sults in an overdesigned nacelle which renders the drive system in- dependent of the receiving airframe. The nacelle contours and structural arrangements are simple and the nacelle contours developed apply to either a "Pinion-high" or "Pinion-low" arrangement. The drive system recommended for testbed application is a "Pinion-high" configuration consisting of the DDA XT701 in combination with a DDA T56-A-14 gearbox driving an eight-blade prop-fan 2.89 m (9.5 ft) in diameter. The "Pinion':"high" configuration was selected as representa- tive of commercial aircraft overwing nacelle installations. The study shows that the XT701 supervisory electronic control and the Hamil ton Standard 54H60 propeller control can both be modified for testbed air- craft application.

o Aircraft suitable for conversion to prop-fan testbed configurations exist and of those examined the GAC GIl and the Boeing KC-135A offer the best chance of achieving success. In the case of the GAC GIl the airframe and the prop-fan propulsion system are very closely matched.

The modification of the GIl wing is, however, much greater than that required for the KC-135A. Although the GII flutter characteristics appear marginal for the two prop-fan locations examined, further analysis at other locations between the two investigated is expected to show that a location exists at which the wing flutter charac- teristics are acceptable. Either of the prop-fan testbeds are capable of performing to the test requirements, first to qualify the prop-fan, and secondly to establish the near- and far-field noise charac- teristics and the design of suitable cabin noise attenuation concepts.

The design of the testbed aircraft configuration is limited by the lack of adequately sized propulsion, i.e., it is not possible to substitute a prop-fan unit capable of generating the thrust of a turbo-fan for an existing primary engine. The twin prop-fan approach adds an element of safety to testbed aircraft since the basic performance of the aircraft is not degraded. Airport operations would, therefore, be conducted using the primary propulsion.

The study further shows: o The data base for contouring prop-fan nacelles to operate at Mach numbers up to 0.80 to be inadequate. The study nacelles were generated from data from an axi-symmetric non-flow through propeller test rig. To design a highly non-symmetric nacelle testbed nacelle for integration with a wing requires more and continued research to establish design parameters and nacelle configurations.

o A prop-fan of larger scale than the present experimental hardware is needed to demonstrate manufacturing feasibility of the spar-shell structural concept and achievement of the proper mass and stiffness distributions.

o There is considerable doubt about the magnitude of the swirl effects and of the means by which these effects can be improved.

o Wind tunnel testing of the full size drive system prior to flight test is not recommended because: • The available wind tunnels are not capable of simulating the prop-fan design point environment in terms of dynamic pressure, Mach number and temperature.

• In most tunnels the tunnel flow solid wall blockage limits are found to be exceeded with the testbed drive system nacelle, prop-fan and a section of wing installed.

• Low speed wind tunnel testing does not directly address the design point of the prop-fan.

• The costs of wind tunnel tests would be high relative to the usefulness of the data obtained.

Overall, the Advance Turboprop Testbed Program described in this report is shown to be an effective means by which the Technology Readiness of Prop-Fan Propulsion can be established.

CONCLUSIONS A number of conclusions have been reached from the results of this study, and each is presented in a single highlighted statement followed by a brief discussion of justification.

o Review of the Testbed Program Objectives and Priorities shows the Testbed Aircraft as the best means of expediting problem solutions.

Thirty program objectives in four areas of technological concern require attention to establish prop-fan technology readiness and to enhance industry acceptance of the prop-fan concept. The majority of the objectives are found to be best addressed through the medium of a flying testbed aircraft.

o A Suitable Powerplant and Gearbox - DDA XT701/T56-A-14 - can be assembled to drive a prop-fan about 3.0m (9.5 ft) in diameter Three power sections in excess of 2984 kW (4000 shp) at sea level were examined for testbed application. Of the three, two were found to be unacceptable or marginal for prop-fan diameter which was fixed at a minimum of 2.42 m (8 ft). In addition to the larger prop-fan diameter attainable wi th the DDA XT701 combination, a further ad- vantage is the free turbine design which allows a large measure of flexibility during flight testing. In the event that the U.S. Army owned DDA XT701 engines are not available for the testbed program, the industrial version of this engine, the DDA Model 570, could easily be converted to flightworthy status.

The drive system integration with the testbed aircraft need not be a completely optimized design aerodynamically so that the installation could be designed as a Quick Engine Change unit.

The nacelle installation can be designed to be independent of the receiving air- frame which is accomplished by over-designing the drive system support structure through the use of Lockheed P-3C "Orion" V-Frame structures.

o Two Aircraft - The Boeing KC-135A and the GAC GIl - are attractive vehicles for conversion to prop-fan testbed aircraft.

Six candidate aircraft were examined for testbed application - the Lockheed C-141A "Starlifter" and "JetStar" -6, the Boeing 737-100 and KC-135A tanker, the Convair 990 and the Gulfstream American "Gulf- stream II." Of these the Boeing KC-135A and the GAC GIl, the two selected aircraft, were the aircraft that best met the design condi- tions, provided sufficient test mission duration and a stable platform for testing, were the most sui table for acoustic investigations and easily modified to the testbed configuration. In the case of the GAC GIl, a wing could be made available for modification and propulsion system installation ahead of the aircraft modification. This would reduce the time during which a GIl aircraft would be out-of-service awaiting modification. The modified spare wing complete with the propulsion system would replace the wing on the aircraft to be modi- fied. At the end of the program the original wing would replace the modified wing and the GIl restored to the original configuration.

o Program should proceed without delay if prop-fan propulsion is, to be considered for inclusion on the next generation commercial passenger short/medium range transport aircraft The testbed aircraft pro~ram requires an elapsed time of 6-3/4 years. During this time the 2.84 m (9.5 ft) diameter prop-fan and the XT701 drive systems are to be manufactured and developed, an airframe is to be acquired and modified, the aircraft flight and acoustic tests are to be completed and the test data reduced. A new short/medium range aircraft for transporting 100/150 passengers with an IOC date of 1990 is foreseen. This would require a design commitment to prop-fan propulsion about 1987 if Advanced Turboprop Propulsion is to be considered as an alternative to turbofan propulsion. The Prop-fan Testbed Program should therefore proceed without delay. This program is of such great importance to potential users in terms of economic benefit that the schedule should in fact, be accelerated.

o Testbed aircraft program outlined provides a cost effective means of verifying prop-fan performance The use of existing power plants, gearboxes, nacelle support structures and airframes and the minimization of modifications ensures No costly a program total cost at the lowest possible level.

turbo-machinery or airframe development is required and both a Boeing KG-135A or a GAG GIl could be made available for the program. The GAG GII is particularly attractive from the cost stand-point because of the availability of a spare wing.

o Wind Tunnel tests of limited value Wind tunnel model tests of testbed aircraft configurations for airworthiness and flutter evaluation can be performed. The wider range of prop-fan experimentation is not amenable to full scale drive system tests in wind tunnels. Such tests can only be accomplished in a limited number of facilities none of which can simulate the prop-fan flight environment. These tests would be restricted by tunnel size limitations which would give rise to severe blockage problems, and by

lack of proper simulation of the flight environment, i.e., M = 0.80 at

an altitude of 10668 m (25,000 ft). Furthermore the lack of anechoic facilities required for near-field acoustic testing in the flight en- vironment also precludes wind tunnel testing for noise charac- teristics.

The effect overall is to limit the usefulness of wind tunnel testing to performance data and provide a small amount of validation of the prop-fan structural characteristics.

APPENDICES

APPENDIX A - TESTBED PROGRAM OBJECTIVES AND PRIORITIES - TASK I The scope of the Testbed Program Objectives and Priorities study was confined to the following four areas of technological concern: o Integrity of the Structure o Acoustic Environment o Aircraft Performance o Functional Systems Operation and FOD Vulnerability INTEGRITY OF THE STRUCTURE Although considerable progress has been made in the last five years it is recognized that the largest prop-fans made so far have diameters of 0.62 m (2.04 ft) and have never been tested in a realistic flight environment and. further- more. have never been tested with an actual turboprop drive system. Before manufacturers and users can commit prop-fan propulsion systems to aircraft design the Integrity of the Structure. both of the prop-fan and of the airframe.

must be verified to establish that large scale prop-fans can be built light enough for flight hardware and can be made stiff and strong enough to sustain the dynamic loads to which they will be subjected.

Propeller Structural Integrity and Dynamics Three technological aspects that concern the propeller and need attention are: (a) the blade dynamic response to the aerodynamic flow field. (b) blade stall and classical flutter characteristics. and (c) the critical speed deter- mination and verification of hub stiffness.

Blade Dynamic Response - Blade dynamic response is a function of the aero- dynamic flow field. the blade aerodynamic characteristics and the blade struc- tural dynamic characteristics. Wind tunnel tests of the 0.62m (2.04 ft) dia- meter prop-fan models should provide data on the first two considerations but would lack proper simulation of the structural dynamic characteristics of large spar/shell blades. To generate the proper flow field. the vibratory response testing must be performed in the presence of a swept wing, nacelle and fuselage combination, sized to be representative of a testbed aircraft.

OBJECTIVE 1: CONFIRM THE SMALL-SCALE TEST EXCITATION LOADINGS IN THE PRESENCE OF A REALISTIC FLOW FIELD AND MAKE AN ASSESSMENT OF THE STRUCTURAL RESPONSE OF LARGE-SCALE BLADES.

APPROACH: Perform tests using a testbed vehicle to evaluate the excitation and overall re- sponse of large-scale prop-fan blades.

Testing should include a full range of speeds and altitudes from static to Mach 0.8 cruise, a full range of ground wind veloci ties and directions in conjunction with representative thrust levels, a full range of wing angle-of-attack with and without flaps and a range of yaw angles.

Blade Classical Flutter - The possibility of classical flutter of prop-fan blades 1s of concern because of the high degree of modal coupling due to the sweep and low aspect ratio, the relatively low first torsional mode frequency and the high operating tip speeds. The susceptability of a blade to classical flutter is dependent both upon the aerodynamic and structural characteristics of the blade. Blade aerodynamic characteristics can be represented by small blades, but structural characteristics can only be accurately simulated at large scale so that flutter testing must be conducted using large scale spar/shell construction blades.

OBJECTIVE 2: CONDUCT BLADE CLASSICAL FLUTTER VALIDATION USING A TESTBED VEHICLE.

APPROACH: Perform tests at high Mach numbers over the full range of operating conditions monitoring stresses and frequencies for indications of the approach of classical flutter.

Blade Stall Flutter - During takeoff and reverse thrust operations, the highly loaded prop-fan blades are largely in a stalled condition and are, there- fore, susceptible to stall flutter. Duplication of the true aeroelastic and geometric characteristics as well as torsional frequency requires the use of large-scale blades. Since stall flutter is most likely to occur at static or low speed at high power conditions, a testbed vehicle would be the best means for conducting tests.

OBJECTIVE 3: EVALUATE PROP-FAN STALL FLUTTER CHARACTER- ISTICS THROUGHOUT THE CRITICAL OPERATING RANGE TO ESTABLISH STALL FLUTTER BOUNDAR- IES.

APPROACH: Perform blade stall flutter validation using a large-scale prop-fan. Monitor blade torsional stresses for a range of operating conditions and estimate the stall flutter boundary.

Cri tical Speed and Hub Stiffness - Any blade rotating assembly should be examined for speed criticality, since propeller blades exhibit several modes of resonant vibration over their operat:i,ng range. The range of blade frequencies of interest is determined by the number of periodic forcing functions possible and by the strength of the excitations. Prop-fans of eight or more blades may have as many as five significant excitations per revolution (5P) so that excita- tions beyond 5P need not be considered. The possibility of excitation at reson- ance within the operating range will be indicated by the intersection of blade natural frequencies of the first,. second, and third modes with the integer order excitation lines.

OBJECTIVE 4: VALIDATE PROP-FAN CRITICAL SPEED AND THE HUB AND RETENTION STIFFNESS OVER THE FULL RANGE OF OPERATING RPM.

APPROACH: Perform critical speed determination using a testbed aircraft prop-fan installation by establishing the blade frequencies for excitations up to 5P. Determine hub and retention stiffness for these conditions.

Propeller-Induced Vibrations and Dynamics All propellers. whether of conventional or advanced design. when mounted in front of a wing experience cyclical loadings due to the flow field generated by the presence of the loaded wing and of the adjacent components such as the fuse- lage and other nacelles. When mounted on a swept wing and operated at high sub- sonic Mach number. the flow field becomes complex and unsymmetrical and induces unusual dynamic loadings on the propeller and. therefore. on the power plant and aircraft structure. Unsteady random and periodic forces can be imposed on the prop-fan by the ground plane. fuselage exterior. swept-wing leading edge. na- celle and engine air inlets. adjacent propellers and nacelles. oblique flow due to yaw. angle-of-attack. crosswind. and other factors. Unless taken into ac- count during design and development. these forces could cause structural failure of the prop-fan and/or the airframe.

Flutter and Dynamic Loads - The· use of prop-fans on high-sub sonic-speed transports introduces the possibility for two types of wing flutter problems: (a) whirl flutter. and (b) a reduction of wing flutter stability. Although both these phenomena can occur on conventional propeller driven aircraft. the higher operating Mach numbers for prop-fans are expected to adverselY al"[t::Cl, ~ill:: 3ta- bility of these modes. Although these two flutter phenomena strictly cannot he separated. whirl flutter stability problems can be avoided by providing adequate mounting rigidity for the propulsion installation.

Propeller aerodynamic and gyroscopic forces and moments are known to reduce wing flutter speeds significantly. The degree to which this occurs is strongly dependent upon configuration. but powerplant spanwise and chordwise location re- lative to the wing are important parameters affecting the degree of coupling between the wing and propulsion system. Performance considerations may require the prop-fan installation to be located farther forward on the wing than exist- ing turbofans, which would tend to increase the coupling with the flexible wing loads.

The whirl flutter and wing flutter coupling are both dependent upon propel- ler unsteady normal forces and moments associated with angle-of-attack changes.

No steady or unsteady normal forces and moment data have been measured for prop- fans, but the coefficients are expected to be significantly higher than those of conventional propellers due to the higher Mach numbers at which the prop-fans operate.

OBJECTIVE 5: OBTAIN AERODYNAMIC DATA BY WIND TUNNEL TESTING TO EVALUATE AND PROVIDE A BASIS FOR FLUTTER, PERFORMANCE AND STABILITY AND CONTROL ANALYSES.

APPROACH: Perform wind tunnel tests using scale models to obtain side-force and moment variations with angle-of-attack. Although unsteady aerodynamic derivatives are need- ed for flutter analyses, the reduced fre- quencies associated with potential for whirl flutter and wing flutter instabili- ties are· qui te low, and a quasi-steady application of steady deri vati ves should provide sufficient accuracy.

Side force and moment data will be measured for a range of Mach numbers, advance ratios, and angles-of-attack suf- ficient to cover the predicted operating envelopes of future prop-fan aircraft, in- cluding overspeed conditions required for flutter evaluation. The instrumentation should be capable of isolating two-axis forces and moments on the prop-fan, ex- cluding those on the nacelle and wing section, if used.

Perform flight tests of the prop-fan testbed aircraft to verify the data mea- sured in the wind tunnel tests. Since the wind tunnel test data will be affected by wall reflections, flight test data should be measured for a few selected conditions and used to verify or adjust the wind tun- nel test data. Instrumentation sufficient to measure side force and pitching moment due to aircraft sideslip would be provid- ed.

Propeller-Induced Vibration - Sound pressures radiating from the prop-fan disc and fluctuating pressures in the prop-fan wake will excite resonances in the airframe structure, and/or drive the structure at non-resonant conditions at potentially destructive amplitudes that will require preventive design. Al- though the technology is available to deal with the design problem, the analyti- cal tools for defining the environment and quantifying the vibratory strain and acceleration amplitudes are not precise enough to avoid large-scale testing.

Evidence is required to confirm that an acceptable vibratory fatigue life can be obtained for aircraft structures fabricated from state-of-the-art materi- als placed near a prop-fan.

OBJECTIVE 6: DETERMINATION OF THE FUSELAGE, WING, AND EMPENNAGE STRUCTURAL VIBRATION SPECTRA, AND THE SKIN AND SUBSTRUCTURE DYNAMIC STRAIN SPECTRA.

APPROACH: Analytically or empirically derive the resonant structural response amplitudes that are induced by the fluctuating pres- sure environment generated by the prop- fan. Validate the derivations with vibra- tion and strain measurements from the testbed aircraft.

Prop-fan Drive System Dynamic Loads and Induced Effects - A structurally safe testbed aircraft is a prime consideration in establishing the testbed air- craft program objectives. Although the structural integrity of the prop-fan is mainly the concern of the prop-fan manufacturer, the influence of the flow field of the installed testbed prop-fan propulsion system must be considered.

Unsteady random and periodic forces can be imposed on the prop-fan by the ground plane, fuselage wall, swept-wing leading edge, nacelle and engine air in- let, adjacent propellers and nacelles, oblique stream due to yaw, angle-of- attack, crosswind and other factors. Unless taken into account during design and development, these factors could cause structural failure of both the prop- fan and the airframe.

The problems associated with accurate prediction of blade and shaft stress- es, corr~lation of analysis with measured data and of the test instrumentation and conditions must also be considered.

OBJECTIVE 7: DETERMINATION OF PROP-FAN DYNAMIC AND IN- DUCED LOADS TO ASSURE STRUCTURAL INTEGRITY OF THE TESTBED SYSTEM.

APPROACH: Data from high- and low-speed wind tunnel tests, static tests and from tests already completed will be used to perform analyses of the testbed installation to validate structural integrity.

The results of the analysis will be corre- lated with measurements from the testbed aircraft in an operational environment.

Scale Effects One of the questions of great significance in developing the prop-fan pro- pulsion system relates to scale effects. So far, it has been demonstrated that a prop-fan 0.62 m (2.04 ft) in diameter when tested in a wind tunnel develops sufficient thrust to lend confidence that the design goals for propulsive efficiency can be met. Other test programs, using prop-fans at the same scale, will continue to provide data on the effect of forward velocity on near-field noise and on propulsive efficiency of a representative assembly of wing, na- celle, and prop-fan when tested in a transonic wind tunnel.

Current studies of the feasibility and economics of prop-fan propulsion in- dicate that large-scale prop-fans will be in the size range of 3.66 to 4.87 m (12 to 16 ft) in diameter. The testbed prop-fan, on the other hand, will be constrained to a diameter of 2.44 to 3.048 m (8 to 10 ft) by the power available from available drive systems. This prop-fan size, relative to the 0.62 m (2.04 ft) diameter prop-fan previously tested, is expected to provide a good basis for the evaluation of scale effects.

Technology areas in which scaling effects are likely to be encountered in prop-fan design can be subdivided into the following general and specific areas: 1. Propeller structural integrity and dynamics scale effects a. Blade strength and stiffness b. Blade structural dynamics 2. Propeller generated near and far field noise scale effects a. Near- and far-field noise prediction b. Cabin-noise attenuation 3. Installed efficiency and interaction scale effects a. Prop-fan/spinner/nacelle interaction b. Slipstream/wing flow interaction 4. Large scale drive system scale effects.

Propeller Structural Integrity and Dynamics Scale Effects To establish the most accurate test data possible and to avoid additional analytical correla- tion studies, the large-scale.prop-fan diameter should be of the order of 2.44 to 3.048 m (8 to 10 ft). The selection of a 2.44 to 3.048 m (8 to 10 ft) dia- meter for the testbed aircraft arises from two considerations: 1. Accurate representation of the total blade airfoil mass and stiffness distributions in the spanwise and chordwise directions, as well as the proportioning of the mass and stiffness contributions of the elements making up any given cross-section of blade airfoil.

2. Accurate representation of size, shape, and thickness of the blade con- struction elements, so that a clear verification of full-size fabrica- tion feasibility will be made.

The results of the SR-3, SR-5, and SR-6 aeroacoustic model designs have demonstrated that the thin, swept blade shape increases the degree of mass- stiffness interaction due to rotation and vibration. The response of a blade to integer-order excitation is related to its frequency and damping. The frequency is determined by the mass and stiffness distribution, and the damping is related to the deflection amplitude and, therefore, the stiffness. The probability of non-integer order response is related to the relative magnitude of the airloads and blade inertia, blade and mode shapes, and the separation of torsional and bending frequencies. The blade inertia, relative location of the blade fre- quencies, steady deflections of a rotating blade caused by body forces, and aerodynamic forces are all determined by the mass and stiffness distribution.

The integer-order response, freedom from non-integer-order response (flutter), and predictable deflection characteristics are essential elements of a full- scale evaluation.

The accuracy of the simulation of a large scale prop-fan blade is size- dependent, since the blade will be made of several materials of different den- sities to reduce weight. Since there are practical limitations on the thickness of blade parts, both from the fabrication and durability standpoints, it is not possible to simulate full-size, . cross-sectional properties in sub-scale size.

For example, in order to withstand airloads, buckling, panel flutter, and FOD with a hollow-blade tip cross-section, the minimum required skin thickness on the pressure side would be 0.152 cm (0.060 in) to 0.203 cm (0.080 in). Scaling this thickness directly with diameter from 3.048 m (10 ft) to 0.62 m (2.04 ft), the thickness would be 0.0305 to 0.0381 cm (0.012 to 0.015 in). Since most com- posite laminates are about this thickness, multilayer laminates, which are ne- cessary to achieve required strength and stiffness properties, are rUled out.

Fabricating a blade skin from thin sheet metal would require completely differ- ent techniques than would be applied to a full-scale blade.

In the area of blade retention, similar scaling limitations are encoun- tered. In order to reverse thrust, blade pitch must be variable. An antifric- tion bearing is required for variable pitch. The area available for the reten- tion and pitch control mechanism is fixed by the hub-to-tip diameter ratio re- quired for aerodynamic performance. The cross-section of antifriction bearings and pitch control elements, gears, ballscrews, links, rod ends, and slider blocks cannot be scaled down below a certain point because of fabrication and durability characteristics.

From design work on SR-3, SR-5, and SR~6 model prop-fans, all of which had solid metal blades without antifriction retention bearings, it was concluded that an accurate demonstration of the dynamic behavior and fabrication feasibil- ity of a large scale prop-fan could not be achieved in a diameter of less than 2.44 to 3.0a4 m (8 to 10 ft).

Propeller-Generated Near- and Far-Field Noise Scale Effects - No data are currently available to demonstrate that propeller noise data can be accurately scaled from one diameter to another. Analytical scaling techniques have been developed, but have not been adequately validated. The testbed program will provide valuable data toward that end. The near-field noise data comparison of a nominally 2.44 m (a ft) diameter prop-fan with data from the upcoming Jetstar testing of the 0.62 m (2.04 ft) diameter prop-fan will help to validate scaling effects methods. This validation, in turn, will permit more confident pre- diction of noise from full-scale prop-fans.

Installed Efficiency and Interaction Scale Effects - While it has been de- monstrated that small scale prop-fans closely approach the aO-percent propulsive efficiency set as a performance goal, the same tests also show that installed propulsive efficiency is strongly dependent on the optimization of the propul- sion system and the wing interaction. As will be detailed in later sections, the testbed concept does not lend itself to aerodynamic/propulsion optimization.

The aerodynamic/propulsion optimization work can best be done in the more flexible environment of a wind tunnel test program, and it is anticipated that results from such a wind tunnel test program could be scaled with a high level of confidence.

In the area of scale effects, it is therefore concluded that: o The basis for determining the minimum prop-fan diameter needed for the testbed aircraft is that the blades be large enough to allow the same structural design concepts as are anticipated for large-scale applica- tions.

o Noise data from the testbed prop-fan with a diameter in the range pro- posed 2.44 to 3.048 m (8 to 10 ft), will be a most valuable addition to the state-of-the-art. Correlation with the 0.62 m (2.04 ft) diameter prop-fan data and theory will provide a basis for scaling to larger diameters.

o The aerodynamic/propulsion optimization that is needed should be ob- tained from wind tunnel tests. No major problems are anticipated in scaling such data to larger-scale designs.

OBJECTIVE 8: VALIDATE AND/OR DEVELOP AERODYNAMIC, ACOUSTIC AND STRUCTURAL SCALING LAWS FOR LARGE-SCALE PROP-FAN INSTALLATIONS.

APPROACH: a. Analyze aerpdynamic, acoustic and struc- tural data from all the technical inves- tigations described in subsequent sections and correlate with data from other experi- mental programs and with existing analyses in order to develop the methodology for the design and characteristics of prop- fans in the diameter range of 3.66 to 4.87 m (12 to 16 ft).

b. Perform flight test of the testbed instal- lation in an operational environment and correlate flight test data to validate scaling laws.

OBJECTIVE 9: DETERMINATION OF BLADE MASS AND STIFFNESS DISTRIBUTIONS AND AEROELASTIC CHARACTERIS- TICS. VALIDATE BY CONDUCTING TESTBED AIR- CRAFT FLIGHT TEST.

APPROACH: Conduct design studies for a range of prop-fan diameters and obtain blade mass and stiffness distributions that satisfy the design requirements. Verify by corre- lating the an.alytical results with flight test data using a testbed aircraft.

OBJECTIVE 10: VALIDATE THE MANUFACTURING FEASIBILITY OF THE SPAR/SHELL CONCEPT FOR PROP-FAN BLADES.

APPROACH: Conduct manufacturing investigations of blade construction for various prop-fan diameters. Establish the practical size limitations and verify manufacturing fea- sibility by fabricating and testing the blades.

Large-Scale Drive System Scale Effects - The largest currently available propeller drive systems generate in the region of 3128 to 4413 kW (5000 to 6000 shp). Studies of the power requirements for future aircraft indicate a need for core engines developing 11,184 kW (15,000 shp) and reduction gearboxes capable of transmitting high levels of torque. The size of the core engine and config- uration of drive system relative to the type of gearbox used affects the aero- dynamic and structural design of the nacelle and gearbox reliability affects the economic viability of an aircraft.

OBJECTIVE 11: ESTABLISH THE FEASIBILITY OF DRIVE SYSTEMS OF 11,184 kW (15,000 SHP) AND ABOVE FOR FUTURE AIRCRAFT IN A COMMERCIAL ENVIRON- MENT.

APPROACH: Perform design studies of engine cores and gearboxes to establish the feasibility of drive system design and the impact of system reliability on aircraft economics.

Formulate drive system sizing laws for use in aircraft system studies.

ACOUSTIC ENVIRONMENT In spite of the fact that acoustic tests have been made on several prop-fan configurations uncertainty still exists about prop-fan noise. Existing test facilities cannot simulate the high transonic cruise speed in an anechoic en- vironment and the conclusion that prop-fans can operate with about 136 dB SPL at the fuselage side is based on extrapolation from low speed tests. The questions to be resolved as:':far as near-field noise is concerned is whether the desired cabin interior noise attentuation can be achieved without incurring weight penal ties that would offset the performance gains of the prop-fan. Far-field noise characteristics are also of concern since compliance with FAR Part 36 requirements for community noise levels must be verified if the prop-fan concept is to gain acceptance.

Propeller-Generated Near-Field Noise Since the prop-fan will be a major cause of cabin noise and vibration, a comprehensi ve understanding of both the noise generated by prop-fans and the relationship of this generated noise to the prop-fan principal parameters is essential.

The determination and evaluation of the prop-fan noise characteristics should use data obtained from sources such as the JetStar model tests, wind tunnel tests, and testbed model and ~ull scale tests, in order to obtain comparisons of analytical predictions.

Near-field noise level prediction methodology is inadequate in several respects such as accounting for wave propagation over a curved surface, cancel- lation and reinforcing from multiple sources, synchrophasing, effects of forward motion on surface reflection, angle-of-incidence and propagation path.

The inadequacy of current theoretical and analytical prediction techniques require the use of a testbed to quantify the near-field noise environment as well as to validate or modify analytical methods. The objectives of the testbed program will therefore include consideration of the near-field noise technology.

OBJECTIVE 12: DETERMINE THE SOUND-PRESSURE DIRECTIVITY AND THE SPECTRA VARIATION AT EACH MULTIPLE OF THE BLADE PASSAGE FREQUENCY FOR EACH PROP-FAN NOISE SOURCE RADIATED LATERALLY ONTO THE FUSELAGE AND FOR FLUCTUATING PRESSURES CONVECTED REARWARD IN THE PROP- FAN SLIPSTREAM IMPINGING ON THE WINGS AND EMPENNAGE.

APPROACH: Measure the sound pressure spectra in a spatial array using a prop-fan testbed Derive analytically, where aircraft.

methods exist, the variation in sound pressure spectra with blade load and thickness, number of blades, helical tip Mach number, blade stall characteristics, in-flow angle in pitch and yaw, and inter- ference effects from wings, nacelle in- lets, and adjacent components of the air- craft. Obtain the same spectra from the testbed aircraft by direct measurement and correlate the analytical and experimental results and demonstrate the adequacy of the analytical methods.

OBJECTIVE 13: DETERMINATION OF THE SOUND PRESSURE LEVELS OVER THE FUSELAGE PRESSURIZED SURFACE AREA.

Measure sound pressure spectra on the APPROACH: : testbed aircraft fuselage surface for the first 10 multiples of the blade passage frequency, while varying operating condi- tions such as speed, altitude, horsepower, and synchrophasing.

Correct the data to large-scale prop- fan applications using the scaling rela- tions from Objective 8 and the parametric relations determined in Objective 12.

Analytically account for multiple prop- fans rotating in a fixed optimum phase relation OBJECTIVE 14: DETERMINATION OF THE STRENGTH AND DIREC- TIVITY OF THE NOISE FROM BLADE VORTICES, SEPARATED FLOW TURBULENCE, AND BLADE THICKNESS AND LOADING AT SELECTED LOCA- TIONS OF THE WING AND FUSELAGE FOR DESIGN MODIFICATION TO TESTBED SIZE PROP-FANS.

Derive the .quantities analytically, and APPROACH: substantiate the derivations with measure- ments from the testbed aircraft.

OBJECTIVE 15: DETERMINATION OF THE FULL-SCALE AIRCRAFT FLUCTUATING PRESSURE SPECTRA WHERE THE PROP-FAN SLIPSTREAM IMPINGES ON THE WINGS AND EMPENNAGE.

APPROACH: Correct the fluctuating pressure spectra obtained on the testbed aircraft in Ob- jective 13 to full-scale aircraft con- ditions using the scaling relations de- III Quantify the termined in Objective 8.

variation of the fluctuating pressure spectra with forward airspeed and rearward convection distances for the prop-fan vortices, the blade thickness and loading noise, and the blade flow separation tur- bulence during takeoff roll.

OBJECTIVE 16: DETERMINATION OF THE EFFECTS OF FUSELAGE SURFACE CURVATURE ON THE STRENGTH OF THE SOUND PRESSURES PROPAGATING OVER THE FUSE- LAGE, AND ON THE PRESSURE DOUBLING AT THE SURFACE.

Establish the effects analytically, and APPROACH: obtain substantiating measurements from the testbed aircraft.

OBJECTIVE 17: FOR A FULL SCALE AIRCRAFT WITH ALL PROP- FANS OPERATING, DETERMINE THE GEOMETRY OF THE FUSELAGE SURFACE AREA WITHIN WHICH SOUND PRESSURES ARE SPATIALLY CORRELATED AT MULTIPLES OF THE BLADE PASSAGE FRE- QUENCY.

Derive the area geometries analytically APPROACH: for the prop-fan testbed, then measure spatial correlation at selected critical locations on the testbed and compare to the analytical derivations. Establish suitable accuracy of the analytical methods; then derive the geometries of the areas of correlated pressures for a full scale aircraft with all prop-fans Propeller-Generated Far-Field Noise The far-field noise characteristics of a prop-fan powered aircraft consti- tute an area of technological concern. Current prop-fan noise prediction meth- odology is based upon an extension of propeller theory and on the noise measure- ments from small scale prop-fans operating in a low forward speed environment.

Propeller and prop-fan noise signatures cannot be measured with accuracy on a static test rig since the blades are usually stalled at useful disk loadings and the noise emissions change drastically with forward speed.

While forward speed requirements can be simulated in large-scale wind tun- nels, other constraints limit the usefulness of the noise measurements. Reflec- tions from the walls of the tunnel interfere with noise measurement by reinforc- ing or cancelling the signals at certain frequencies. It is usually not feasi- ble to line the tunnel walls with sound-absorbent material, especially if high speeds are involved. In addition, the tunnel generated noise is high in the low-frequency range which would seriously interfere with measurement of the prop-fan fundamental and low harmonic orders.

OBJECTIVE 18: VERIFY THE PROP-FAN FAR-FIELD NOISE RE- QUIREMENTS AS STATED IN FAR PART 36 CERTI- FICATION TEST.

APPROACH: Perform measurement of the noise of a large-scale prop-fan installed on a test- bed aircraft at representative flight speeds. The testbed aircraft and drive system should be such that there will be no excessive interference with the noise emitted by the prop-fan.

To ensure reliable noise measure- ments, the following requirements must be met: a) The testbed aircraft should be capable of operation with the non-prop-fan engines throttled back.

b) The non-prop-fan engines should not contribute significantly to noise in the frequency range under considera- tion, Le., 150 to 500 Hz. In this respect, a typical turbofan generates fan noise above 2 KHz, while low-fre- quency jet noise is of low magnitude.

c) The prop-fan drive system must also comply with the requirement of non- interference with the prop-fan noise signature.

Passenger Cabin Noise and Vibration Passenger cabin noise and vibration levels are among the principal areas of concern in the application of the prop-fan to commercial passenger transport aircraft. Passenger comfort and public acceptance of the advanced turboprop require that the levels of noise and vibration in the passenger cabin of a prop- fan-powered aircraft be no greater than those in contemporary turbofan aircraft.

Cabin noise and vibration levels are strongly influenced by: o The frequency, strength, and incidence of the propeller sound pressures.

o The degree to which the structure resonance conditions coincide with propeller excitation.

o The extent to which the structure/soundproofing/trim design has been optimized to counteract noise impinging on the exterior and to minimize the interior noise and vibration.

Evidence is required to show that passenger cabin noise can be controlled to desirable levels without incurring weight penalities that would offset the prop-fan fuel economies.

OBJECTIVE 19: DETERMINE THE MAGNITUDE OF THE NOISE RE- DUCTION ACHIEVABLE IN THE PASSENGER CABIN BY CREATING A "MISMATCH" BETWEEN PROPELLER BLADE PASSAGE FREQUENCIES AND THE CABIN ACOUSTIC MODE FREQUENCIES.

APPROACH: Measure the interior noise level in the testbed aircraft while varying the fre- quency of the exterior noise, using the same experimental setup as used for Objective 20.

OBJECTIVE 20: CONDUCT A RESONANCE FREQUENCY MODAL SURVEY AND DETERMINE THE FUSELAGE SHELL FREQUEN- CIES AND MODE SHAPES IN THE FREQUENCY RANGE OF 30 TO 500 HZ.

APPROACH: Use computerized transfer function analy- sis techniques with electroacoustically simulated prop-fan noise excitation of the testbed fuselage structure to experimen- tally determine the shell mode shapes for the complete fuselage. Repeat the experi- ments in the vicinity of the prop-fan, us- ing actual prop-fan noise excitation of the structure, to evaluate the techniques and validate the results.

OBJECTIVE 21: DETERMINE THE RELATIONSHIP BETWEEN THE FUSELAGE SHELL RESONANCE FREQUENCIES AND MODE SHAPES AND 1) THE SPATIAL CORRELATION OF THE IMPINGING EXTERIOR NOISE, AND 2) THE INTERIOR VOLUME ACOUSTIC MODES.

Conduct analyses, and limited experiments APPROACH: on the testbed fuselage, to obtain the interior volume acoustic modes. Then use the spatial correlation data obtained in Objective 15 and the modal survey results of Objective 22 to establish the above relationships.

OBJECTIVE 22: VERIFY MINIMIZATION OF THE STRUCTURE MODAL RESPONSE BY OPTIMIZING PROP-FAN SYNCHRO- PHASING, AND DETERMINE THE MAGNITUDE OF THE NOISE REDUCTION OBTAINABLE BY USING SYNCHROPHASING TO "MISMATCH" AREAS OF COR- RELATED EXTERIOR SOUND PRESSURES WITH SHELL MODES.

Using the same techniques for analysis and APPROACH: electroacoustic simulation of prop-fan noise as used in Objective 20, simulate four synchrophased prop-fans on the test- bed airframe. Vary the phase relation of the noise sources while repeating the mode determinations so as to systematically minimize the modal response and identify the optimum phase relations.

Concurrent with these experiments, measure interior noise level in the test- bed aircraft and obtain the data for re- lating noise level with modal response and noise source phase relations.

OBJECTIVE 23: DETERMINE THE IMPROVEMENTS ATTAINABLE BY STRUCTURAL TAILORING SO AS TO OPTIMIZE SHELL "MISMATCH" MODE SHAPE AND LOCATION RELATIVE TO AREAS OF CORRELATED EXTERIOR SOUND PRESSURE.

APPROACH: Modify the fuselage shell structure so as to use the restraint and/or stiffening effects of the floor, ceiling, and inter- ior partitioning to alter the shape and location of the shell modes. Exper imen- tally determine the mode shape changes; measure the interior noise level changes.

OBJECTIVE 24: OPTIMIZATION OF THE MATCHING OF THE DY- NAMIC PROPERITES OF THE FUSELAGE SHELL STRUCTURE AND THE INTERIOR TRIM PANEL/AIR SPACE/INSULATION ASSEMBLY APPROACH: Conduct analyses using existing and newly formulated theory as necessary, systemati- cally varying the mass, stiffness, damp- ing, absorption, and air space coupling of the structure and soundproofing/trim assembly, to perform a parametric optimi- zation study. Install 2 variations of the soundproofing/trim design in the testbed, obtain measurements of sound transmission loss for the design variations, and eval- uate the analyses methods. Demonstrate the effectiveness of an optimized fuselage shell/trim/insulation design.

AIRCRAFT PERFORMANCE Optimum installation of the prop-fan propulsion system into a practical aerodynamic environment represents concern for at least two reasons. First, the high solidity and blade Mach number of the prop-fan creates core engine inlet problems which are compounded by the configuration of the engine air duct which must be arranged to account for the gearbox and drive shaft. Secondly, the integration of the prop-fan, nacelle, and wing into an efficient aerodynamic design must be arranged to minimize losses due to swirl and scrubbing effects which would tend to reduce the benefits of prop-fan propulsion.

Installed Propulsive Efficiency Wind tunnel studies of small-scale prop-fans indicate that prop-fan net efficiency will reach the BO-percent goal at a cruise Mach number of O.B. One objective of the technology development program will be to demonstrate that this goal can be achieved for a large-scale installation. This may, however, be very difficult to do on the testbed aircraft. Although flight test measurements of propulsion system thrust can be obtained under proper conditions, the accuracy of the thrust from the test installation determined by conventional flight test techniques may be low for a testbed aircraft powered by more than one type of engine. Some consideration has been given to the measurement of the thrust from the test engine by strain-gaging the engine support system. This has been done for pylon-mounted engines but would be very difficult, if not impossible, for the type of wing-mounted installation anticipated for the testbed.

Among other things, engine inlet performance can be significantly affected by the shape of the prop-fan spinner and hub and by the nacelle immediately be- hind the prop-fan. Engine inlet performance will, therefore, require verifi- cation both in wind tunnel tests and by flight tests to ensure that the propul- sive efficiency is not impaired by duct design and lip location.

OBJECTIVE 25: PERFORM HIGH-SPEED SMALL-SCALE WIND TUNNEL TESTS TO OBTAIN DATA TO VERIFY PROPULSIVE EFFICIENCY.

APPROACH: Expand .NASA-Ames prop-fan/nacelle/wing test program to include more realistic nacelles. Augment with tests from core inlet test rig.

Interaction Effects: Aerodynamic/Propulsion Integration and Optimization Two important aspects are associated with the integration and optimization of the prop-fan/nacelle/wing installation: 1) the need to maintain the aerody- namic efficiency of a high-speed airfoil immersed in the prop-fan slipstream, and 2) the need to recover the propulsive thrust from the slipstream swirl. Ref- erence 1 reports a wind tunnel study of a supercritical wing immersed in a sim- ulated prop-fan slipstream. Swirl was found to effect wing drag, and the slip- stream power additions affected wing shock location, but generally, these ef- fects were less than anticipated. Some anomalies in these data, however, raise questions about the test techniques employed and the need for further study has been recognized. Upcoming tests of the 0.62 m (2.04 ft) diameter prop-fan in the NASA Ames 11-ft Transonic tunnel should provide additional verification data.

The wing can also significantly affect propulsive efficiency. Recovering residual swirl from the prop-fan slipstream, more than any other single factor, has the potential for providing large gains in efficiency. This swirl might be recovered by proper local tailoring of the wing as shown in the analytical study of Reference 1, but complete swirl removal could result in an impractical wing structure with sheared front and rear spars. More work is certainly needed in the area of optimizing the prop-fan/nacelle/wing interaction region. Although test data will be needed to validate and augment such analytical work, the test- bed aircraft will not be the proper s'ource for such data because the test wing will not be supercritical. Even if gloves are used to simulate a supercritical wing section locally, the installation and modifications together with any sub- sequent modifications would be very expensive.

It is, therefore, recommended that aerodynamic/propulsion integration and optimization studies should be a major part of a wind tunnel test program. This does not imply that the testbed aircraft will not contribute to solutions to this problem area. With proper instrumentation, e.g., wake and swirl rakes, distributed static pressure orifices, a large amount of useful data can be ob- tained which will aid the optimization of the prop-fan installation.

OBJECTIVE 26: PERFORM HIGH- AND LOW-SPEED WIND TUNNEL TESTS OF SCALED TESTBED (SPECIFICALLY PROP-FAN. NACELLE AND WING) THROUGH A RANGE OF FLIGHT AND PROPULSION SYSTEM VARIABLES TO OBTAIN DATA ON THE FLOW FIELD CHARACTERISTICS.

APPROACH: These data will obtained by placing be rakes in several azimuthal chordwise posi- tions behind the prop-fan. The wing upper and lower will fitted surfaces also be with several chordwise rows of static pressure orifices. The flow field charac- teristics for the wing alone. prop-fan alone. and for the wing/prop-fan with both metric and non-metric nacelles/prop-fans.

will be obtained. Definition of the flow characteristics will include swirl angles.

axial velocity increments. surface pres- sure distributions. effect of local con- touring. and effects of blockage.

Although the testbed aircraft will not be equipped with a supercritical air- foil section. the wind tunnel results will be analytically applied to such sections.

Tailoring of the nacelle can then be ac- complished and the results verified.

OBJECTIVE 27: DETERMINE ENGINE INLET PERFORMANCE FOR THE TESTBED PROP-FAN INSTALLATION THROUGH A RANGE OF FLIGHT CONDITIONS.

APPROACH: Install pressure rakes in the scaled test- bed installation inlet and perform high- speed wind tunnel tests to obtain data.

Verify by flight test of the large scale propulsion system over a range of condi- tions.

FUNCTIONAL SYSTEMS OPERATION AND FOD VULNERABILITY A number of propulsion subsystems also present areas of concern. Among these are the prop-fan pitch control system and the effectiveness of the thrust reverser, and the prop-fan blades vulnerability to foreign-object damage.

Prop-Fan Pitch Control System Although the prop-fan pitch control system is expected to require only the normal functions of a conventional turboprop system, the testbed installation will provide verification of the assumptions that state-of-the-art systems are adequate.

OBJECTIVE 28: DETERMINE THE CHARACTERISTICS OF THE TEST- BED PROP-FAN INSTALLATION CONTROL SYSTEM.

APPROACH: Perform flight tests in an operational environment to obtain data on the trans- ient behaviour of the control system.

thrust-Reversing Effectiveness Application of prop-fans to transport aircraft will require knowledge of the thrust-reverser effectiveness. Measurement of reverse thrust, although pre- senting difficulties, has the advantage that the prop-fan propulsion system would be operated in a flight or taxi condition in which other propulsion units on the testbed aircraft would be shut down or at idle conditions.

OBJECTIVE 29: DEMONSTRATE THE EFFECTIVENESS OF THE PROP- FAN IN THE REVERSE THRUST MODE.

APPROACH: Perform measurements of reversed thrust in a flight or taxi condition with engines other than the prop-fan unit at flight idle or shutdown. Measurement of reversed thrust will be obtained in a manner simi- lar to that for obtaining flight thrust.

Vulnerability to Foreign-Dbject Damage Foreign-object damage (FOD) analytical methods have been correlated with fan-blade development test data. No tests have been performed using prop-fan blades constructed using the spar /shell concept, but analysis of the blades using the available methods shows that the blades can sustain large-bird strikes wi thout affecting the structural integrity of the blades. No criteria exist that specifically relate to prop-fan FOD tolerance. The ability of the prop-fan to sustain foreign-object impacts is found to be in excess of a criteria estab- lished for turbofans (FAA Advisory Circular 33-1B). However, tests on large- scale blades are needed to verify impact resistance.

OBJECTIVE 30: VERIFY FOD TOLERANCE FOR ADVANCED LARGE- SCALE PROP-FAN BLADES FABRICATED USING THE SPAR/SHELL CONCEPT TO ESTABLISH VULNERA- BILITY.

APPROACH: Perform FOD tests on large-scale swept prop-fan blades in static tests simulating bird-strike and other objects such as nuts, bolts, small pieces of metal, and other materials such as dirt and sand that might cause blade erosion.

TESTBED PROGRAM PRIORITIES Priorities for the testbed program objectives are based upon the relative importance of the integrity of the structure, acoustic environment, aircraft performance, and systems operation. Areas of technological concern are ranked according to priority as outlined below.

Program Priority 1 - Integrity of the Structure The most important objectives are those related to the integrity of the structure, which includes both the prop-fan and the airframe, as well as scale effects.

Three areas of concern fall into this category: a) Propeller structural integrity and dynamics b) Propeller induced vibrations and static and dynamics loads c) Scale effects Program Priority 2 - Acoustic Environment Public acceptance of the prop-fan will be dependent upon the far-field im- pact on community noise environment and of the near-field effect on the travel- ing public. Two acoustic areas must, therefore, be given second priority for testbed program objectives: a) Propeller-generated near- and far-field noise b) Passenger cabin noise and vibration Program Priority 3 - Aircraft Performance Those technological concerns that affect aircraft performance are placed third in order of priority. These objectives concern installed propulsive efficiency and the interaction effects that, to some extent, can be controlled by proper design of the powerplant nacelle and nacelle/wing integration.

Two technology areas fall into this category: a) Installed propulsive efficiency and interaction effects b) Engine inlet performance Program Priority 4 - Functional Systems Operation and FOD Vulnerability Those items that are essential to the operation of the testbed, but which are related to the functional systems of the testbed installation and which can be approached by functional test and development, are ranked lowest in order of priority. Three items fall into this category: a) Prop-fan control system b) Thrust-reversing system c) FOD vulnerability Program Objectives Categorization The program objectives are sUbdivided into task size units within each level of program priority and are further ranked in importance on a subpriority basis.

A review of the technology concerns, objectives, and priorities is given in Table A-I. Also shown is the identification of the subpriorities within each technological area and the methods of solution available to satisfy each objec- tive.

Although more than one method may be necessary to obtain the required solu- tions, the preferred methods are indi~ated by the circles in Table A-I.

TABLE A-l. PROGRAM OBJECTIVES AND PRIORITIES P~OBLEM SOLUTION METHOD SUB TECHNOLOGY PRIORITY OBJECTIVE PRIORITY AREA TESTBED WIND TUNNEL STATIC AIRCRAFT TES. ANALYSIS HS LS 1 INTEGRITY OF 1 Blade dynamic response 1 ® THE STRUCTURE validation 0 Propeller 2 Blade classical flutter 2 X X ® structural validation integrity & 3 Blade stall flutter validation 3 X X dynamics 4 Critical speed & hub stiffness 4 X

~

validation 0 Propeller 5 Determine aerodynamic data 1 X ® ® induced for flutter analyses vibration & 6 Determine structural vibration 2 X X X X ® dynamics spectra magnitude 7 Drive system dynamic loads 3 X X X ® & induced effects 0 Scale 8 Validate or develop scaling 1 X X ® effects laws 9 Blade mass & stiffness 2 X X X ® distribution determination 10 Demonstrate full size prop-fan 3 X fabrication feasibility 11 Establish drive system feasi- 4 X bility for 15,000 SHP & above 2 ACOUSTIC 12 Sound pressure directivity 1 X X ® ENVIRONMENT and spectra variation 0 Propeller 13 Sound pressure levels on 2 X ® generated pressurized surfaces near-field 14 Noise strength & directivity X X ® noise determination -X Fluctuating pressure spectra 4 X 16 Effects of fuselage curavture 5 X 17 Geometry of correlated sound 6 X

~

pressure area 0 Propeller 18 Verify prop-fan compliance 1 ® generated with FAR Part 36 far-field noise 0 Passenger 19 Minimization of sound trans- I X ® cabin noise mission & vibration 20 Resonant frequency modal 2 X ® survey 21 Fuselage modes and external ® noise relation 22 Noise reduction & structural 4 X ® response minimization by synchrophasing 23 Improvement thru optimization 5 X X ® of shell modes 24 Noise reduction thru cabin 6 X X ® dimension changes 3 AIRCRAFT 25 Verify propulsive efficiency 1 X X PERFORMANCE 26 Determine flow field effect 2 X X

~

on wing 27 Verify engine inlet performance 3 X X X ® 4 SYSTEMS 28 Verify drive system control 1 X ® OPERATION system <!) 29 Verify reverser effectiveness 2 X 30 Determine prop-fan vulner- 3 X ® abill"ty to FOD

o PREFERRED METHOD OF SOLUTION

APPENDIX B - CANDIDATE PROPELLER DRIVE SYSTEMS - TASK II The rapid advance of turbofan technology for high-speed cruise during the 1950/60s resulted in a reduction in the demand for turboshaft engine cores for propeller-driven aircraft application. Turboshaft engine development was, therefore, reduced to a level consistent with the requirements for rotary wing aircraft.

A survey of the available turboshaft core engines was conducted wi thout regard for the purpose for which the engines were developed, i.e., either for propeller application or for rotary-wing use. Available gearboxes suitable for the drive system application were also investigated. A prime consideration in the selection of a drive system was to avoid costly turbo-machinery and gearbox development.

Typically, the bare drive system consists of a core or power section, a torquemeter, interconnecting struts, and a reduction gearbox. The drive systems utilize available gearboxes which have offset power input pinion gears, and can be configured with the gearbox either in the "pinion-high" or "pinion- low" arrangement, as shown in Figure B-1. The choice depends on the type of installation required for the airframe. The "pinion-high" configuration would generally be representative of an engine nacelle over-the-wing drive system installation, whereas the "pinion-low" arrangement would be consistent with the engine nacelle under-the-wing arrangement.

POWER SECTION PINION GEAR ~-_~ GEARBOX TORQUE METER, PINION-LOW CONFIGURATION PINION-HIGH CONFIGURATION Figure B-1. Typical Drive System Configuration DRIVE SYSTEM DESIGN REQUIREMENTS The drive system for the testbed aircraft, as a minimum, should have the capability to satisfy the following design requirements: o Power an advanced propeller for flight research from static sea-level conditions to Mach 0.8 at altitudes of 10,668m (35,000 ft) or higher.

o At the design cruise conditions of Mach 0.8 at 10,668m (35,000 ft), the drive system should be capable of powering an advanced propeller over a range of conditions given by: 2 2 Power loading - 209 to 301 kW/m (26 to 37.5 shp/d ) Propeller tip speeds - 183 to 244 m/s (600 to 800 fps) Specifically, the drive system should have the capability of powering a given propeller at design cruise conditions in each of the following three operating combinations: 2 2

Case 1: 209 kW/m (26 shp/d ) @ V = 183m/s (600 fps)

T 2 2 Case 2: 241 kW/m (30 Shp/d ) @ V = 213m/s (700 fps) T

Case 3: 301 kW/m (37.5 Shp/d ) @ V = 244m/s (800 fps)

T where V is the prop-fan tip speed •.

T o Drive system minimum power level at sea level to be 2983 kW (4000 shp) In addition to these design requirements, the drive system should be: o Readily available or easily derivable from existing hardware and should include the core engine, gearbox, nacelle, controls, and accessories.

o Configured so that the internal and external flow lines give low installation performance losses.

o Capable of providing acceptable operation of all components throughout the flight envelope.

POWER SECTION AND GEARBOX SURVEY Power Section Survey A survey of domestic turboprop/turboshaft engines showed the number of engines in the approximate power level and performance range to be very limit- ed, to the extent that only five were identified as capable of satisfying the minimum power level requirement. The power sections identified were: o Detroit Diesel Allison T56 Single Shaft Turboprop o Detroit Diesel Allison XT701 Free Turbine Turboshaft o General Electric GE T64-10-415 Free Turbine Turboshaft o Lycoming T55-LTC4B-12 Free Turbine Turboshaft o Pratt Whitney JFTD12A Free Turbine Turboshaft Following closer examination of the characteristics of each of the engines, the P&W JFTD12A was found to be unsuitable for testbed aircraft application be- cause the drive shaft was arranged to extend rearward yielding an engine intake and gearbox configuration unsuitable for a tractor-type propeller application.

The Lycoming T55-LTC4B-12 turboshaft engine data indicated a capability of operation up to an altitude of 7,62Om (25,000 ft). No data were available for higher altitudes or for changes required to increase the altitude capability.

The P&W JFTD12A and the Lycoming T55-LTC4B-12 were, therefore, eliminated as candidate power sections for testbed aircraft application.

The performance characteristics of the remaining power sections are given in Table B-1.

TABLE B-1. CANDIDATE POWER SECTIONS POWER AVAILABLE POWER CRUISE SLS SECTION M = 0.8/10,668 m (35,000 ft) kW (shp) ~W(Shp) *DDAXT.701 2520 (3380) 6018 (8071) DDA T56 1819 (2440) 3423 (4591) **GE T64-10-415 1350 (1810) 3£ 66i.438Ql.

*Detrolt Diesel Allison **Generol Electric Gearbox Survey Examination of the available gearboxes indicated that the following units possessed the capability of matching the output of the candidate power sec- tions: o Detroit Diesel Allison T56-A-14 o Detroit Diesel Allison T56-A-15 o Ishikarapima-Harim Heavy Industries IHI T64-2 SDG DDA T56-A-14 Gearbox - This gearbox is a "pinion-high" configuration as used on the Lockheed P-3C "Orion" aircraft but can be adapted for prop-fan application using either the XT101 or T56 power sections. The modification required to match the XT101 is complicated by the fact that the rotation of this engine is opposite to that of the T56 power section, and by the signifi- cantly lower RPM of the XT101. The clockwise rotation and the 11,500 RPM of the XT101 require changes to the main drive sun gear and pinion and to the accessory drive train to provide corr'ect rotation for the oil pump and tach- ometer speed.

DDA T56-A-15 Gearbox - This is a "pinion-low" gearbox used for the Lockheed C-130 drive system. Because of the design of the gearbox lubrication system, which requires baffles located adjacent to the pinion and which cannot be re- located, the maximum diameter of the pinion is restricted. This in turn would cause a small reduction in the diameter of a prop-fan used for a "pinion-low" arrangement to reach a tip speed of 244m/s (800 fps).

IHI T64-2 SDG Reduction Gear - This gearbox is rated at 2535 kW (3400 shp) and has a reduction ratio of 14.31. The gearbox could be used by modifying the pinion and bullgear in the same way that the T56 gearbox is altered.

DRIVE SYSTEM ASSEMBLY The bare drive systems are assembled by combining the power sections and The drive the appropriate gearboxes by means of a connecting torquemeter.

system assemblies can be configured as either "pinion-high" or "pinion-low", depending upon installation requirements. The length of the torquemeter is, to some extent, dictated by the engine intake requirements if scoop-type inlet short-coupling with abrupt duct curvature is to be avoided.

GE T64-10-415 The assembly of this drive system for "pinion-high" and "pinion-low" con- figurations is shown in Figure B-2, together with the principal dimensions and characteristics.

DDA T56 The "pinion-high" and "pinion-low" assemblies are given on Figure B-3. This assembly is based upon that of the Lockheed C-130. The principal dimensions and data are also included in Figure B-3.

DDA XT701 The XT701 drive system assembly for "pinion-high" and "pinion-low" are illustrated in Figure B-4. Also included are the dimensional data and the principal characteristics.

TORQUEI.1ETER GEAR BOX.

r1.,--- __ P_O_W_E_R_1'-.E_CT_I_O_N ___ _

r

2.19m (83.43 ins)-------~

k 2.52 m

(99.16 ins) ------------1 2.79 m

1------------- (J.09. 86 ins) -------~

PINION,HIGH CONFIGURATION OVERALL (,G.

POWER SE':::TIOtJ r " '_. '~.

- i....- ___ --I~ 0.036 m

(1.40 ins)

1(15.73 ins) 1: -----'

I. ....J t-- O.?m----l

J (26.43 ins) I J . 0.48_m

. 0 632 J (1.89. ins)

'!(i4:88 ~s)l ~i

I.. 0.89 m (35.04 ins) f----- 1. 37m (53.86 ins) 1-- ____ 1.41 m-,- ___ ~ (55.43 ins) P!NION'LOW CONFIGURATION Figure B-2. T64 Drive System Assembly GEARBOX 0.45 m TORQUE METER POWER SECTION (21.4 ins) 0.711m (28.007 ins) 2_18 m (85.755 insl.

,---J----v A~ __ ~~ ______________ ~A~ ______________ ~

I 1

F.:;==:::~I-------- ---~J.- ------- --_.

1.7 m + .0016 !II HOT (67.198 ins + 0.065 ins HOT) f----------:- __ - 3.04 !II + 0084 m --:=~-------_l (119.698 ins + 0.332 ins HOT) f---r'------....::-....;3' .44 !II (135.194 ins) -~-------l ~-----43.71m + .0084 (145.979 ins + 0.332 ins HOT) _______ --+

)

I 0.28 !II (11.00 ins) PINION-HIGH CONFIGURATION 0.21 !II (8.31 ins) 0.048 !II 0.25 !II I (1.9 ins) (9.7 ins) , 0.28 !II (11.00 ins) 0.402 !II .,.---+---4----l 0.038 !II (15.83 ins) (1.5 insl ins)-----l 1----1.85 !II (72.8 ins)--------l PINION-LOW CONFIGURATION Figure B-3. T56 Drive System Assembly 3.6 m 1-------------- (141. 85 ins)-------------.; r----- TO P.S. C.G.--------I /4---iO TOTAL C.G.-_'" POWER SECTION C.G.

OVERALL C.G.

I

0.28 m I

(H.OO ins) I I 0.4 m --+---i---.., <15.71 ins} I 1. 031 m 0 . 66 1- m -";--(40.75 (26.20 ins) t-- _________ 2.72 m (107.34 ins) PI NON- HIGH CONFIGURATION 1.

0.28 m (H.Od·ins) P!NION-LOW CONFIGURATION Figure B-4. XT701 Drive System Assembly PROP-FAN SIZING AND DRIVE SYSTEM CHARACTERISTICS The prop-fan diameter for testbed application should be as large as pos- sible if the size of prop-fans for aircraft of the future is to be properly represented. Prop-fan and nacelle diameters as a function of cruise power and prop-fan RPM at an altitude of 10,668m (35,000 ft) are shown in Figure B-5. A turboshaft drive system for testbed application at the minimum acceptable dia- meter of 2.43m (8 ft) must be capable of generating 1789 kW (2400 shp). The deSign point for the prop-fan at cruise conditions, i.e., M=0.8 at 10,668m (35,000 ft), is a disk loading of 301 kW/m (37.5 shp/d ) and a tip speed of 244m/s (800 fps). These data together with the data for disc loading of 241 2 2 and 209 kW/m (30 and 26 shp/d ) are also shown in Figure B-5.

Prop-fan diameter for each candidate drive system is also shown in Tabl,= B-II and the principal characteristics of the propeller drive systems are shown in Table B-III.

The requirement for a minimum diameter of 2.43m (8 ft) would tend to elimi- nate the GE T64 drive system from consideration. Drive system availability is, however, an important factor so that it is considered expedient to carry the GE T64 drive system as a candidate until availability of all candidate drive systems is verified.

The principal candidates for the drive system are the free turbine DDA XT701/T56-A-14 and the fixed-speed single shaft DDA T56-A-14. Because of drive shaft RPM flexibility offered by the free turbine power sections, and th,= advantages arising from that feature in flight research activities, a drive system utilizing either the DDA XT701 or the GE T64 would be desirable.

DRIVE SYSTEM/NACELLE INSTALLATION The drive system installation to form a Quick Engine Change (QEC) unit was accomplished by designing the nacelle contours to a NASA supplied area distri- bution curve, Figure B-6. The spinner and nacelle shapes were configured to retard the airflow to alleviate blade-root choking. The data of Figure B-6 were derived from NASA tests of axisymmetric nacelles without air inlets. They can, therefore, only be considered as guidelines in the design of configura- tions that are highly unsymmetrical and require internal flows for the engine TIP SPEED = V m/s (F.P.S.)

m T Ft 183 (600) 213 (700) 244 (800) 20 6 .

~ H o

o O~----~----~------~----~

o 2000 4000 6000

PROP-FAN RPM D - PROP-FAN DIAMETER P D - NACELLE DIAMETER N m Ft 8 24 7 kW/m2 ( SHP/D ) (26) (30)

. (37.5)

16 5 ~ H o ( SHP/D2) 3 (26) (30) (37.5) o~--~--~----~--~--~--~~--~--~

o

o 1000 2000 3000 4000 5000 6000 7000 8000

CRUISE POWER - kW 10,000 o 2000 4000 CRUISE POWER - SHP Figure 6-5. Prop-Fan Sizing - 1O,668m (35,000 ft) Altitude TABLE B-II. DRIVE SYSTEM PROP-FAN DIAMETER * DRIVE SYSTEM PROP-FAN DIAMETER 2.89 m DDA XT701/T56-A-14 (Mod) (9.5 Ft) 2.47 m DDA T56/T56-A-14 (Mod) (8.1 Ft) (Mod) 2.l3 m GE T64-10-415/IHI T64-2SDG (7.0 Ft) *BASED ON: MACH = 0.8 ALT. = 10,668 m (35,000 ft)

SHP/D = 301 kW/m2 (37.5 SHP/ft )

V = 244 m/sec (800 ft/sec)

T TABLE B-III. CANDIDATE DRIVE SYSTEM SUMMARY ENGINE GE T64 DOlo T56 DOlo XT70I ~ F:M:INF Tvrr. FREE ruRBIHE FUED SPEED FREE ruRBIN!

~ COONTERCLOCJ(WISE RO'I Al ION -ALF" COUNTF.R('·UlCKIII SF. CLOCKWISE PERFORMANCE roWER SLS ... 11 (SHP) 3266 (43S0) 3423 (4591) 6018 (S071) 1068S • (35,000 IT) H - O.S 1350 (1S10) IS19 (2440) 2520 (3380) RPH HAll CONTUltlOIJS 13600 13820 11500 PROP-fAN SIZING DISK I.OAOING kW/,.2 _ (SIIP/D ) JOI (37.5) JOI (37.5) JOI (37.5) JOI (37.5) (37.5) JOI (37.5) JOI (37.5) JOI (37.5) 301 (37.5) (700) (600) V .. I. (fr.) 244 (800) 21) tS3 (600), 244 (SOO) 2H (700) 18) (bOO) 244 (SOO) (700) ISS TP RI'II 2140 1870 1600 1900 1660 1430 HIO 1410 1210 .. ..

rROP-FAN DIAMETER .. (fT) 2.13 (6.97) 2.47 (8.1) 2.89 (9.5)

-

- - -

REIlUCTION GEAR 1 I 2 J I 213 (700) Vyp ./SEC (FT/SEC) 244 (800) 211 (700) 181 (6UO) 244 (800) IS3 (600) 244 (800) (700) 181 (600) 2JJ PINION GEAR .. 6sT/sr 6)T/8P 56T/8P ... 68T/SP ..

.. NIA

... .. .. ..

MAIN DRIVE GEAR 108T/8P llJT/SP 120/8P 108T/8P N'A

A1.TF.RNATOR GEAR NIA ... 6ITIlOP

... .. NIA ..

OIL PUMP DRIVE GroAR HIA ... NIA 78T/IOP

... 4 ..

OIL PUHP DRIVEN GEAR .. NIA 3)T/lOP

.. NIA ... ..

MAS!; PROPERTU':S WEJGHTS REDUr.TI(lN GFoAR kg (U) 194.1 (428) 249.6 (550.5) 249.6 (550.5) TOR~F.NETER kg (LB) 18.14 (40) 27.21 (60) 27.2l (60) (700) rowER SECTION kg (La) )26.5 550.7 (1214.5) Sl4.7 (1179) TOTAL kg (LB) 538.74 (1188) 827.51 (1B25) SI1.48 (1789.5) ROI,I. YAW PITCH ROJ.l. YAW rnCH ROLL YAW rnclI - - MOMENT OF INRRTIA kg/ .. I.R/FT 130.7 ( 638) - - - - - - .00S (1.9) IoBllVE roWER SEenON 'i.

C': UH:AT((lN PINIflH I,OW 1ft (Jns) W.L. 102.2 BL 95.7 ... MIA ..

1.62 (63.83) lofT OF TIIRUST NUT "AI.F - AFT LOOKING FORWARD NASA AREA DISTRIBUTION 0.1 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0 X/L Figure 8-6. NASA Nacelle Area Distribution Curve and oil coolers. The significant part of Figure B-6 for nacelle design is the portion up to the location of the maximum area.

DRIVE SYSTEM NACELLE DEVELOPMENT Development of the nacelle designs for the various drive systems was based upon Hamilton Standard recommendations, Reference 2. The principal design con- ditions were: o Nacelle Length - 1.0 Propeller Diameter from the Wing Quarter Chord to the Prop-Plane.

o Nacelle Diameter/Propeller Diameter DN/Dp = 0.35.

When applied to practical designs. the nacelle length (prop-plane to wing C/4 @ 1.0 Dp> was found to place the rear portion of the power section in such a position, relative to the wing, that the propeller thrust line/wing reference plane separation was unnecessarily increased and fairing between the nacelle and wing leading edge rendered difficult. By changing this dimension to 1.0 Dp from the prop-plane to the wing leading edge, the turbine portion of the power

section could be moved forward away from the wing maximum thickness, permitting minimization of the thrust line offset and improved integration of the nacelle and wing.

Nacelle layouts and contours were developed for the three selected engines for "pinion-high" and "pinion-low" configurations. Initially, the similarities between the T56 and the XT701 were thought to be such that a set of common con- tours could be used for either engine. The possibli ty of using an existing nacelle such as the C-130 and P-3C was also investigated. The nacelle develop- ment investigation considered the following cases: o Modified C-130 and P-3C T56 nacelles.

o Common T56/XT701 nacelle "pinion-high" and "pinion-low" configurations.

o T64 nacelle contours for "pinion-high" and "pinion-low" configurations.

o T56 nacelle common contours for both "pinion-high" and "pinion-low".

o XT701 nacelle contours for "pinion-high" and "pinion-low" configura- tions.

o Revised XT701 nacelle common contours for "pinion-high" and "pinion-low" configurations.

Modification of Existing Nacelles As a low-cost approach to nacelle design, the nacelle contours for two existing designs, the Lockheed C-130T56 and the Lockheed P-3C T56 nacelles, were investigated for conformance to the NASA area distribution curve, Figure B-6.

Lockheed C-130 Modified Nacelle - Modification of the area distribution and the modified contours for the Lockheed C-130 "pinion-low" nacelle are shown in Figure B-7. Improvement of the distribution would require a prop-fan hub of greater diameter than that of the C-130 propeller. This would require re- location of the engine intake downward. It is possible to add area to conform to the NASA distribution in the region of the maximum cross-sectional area, but the actual and modified distributions forward of the maximum area are so far removed from the NASA curve that the modified nacelle would not present a sat- isfactory representation of the true conditions for the prop-fan.

2 3

! ! I ~

-+-r----t---7--\ "rr" I i

I I I

I NACELLE STATIONS

NACELLE SECTIONS C-130 NACELLE MODIFIED FOR PROP-FAN 1 .00 r------77'i~-__,r__j~=::=:=:3 .80 .60 A/A FAIRED C-130 MAX .40 I .20

I

.80 1.00 .40 .60 .20 X/L Figure B-7. Lockheed C-130 Modified Nacelle Lockheed P-3C Modified Nacelle - A situation similar to that of the C-130 exists when the modification of the Lockheed P-3C nacelle is considered. This nacelle, shown in Figure B-8, is configured for a "pinion-high" arrangement.

When the actual and NASA area distributions, Figure B-8, are compared, the mis- match can be seen to be far greater than that of the C-130 nacelle due pri- marily to the larger base area of the P-3C nacelle. In this case, adding area to the nacelle at the location of the maximum cross-sectional area of the NASA distribution does not change the actual distribution enough to provide an ade- quate representation of a prop-fan nacelle. As in the case of the C-130, the area distribution modification is constrained by the location of the engine and oil cooler inlets.

Revised T56 Nacelle Contours - The T56 nacelles, described in the preceding

text, were contoured by normalizing on the base area at an X/L = 1.00 and de-

riving the maximum cross-sectional area, A ' at X/L = 0.414 for the nacelle.

MAX The resulting nacelles produced values of nacelle equivalent diameter to pro- peller diameter of 0.49 for the C-130 and 0.62 for the P-3C, with the actual distribution having little correspondence to the NASA curve.

In an effort to improve the area distribution of the existing C-130 T56 nacelle, the reference cross-sectional area used for normalizing was changed

from that at X/L = 1.0 to that at X/L = 0.414. The resulting nacelle and cor-

responding area distribution, shown in Figure B-9, conform closely to the NASA curve over the spinner and prop-fan hub, and can be achieved without changing the nacelle lines behind the hub.· The effect of adding a slight nacelle build-up behind the hub is also shown in Figure B-9. Using this technique, the C-130 T56 nacelle with or without contour build-up could provide a minimum modification arrangement.

Applying the same technique to the P-3C nacelle does improve the area dis- tribution, as shown in Figure B-8, but would require contour build-up to match the NASA curve over the hUb/intake region. Since the base area of the nacelle at the propeller hub is fixed by the dimensions of the existing propeller, any increase in cross-sectional area by contour build-up would affect the engine intake region of the nacelle. This would require a considerable reconfigura- tion of the nacelle shapes amounting to a new nacelle.

T56/XT701 "Pinion-High" Nacelle - The nacelle layout, contours and area distribution for the "pinion-high" configuration T56/XT701 common nacelle are shown in Figure B-10. The nacelle contours follow the desired area distribu- MODIFIED CONTOURS , I , I I 1 2

.. t·..-·---t

- -:~: --I-

--:.~._ .J-----l -lJ--- --fH+-++-+- .... -H-. -_. ._-

r

I

T'~

NACELLE SECTIONS P-3C NACELLE MODIFIED FOR PROP-FAN NACELLE 1 .6 NORMALIZED TO ARE!:@

X/L = U.414

1.0 A/A MAX 0.8 0.6 EXISTING 0.4 01 STRI BUTt ON 0.2 0.2 0.4 0.6 0.8 1.0 X/L Figure B-8. Lockheed P- 3C Modified Nacelle, Pinion-High MODIFIED CONTOUR NACELLE SECTIONS

II

NACELLE STATIONS ·8 1.4,---------- __ ...., , .. -- - ...

1.3 "''''' 1.2 .............. , ,- / / NACELLE NORMALIZED 1.1 / TO AREF@ X/L = 0.414 1. 0 t------ri',.-. - SPINNER/HUB CONTOURS " MODIFIED .90 .-J ., /1

.80 It--- "'-- NASA DISTRIBUTION

A/A .70 V ~ I

MAX .60 " NACELLE CONTOURS I MODIFIED LOCALLY .50 .40 .30 .10 .20.30.40.50.60.70.80.901.0 X/L Fibure B-9. Lockheed C-130 Modified Nacelle Revised Reerence Area

I

I I I NACELLE STATIONS I I I 2 3 4 5 I

I I I

I I I NACELLE SECTIONS PROP-FAN INSTALLATION, TS6/XT701 PINION - HIGH 1.00 0.90 0.80 , 0.70 0.60

I

, 0.50 A/~

)

0.40 0.30 TS6/XT701

/

0.20 -·-NASA DISTRIBUTION o~~ __ ~ ____ ~~ __ ~~ __ ~~~ o 0 10 0 . 20 0.40 0.60 0.80 01.00 • 0.30 0.50 0.70 0.9 X/L Figure B-lO. T56/XT701 Common Nacelle, Pinion-High tion and are highly unsymmetric. The arrangement of the scoop-type air intake above the engine and the oil cooler inlet below gives essentially parallel top and bottom lines for the nacelle, a result of the offset between the propeller Adherence to the area distribution shaft and the power section centerline.

curve, as shown in the plan view of Figure B-10, produces a pronounced bulge in the nacelle shape at the maximum cross-sectional area.

Since the T56 engine is slightly smaller in diameter than the XT101, al- though longer, the T56 power section fits into the contours configured for the XT101. The nacelle shapes are based on the prop-fan diameter of 2.90m (9.5 ft) for the XT101 and the maximum cross-sectional area is determined by the ratio

DN/Dp = 0.35. Changing from the XT101 to the T56 engine at a disc loading of

2 2 301 kW/m (31.5 shp/d ) would reduce the prop-fan diameter to 2.5m (8.1 ft) and would require recontouring of the nacelle to satisfy the DN/Dp = 0.35 ratio.

To maintain the same contours for both drive systems would, therefore, require a constant diameter for the prop-fan. This, in effect, means that the XT101 drive system would represent a maximum cruise disc loading of 301 kW/m (31.5 2 2 2 Shp/d ) while the T56 would operate at approximately 211 kW/m (21.0 shp/d ).

Downstream of the maximum cross-sectional area, the nacelle contours can be modified for particular airframe installations without compromising the overall area distribution.

T56/XT701 "Pinion-Low" Nacelle - The nacelle layout, contours, and area distribution for the T56/XT101 "pinion-low" configuration common nacelle are shown in Figure B-11. The nacelle bottom line is controlled by the depth of the drive system power section accessories, which are mounted at the front of and below the XT101 engine intake. The same considerations relative to the nacelle size, prop-fan size, and disc loading as discussed in the preceding text for the "pinion-high" configuration apply to the "pinion-low" nacelle.

Because the XT101 accessories control the depth of the nacelle, when using the T56 power section, the bottom line could' be raised slightly relative to the position shown in Figure B-11. In general, the "pinion-low" nacelle with a "chin" type of air intake on the underside of the nacelle and with the oil cooler ducting arranged on the upper portion increases the cross-sectional area, as can be seen from the comparison of the NASA and actual area distribu- tions shown in Figure B-11.

I

~-t

NACELLE STATIONS

I

I I

I

3 4 NACELLE SECTIONS PROP-FAN INSTALLATION, T56/XT701 PINION-LOW 1.00~----------~~~------------~ 0.90 0.80 , 0.70 0.60 A/~.-- -~ 0.50

J

, 0.40 __ T56/XT701 0.30

/

0.20 ____ NASA DISTRIBUTION 0.10 O~~--~~--~~--~~--~~--~

o 0.100.200.3°0.40°.500.600.700.800.901.00

X/L Figure 8-11. T56/XT701 Common Nacelle, Pinion-Low GE T64 Nacelle Contours - The nacelle contours and layout developed for the GE T64 drive system and the corresponding area distribution data are shown on Figure B-12 for the "pinion-high" configuration.

The nacelle area distribution shown in Figure B-12 corresponds to the NASA

curve up to X/L = 0.50. Although the curve is slightly atypical beyond X/L =

0.50 due to the various inlets, generally the curve is smooth. The nacelle configuration shown in Figure B-12 has the scoop inlet located on the upper part of the nacelle and the oil cooler below the engine.

In the case of a "pinion-low" configuration, the base cross-sectional area

at X/L = 1.0 is the same as that of the "pinion-high" configuration so that the

area distribution of Figure B-12 applies to both configurations. The portion of the spinner/hub/nacelle up to Nacelle Station 1 is circular in cross-section.

The T64 nacelle was further refined to provide a DN/Dp = 0.35. The refined

nacelles were contoured by reducing the maximum cross-sectional area, ~AX' to

the value required to give DN/Dp = 0.35 at X/L = 0.414. The actual base area

at X/L = 1.0. was not changed in the revised nacelles and although the spinner/hub/nacelle contours follow the ideal distribution up to values of X/L = 0.50, beyond that point the contours are allowed to depart from the curve.

Since the testbed installation is not an aerodynamically optimized nacelle/

wing integration, the departure behind X/L = 0.50 is not expected to adversely

affect aerodynamic performance. The refined "pinion-high" nacelle layout, con- tours, and area distribution are shown on Figure B-13.

The "pinion-low" arrangement, Figure B-14, is also reduced in size from that illustrated on Figure B-12. The actual area distribution follows the

NASA curve closely up to XiL = 0.10. Beyond that point, the area is permitted

to vary to provide a faired nacelle.

In both refined nacelles the important regions of the hub-spinner/nacelle

contours conform to the NASA distribution. Because the base area at X/L = 1.0

is likely to vary from one installation to another, some variation in the area

distribution can be expected in the region of the nacelle from X/L = 0.50 to

X/L = 1.0.

T56 Nacelle Contours - The original T56 nacelle contours were generated on the assumption that a single nacelle could be designed for both the XT101 and T56. Since nacelle size is a function of the prop-fan diameter, the nacelle envelope for T56 application was too large. The T56 nacelle "pinion-high" and "pinion-low" variants were, therefore, revised to accommodate the T56 based on .2 3 10 1'1 NACELI E SECTIONS 1. ~----------~~~~,'~,_-_~_ 1.0 .90 .80 .70 A/A MAX .60 .50 .40 .30 NASA DISTRIBUTION 00 .10.20 .30.40.50.60.70.80.90 1.0 X/L Figure B-12. T64 Pinion-High Nacelle

III I I I

I 1 2 , I I NACELLE SECTIONS NACELLE STATIONS AREA DISTRIBUTION 'GE T64 1.0 0.90 0.80 0.70 0.60 A/"\wc 0.50 0.40 0.30 0.20 NASA DISTRIBUTION 0.10 o~ __ ~~~~~~~~~~~~~ o 0.20 0.40 0.60 0.80 1.0 0.10 0.30 0.50 0.70 0.90 .,{/L Figure B-13. T64 Pinion-High Refined Nacelle .

I

.~

. ,.".....L--~,.... +->1-1 J ~ ~

I . ~. -) I

3 4 5

-~r- ,~ 11 _,1 ~/-[·-·.r-r-·1-_~·r

.

1'- L. -~ .-.... i

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I I -' I

N ~~ ,'·--1/

2 3 4 5~ 6 7 8 9 10 11

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!rv I I I I 7 8

NACELLE STATIONS

I

I--"' !,..-oL-- V ~

~

......

T- ~ 11 13 ~ACELLE SECTIONS I.! .-----------.., ,.".-- ....

, 1.0 AREA DISTRIBUTION GE T64 .9'0 .70 A/A MAX .60 .50 40 NASA DISTRIBUTION .3C ~~~~~~~~~~~~~

va .\0 .2030 40 50 60 70 SO 90 !O

X/L Figure B-14. T64 Pinion-Low Refined Nacelle a prop-fan diameter of 2.5m (B.1 ft). The nacelle maximum cross-sectional area

located at X/L = 0.414 was determined using the ratio DN/Dp = 0.35, with Dp =

2.5m (B.1 ft). During the investigation of the nacelle contours, it was found that a common set of contours could be used for the "pinion-high" and "pinion- low" configurations. The nacelle layout and contours are shown in Figure B-15.

The contours shown are for the "pinion-low" arrangement. Rotation of the con- tours 3.14 radians (1BO degrees) gives the arrangement for the "pinion-high" configuration. Al though the envelopes are the same, the mounting and struc- tural arrangement for the "pinion-high" and "pinion-low" arrangements would be different.

I~

f1t7

I 1 2 3 4 5 6 PINION-lOW CONFIGURATION Figure B-15. T56 Pinion-High and Pinion- Low Common Contours XT701 Nacelle Contours - The departure from a universal T56/XT701 nacelle and the development of common contours for the T56 "pinion-high" and "low" nacelle led to the development of a similar set of contours for the XT701 drive system. A number of important differences in the nacelle arrangement occurred as the result of the contour refinement. First, because the nacelle/ wing in- tegration was not an optimized arrangement, the shaping of the maximum beam of the nacelle beyond the location of the maximum cross-sectional area was changed to allow the maximum beam dimension to remain constant over the aft portion of the nacelle. Second, in the previous designs, the engine and oil cooler ducts were arranged to be opposite each other; in the revised design, the ducts are arranged in a stacked and staggered configuration to best use the available space wi thin the nacelle envelope. The nacelle layout, contours, and area distribution are shown in Figure B-16. Further refinement of this nacelle was performed to change the shapes for structural and manufacturing simplicity.

The major change was the slight increase in nacelle diameter to permit the use of major structural elements from the Lockheed P-3C nacelle and the introduction ofa constant shape for the upper portion of the nacelle, allowing straight sides on the nacelle. The final design for the XT701 contour sand area distribution are shown in Figure B-17.

Prop-Fan and Nacelle Ratios - The nacelle/prop-fan diameter ratios are shown in Table B-IV. The nacelle diameter, D , is the equivalent diameter N

based on the maximum cross-section area, or reference area occurring at X/L =

0.414. These data show that new nac.elles can be configured with the desired diameter ratio but that some compromise is necessary if existing nacelles are to be used.

CANDIDATE PROPELLER DRIVE SYSTEM INSTALLED PERFORMANCE Installed performance for the prop-fan drive systems using eight bladed prop-fan data are provided in Figure B-18 for takeoff conditions and at al- titudes of 4572m (15,000 ft), 7620m (25,000 ft), 10,668m (35,000 ft) for the XT701 and GE T64 drive systems and 4572m (15,000 ft), 7620m (25,000 ft), 11,OOOm (36,089 ft) for the T56 drive system. Jet thrust data are also given.

NACELLE STATIONS o 20 405060 80 100 120 140

10 3p' T I Y lInT I

I I I I I I I I I I PINION HIGH CONFIGURATION

I I . I II

*B~~I

I I

90 120 to 140 NACELLE SECTIONS

A/~'l.O~r--A

0.8 A/A MAX o ., !l.t. O.b !l.d 1.0 Figure B-16. XT701 Nacelle Common Contours o 70 SO 90 100 110 NACELLE SECTIONS 1.2 A/Amx = 1.0 Figure B-17. XT701 Nacelle Final Contours TABLE B-IV. SUMMARY OF NACELLE/PROP-FAN RATIOS J? D~l '':-!Ax i Z ,rr) i D~I/D?

~.u:!u'E 3 2 (Y':') 'FT)

I ~

I I (9.;) :(701 :;tinioD-Low 1.00 (l0.77) 2.89 1.U '3.;0) 10.Jd8 I i *T56 P1Qioa-Lov 1.00 (10.77) 2.:'7 (8.1) 1.13 '3.70) 10 •• 5 I (9.5) lC!701 ?iD.1oa-iU,h 0.122 <9.93) 1.39 l.08 0.36) i 0.;58 I *TS6 ?11l1oa. ... H1~h 0.922 (9.'n) 2.:t7 LJ8 lJ.56) ! G.!.';'

'.8.1) I

, , ..

~~. ?1nioa-Lov 0.515 (5.55) (6.97) ').Sl ~::.b6) : O •. HI I (5.;;) 1'64 Pinion-fUga. ·).515 2.1: 1.31 l2.oe, 10.J8

(6.H) I

I I ?inl,011 ... Lov (l2.40) C-130 - Tl6 l.ll 2 •• 7 \a.l) 1.~l i3.97) I 'J.~9 i I C-130 - ~56 P1a.i011-Wv 0.d7 (9.3;) Z.:'; (S.l) l.O5 (J ... Sf ! 0.:'2

I

(Revts.eI) P3C - :56 ~141o'Cl ... Ulb 1.88 (:0.:0) 1 ... 7 ,3.1) ~.!6 15.071)10,1)2 (3.53) 1 .:'7 P3C - !56 l'in1oa-K1gh. 4001 (ll.31) :. .. ;.] (8.1) !..l7 0 ~R.evtaecl) :cr701 P1a.1on-Hi,b/Lav 0 •• 45 (9.1) :.89 (9.5) i..u4 0.3.5 'l.H)

I

.~ • tteteraac.. ero .. Sectional .\rea at I./L • O ... 1 ..

)1 • ?ro1)-Faa. Ou-car O~~ • ~.c.ll. Equivalent il1aaetC'£' ~ :t../L • 0 ... 14 *tha •• tlac_Ue .:onfiaurac1oas were a .. UDad :0 o. ch • .lame a.s the .:cr;Ol 'eca.use of the si.':llllar1t1 .. ot ctl. X't101 and. :,,6 enlin... :!'l. rS6 :taceU.s ':oUl ~. :-e- tined to ~1v. O~t/0l' • 0.J5 .:on.ahUllt ntb tM Z.:'7 0; (8.1 C'n ,U_tar ,rap-ian.

Ts6-A-1s XT701 LB ~ 10- LB X kN kN 10- -3 ALT X 10 ALT X 10- m (FT) m (FT) ~ VJ SL } ::::> 20 T.O. POWER 4.57 ( 15) 1.52(5) 40 $4 SL } INTERMED lATE E-< E-< 8 4.57(15) POWER VJ !;l MIL POWER M =.8 §l z 3 30 10.97(36) ~6 E-< ~ z } NORMAL 7.62 (25) } POWER ____ --...:IL......- MAX 4.57 (15) 10.67 (35) CONT JET THRUST ;;:.SL 7.62 (25) JET THRUST 1.52(5) 10.97(36) DATA NOT AVAILABLE ) ( ASSUME 10% PROP THRUST 0 0 0 0 100 200 300 m/s 100 200 300 m/s , , , , 0 i 600 KTS o 200 400 600 KTS TRUE AIR SPEED TRUE AIR SPEED Figure B-18. Drive System Installed Performance T64-GE-415 LB x kN 10- -3 ALT x 10 6 m (FT) t; SL ~ } MAX POWER 0:: ~4 1.52 (5) E-< ~ Z 3

4.57~

7.62~}NORMAL

POWER 10.67 (35)- 4.57 (15) SL & 1.52 (5) JET THRUST 7.62 (25) 10.67 (35) 200 m/s 0 100 300 , » 600 KTS o 200 400 TRUE AIR SPEED Figure B-18. Drive System Installed Performance (Cont'd) REDUCTION GEARBOX MODIFICATIONS FOR PROP-FAN APPLICATION Application of the available gearboxes to the prop-fan drive systems re- quires modification of the gearboxes. to match the power-section performance, direction of rotation, and the sizing requirements for the prop-fans. Free turbine power sections have an advantage over fixed-speed, single-shaft cores through the ability to vary engine speed and power level to change prop-fan tip speed. Only one modified gearbox is required for the free turbine units, whereas a separate gearbox would be required for each tip speed for the single-shaft, fixed-speed engine.

T64-2 SDG Reduction Gearbox The gearbox, shown in Figure B-19, consists of a planetary section and an offset section. Conversion to prop-fan application requires modification of the offset section only. The current planetary gearbox, rated at 2535 kW (3400 I I J '-§.: Figure B-19. T64-2 SDG Gearbox shp) , could be used with a modified offset gearbox to give the required tip speed. Modifications to the offset section do not require major changes to the housing castings but do need: o Minor modification to the offset section castings o New offset gears o Modified gear mesh lube o New lube tube/seal/spider The gearbox can be used either in the "pinion-high" or "pinion-low" con- figuration without changing output.

The gearbox for reverse rotation would require a completely new design for the offset section.

DDA T56-A-14 Gearbox This gearbox is a "pinion-high" configuration as used in the Lockheed P-3C "Orion"; it can be adapted for prop-fan application using either the T56 or XT101 power sections, although the modifications required to match the XT101 to the prop-fan are complicated by the rotation of the XT101, which is opposite to that of the T56 and by the significantly lower RPM of the XT101.

Gearbox Modification for the DDAT56 - The following modifications are re- quired to match the T56 power section for tip speeds of 1S3, 213, and 244 m/s (600, 100, and sao fps): Rework Required- o Machine rear housing for larger pinion gear o Reroute oil supply to pinion bearing externally o Machine main diaphragm to ~eroute oil supply New Parts Required: Operation of the prop-fan at tip speeds of 1S3, 213, and 244 m/s (600, 700, and SOO fps) will require three gear set configurations: o 6ST. SP Pinion Gear 244 m/s (SOO fps) o 10ST, 8P Main Drive Gear

I

o 63T, 8P Pinion Gear 213 mls (700 fps) o 113T, 8P Main Drive Gear o 56T, 8P Pinion Gear 183 mls (600 fps) o 120T, 8P Main Drive Gear o Offset Pinion Gear Lube Nozzle Gearbox Modifications for the DDA XT701 - Since the XT701 is a free turbine power section, only one set of gears is required. The modifications to the gearbox for use with the DDA XT701 are as follows: Rework Required: o Machine rear housing for larger pinion gear o Machine rear housing for added idler gear spindle o Reroute oil supply to pinion bearing externally o Machine main diaphragm to reroute oil supply o Machine nose bearing plate for nose oil pump o Machine clearance on inner diaphragm for idler gear New Parts Required: o 68T, 8P Pinion Gear o 108T, 8P Main Drive Gear o Offset Pinion Gear Lube Nozzle o 61T, 10P Alternator Gear o 78T, 10P Nose Oil Pump Drive Gear o 33T, 10P Nose Oil Pump Driven Gear o Idler Gear Spindle o NTS Helical Spline Coupling o Lockout Spacer in Prop-Fan Brake In addition, the XT701/T56-A-14 combination would also require: o New torquemeter and housing o New compressor inlet adapter ring for interconnecting strut attachment o New interconnecting struts The modified T56 gearbox is illustrated on Figure 8-20.

MAIN DRIVE GEAR --REAR CASE INNER DIAPIIRAGM OIL SEAL pmr SIIA~fT~~~~m~~~~~~~~~trru~-SUN GEAR IIUB

~;~st~~-REAR CARRIER

t:;;~~~ ~~~~~~~~~l_PROP BRAKE

'\ SUN GEAR NOSE SCAVENGE PUMP SCAVENGE PHESSURE RELIEF VALVE---- .... /'-~ ACCESSORIES DRIVE GEAR Figure B-20. T56 Gearbox Modifications DRIVE SYSTEM QUICK ENGINE CHANGE (QEC) UNIT The design approach to the drive system was to consider the drive system as an independent unit and, therefore, independent of the subsequent receiving airframe. Since QEC unit weight was not a critical item, the design approach included over-design of the drive system mounting structure to permit universal application of the unit.

In the case of the DDA T56/XT701 drive systems, the QECs would use support structure from the C-130 "pinion-low" nacelle and P-3C "pinion-high" nacelle.

A typical QEC unit, shown on Figure B-21, consists of the bare drive system housed in a nacelle complete with mounting structures, air induction systems, exhaust systems, subsystems such as starting and electrical, the prop-fan and engine controls, and the lubrication system. The unit is designed for ease of assembly/disassembly at the parting plane, which is the juncture between the

trlo_OOj

0,--- 12 7 :

o Loo_ooJ

7 LUBRICATION SYSTEM

1 POWER SECTION

8 CONTROLS

2 GEAR BOX

9 STARTING SYSTEM

3 TORQUE METER

10 NACELLE STRUCTURE

4 PROPFAN

11 DOORS AND PANELS

5 AI R INDUCTION SYSTEMS

12 RES IDUAL OIL AND GREASE

6 EXHAUST SYSTEM

Figure B-21. Typical QEC QEC and the fixed portion of the nacelle on the airframe. The QEC design ca~ be varied for either an underwing or an overwing installation. Investigation of an overwing arrangement shows that the nacelle structure can be assembled from P-3C nacelle parts as follows: P-3C Parts Used in QEC Quantity Name P-3C Part No.

Forging Machined 918468 1 Assembly U-Frame 839475-1 Assembly U-Frame 839475-2 1 Mount 632238-7 2 Mount 632238-9 Mount 632238-51 Mount 918611-3 2 Mount 918611-1 2 The lubrication system can be designed to use the oil cooler from the C-130 identified as part No. 697226.

Drive System QEC Weights - The weights for each QEC unit for the three candidate drive systems are shown on Table B-V.

TABLE B-V. QEC UNIT WEIGHTS QEC UNIT WEIGHTS Configuration Drive System Underwing Overwing 1669 (3680) GE T64 kg (LB) (4366) 1827 (4030) 1980 T56 kg (LB) 1953 (4307) (LB) 1800 (3971) XT701 kg Drive System Controls Prop-Fan Control System Description, T56 and XT701 - A feasibility study was conducted by Hamilton Standard to determine the suitability of the control used for 54H60 propellers on the Lockheed C-130 and P-3C aircraft. This control readily fits a 60-spline shaft such as that of the T-56 engine (a fixed-speed engine) and also has a high pumping capacity. Compatibility of the control has also been established for the DDA XT701 free turbine power section, but some modification is necessary to achieve variable-speed capability. The control, which currently operates at 1020 RPM is designed for pump flows of 0.057m /min (60 quarts/min). Following a whirl test on a modified 54H60 con- trol and propeller hub at 1800 rpm, it was concluded that the control could operate at this speed and was capable of withstanding the loads imposed if minor modifications were performed. These modifications would consist of: o Replacement of standby pump drive gear o Increased clearance for the transfer bearing o Speed bias and linkage removal o Redesign of governor flyweights and speeder spring o Removal of Beta control differential gear train o Brushlock removal or revision o Addition of a heat exchanger for transfer bearing cooling In the study, prop-fans having diameters of 2.47 and 3.05m (8 and 10 ft) with eight or ten blades were examined. Although it is feasible to use the control for these conditions tit was found that the pitch change rates for 2.44m (8ft) 8 blade and for the 3.05m (10 ft) 8- and 10-blade combinations were slow and well below the rates considered acceptable for rapid transients.

Control functions examined include negative torque sensing (NTS), overspeed prevention, normal governing, feathering, and reversing.

It has already been established that the control is compatible with either the T-56 or the XT701; 1'n the case of the T-56 t however, a negative torque sensing system is required and is the only control hardware difference between the two engines.

Normal governing can be accomplished with the modified control and is independent of the type of engine.

Feathering is likely to be slow because of the low pitch change rates, as shown in Table B-VI and feathering out of an overspeeed condition where higher pitch change loads exist may not be possible with the modified 54H60 control.

With the provision of adequate overspeed protection, however, this may be in- consequential.

Unfeathering using the modified control does not appear to present problems since an electrically driven auxiliary pump already on the control will be used.

Since the use of the 54H60 pitch-lock is not feasible in the prop-fan actuator, and conversely, the prop-fan pitch-lock concept is not compatible with the 54H60 control, some form of pitch-lock device should be incorporated into the prop-fan rotating hardware to prevent overspeeding in cases of in- advertent decrease in blade angle. A number of arrangements have been con- sidered, and an electrically operated in-flight stop is considered to be feasible. This device would provide testing flexibility, although it would be necessary to be certain of the stop location at all times if overspeed pro- tection is to be provided.

Engine Control -- T56 and XT701 A hydro-mechanical engine control system having an electronic supervisory system will be used in conjunction with the 54H60 prop-fan control.

TABLE B-VI. PROP-FAN PITCH CHANGE RATES PITCH CHANGE RATE RADS/SEC (DEG/SEC) DIAM.

NO. BLADES MAIN & MOD MAIN PUMP m (FT) ONLY STANDBY PUMPS (14.5) 8 0.158 ( 9.05) 0.253 2.44 (8) 0.457 0.285 (26.2) 10 (16.31) 2.44 (8) ( 7.8) (10) 8 0.069 ( 3.92) 0.136 3.05 3.05 (10) 10 0.124 ( 7.11) 0.248 (14.2)

APPENDIX C - CANDIDATE TESTBED AIRCRAFT - TASK III The procedure adopted for the identification and selection of the candidate testbed aircraft consisted of a two-level screening process. A survey of all NASA aircraft was conducted, and a list of those aircraft capable of meeting the specific design requirements was compiled. In addition, aircraft not in the NASA inventory were also included where suitability was established.

An initial screening of the list of these aircraft was conducted, and those that appeared unsatisfactory or marginal were eliminated. Such criteria as lack of compatibility with commercial passenger transport configurations, air- craft and prop-fan scaling mismatch, adverse location of the prop-fan, marginal aircraft performance, insufficient ground or component clearances, and lack of potential for modification, provided the basis for elimination. Following the ini tial screening, the aircraft remaining consituted the list of candidate testbed aircraft and were subjected to further and more detailed analysis.

CANDIDATE TESTBED AIRCRAFT DESIGN REQUIREMENTS The design requirements for a testbed vehicle for flight research testing of a propeller of advanced design were as follows: o Speed/Altitude - Mach 0.8 @ 9144m (30,000 ft) and above o Capable of operating safely at normal flight conditions with the prop-fan powered or unpowered o Takeoff and landing restrictions for the prop-fan operation acceptable o Vehicle to be configured with one prop-fan drive system o Sufficient primary propulsion to be retained to permit operation of the vehicle with the prop-fan powered or unpowered o Non-optimum drive system installation acceptable In addition, the testbed vehicle was required to provide a stable platform for accurate measurement of flight test data and proper simulation of the environment in which the prop-fan could be tested to satisfy the program objectives, and be large enough so that the aircraft geometric ratios would be representative of large-scale propulsion. The selected vehicles should also be capable of modification to multi-prop-fan testbed configurations.

AIRCRAFT SURVEY A preliminary list of suitable aircraft was compiled from the approximately 110 aircraft in the NASA inventory, consisting of: Lockheed C-141A (L-300) Lockheed JetStar -6 Convair 990 Gulfstream American Corporation "Gulfstream II" Boeing 737 Boeing KC135A (707 - 100) Boeing B52H Other aircraft considered included: McDonnell-Douglas DC9-10 Boeing 727 BAC 111 These three aircraft were not pursued as testbed configurations for the following reasons: o DC9-10 - McDonnell-Douglas. were under contract to the NASA-Lewis Research Center to examine the aircraft as a testbed vehicle. Inclusion in the study would have resulted in some' duplication of effort.

o Boeing 727 - Omitted as a candidate since this aircraft does not appear in the NASA inventory.

o BAC 1-11 - Does not meet speed/altitude design requirements and is a foreign aircraft.

These three aircraft (as are the JetStar and GIl) are all aft-mounted propul- sion configurations each presenting a clean wing for prop-fan application.

The survey included the physical location of each aircraft, the current or This information is given in planned configuration, and the availability.

Table C-I.

TABLE C-I. AI RCRAFT SURVEY CURRENT OR PLANNED AIRCRAFT MODEL NO. LOCATION CONFIGURATION AVAILABILITY COMMENTS Lockheed C-141A C-141A Ames RC Telescope Program Not Avail.

Dryden FRC Lockheed JetStar -6 Prap-Fan Acoustic Tests Mid '81 Slipper Tanks Will No.3 '83 be Removed LFC Program Convair 990 10-37 Ames RC Airborne Instrument Lab Nat Avail.

High Operating Costs AF Program 11-29 Ames RC

ffi

G"'."~m "ffi

Johnson RC Shuttle Simulator Trainers Not Avoil.

Langley RC Terminol Area Not Avail.

Boeing 737 -100 Configured Vehicle 19437 Program

-

Boeing KC-135 Returns to AF at High Operating Costs Edwards AFB On Laon - Wing let Program End of Phase I

CD

Zera G Johnson RC Not Avail.

CD

Service life Expended In '85 Edwards AFB X-IS, RPV Hymet Boeing B-52 B Available Joint Program AF F III TESTBED AIRCRAFT/PROPULSION SYSTEM CONFIGURATION The installation of the prop-fan drive systems on the testbed aircraft falls into two categories: o Prop-Fan Propulsion System Substitution - This type of propulsion system arrangement requires the removal of an existing primary propulsive unit and the substitution of a prop-fan propulsion system at the same location.

o Prop-Fan Propulsion System Addition - All existing primary propulsion is retained and the prop-fan propulsion system is added to the aircraft con- figuration.

The prop-fan propulsion system installation can be further defined by the location on the aircraft wing, i.e., for a "pinion-high" drive system config- uration, the installation would generally be an overwing configuration, whereas the "pinion-low" arrangement would usually correspond to an underwing location.

The prop-fan substitution arrangement, from the structural standpoint, would provide the best potential for modification, since wing structure to support the power plant would already be available to accommodate any candidate drive system with minimum modification. In those cases where the prop-fan installa- tion would be an addition to a wing, the structural changes would probably be much more extensive.

The location of the prop-fan on the aircraft should be selected such that an environment exists that would permit testing the system throughout the full range of operating condl tions. The configuration design must, therefore, be conducted so that the testbed aircraft will achieve the testbed program objec- tives.

The primary objective of the testbed is the verification of structural integrity, first of the prop-fan and second of the nacelle/airframe structure.

It is, therefore, desirable that some means of changing excitation factor should be included in the design, and several means of accomplishing this have been considered, including variable toe-in and droop angle for the nacelle and leading-edge extensions to increase blade proximity. The non-symmetry of the nacelle due to the presence of unsymmetric air induction systems is also of concern, since the area distribution of the spinner/hub/nacelle may be affected.

Because the prop-fan system is to be installed on an existing aircraft, the degree of nacelle/wing integration optimization is limited. It is expected, however, that some contouring of the nacelle/wing interface can be included in the configuration design.

Investigation of near-field noise can be conducted in an environment that closely simulates that of a large-scale propulsion system with proper suppression of the prop-fan drive system noise. The prop-fan fundamental signal can be isolated and the higher frequencies made to dominate the noise spectrum so that clear signals can be obtained over the entire spectral range of frequencies. Provision must also be made to include testing of various noise attenuation concepts in the fuselage.

Testbed Aircraft Configurations Testbed aircraft configurations were developed for the following aircraft and drive system combinations: Drive System Aircraft XT101 pinion-high and low Lockheed C-141A (L300) T56 pinion-high and low T56 pinion-high Lockheed JetStar -6 T64 pinion-high XT101 pinion-low Boeing 131-10 T56 pinion-low T64 pinion-high and low XT101 pinion-high and low Boeing KC-135A (101-100) T56 pinion-high and low XT101 pinion-high and low Boeing B52B T56 pinion-high and low XT101 pinion-high and low Convair 990 T56 pinion-high and low T64 pinion-high and low XT701 pinion-high GAC "Gul fstream II" T56 pinion-high C-141A Testbed Configurations - The C-141A configured as a testbed using the T56 drive system is shown in Figure C-1. This testbed configuration re- quires the removal of the left-hand inboard turbojet engine and the sub- sti tution of the prop-fan drive system. Two arrangements are shown for the T56; three views show the installation for the "pinion-low" gear box arrange- ment, and the auxilliary views illustrate the "pinion-high" overwing installa- tion. The arrangements for the XT701 drive systems are similar, except that the installation would be substituted for the right-hand inboard turbojet. Of the two installations, i.e., "pinion-high or low," the "pinion-low" arrangement leading to an underwing configuration is preferred because of the reduced length of the exhaust duct nacelle, no interference with the trailing-edge flaps or spoilers, and because this installation is more favorable for conducting acoustic tests.

JetStar Testbed Configuration - JetStar -6 testbed configurations were gen- erated by adding the prop-fan installation to the left-hand wing at the loca- tion of the external fuel tank, which is removed. The T64 "pinion-high" test- bed configuration is shown in Figure C-2. In this configuration, sufficient ground clearance exists for the 2.13m (7 ft) diameter prop-fan and the prop-fan sweep does not overlap the main aft mounted propulsion system. Ground clearance in the normal attitude is marginal for the T56 mounted at the same location, Figure C-3. In the rolled attitude, however, a tip/ground interference occurs.

The XT701 installation was not included in the JetStar stud ies, as the 2.89m (9.5 ft) prop-fan diameter would result in a ground interference condition.

Boeing 737 Testbed Configuration - The Boeing 737 configured as a prop-fan testbed is shown in Figure C-4 for .the T56 "pinion-low" drive system configura- tion. Since the prop-fan drive system is an addition to the configuration, it is located outboard of the left-hand turbofan at BL 293 LH. In the case of the XT701, the " pinion-low" drive system arrangement, Figure C-5, would be located on the right-hand wing at BL 293 RH. Both the T56 and the XT701 "pinion-high" installations could be located above the wing at the 'same but opposite spanwise location. The T64 installations for both "pinion-low" and "pinion-high" at BL BL 281.8 -~:=:==~ T56 OVERWING INSTALLATION BL 281.6 ~:E=?- T56 UNDERWING INSTALLATION Figure C-l. C-141A Testbed Configuration B. L. 154. 68 --1'--++1-- I F.S. 376.69 Figure C-2. JetStar -6, T64 Testbed Configuration ~""'----"\-1~:-----"'~ - -- F . S. 365.89- TIP !FUSELAGE ' CLEARANCE

1.56mJ

B.L. 154.68 (5.1 FT) 'CLEARANCE 0.27 m (.9 FT) L __ t::::::====~(8)=::::::" ::::::::=---;

t

0.17 m (6 • .6 in.) INTERFERENCE ROLLED ATTITUDE 0.269 m (10.6 in) CLEARANCE NORMAL ALTITUDE Figure C-3. JetStar -6, T56 Testbed Configuration PROP DIA : 2.46 m (8.1 FT) B.L. 293 .0- ~~~~~..\- 1. 81 m (5.9 FT) ---+---...-1 2.46 m (8.1 FT)-i---o-!

,~ ~ .. __ ._. ___ [l~

o

. W.L. 197.0 _.

ENGINE 1. WL 186.0

Figure C-4. Boeing 737-10, T56 Testbed Configuration ENGINE ~ (W.L. 184.0)

I

B.L. 293.0 W.L. 195.0 PROP DIA = 2.89 m (9.5 FT) INTERFERENCE 0.26 m (n.87 FT) Figure C-5. Boeing 737-10 Testbed Configuration XT701 Pinion-Low 293 are shown in Figures C-6 and C~7. The "pinion-low" installation, Figure C-6, is slung beneath the wing and does not interfere with the trailing-edge devices as is the case with the "pinion-high" installation, Figure C-7.

KC-135A Testbed Configuration - The KC-135A configured as a prop-fan test- bed aircraft is shown in Figures C-8 and C-9 for the T56 underwing installa- tion, and in Figure C-10 for the overwing.

The corresponding installation of the XT101 on the right-hand wing is shown in Figures C-11 and C-12 for the underwing installation and in Figure C-13 for the overwing installation.

Ground clearance is not a problem with any of the installations, since the lateral clearance angle is determined by outer engine ground contact.

Variation in prop-fan si ze, however, would be constrained with the under- wing installations to 3.12m (12.2 ft) and to 5.18m (11 ft) for the overwing in- stallations.

!lIGI!iE i. "II.!.. 186.0" B.!,. 293.0 Figure C-6. Boeing 737-10 Testbed Configuration T64 Pinioo-Low "ac=~ II.!.. 230.0

I

B.!.. 293.0 II.!.. '222.0 I \ \ PROP DIA 2.19 .. (7.2 F"I)

o

Figure C-7. Boeing 737-10 Testbed Coofiguration T64 Pinioo-High PROP-FAN DIA B.L. 2.47 m 323.295 (B.1 FT) ~::::::~ I -W.L. B4------ __ ~ __ ~tlt~~~~~~~==--------~~-- CLEARANCE ANGLE 0.14B RADS (B.5 DEGS) WING MAC 6.14 m (20.16 FT) Figure C-8. Boeing KC-l35A Testbed Coofiguration T56 Pinion-Low NACELLE ~ Vi.B.L. 315.0 IN W.L.203.83 /W.L. 203.8: '-----I-+-_++_ +---I-,d/C----------- PROP-FAN DlA.

2.46 m (8.1 FT) F.S. 784.35 I B.L. 321.64

w.L. 203.83

F. S. 821. 72 Figure C-9. Boeing KC-l35A T56 Pinion-Low Installation PROP-FAN DlA.

:? .46 m " 746.98 NACELl.r Ii.

\\ .B.L. 315. n IN WING CHORD PLAKE 3:::~.r" F.S.

W.L. 2L.3.83 Figure C-I0. Boeing KC-135A T56 Pinion-High Installation W.L. 230 I I ,

111'

\ I I

\ I ' ,I

I ; \1

! i-r'

I I I r ~.

i I

I W.L. 190.432 2.89 m (9.5 FT) PROP DIA B.L. 323.295 . --W.L. 230 ---e;=n~=:::::--r=,~::::;====::::t-+--~--+ CLEARANCE ANGLE 0.148 ~~S

on

i OlL.

Figure C-ll. Boeing KC-135A Testbed ConfigurationXT701 Pinion-low NACELLE i W.B.L. 315.0 IN WING CHORD PLANE 203.83 PROP-FAN DIA. / 2.89 m (9.5 FT) F.S. 828.19 784.36 321.65 203.83 740.54 Figure C-12. Boeing KC-135A XT701 Pinion-Low Installation NACELLE ~ WBL 315.0 IN WINC CHORD PLANE

'. .. ~:I II "Lt'-' .-- --- -'-'-'-'-'~=--==--=::::::~-+--I---Q:.4~F===

__ "'--L_. II ~ . t!: -- - --"-'-"_.. .

. . I I --.---

l' il--II:7. __ . __ .-. __ . ___ . ---

....... ._----.-

[- F .5. 828.18 W. L. 203. 83 ~I ___ .......... _____ - ___ -, W.L.

231.62 PROP-FAN DIA 2.89 m (9.5 FT) B.L. 318.23 F.S. 740.53 W.L. 218.62 Figure C-13. Boeing KC-l35A XT701 Pinion-High Installation In the case of the overwing nacelles, the nacelle interferes with the high- speed aileron to the extent that this control would be eliminated if a long jet pipe is required. Lateral control at high speed may, therefore, be a problem for this configuration.

Boeing B-52B Testbed Configuration - The Boeing B-52B Testbed Configuration featuring the T56 underwing installation is shown on Figures C-14 and C-15, and These in- the corresponding overwing installations on Figures C-16 and C-17.

stallations are located on the wing at BL 217, which is the existing pylon mount for equipment test purposes.

The installations illustrated on Figure C-18 and C-19 are for the XT701 underwing, and on Figures C-20 and C-21 for the overwing installation.

These installations suffer from the following disadvantages: o In all cases, the depth of the nacelle is approximately the same as the depth of the wing. This, in conjunction with the high incidence angle and leading-edge sweep of the wing, complicates the integration of the wing and nacelle and the arrangement of the engine support structure. An additional O.3m (12 in ) was added to the nacelle behind the QEC parting line to simplify nacelle/support structure/wing integration.

o The magnitude of the wing chord at the pylon mount is very large compared with the diameter of the prop-fan, with the result that the scale effect will be such that the prop-fan will not significantly affect the wing flow field and would, therefore, not present a realistic situation for assessing the aerodynamic influence of the prop-fan on nacelle/wing com- binations. Furthermore, the wing blockage particularly for the underwing installations almost obscures the prop-fan swept area.

One other effect of the large chord is the inordinately long jet pipe re- quired for the overwing installations leading to increases in installed weight and additional losses in jet thrust.

Since the wing section data for the B-52B were not available, a repre- sentative section, NACA 0012-64 base thickness form, was scaled to the appro- priate thickness for the wing. Because of the high incidence angle of the wing F.S. 100 F.S. 535.19 W.L. 100~;L----~~==~====::::~~------------ --------------~~--4'~------~~----~~------------- B.L. 217.06 Figure C-14. Boeing B-52B Testbed Configuration T56 Pinion-Low v..'.L.

236.00 PROP-FAN DIA.

2.46 m (8.1 FT) F.S. 595.89 NACELLE 'i 218.20 F. S. 644.38 B.L. 218.00 F.S. 692.87 W.L. 236.00 Figure C-15. Boeing B-52B T56 Pinion-Low Installation .s:::.;~~--F.S. 535.19 B.L. 218.21 -ILL. 250.50 Figure C-16. Boeing B-52B Testbed Configuration T56 Pinion-High w.L. 236.00 F.S. 595.89 PROP-FAN DIA.

2.46 ID (8.1 FT) F. S. 644.38 F.S. 692.87 B.L. 218.00 W.L. 236.00 Figure C-17. Boeing B-52B T56 Pinion-High Installation F. S. 5 17 • 89 -S;:!;:;:;:::2..

B.L. 217.06 IF'

.'< U

\.l.t. ~14.51

~-'I-"')--

~

--j 00 00

Figure C-1B. Boeing B-52B Testbed Configuration XT701 Pinion-Low W.B.L. 218.20 I' .L.

236.0(', F.S.

PROP-FAN DIA 2.89 m (9.5 FT) B.L.

F.S. 591.82 W.L.

Figure C-19. Boeing B-52B XT701 Pinion-Low Installation I I .

F.S. 536.39 ....s::;;~t:;2 B.L. 218.21 VI.L. 252.99 Figure C-2O. Boeing B-52B Testbed Configuration XT701 Pinion-High W.B.L. 218.20 i NACELLE PROP-FAN DIA F.S. 687.18 236.00 2.89 m (9.5 FT) F.S. 644.39 B.L. 218.21 F .5. '601.58 W.L. 236.00 Figure C-21. Boeing B-52B XT701 Pinion-High Installation and the high-speed capability of the aircraft, the section was assumed sym- metrical.

Convair 990 Testbed Configuration - The Convair 990 configured as a testbed is shown in Figure C-22 for the DDA T56 engine underwing installation. As a , , , B.L. 266.0-~~~~~ REMOVE ANTI-SHOCK BODIES

W. L • 44. 0 ~~~~:::;::;;;:::::::::>

co

Figure C-22. Convair 990 Testbed Underwing Configuration T56 Pinion-Low prop-fan testbed using the T56 engine, the inboard left-hand primary engine is removed, and the T56 nacelle is substituted either as an underwing pinion-high installation, Figure C-22 , or an overwing pinion-low arrangement, Figure C-23.

This configuration also requires removal of the anti-shock bodies so that test equipment, i.e., pressure rakes, can be fitted to the wing and to ensure that the prop-fan wake over the wing is not influenced by existing components of the aircraft.

\ \ \ \ \ \ \

-.-

-"

.-.-

B.L. 266.0 '.

\.

Figure C-23. Convair 990 Testbed Overwing Configuration T56 Pinion-High Similarly, in the case of the XT701 testbed configuration, the right-hand inboard engine is removed and the prop-fan is substituted either as an under- wing installation, Figure C-24, or an overwing installation, Figure C-25.

Of the two principal configurations the underwing arrangements are pre- ferred since there is no effect upon the trailing edge devices.

1"":2.89 .. ...J I (9.S FT)

.J.~ • • O.O-E-~.;!.--'=::::=:::~~==:::==:::::'"

.,. ".,11.... __ /ENG_r:_::::._C::.~:_::~_._E[\1.:L.:8:4_.:....0_~-=~~

Figure C-24. Convair 990 Testbed Figure C-25. Convair 990 Testbed Underwing Configuration XT701 Pinion-Low Overwing Configuration XT701 Pinion-High GAC Gulfstream II (GIl) Testbed Configuration - The T56 GIl testbed is shown in Figure C-26. This aircraft has the advantage that the prop-fan pro- pulsion unit is an addition to the configuration rather than a substitution.

Furthermore, the wing leading edge has no high-lift devices, which simplifies the nacelle/wing integration.

1. 8 m (5. 91 FT

I

.F. S.

I

F.S. 389.40

I

W.L. 120-t~----

W.L. 85.8~--~~==~====Sfr~~~~::~~=: STATIC GROUND LINE \\ II Figure C-26. GAC Gulfstream II Testbed Configuration T56 Pinion-High Because of the size of the aircraft, overwing installations only are possi- ble and are shown on the left-hand side of the aircraft for the T56, Figure C-26 , and for the XT701 on the right-hand side of the aircraft, Figure C-27.

Since the wing thickness increases from BL 145 to the center, the engines are located at BL 145 to take advantage of the structural characteristics of the inboard wing.

F.5. 318.60 . =:=:: - .

i-==~

, I Figure C-27. GAC Gulfstream II Testbed Configuratioo XT701 Pinion-High Testbed Aircraft Performance Testbed mission analysis is based upon the design requirements of a flight Mach number of 0.80 at an altitude of 10,668m (35,000 ft), at standard atmos- phere conditions.

For purposes of comparison, the following assumptions have been made: o The prop-fan operates at zero net thrust except during test operations.

o For those aircraft where substitution of an original engine with a prop-fan occurs takeoff distance is computed using three-engine ferry rules.

o For large aircraft, e.g., C-141A, KC-135, Convair 990, a fuel allowance of 453 kg/hr (1000 lb/hr) has been included for operating the prop-fan, regardless of the drive system type.

o Although the test mission profiles have not been defined, it has been assumed that it will be necessary to obtain data over a wide range of prop-fan thrust values. The start test weight has, therefore, been limited to that which will provide the capability of achieving a Mach number of 0.8 at 10,668m (35,000 ft) with zero net thrust from the prop-fan.

o Reserve fuel allowance is sufficient for approximately one hour of flight time at low speed and low altitude.

o For the smaller aircraft, e.g., JetStar -6, and the GIl, ballast to correct the lateral imbalance is included in the zero fuel weight, ZFW.

Time to complete a flight test at a Mach number of 0.8 at 10,668m (35,000 ft) with several testbed aircraft/drive system configurations is presented in Figure C-28. The corresponding test mission profile is also shown in the figure.

BEGIN TEST AT

II 10668 m (35,000 FT.) ,8M

E~ID TEST AT 1066B m (35,000 FT.) .BM

l

~RUNWAY CONDITIONS

ELEVATION - 701 m (2300 FT.)

o

II TEMPERATURE - 27 C ( 80 F)

~IO WIND NO GRADIENT

I

t/ TAKEOFF ~ LANDING DURA- DRIVE START AIRCRAFT RAMP WT Kg (LB) ENG TEST WT Kg (LB) TION (LB) .

TEST WT Kg SYSTEM (RRS) 95,254 (210,000) XT701 C-141A 99,271 (218,900) 66,224 (146,000) 4.33 GIl 27,341 (60,288) 26,303 (58,000) XT701 19,484 (42,960) 3.38 KC-135A (183,000) 75,734 (167,000) 156 82,990 48,615 (107,200) 4.69 CV-990 87,072 (192,000) 82,537 (182,000) T56 60,769 (134,000) 4.2 114,452 (252,376) (246,476) XT701 B-52H 111,777 81,982 (180,770) 4.5 T64 JETSTAR-6 18,367 (40,500) 16,779 (37,000) 15,372 (33,900) 0.65 33,407 (73,500) 31,818 (70,000) T56 B-737 ZFW30,339 (66,900)

-

Figure C-28. Testbed Aircraft/Drive System Configuration Performance C-141A Testbed Configuration Performance - Capability of the C-141A testbed configuration to meet the design conditions is shown on Figure C-28. A start test weight of 95254 kg (210.000 Ib) was selected for the testbed configura- tion. At takeoff the ramp weight is 99271 kg (218.900 Ib) and the takeoff distance is 1890m (6200 ft) at an airport elevation of 701m (2300 ft) and at 0 0 temperature of 300 K (80 F). Test mission duration is 4.33 hours.

JetStar -6 Testbed Configuration Performance - Capability of the JetStar -6 configured as a testbed is shown in Figure C-28. for the GE T64-powered prop- fan. Starting at a ramp weight of 18.367 kg (40.500 Ib) and climbing to test altitude. the start test weight is 16.719 kg <31.000 1 b). At a zero fuel weight of 13,741 kg (30,300 Ib), which includes 1315 kg (2900 Ib) of ballast for lateral balance, and adding 1632 kg (3600 Ib) of reserve fuel gives an end test weight of 15,373 kg (33,900 Ib). The test duration for 1406 kg (3100 Ib) of fuel is 0.65 hours.

The installation with the T56 engine results in a further decrease in the flight duration due to the increase of the zero fuel weight to 14104 kg (31,100 Ib) leaving 1038 kg (2290 lb) of fuel for the test mission. This amounts to 0.48 hours of test time.

Boeing 737-10 Testbed Configuration Performance Investigation of the Boeing 131-10 performance, also shown on Figure C-28, indicated that an un- modified aircraft at an altitude of 10668m (35.000 ft) would be capable of a speed of Mach 0.802 at a weight of 31,751 kg (70,000 Ib). The ramp weight corresponding to this start test weight is 33,409 kg (73,500 Ib). Modifying the aircraft to a testbed configuration with one prop-fan unit on the wing and the addition of pressure rakes would produce increases in drag that would re- duce the Mach number below that set by the testbed design requirements.

Boeing KC-135A Testbed Configuration Performance - The T56 powered KC-135A testbed configuration performance is shown on Figure C-28. At a ramp weight of 82990 kg (183,000 lb), the 3-engined ferry take-off distance is 2774m (9100 ft) over a 15m (50 ft) obstacle. Start test weight following a climb to 10668m (35,000 ft) is 75734 kg (167,000 Ib) and the end test weight is 48615 kg ( 107,200 1 b) • Test duration at these weights is 4.7 hours. The zero fuel weight for this configuration is 44542 kg (98,200 Ib) which includes the weight of the test equipment.

The configuration with the XT701 drive system has slightly reduced zero fuel weight, which together with the improved specific fuel consumption (SFC) for the XT701, would give slightly more test duration time than the T-56 con- figuration.

Boeing 8-52 Testbed Configuration Performance - The B-52 testbed perform- ance is shown in Figure C-28. Since performance data for the B-52B were not available, the testbed performance was generated using available B-52H data.

The B-52H, which is powered with P&W TF33 turbofans provides improved perform- ance over the J57 powered B-52B version.

At a ramp weight of 114452 kg (252,376 lb), the takeoff distance is 1033m (3390 ft) over the 15m (50 ft) obstacle. Climbing to test altitude reduces the weight to 111717 kg (246,476 lb) the start test weight, and at an end test weight of 81982 kg (180,776 lb) the test mission duration is 4.5 hours.

Because of the large speed margin above the test cruise Mach number, higher ramp and start test weights can be achieved with large increases in test duration. The ramp weight selected provided a test duration compatible with other candidate testbed aircraft.

Convair 990 Testbed Configuration Performance - The capability of the Con- vair 990 configured as a testbed aircraft, shown in Figure C-28, is for the T56 powered prop-fan. Ramp weight in this configuration is 81,012 kg (192,000 lb) and the start test weight following climb to test altitude is 82,537 kg (182,000 lb). At a zero fuel weight of 56,000 kg (123,500 lb), which accounts for the removal of the anti-shock bodies on the wing, the addition of test equipment and adding reserve fuel of 4761 kg (10,500 lb), gives an end test weight of 60,769 kg (134,000 lb). The test duration for 21,768 kg (48,000 lb) of fuel is 4.2 hours.

The installation of the XT701 engine decreases the zero fuel weight to 55,980 kg (123,441 lb) and the lower specific fuel consumption of the XT701 would increase the test duration over that of the T56-engined configuration.

GAC GIl Testbed Configuration Performance - The performance of the GIl as a prop-fan testbed powered by an XT701 engine over the test mission profile is given on Figure C-28. Beginning at a ramp weight of 21341 kg (60,288 lb) the start test weight at 10668m (35,000 ft) altitude is 26303 kg (58,000 lb). The zero fuel weight is 17722 kg <39,079 lb), including 680 kg (1500 lb) ballast for lateral balance, and the end test weight is 19484 kg (42,964 lb). Mission test time is 3.38 hours.

The lower powered DDA T56 engine will produce slightly less test time at al ti tude because the zero fuel weight is slightly greater than that of the XT701 configured testbed.

CANDIDATE TESTBED AIRCRAFT - INITIAL SCREENING AND SELECTION An initial screening was conducted using criteria such as mission perform- ance, clearances, scale mismatch, acoustic test sui tabil i ty, and commercial passenger transport configuration compatibility to establish testbed suitability. As the result of this screening, the Lockheed -6 JetStar and the Boeing 737 were eliminated from the list of candidates for the following reasons: o Lockheed -6 JetStar One testbed configuration only - that with the GE T64 engine provided an aircraft configuration with sufficient prop-fan tip/ground clearance, as shown in Figure C-29 , with the aircraft in a rolled attitude. Minimum clearance would also exist for the combined condition of two flat tires and a landing gear strut fully compressed.

The installation with the T56, Figure C-30, has a tip clearance in the normal ground attitude of 10.8 inches; in the rolled attitude, however, a tip/ground interference of 6.6 inches occurs.

The mission test time available, 0.65 hour, is unacceptable from a flight test standpoint, since very little data could be accumulated in such a short test time and the cost of acquiring such data would be high. In addition, this configuration is considered to present a moderate risk as far as wing flutter is concerned.

o Boeing 737-10 The Boeing 737-10, Figure C-31, was eliminated as a testbed candidate be- cause the unmodified aircraft performance at a weight of 31,751 kg (70,000 lb) and an al ti tude of 10668m <35,000 ft) has a Mach number capability of only 0.801.

When in the testbed configuration with the TIP/FUSELAGE CLEARANCE 1.78 m (5.85 FT) I /

1 -

W.L. 100 0.076 m (0.25 FT) CLEARANCE ROLLED ATTITUDE PROP PLANE F.S. 376.69 GROUND CLEARANCE 0.56 m (1.83 FT) Figure C-29. JetStar -6 T64 Testbed Clearances J ..- / ./ TIP/FUSELAGE CLEARANCE CLEARANCE 0.27 m

(0.[ ____ '~======~=(§)=~-.-J

I

0.16 m (0.55 FT) INTERFERENCE ROLLED ATTITUDE Figure C-30. JetStar -6 T56 Testbed Clearances PROP-FAN DIAMETER 2.46 m (8.1 FT) T56 ENGINE NO CLEARANCE BLOCKAGE IN PROP-FAN PLANE DUE TO ENGINE , , AND WING ; L B. L. 293.0 - e::::=~:.:::~-\- PROP-FAN PLANE

Ii

>zrnrrl .---'----::::::;0 ......... --

-""~t..~,.,.-- ',-;~-' -' -~-:--~J -

------------c' ~ .. -------:-.:-.----_.J';.....----

,,01 "., .. ~ TESTBED WITH DDA T56 ENGINE LEFT SIDE UNDER WING INSTALLATION Figure C-31. Boeing 737-10 Prop-Fan Testbed addition of trim and test equipment drag, the speed/altitude performance will fall short of the M=0.8/10,668m (35,000 ft) desired for the testbed aircraft.

The location of the prop-fan propulsion system on the wing, Figure C-31, is such that a moderate element risk would be incurred from the wing flutter standpoint.

As a vehicle for gathering acoustic data, the configuration is unsuitable because of the proximity of the basic aircraft jet engine and of the shielding effect of the engine nacelle and inboard portion of the wing as shown in Figure C-31. Ground clearance would be inadequate with the 2.89m (9.5 ft) prop-fan of the XT101 underwing installation, since tip/ground interference would occur in the rolled attitude as indicated on Figure C-32.

The following aircraft remain as candidate aircraft subsequent to the application of the inital screening criteria: o Lockheed C-141A o Boeing KC-135A o Convair 990 o Gulfstream American Corporation "Gulfstream II" The Boeing B-52B, although a purely military aircraft and therefore not representative of commercial transport aircraft, was also retained as a special class of testbed vehicle with limited potential as a prop-fan testbed vehicle.

Candidate Testbed Aircraft Analyses Testbed Aircraft Performance and Buffet Limits - The performance and buffet limits as a function of Mach number of each candidate testbed aircraft with XT101 engines are shown in Figure C-33. These data show the relationship of the weight/altitude curves at start and end test weights at the design conditions of Mach 0.8 at altitudes above 9144m (30,000 ft). The 19 buffet limit is superimposed on each plot.

ENGINE ct

2.89 m (W.L.184.0) (9.5 FT) 2.13 m (9.5 FT) B.L. 293.0 W.L. 195.0 PROP DIA = 2.89 ~ (9.5 FT) INTERFERENCE 0.26 m (0.87 FT) TIP/NACELLE CLEARANCE 0.48 m (1.6 FT) Figure C-32. Boeing 737-10 XT701 Underwing Configuration Interference o XT701 PROr-f'AN lNSTAlLAnON GIl CONVAIR 990 o PROrFAN QP!RAnNG 19 BUFFET LIMIT FT . 'gl /MX START caulU WT.

19 BUFFET lIMH~1'I

m GW • 6/J,769 Kg 034,000 U) 12 ·82,537 Kg (182,000 LB 38 V MAX WITH 55% PWR ON MAIN ENGINES + FULL PWR ON PROPFAN AlP '3l MAX START CRUISE 36 11 WEIGHT. 26,303 Kg (58,000 La) \ II END WEIGHT • 6/J,769 Kg (134,000 lB)1 I

J

I START TEST WEIGHT ·82,537 Kg (182,000 U

/

7~ __ ~ ____ =- __ ~ _____ ~ ___ -= __ ~

8 .76 .80 .84 .88 .92 .78 .80 .82 .84 .86 .8a .;>0 22 MACH NO.

MACH NO.

KC-135A .lg BUFFET UMiT C-141A GW • .a,615 Kg (107,200 U FT FT f 4(l 4(l 12 END TEST WEIGHT· .a,615 Kg (107,200 La) \ 66,4J7 kg (1~,5oo La) 38 J8 75,7l4 Kg (167,000 lil) PROP-f~ OFF 19 BUFFET UMiT 36 1\ 11 REaD TEST I w..._-, ~HI

~

I I 32 32 7 7 I o I I 30 30 -,p,,;.,.,""' ........... "".,.,.,,. MINIMUM TEST I AlnTUDE I I 28 28 I I I a I 26 26 I I 7~ __ ~ _____ ~ __ ~ _____ ~ __ ~~~ ~~78;---.~80~---.~82~---.~84~---.a~6~--~.88=---~.;>O· .78 .ao .82 .84 .86 .as .90 MACH NO.

MACH NO.

Figure C-33. Testbed Aircraft Performance and Buffet Limits These data are required for the Task IV Evaluation, Appendix D, to compare the design capability and to rank the aircraft/drive system combinations.

Data for the Lockheed C-141A are shown in Figure C-33 and indicate a small speed margin over Mach 0.8 and a buffet cut-off in the region of 10668-11277m (35,000 - 37,000 ft).

The Boeing KC-135A data, of Figure C-33, show that the speed margin at the start and end test weights and the altitude range are large and not constrained by the Ig buffet limit.

In the case of the Convair 990, Figure C-33, a wide speed margin is achievable over the range of test weights, although the altitude range is re- stricted to a maximum value in the region of 10668m (36,500 ft). The speed/altitude capability falls inside the Ig buffet limit.

Figure C-33 also presents similar data for the GAC GIl, which indicates a substantial speed and al ti tude margin over the desired conditions. The onset of buffet, however, is the limiting condition and causes a very slight re- duction in the maximum achievable Mach number.

Testbed Configuration Weight Summary - The weight summary for the six aircraft comprising the initial list of possible testbed aircraft is shown on Table C-II. These weight data were used to establish test mission profile data with various drive systems.

Preliminary Appraisal of Candidate Testbed Flutter Characteristics Preliminary appraisals were made of the candidate testbed aircraft relative to the risk of encountering wing flutter problems that might place the testbed program in jeopardy. The appraisals that follow were primarily based on the location and on the extent of the changes in the mass and inertial properties of the wing-engine system. In some cases, flutter parametric analysis results were also used. Since the appraisals were not based on specific flutter anal ysi s, they are qual i tati ve in nature and are intended only for use in the Task III screening to establish the suitability of candidate testbed aircraft.

C-141A Testbed Configuration Flutter Appraisal - No flutter problems are anticipated with the C-141A testbed configuration. The substitution of a prop-fan power plant in place of an existing inboard P&W TF33 power plant results in a weight reduction of approximately 1451 kg (3200 Ib), which is almost equi valent to a weight reduction for the existing powerplant of 43 percent.

TABLE C-II. TESTBED AIRCRAFT WEIGHT SUMMARY WITH VARIOUS DRIVE SYSTEMS XT70J XT701 T56 T56 T64 UNDER OVER OVER OVER UNDER WING WING WING WING WING 60,650 60,677 ZERO FUEL WT

*

(133,770) (133,711) FUEL 68,048 68,048

*

C-141A (150,020) (150,020) GROSS WT 128,725 128,698

*

(283,790) (283,731) ZERO FUEL WT 44,543 44,516 (98,200) (98,141) FUEL BOEING 51,202 51,516 KC-l35 (112,880) (112,880) GROSS WT 95,744 95,718 (211,080) (211,021) 55,992 ZERO FUEL WT 56,019 (123,441) (123,500) CONVAIR FUEL 47,301 47,301 (104,280) (104,280) GROSS WT 103,319 103,293 (227,721) (227,780) ZERO FUEL WT 17,763 17,726 (39,160) (39,079) FUEL 10,491 10,491 Gil (23,128) (23,128) GROSS WT. 27,346 27,309 (60,288) (60,207) ZERO FUEL WT 11,535 11,902 (25,430) (26,240) FUEL 5,942 5,942 JETSTAR (13,100) (13 ,100) GROSS WT 17,477 17,844 (38,530) (39,340) ZERO FUEL WT 30,346 30,193 30,167 30,319 (66,901) (66,565) (66,506) (66,842) BOEING 8,661 8,661 FUEL 8,661 8,661 (19,095) (19,095) (19,095) (19,095) GROSS WT 39,007 38,854 38,828 38,980 (85,997) (85,660) (85,937) (85,601) *UPPER ENTRY 13 IN kg, (LOWER ENTRY IS IN LB) Because of the inboard location, at 30 percent of the wing semi-span, this change is not expected to affect wing flutter speed adversely.

Convair 990 Testbed Configuration Flutter Appraisal - The substitution of prop-fan propulsion system in place of an existing inboard engine is approx- imately equivalent to a weight reduction of 907 kg (2,000 lb) or 33 percent over the weight of the original powerplant. It is considered that this change is not likely to alter the wing flutter characteristics unless the flutter speed is unusually sensititve to the weight of the inboard engines. The re- moval of the two adjacent anti-shock bodies is considered of little consequence from a flutter standpoint. This configuration is not likely, therefore, to encounter flutter problems.

GIl Testbed Configuration Flutter Appraisal - Addition of a prop-fan propulsion system weighing roughly twice as much as the wing semi-span will drastically alter the wing flutter characteristics.

Parametric stUdies of wing flutter of wings of similar planform indicate that the addition of a large concentrated mass located at 36.3 percent of the semispan may increase the flutter speed over that of the base wing.

Since these studies do not account for variations in wing fuel, flexibility of attachment structure, and other variables that may be important on the testbed aircraft, they can be used only as a preliminary indication that the prop-fan installation may not cause flutter problems. The proposed installa- tion of 680 kg (1500 lb) of ballast on the wingtip of the side opposite the prop-fan installation, to provide lateral balance, is not expected to cause flutter problems, since the weight is approximately equivalent to one of the 0.95m (250 gal) wingtip tanks with which this aircraft has been certified.

The risk of encountering a serious flutter problem with the testbed configuration is considered to be low, but flutter analyses will be required to verify this position.

KC-135A Testbed Configuration Flutter Appraisal - The replacement of an in- board nacelle with a prop-fan propulsion system results in a net weight change of 635 kg (1400 lb) or 26 percent of the weight of the existing P&W J57 power- plant. This weight change, located at 41 percent of the wing semi-span is not sufficient to change the flexible wing fundamental modes significantly and is, therefore, not expected to adversely affect the wing flutter characteristics.

Boeing 8-52B Testbed Configuration Flutter Appraisal It is considered very unlikely that this configuration will encounter flutter problems as a prop-fan testbed system. The inboard location of the prop-fan installation, together with the small inertia properties compared wi th those of the wing- -engine system, should not change the dynamic and flutter characteristics of the wing. Since the location has already been used to carry a variety of pylon-mounted stores and equipment, many of which had greater weight and in- ertia properties than the proposed prop-fan installation, the risk of flutter problems arising with this testbed installation is estimated to the lowest of any of the candidate testbed aircraft.

Stability and Control Analyses Estimates of the stability and control changes due to prop-fan application have been made for the four candidate testbed aircraft. The analyses show that there are no significant changes in stability when the XT701 drive system is installed on any of the testbed aircraft.

The stability changes analyzed were those considered to be of greatest importance, and consisted of the pitching and yawing moments caused by the installation of the prop-fan. The changes in the control derivatives C and m a C n{3 are caused by the prop-fan normal force, and an estimate of this force provided by Hamil ton Standard is shown on Figure C-34. The effects of the prop-fan on total yawing and pitching moment are shown in Figures C-35, C-36, C-37 and C-38 for the C-141A, KC-135A, Convair 990 and the GIl, respectively.

Yawing and pitching moment data for the C-141A were obtained from Lockheed- Georgia data files and estimates of the KC-135A, Convair 990 and GII were obtained by the use of DATCOM.

The changes in stability are also shown on Figures C-35, C-36, C-37 and C-38 for the C-141A, KC-135A, Convair 990, and GIl, respectively.

These data show the following trends due to prop-fan installation: 1) The change in the stability derivatives decreases as Mach number increases.

2) The effect of the prop-fan on the large airplanes, C-141A, KC-135A, and Convair 990, is small.

xr701 PROPELLER NORMAL FORCE COEFFICIENT SEA LEVEL MAX POWER PROP-FAN DIAMETER 2.89 m (9.5 FT) .5 ~ .4 ~ ~ u '""' z: :; .3 '" '" '" u u '" '" => .2 '" -l ~ C Z ..

'" -l .1 -l 0..

'"

~ '" '" .8 .4 .6 o .2 MACH NUMBER Figure C-34. XT701 Prop-Fan Normal Force PI'tCHING XOMEN'1' - WITHOUT """...fAN --- iNlTH 'ROP..fA,."

.J.O roTAL C 1.6 ".

'!OTAL C , -2.6 "13 I • ________ : lO"/DEC , , ----- , ,

J

v -2.2 ..

. l • 4 ..

l!ACIi ,nlMl\E1 -Z.5 lc .lCOI! -2.0 lIa 1.4 lO"'OFJ; -I.!

-1.0 I.J ., ..

.'

.'

MACH :ruMBER Figure C-35. C-141A Prop-Fan Effect on Pitch and Yaw PI!CRL'IG MOIIE>T - Wl'nllA.ff tttOnAil --W[TII PIOPt'AIiI J • .l TOTAL I.U C °fJ J 10 /DEG -1 ..

I.'

-1.4 1.5 -I.'I:-J-_-.~J---'!--~;::;;"""""\ .J ..

Figure C-36. KC-l35A Prop-Fan Effect on Pitch and Yaw - IIITOOUT PROP-FAll - -lltTn PROP-FAll PIrCHL'IG l!OKElIT I // .'/ ~/ --.-. ~:;;.::/ ':'OTAL 1.'.

TOTAL S!"

COli !."

J (10 /OEGl 1.') \ \ \ 1.6

\

\ <1c !,' 'fJ "", (lO':OEG)

'''-....----

Figure C-37. Convair 990 Prop-Fan Effect on Pitch and Yaw P I!CHL~G >1OIIEm" YAIIISG >!O!!ElIT _....,THOUT Plor..fAN _WITHOUT 'RO,..,AN ---WITH pttOP";AN --- wITH 'ROfI..fAN 1 a I 2.2

--------::-.\1. C -. I

TOT.\I. e "4 I nfJ , , , 3 102/DEG , to IDEe _J.t 2.1 ., "

------ .-

.,

----

.-

----

,.- " .-~

-

-3.0 2.0 .2 .4 ..

.4 • 0 ..

~ACH ~:UMB ER >iACH '''lMllER 7.0 -I.J <lC ole 0 .. (\ ..

nfJ -1.1 lO .. /Oi....

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4.l ·.i • 4 . .

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--.A.CH ~BEJ.

'lACK _BD.

Figure C-38. Gulfstream II Prop-Fan Effect on Pitch and Yaw 3) The GIl exhibits the greatest change in stability due to the prop-fan.

At low Mach number C changes by seven percent. Although not significant. some minOrn~egradation of the flying qualities may occur.

Testbed Aircraft Suitability for Acoustic Test Near-Field Acoustic Analyses - The four testbed configurations were re- viewed for sui tabili ty as acoustic test and data gathering vehicles by con- sidering the common features as far as near-field acoustics are concerned, as described below.

It was determined that all of the configurations have: o Sufficient fuselage volume o Cabin pressurization o Representative fuselage structural configurations o Potential for modification for test of acoustic suppressive concepts o Space for acoustic.test equipment In addition, each configuration has a number of advantages and disadvan- tages, as follows: Disadvantages Advantages Testbed Large power plant on High position of prop-fan rel- Lockheed C-141A same side as prop-fan, ative to the fuselage center line.

(L300) with lower noise frequencies from large discharge which may interfere with prop- Large ratio of fuselage diame- fan noise measurements ter to prop-fan diameter Large separation of prop-fan, fuselage and wing leading edge Separation of adjacent nacelle - different water line location Fuselage separation/ Prop-fan height relative Boeing KC-135A prop-fan diameter to fuselage centerline com- (101-100) ratio too large patible Portion of the in- Fuselage diameter/prop-fan board wing obstructs diameter ratio indicates a prop-fan noise pass- good match age to fuselage Wing obstructs noise Prop-fan height relative to Convair 990 passage to fuselage fuselage centerline compatible Fuselage separation/ Convair 990 (cont'd) prop-fan diameter ratio larger than would occur in actual design Fuselage is small re- Prop-fan height relative GIl lative to the prop-fan to fuselage centerline diameter compatible Wing blockage is well aft of the prop-fan plane Far-Field Noise Analysis - The prop-fan testbed aircraft will have noise sources other than the prop-fan as follows: o Prop-fan Drive System o Testbed Primary Engine o Testbed Airframe Noise These additional sources generate the background noise for which an acoustic analysis was performed to identify all of the noise characteristics.

The airplane reference conditions chosen for the analysis were: Level Flyover, Altitude 308m (1000 ft) Speed 72m/sTAS (140 KTAS) ISA + 10 C Conditions 70% Relative Humidity The acoustic comparisons for all sources were made for the following condi- tions: o For the aircraft directly over the microphone (which is very close to the peak noise from the prop-fan).

o With atmospheric attenuation effects included.

o For a microphone mounted with its diaphragm close to the ground (the noise levels thus include a 6dB increase over free-field noise levels).

o In one-third octave band levels.

Noise Source Characteristics Prop-Fan Alone - The noise levels were predicted for a single prop-fan with the following characteristics: Prop-fan Diameter 2.89m (9.5 ft) Number of Blades V , Rotational Tip Speed 244m/s (800 ft/sec) T Power 1491/2982/4474/5965 kW (2000/4000/6000/8000 shp) The predicted one-third octave band spectra are shown in Figure C-39.

Prop-Fan Drive System - The drive system data are based on test cell measurements of the DDA XT701. The noise data are shown on Figure C-40 for power levels of 1491/2982/4474/5965 kW (2000/4000/6000/8000 shp).

Aircraft Flyover Noise - Measured flyover noise data from the Lockheed C-141A, Boeing KC-135A, Convair 990, and the GAC GII with hardwall nacelle and "Hush" kit, are shown on Figures C-41, C-42, C-43, C-44 and C-45 for the i n- dicated power settings.

The primary engine noise dominates the spectra for these aircraft.

Testbed Airframe - The predicted "clean" airframe noise levels with the gear and flaps up for the Lockheed C-141A, Boeing KC135A, Convair 990, and the GAC GII are shown in Figures C-46 through C-49.

SINGLE PROP-FAN, DIRECTLY OVERHEAD ~ "'d I ~ ~ kW (SHP) H (8000) p z (6000)

;a

(4000) ~ (2000) ~ E-l U § H :::t: E-l 60 ~ Z 5 6 7 R 9 1 2 6 7 8 9 1 3 4 6 7 8 91 2 4 5 4 5 2 3 10000 20000 FREQUENCY - HERTZ Figure C-39. Prop-Fan Alone Noise XT701 DRIVER ~ 120 "'d I ~ ::> kW (SHP) ;j 100 5965 (8000) (6000)

~

~ (4000) ~ 80 (2000) ~ E-l U o § 60 H :::t: E-l ~ Z o 40 2 I, <;67891 4567891 3456 20 100 1000 20000 FREQUENCY - HERTZ Figure C-4O. XT701 ProP-Fan Drive System Noise ~ "C I !il 3 ENGINES OPERATING &i ...l Sl ~R ---------------.~ < "...- I =:l T/Ol. • _---------_" I>l CRUl.S2/ .""",- ~ f-< " U

0 " APP 1.2""

CRUISE APPROACH iil ....

:= f-< I>l Z 4.567891 45t.7891 4567891" 100 1000 10000 20000 FREQUENCY - HERTZ Figure C-41. Lockheed C-141A Flyover Noise 12C 3 ENGINES OPERATING W 100 &.i ...l ~".:.~, /' APP 1. 75 ell! 1 • .50 I>l Z o 3456781 J • :'~78~1 1000 100()(1 FREQUENCY - HERTZ Figure C-42. KC-l35A Flyover Noise ~ "C I 3 ENGINES OPERATING ...l I>l 100 &i ...l ~

ff~

~

------ -. '\ '\ ___ •• _.. -_. J \\ _ 1;0

~ I>l --- --·. __ .. 7,J 'X""~C/8 > -- ,\APP 1 < U

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Q <>: ....

::: f-< I>l Z 6 1 891 3 , 5 tt 7 8 91 J 4567891

to 100

10000 FREQUENCY - HERTZ Figure C-43. Convair 990 Flyover Noise 2 ENGINES OPERATING '" "Y ....

..

~ ,.J 100 THRUST SETTING, LBS.

"" ~ NET THRlIST ..

~ ..

<J o 80 ... 9771/172 "" ....

.c 'i' ..

o

"

-----

2644/281

~-- ---

2883/2979

/' --

/ ~~

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~ight Idle 300 Lbs.

2 34567891 34567891 2 4567891 20000 10000 FREQUENCY HERTZ Figure C-M. GAC Gulfstream II Flyover Noise - Hardwall Nacelle DO ."

I ....

..

> ~ 2 ENGINES OPERATING " ~ 100 ..

~ ..

'J o ~ 80 ....

.c f-< I ..

9673/124 o " 2870/2430 5524/326 345678'11 2 345h78Ql 2 4 5 b 7891 20000 10000 20 lUO FREQUENCY HERTZ Figure C-45. GAC Gulfstream II Flyover Noise - Hush Kit Nacelle

I

~ 100 ~

-

....

Z ~ 80 '"-l > < :.

U o 60 § ....

-

-

=-- !::l Z o .. j&7d91 2 " 5&i391 2 J 45&;~9l :0 LOOOO 20000 LOa LOOO FREQUENCY - HERTZ Figure C-46. Lockheed C-141A Airframe Noise L20 ::Q ~ I ~ ~ ....

~ z < ::Q ~ ~ =-- :.J Q ::I: ....

-

-

=-- ~ z ;: J 4'6]81l Z J 36.43l .3611491 100 1000 10000 20000 FREQUENCY - HERTZ Figure C-47. Boeing KC-l35A Airframe Noise ...

;a

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-

- :...

~ :z: 40 o J ~ ~&7a9L 2 J 4 Sa7a91 L S~;d9l 20000 100 LOOO LOOoo FREQUENCY - HERTZ Figure C-48. Convair 990 Airframe Noise UOI.-----------------------------------------------------------------~ ~ :z: 40 C ~ 4 ~o7S9l J 4 ~61a9l 4 S6739!

zoooo LOO LOOO 10000 FREQUENCY - HERTZ Figure C-49. GAC Gulfstream \I Airframe Noise Prop-Fan Drive System Muffler - The data of Figure C-50 shows that the prop-fan noise is subject to considerable masking from the noise radiated by the XT701 drive system, principally from the exhaust.

Achievement of a cleaner noise signal from the prop-fan requires reduction of the drive system exhaust noise. This could be done by either locating the drive system exhaust over the wing well upstream of the trailing edge to take advantage of wing shielding or by adding a larger muffler to the exhaust.

Figure C-50 also shows the drive system noise with 15db suppression throughout the spectrum.

uo == "0 I Prop-Fan 5965 kW • 100 ~ (8,000 shp) 5965 kW ~ Driver ~ (8,000 shp) 5965 kW Q Driver With (8,000 shp) ~ 80 15 dB Exhaust ;:Q Suppression ~

~

c.J 60 Q .

It )6789L l 1 10 S678~l l 1 2 :0 100 looo 200000 FREQUENCY - HERTZ Figure C-50. Prop-Fan Driver Suppression Requirement The need for this amount of suppression is applicable to all the candidate aircraft and is shown on the aircraft component noise spectra.

Aircraft Component Noise Spectra - Component noise spectra are shown on Figures C-51. C-52. C-53. and C-54 for the Lockheed C-141A. Boeing KC-135A.

Convair 990 and the GAC GIL These spectra are predicted for peak flyover noise with drive system suppression and with the noise generated by the primary engines at flight idle power. Al though the latter predictions are based on flyover noise measurements. some degree of uncertainty does exist as to their actual value. The flight test for an Acoustic Test Program. however. could be planned to better define these noise levels. In the case of the GAC GIl. a "Hush kit" is available which could further reduce the noise level of the "Spey" engines. The data show that the cleanest prop-fan noise signal is that from the GAC GIl testbed.

EST. Porent Engine Noise A' F @ Fit Idle .r rome ___ ~___ • )1 --.

.a Nois. .-

• 60

(Cleon) ".....' V

.......... """

-- "'"

!

<J Propfon /' /"

............ -.................. "

o Noise ~

.............................. '\

~ (8000 SHp) ./";/" :.c ................. .......... .......... _, ~~ Suppressed ....

:. 40 ........ ~ D~~

o

................ \. \ (8000 SHP) 4Sb7'31 2 J 4567891 34567601 2 2 2 10000 20000 100 FREQUENCY I HERTZ 1000 Figure C-51. C-141A Component Noise Levels ..,

"

I eo EST. Porent ..

Engine Noise > ..

..... @ Fit Idle ';!

Cleon ............... • j Air Frome -......... -.--::: - _ ~ - - .......

~ 60 Noise -. /" I ~ - - .......

.2

,.. -- ""

'J Propfon / ..... /'"' ................. .......

o Noise...,. ................... ........

] ~ (BooO SHP) ./'......... ........ ....... .............. ............_, ;: Suppressed I Driver ~ 40 ................... , (BOoo SHP) .....

34567891 2 J 4567891 J 4567B91 2 20000 10000 FREQUENCY, HERTZ Figure C-52. KC-l35A Component Noise Levels II ~ -' u Air Frome c o .0 Noise \. ----- ..

(Cleon) '\.~--- --oL"""-::: ..........

.e 60

;::7 V'-'-- u Prop-fon ~./ ---- ' ..........

o Noise " ............. -........... EST Porent 1?

(8000 SHP) ........ -...,. ..................... _...... ................. Engine Noise' ...c l- ./' ~ ..... ~ @ Fit Idle I ............... -"\. ~ ./ Suppressed " <5 40 ...... y- Driver .... , ( "', ,8000 SHP) 2 3 4567891 2 4567891 2 3 4567891 2 1000 20000 10000 FREQUENCY - HERTZ Figure C-53. Convair 990 Component Noise Levels .., "0 Qj ~ Prop-fon -' "0 Noise C (BOoo SHP) ~ _ .../ Suppressed II --~ Driver > ,/ ....... (BooO SHP) .2 Air Frome~ ___ u --===----~ Noise ..".,.",,--- / -- ---.

1?

(Cleon)'" ........ ___ -- -. -~ .........

~ ...

--- ... _---- ',,-

I

-:/'"

., .......... ~ c EST Po rent ... ~~

;/

Engine Noise .... ,.... , @Flt Idle ......... ...............

34567891 23 4567B91 7 891 2 2 3 4 5 6 20000 1000 10000 FREQUENCY - HERTZ Figure C-54. G II Component Noise Levels Testbed Aircraft Alternate Configurations The single engine testbed aircraft were examined for possible conversion ::-0 multi-prop-fan configurations. It was found that the large aircraft - C-141A, KC-135A, and the Convair 990, which in the single prop-fan configuration were propulsion substitutions - would require a change to propulsion addition to achieve a multi-prop-fan testbed. In the case of these aircraft, the prop-fan propulsion units would be located on the wings inboard of the existing inboard primary propulsion. As far as possible, the units would be located to provide the desired clearances for structural and acoustic considerations. The three configurations are shown on Figures C-55, C-56, and C-57.

The multi-prop-fan GAC GIl is achieved by adding a second wing-mounted prop-fan drive system as shown in Figure C-58.

B.l.. ~E3J~:=;:~\- 219.679 F.S.

597.659 U70l CRIn SYSTE!!

=

;i~:99S!

~-.~~---~~

Figure C-55. C-141A Twin Engine Figure C-56. KC-l35A Twin Engine Testbed Overwing Configuration Testbed Overwing Configuration ---I ---- - -----~~=-==~ B.L.

266.00 F.S.

695.00 XT701 DRIVE SYST~ W.L.

73.00 ---, I Figure C-57. Convair 990 Twin Engine Testbed Overwing Configuration _<T101 ORrlE ,YSTDf :l.L.

73.;,2 I Figure C-58. Gulfstream II Twin Engine Testbed Overwing Configuration

APPENDIX D - TESTBED SYSTEM EVALUATION AND RECOMMENDATIONS - TASK IV A list of evaluation factors developed in accordance with the NASA State- ment of Work and approved by NASA was used to perform the Advanced Turboprop Testbed System Evaluation from which recommendations were made to the NASA Lewis Research Center regarding equipment requirements for the Testbed Program and for the Program Plan. To simplify the process, the evaluation was divided into a "Drive System Evaluation and Selection" based on Task II results, and an "Air- craft Evaluation and Selection" based on Task III results. This enabled the selection of the Drive System to be made before proceeding with the Aircraft Evaluation, thereby eliminating the Drive System as a variable in the Aircraft Evaluation process. The doubt surrounding the availability of the aircraft considered in the study was also sufficient cause to remove aircraft avail- ability from the evaluation. This came about when a survey of the list of NASA aircraft revealed that none of the aircraft suitable for Testbed application would be available in the near or far terms for the Testbed Program. Aircraft availability was, therefore, made a separate consideration addressed following to the evaluation. This survey indicated that acquisition of an airframe for the testbed aircraft may be possible only by purchasing a suitable vehicle.

This is particularly true if the prop-fan testbed program is to be accelerated.

CANDIDATE DRIVE SYSTEM EVALUATION AND SELECTION Five propeller drive systems were investigated in Task II and three engine/gearbox combinations emerged as candidates for the testbed aircraft drive system as listed below: Power Section Gearbox DDA T56 T56-A-14 DDA XT701 T56-A-14 GE T64-415 IHI T64-2 SDG Of the three drive systems two, the T56/T56-A-14 and the GE T64-4151IHI T64-2 SDG, are in production, whereas the third, the DDA XT701, exists in suffi- cient quantity to support a testbed aircraft program.

Drive System Evaluation Evaluation criteria for the drive system were grouped into the following categories: Operational Characteristics o Shaft horsepower at design condition o Fixed speed or free turbine Prop-Fan Sizing o Disc loading o Structural validation constraints Drive System Modification o Gearbox modification o Power section modification o Normalized cost o Risk Engine and Gearbox Availability o Power section availability o Gearbox availability o Spares availability Prop-Fan Control System Requirements o Modified 54H60 control compatibility o Control functions required Nacelle Structure o Overdesigned structure o New contours Engine Controls o Fuel control These items are listed on Table D-I, where the relative merits are addressed.

Hamil ton Standard has determined that an accurate demonstration of dynamic behavior and fabrication feasibility cannot be achieved with prop-fan diameters of less than 2.44m (8 ft). Since the prop-fan diameter for the General Electric GE T64-415 was only 2.16m (7.1 ft), this drive system was eliminated from consideration. The selection of the drive system for the testbed aircraft, therefore, became a choice between the DDA T56 and the DDA XT701.

Comparing the two drive systems, it is readily apparent that the XT701 pro- vides the largest diameter prop-fan 2.89m (9.5 ft) with a possibility of in- creasing to 3.05m (10 ft) when higher power levels on the XT701 have been demon- strated. This is about 17 percent greater in diameter than the nearest rival, the T56-sized prop-fan 2.47m (8.1 ft) in diameter. The gearbox power limitation at sea level for the XT701/T56 3729 kW (5000 shp) and the T64 2237 kW (3000 shp) will affect ground operations.

The XT701, which has a free turbine power section, has another advantage over the T56, a fixed-speed unit, in that the prop-fan tip speeds can be varied continuously over a wide range. This speed range provides test condition flex- ibility of great value in a flight test program. In addition, the fixed-speed T56 requires a negative torque-sensing system, which is one more control func- tion than is required by the XT701.

Drive system modifications of significance are those required to match the T56-A-14 gearbox to the drive system test requirements. The modification to the gearbox for speed compatibility with the XT701 requires only one set of new gears, whereas the T56 requires three sets of new gears, one for each tip speed.

Because the XT701 rotates counterclockwise and the T56 gearbox is designed for clockwise rotation, additional gearing modifications are also required to rotate TABLE D-I. DRIVE SYSTEM EVALUATION DOA XT701 DOA T56 GE Tl4 GEARIOX TYPE T6~-Z SDG MODIFIED MODIFIED T56-A-14 MODIFIED T56-A-14 (4310) 3423 (~591) KII ISHP) S.L.S. 6016 (8071) 32" Klt (SliP) OPERATIONAL ( 1610) (ZUO) 1350 1819 Z5Z0 (3310) M-O.I 10.7K(35K) ALT.

CHARACTER ..

FIXED SPEED DR ISTICS FREE TURBINE FIXED SPEED FREE TURBINE FREE TURBIME TIP SPEED YES NO YES CONTINUOUSLY VARIABLE DISK LOADING 301 Klt/M~ 2.13 (6.97) 2.47 (e.l) 2.89 (9.5) (37.5 SHP/FT lOlA M(F11 SIZING SIZE FOR STRUCTURAL UNSATISFACTORY MARGINAL SATISFACTORY VALIDATION SINGLE GEAR SET THREE GEAR SETS SINGLE GEAR GEARBOX SET DRIVE TORCUEMETER EXISTING EX:STING NEW STSTEM INTAKE CASE NOT RECUIRED NOT RECUIRED RECUIRED INTERCON. STRUTS MODS NORMALIZED COSTS 1.0 (2 GBOXES) 1.0 (3 GBOXES) <1.0 (2 GBOXES) R!SK 1.0 1.0 >1.0 POIIER SECTION 5 XT701 5 DEVELOP- IN PRODUCTION IN PRODUCTION AVAILABILITY "'ENT UNITS ENGINE ~ GEARBOX IN PRODUCTION GEARSOX AVAILABILITY IN PRODUCTION IN PRODUCTION AVAIL.

LIMITED COMMERCIAL SPARES AVAILABILIT~ IN PRODUCTION IN PRODUCTION SEi A V AILABLE MODIFIED S4H60 CONTROL NOT COMPATIBLE COMPATIBLE COMPATIBLE PROP-FAN OVERSPEED·P~TECTION RECUIRED RECUIRED RECUIRED CONTROL NOT REQUIRED NTS RECUIRED HOT REQUIRED SYSTEM RECUIRED BOvERNING REQUIRED REQUIRED REOUIRED FEATHERING RECUIRED (SLOW) RECUIRED (SLOII) REvERSING FIXED BLADE FIXED BLADE F'!XEO BLADE STRUCTURE OVERDESIGNEO RECUIRED RECUIRED REQUIRED NACELLE CONTOURS NEil CONTOURS NEW CONTOURS NEw CONTOURS .

ENGINE FUEL REQUIRED RECUIRED RECUIRED CONTROL the accessory drives in the proper direction. The testbed program with a single prop-fan configuration could be operated with two modified gearboxes for the XT701 drive system, but utilizing a T56 drive system would require at least three gearboxes to minimize "down time" interference with the testbed program when changing prop-fan tip speeds. However, there is a slightly higher risk associated with the XT701/T56-A-14 gearbox because of the high power level of the XT701, which could place restrictions on operating at high power conditions (low altitude).

No problems are associated with availability of the T56 for the testbed program, since the engine is in production. In the case of the XT701, five engines exist with another five at various stages of development. In addition, an industrial engine, the Model 570, has a large degree of commonality with the XT701, the principal difference is in the compressor case material which is titaniun for the flight weight XT701 and steel for the Model 570. Reliability and availabill ty of spare parts are not expected to present problems for the Furthermore, it is considered that the number of Preliminary Flight XT701.

Rating Test and developmental engines is sufficient to support the testbed program.

The Drive System Selection is summarized in Table D-I1. Of the ten items listed, the XT701/T56-A-14 combination is superior to the T56/T56-A-14 in 5, of equal standing in 2 and is not as good as the T56/T56-A-14 in 3 items.

TABLE 0-11. DRIVE SYSTEM SELECTION 'l--f"'I~ ,\ i llJ.

I

•

•

•

•

•

•

•

• •

•

•

•

The XT701/T56-A-14 Drive System, based on this analysis, is the selected Drive System for the Advanced Turboprop Testbed Aircraft because it: (a) provides the largest diameter prop-fan wi thin the constraints of the available power level, (b) has the flexibility to continuously vary prop-fan speed for test purposes, (c) reduces the number of gearboxes required for this program and eliminates the reliability risk associated with gearbox dismantling and reassembly to change gear sets, and (d) requires less control functions to operate the drive system than the T56.

Because of the uncertainty surrounding the availability of an airframe for the testbed program, the drive system will be designed as a uni ver sal QEC uni t with structural margins high enough to permit installation on any of the candidate testbed aircraft. Over-design of the nacelle structure does not involve a weight increment of great significance.

CANDIDATE AIRCRAFT EVALUATION AND SELECTION Aircraft selected for consideration as Advanced Turboprop Testbeds in Task III were confined to those known to be in the NASA inventory or available to NASA through loan arrangements with the Mil i tary services. The candidate aircraft evaluated were: o Lockheed C-141A o Boeing KC-135A o Convair 990 o Gulfstream American Corporation "Gulfstream II" o Boeing B-52B These candidate aircraft fall into three types for which two classes of propulsion system application are possible and for which two variations of prop-fan installation can be identified.

Candidate Testbed Aircraft Categories The candidate testbed aircr8ft fall into three categories as follows: o Commercial passenger transports o Military transports representative of commercial aircraft designed for FAA certification o Military aircraft non-representative of commercial aircraft but having limited potential for advanced turboprop testbed application by virtue of previous usage as a test vehicle Candidate Testbed Aircraft Propulsion System Configurations The propulsion system configurations of the candidate testbed aircraft were divided into two classes: o Prop-fan propulsion system substitution: This class of propulsion system configuration was characterized by the removal of an existing propulsive unit and the substitution of a prop-fan propulsion system.

o Prop-fan Propulsion System Addition: The existing propulsion system was retained for this propulsion configuration and the prop-fan system was added to the aircraft configuration Candidate Testbed Aircraft Prop-Fan Installation Variants Two variations of prop-fan propulsion unit installations were identified as follows: oPinion-high overwing installation oPinion-low underwing installation Evaluation Criteria The evaluation criteria categorized according to function are as follows: A. Aircraft Safety Requirements o Ground Operational Safety o Flight Operational Safety o Aircraft Structural Integrity B. Operational Characteristics Requirements o Compliance with Design Conditions o Test Mission Duration o Aircraft Stability and Control o Installation Effects

c. Testbed Program Objectives Achievement

o Realistic Environment for Dynamic Loads Validation o Acoustic Data Acquisition o Prop-fan Scale o Installed Propulsive Efficiency and Interaction Effects D. Data Availability o Contractor Access to Aircraft Data E. Potential for Modification to Research Aircraft Configuration o Performance with EXisting and Projected Drive Systems F. Relative Costs of Testbed Systems o Comparison of Testbed Systems ROM Costs Evaluation Criteria Ratings and Procedures Since it is unlikely that anyone of the selected testbed aircraft will have all of the features desired for the testbed aircraft, a number of evaluation criteria ratings have been identified to assist in the selection process. Each evaluation criterion is rated on a scale of 0 to 3 for acceptability, but because of the diversity of the evaluation criteria and their equally diverse degrees of importance, each rating is "weighted" on a scale of 1 to 4 according to the level of priority or importance of the criterion under evaluation.

The ratings used in the evaluation are as follows: Acceptability Rating Unacceptable Marginal Satisfactory Good The weighting factors applied to each of the evaluation criteria listed on Table D-III cover a scale of 1 to 4, with the higher levels of weighting factor indicating higher levels of criterion priority.

A total score is produced for each candidate testbed aircraft by the sum- mation of the products of the Evaluation Criterion Rating (ECR) and the Weighting Factor (WF) as follows:

Total Score = ~ECR x WF

The candidate testbed aircraft are then ranked according to the weighted score for which the higher scores indicate those aircraft sui table for the Ad- vanced Turboprop Testbed System Application.

Testbed Aircraft Evaluation The evaluation process was conducted by dividing the procedure into a number of components and sUbcomponents: o A statement identifying the major concerns or conditions to be satisfied was first formulated.

o This was followed by the identification of specific evaluation criteria and a description of each item evaluated.

TABLE 0-111. EVALUATION CRITERIA IDENTIFICATION EVALUATION CRITERIA AND WEIGHTING FACTORS WEIGHTING FACTOR EVALUATION CRITERIA AIRCRAFT SAFETY A

-

PROP-FAN LOCATION A-l ENGINE-OUT SAFETY A-2 STRUCTURAL INTEGRITY A-3 OPERATIONAL CHARACTERISTICS B

-

DESIGN CRUISE CONDITIONS COMPLIANCE B-1 TEST MISSION DURATION B-2 AIRCRAFT STABILITY AND CONTROL B-3 INSTALLATION EFFECTS B-4 TESTBED PROGRAM OBJECTIVES ACHIEVEMENT C

-

DYNAMIC LOADS VALIDATION C-1 NEAR-FIELD NOISE DATA ACQUISITION C-2 FAR-FIELD NOISE DATA ACQUISITION C-3 PROP-FAN SCALE C-4 INSTALLED PROPULSIVE EFFICIENCY C-5 VALlDATlON INTERACTION EFFECTS VALIDATION C-6 DATA AVAILABILITY D

-

AIRCRAFT DATA AVAILABILITY D-1 POTENTIAL FOR MODIFICATION TO RESEARCH E

-

AIRCRAFT CONFIGURATION POTENTIAL FOR MODIFICATION TO A E-1 RESEARCH AI RCRAFT RELATIVE COST OF TESTBED SYSTEMS F

-

MODIFICATION COST DATA RANKING F-1 o The evaluation rating for each item was then developed and the weighting factor applied.

o The weighted evaluation rating for the testbed evaluation was then determined. Averaging was used when more than one item was involved in the process.

Each of the Evaluation Criteria (EC), identified alphanumerically, is shown on Table D-III, together with the appropriate weighting factors.

AIRCRAFT SAFETY REQUIREMENTS The aircraft must be capable of operation on the ground and in the air wi thout damage to the prop-fan, the installation and the aircraft and without danger to the crew. Requirements include ground operational safety, flight operational safety and structural integrity.

Ground Operational Safety EC A-1 Prop-fan Location The prop-fan location must be such that: Sufficient ground clearance will exist to permit operation of the prop-fan installation without damage under normal operating conditions.

Sufficient ground clearance will exist following the deflation of a tire or tires in combination with full contraction of a landing gear strut.

Sufficient clearance will exist between the prop-fan and adjacent components to permit operation of the prop-fan without damage and interference.

The criteria for clearances recommended by Hamilton Standard are: 1.8m (6 ft) Prop-fan Tip/Ground-Normal Attitude-H Prop-fan Tip/Fuselage-F 0.8D For Acoustics p 0.2D For Excitation p where Dp is the prop-fan diameter Additional prop-fan clearance nomenclature is identified in Figure D-1.

;; ,v:;z;:;. 177.7;»7)/77»)),..//777/7

i I , i H3 Hl ~ - GROUND CLEAR}u~CE NO&~ ATTITUDE HZ - GROU1~ CLEARfu~CE ROLLED ATTITUDE H3 - GRamm CLEARAl.'lCE COMPRESSED STRUT Mm FLAT TIRE F - ~USELAGE/pROP-FAN TIP CL&~Rfu'lCE Figure 0-1. Prop-Fan Location and Clearance Definition The data for the evaluation criteria development are given in Tables D-IV.

D-V, D-VI and D-VII for the Lockheed C-141A, Boeing KC-135A, Convair 990, and the Gulf~tream American GIl over- and under-wing configurations, respectively.

Flight Operational Safety Engine-out Safety - The testbed aircraft must be capable of safe operation following an engine failure.

TABLE D-IV. EC A-l PROP-FAN LOCATION EVALUATION - C-141A WEIGHTI NG FACTOR = 2 C-141 TESTBED AIRCRAFT OVERWING CRI- PROPUL HEIGHT/OIST RATlNG

H = f (0 )

CRITERIA TERION LIMIT m (in) VALUE SYS ... P SCORE NA T64 NA NA NA GROUND CLEARANCE ~.94 (194.65) 2.00 NORMAL T56 6 Hl XT701 5.06 (199.15] 1.750 Q 6 T64, NA NA NA NA GROUND CLEARANCE OUTBO. ENGINE ROL1.EO A ITITUDE 3 .00 (118.0) 1.21 Do 6 T56 3 FI RST CONTACT H2 XT701 .960 2.79 (110.0) 6 p NA NA GROUND CLEARANCE NA T64 NA DEFLA TED TI RES & '6 i4 .79 (188.65] 1.940 3 T56 p CONTRACTED STRUT H3 XT701 1.690 6 ~.91 (193.15 3 NA T64 NA NA NA PRCP-FAN/FUSELAGE CLEARANCE 4.06 (160.0) 1.640 6 T56 p l- F XT701 3.91 (154.0) 1.350 6 p C-141A TESTBED AIRCRAFT UNOERWING CRI- PROPUL HEIGHT/DISl RATING CRITERIA H = f (D ) TERION LIMIT VALUE m (ins) SYS ?

SCORE NA NA NA NA T64 GROUND CLEARANCE 1.200 NORMAL 2.97 {116.85 T56 3 p Hl XT701 3.71 (146.05 1.280" 6 T64 NA NA NA NA GROUND CLEARANCE OUTSO ENGINE ROL1.ED A ITITUDE 2.44 (96.0) T56 .980 6 FIRST CONTACT H2 XT701 1.98 (78.0) .680 p GROUND CLEARANCE NA T64 NA NA NA DEFLA.iED TIRES & 1.090 2.71 (106.85 3 CONTRACTED T56 6 STRUT H3 XT701 3.56 (140.0) 1.230 3 6 NA NA NA NA T64 ?ROP-FAN/FUSELAGE CLEARANCE 3.94 (155.0) 1.590 3 6 T56 p F XT701 3.61 (142.0) 1.240 3 6 p TABLE D-V. EC A-l PROP-FAN LOCATION EVALUATION - KC-l35A

WEIGHTING FACTOR = 2

KC-135A TESTBED AIRCRAFT OVERWING CRI- PROPUL HEIGHT/DIST RATlNG

CRITERIA H = f (0 ) TERION

LIMIT VALUE SYS m (in) P SCORE NA NA T64 NA NA GROUND CLEARANCE NORMAL 2.14 (84.13) 0.87 Dp T56 3 HI XT701 1.92 (75.73) 6 0.66 Dn 3 NA NA T64 NA NA GRCUND CLEARANCE OUTSO. ENGINE ROLLED A ITITUDE 1 .45 (57.00) 6 T56 0.59 De 3 FIRST CONTACT H2 1.33 (52.41) 6 XT701 3 0.46 Dn GROUND CLEARANCE T64 NA NA NA NA OUTSO. ENGI NE DEF~ TED Tl RES & 1.07 (42.0) 0.43D T56 3 FIRST CONTACT CONTRACTED STRUT o H3 XT701 1).96 (37.6) 0.330 6 T64 NA NA NA NA PROP-fAN/FUSELAGE CLEARANCE

T56 5 .07 (199.46 2.06D · 6

e

:0--' F XT701 6 ~4. 83 (190.00) 1.660 KC-135A TESTBED AIRCRAFT UNOERWING CRI- HEIGHT/DIS, PROPUL RATlNG

CRITERIA H = f (0 )

TERION LIMIT VALUE SYS m (ins) P SCORE T64 NA NA NA NA GROUND CLEARANCE 1 .52 (59.82) NORMAL T56 0.61 De 6 HI 1 .26 (49.42) XT701 0.430 T64 NA NA NA NA GROUND CLEARANCE OUTSo ENGINE KOlLED A ITITUDE 0.78 (30.59) 310 T56 3 6 0. p' FIRST CONTACT H2 0.56 (22.19) 0.190 XT701 3 GROUND CLEARANC1: NA T64 NA NA NA OUTSo. ENGINE DEFLA. TED TI RES & 1.27 (50.00) 0.510 3 FIRST CONTACT CONTRACTED T56 6 STRUT n H3 0.360 XT701 1.04 (41.00) 3 6 n T64 NA NA NA NA PROP-fAN/FUSELAGE CLEARANC1: 2.000 p.08 (200.00 3 6 T56 P- ..

,- 14.89 (192.40 1.700 XT701 6 TABLE D-VI. EC A-l PROP-FAN LOCATION EVALUATION - CONVAIR 990 WEIGHTING FACTOR = 2 CONVAIR 990 TESTBED AIRCRAFT OVERWING CRI- PROPUL HEIGHT/DIST RATlNG H = f (0 ) CRITERIA TERION LIMIT m (in) VALUE SYS P SCORE NA NA T64 NA NA GROUND CLEARANCE NORMAL 2.12 (83.5) 0.860 T56 6 Hl XT701 (81 .1) 0.710 6 2.06 3 n T64 NA NA NA NA GROUND CLEARANCE OUTBO. ENGINE ROLLED A ITITUDE 1.84 (72.5) 0.750 T56 6 FIRST CONTACT H2 XT701 1.59 (62.5) 6 0.55° 3 NA NA GROUND CLEARANCE T64 NA NA OUTBO. ENGI NE DEFLA TED TI RES & 1.52 (60.0) T56 0.620 3 FIRST CONTACT CONTRACTED STRUT P 1.32 (52.0) H3 XT701 6 0.46°0 3

-

T64 NA NA NA NA PRCP-fAN/FUSELAGE CLEARANCE 3.71 (146.0) 6 T56 1.500 n

-

F XT701 6 3.49(137.5) 1.210 n CONVAIR 990 TESTBED AIRCRAFT;NOERWING CRI- PROPUL HEIGHT/DIS RATING

CRITERIA H = f (0 )

TERION LIMIT VALUE SYS m (ins) P SCORE T64 NA NA NA NA GROUND CLEARANCE NORMAL 1.44(56.5 ) 0.580 T56 6 Hl XT701 1.22 (48.1) 0.420 T64 NA NA NA NA GROUND CLEARANC: OUTSO ENGINE ROLLED A ITITUDE 0.440 T56 1.08 (42.5) 6 P. FIRST CONTACT H2 XT701 0.76 (30.0) 0.26D 3 6 n GROUND CLEAAANC~ NA T64 NA NA NA OUTSO. ENGINE DEFLA TED Tl RES & 0.310 T56 0.76(30.0) 3 FIRST CONTACT CONTRACTED 6 STRUT H3 XT701 0.44 (17.5) 0.150., 3 6 NA NA NA NA T64 ?ROP-FAN/FUSELAGE CLEARANC: 1.540 T56 3.81 (150.0) 3 6 ;: XT701 3.58(141.0) 1.230 I 3 6 p TABLE D-VII. EC A-l PROP-FAN LOCATION EVALUATION - Gil WEIGHTING FACTOR = 2 GUlFSTREAM II TESTBED AIRCRAFT OVERWI NG· CRI- PROPUL HEIGHT/OIST RAT1NG

H = f (0 )

CRITERIA TERION LIMIT m (in) VALUE SYS P SCORE NA NA NA T64 NA GROUND CLEARANCE NORMAL T56 0.61 (24.0) 6 0.250", 3 Hl XT701 (17.4) 0.44 0.150", 2 4 NA NA NA T64 NA GROUND CLEARANCE PROP-fAN TIP ROL1.ED ATTITUDE 0.27 (10.8) T56 0.11 On 4 FIRST CONTACT H2 XT701 0.11 (4.5) 0.040 NA GROUND CLEARANCE NA NA NA T64 PROP-FAN TIP DEFLA TED TI RES & T56 0.18 (7.2) 0.700 4 2 FIRST CONTACT CONTRACTED STRUT n 0.010 H3 (1 .2) XT701 0 •. 03 0 0 n NA T64 NA NA NA PRCP-FAN/FUSELAGE CLEARANCE T56 1.48 (58.3) 0.600 6 I-- i XT701 6 1 .26 (49.46) 0.440 3 n EC A-2 Engine-out Safety Primary Engine-out Operation Prop-Fan Engine-out Operation The testbed aircraft must be capable of takeoff and landing with a primary engine failed and with the prop-fan at flight idle or full power. This criter- ion is particularly important where primary engine substitution has been made.

The C-141A, the KC-135A, and the Convair 990 fall into this category of air- craft.

Data for the KC-135A and Convair 990 are not available for an assessment of the two-engine operation. However, the data for the C-141A two-engine operation have been analyzed and are presented in Figures D-2 and D-3. The most critical case, that of takeoff with Air Force hot-day conditions prevailing is shown.

The thrust available and thrust required, and the drag increment due to two failed engines are shown in Figure D-2.

Drag at L/D and the thrust at normal rated and military rated thrusts MAX are shown Figure D-3 for two engine operation. The corresponding climb GROSS WEIGHT LB KN I<g.,O· (LB. IO~ MRT 127.0 (280) 108.9 (2.0) 99.S (220) ~ 90.7 (200) 3.

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'-' ~!

60 SO 100 120 l40 i6J " 120 i60 200 240 280 ;:0

P:! i WEIGHT _ • iO·

).t;;--,)"';.l,----,7'J.'i'J --0'0. 7 4--"7';:.5 'MCH NUM8ER Figure D-3.

Figure D-2. C-141A Thrust Available and Required - C-141A Two-Engine Operation Two Engine Operation gradients are also shown on Figure D-3.

The air minimum control speed for two engine operation of the C-141A covers a band of speed of 69.4 to 74.6m/s (135 to 145 knots) true airspeed. Th(>~.: cat'} show that a posi ti ve climb gradient of about 1 percent is available for two engine operation.

No data are available for the KC-135A and Convair 990; however, assuming that similar conditions exist for these aircraft, at testbed weights, two-engine performance should be available to provide a measure of safety.

In the case of the Gulfstream II, the propulsion system is an addition to the existing primary propulsion system so that normal operation is possible and FAR Part 25 performance with one primary engine failed is satisfied.

The data for this evaluation are shown in Table D-VIII.

TABLE D-VIII. EC A-2 ENGINE-OUT SAFETY EVALUATION WEIGHTING FACTOR-4 CRI- RAMI' CUMBGRAD RAnNG TERION WEIGHT 2-£NG AF HOT TEmEDA/C PROPULSION DAY GEAR DOWN RISK VALUE SCORE SYSTEM Kg (UI SUBSTITUnON 99271 HIGH 1 <4 I%@ NRT C-I<4IA (218,900)

I

HIGH 1 <4 NO DATA KC-13SA SUBSnTUT10N 82990 SIMILAR (183,000) TO C-I<4IA 87012 HIGH I SUBsnrunON NO DATA CONVAIR 990 SIMILAR (192,000) TO C-I<4IA

I

LOW 3 12 ADDITION NO REDUCTION

GULFSTREAM 11 I 21210

IN CUMB

I (60,000)

GRADIENT Aircraft Structural Integrity The prop-fan must be installed without incurring problems which could affect the structural integrity of the testbed aircraft. The risk of encountering wing flutter problems and severe changes in balance must be evaluated.

EC A-3 Structural Integrity Wing Flutter Appraisal Aircraft Balance Check Modification of the candidate testbed aircraft must be achieved without adversely affecting the airframe structural integrity. The two principal concerns in this area are wing flutter and changes in the aircraft balance characteristics.

Wing Flutter - Appraisals of the candidate testbed aircraft have been made relative to the risk of encountering wing flutter problems which could jeopardize the testbed program. The candidate testbed aircraft have been appraised based upon flutter parametric analyses, where available, and on the basis of location and extent of changes in the mass and inertial properties of the wing-engine systems.

The propulsion system/wing configurations are shown on Figures 0-4, 0-5, 0-6 and 0-7 for the Lockheed C-141A, Boeing KC-135A, Convair 990, and Gulfstream II and the data for the evaluation are shown on Table D-IX.

Iliol W'I' in] IA (UOl \.1) .

r-- .),. (140.) ... ) ~ __

l

(:::'-C10.0 rw ZI· -:O"~~:"?

-t- - . ----ol-+---

L rwa rLY 1/1' lll' Ie (7otl LI) .4.~.1

. o_o-+o~-·

____ ----~~O ~---,--..

O_. __ ._._.!

,~ ______ ~ool _____________ ~~rl

;_".__ . _....,.-- _. __ 0 _-+--+!ftl- Figure D-4. C-141A Propulsion System Changes .'-....

!i.ASTle AilS Nt: "-1'" 1""1. tnuLl) f Figure D-5. KC-l35A Propulsion System Changes Figure D-6. Convair 990 Propulsion System Changes

I

_ 4==-. ---,-=/=-~_~---,-r- --:-~-fIJ~

'I - /

\\ !

f~ ---'---~:----""--- Figure D-7. Gil Propulsion System Arrangement Aircraft Balance - Modification of the aircraft to the testbed config- uration must not cause undue restriction of the useable range of center-of-gravity location or cause the center-of-gravity to move beyond the existing boundaries of the aircraft center-of-gravity envelopes.

Center-of-gravity changes must not cause aircraft flight restriction within the existing structural envelope. Longitudinal imbalance may be corrected by the addition of ballast, which may include a fixed amount of fuel. Latera 1 imbalance may be corrected by fuel management procedures and by the addition of ballast where necessary.

TABLE D-IX. EC A-3 STRUCTURAL INTEGRITY EVALUATION WEIGHTING FACTOR· 4 CRI.

RATING RATING

CRI- I!

SCORE iERlCN FlUmR VALue A/C BALANC. VALUE TESTBED A/C AVG. SCORE I Within Current ~R/llc 3 3 3 C-I~IA Env.lope

I

Wlthin Current 3 3 12 ~RlIIc 3 • Env.lo ..

KC-135A

I

Wttnin CUrTWIt l 3 ~Rllk 3 Env.1ope CONVAIR 990 ,- Lot_I ond Long;-I tudinal Balance 2 2 a ~teRllk 2 GULFSTREAM /I Affected. So lIolt Required I C-141A Balance - The C-141A aircraft as a prop-fan testbed has no proble~3 Trom the standpoint of aircraft balance. The substitution of the prop-fan propulsion system for the inboard TF33-P-7 engine and nacelle group results in negligible change in the balance characteristics of the aircraft.

There is no significant difference in the overwing versus the underwing installation of the prop-fan propulsion system from the standpoint of aircraft balance.

KC-135A Balance - The balance characteristics of the KC-135A aircraft as a prop-fan testbed vehicle are not significantly changed by the prop-fan installation, since the location of the horizontal axis of the testbed propulsion system is very close to that of the inboard nacelle. No detailed balance data are available for the KC-135A, but it is unlikely that the aircraft balance wi 11 be adversely affected by the substitution of the prop-fan propulsion system. There will be no significant difference in the longitudinal balance effects for the overwing or underwing prop-fan installations.

Convair 990 Balance - The Convair 990 aircraft as a prop-fan testbed has excellent balance characteristics. The prop-fan installation, which is lighter than the CJ805 engine installation it replaces, is mounted so that the center of gravity of the total installation is behind that of the CJ805 installation. The total change in aircraft longitudinal moment is negligible. There is no significant difference, from the standpoint of aircraft horizontal balance, between the overwing and the underwing installation.

Gulfstream II Balance - The Gulfstream II encounters some balance problems as a prop-fan testbed because of the small size and geometry of the aircraft so that the installation of the prop-fan has a greater influence than occurs on the other, larger candidate airplanes. Since the prop-fan propulsion system is an add-on rather than a substitution, the total zero fuel weight is increased rather than decreased, and since the prop-fan installation is mounted on the wing, the balance characteristics of the aircraft are affected both laterally and longitudinally. The lateral unbalance can be corrected by the addition of lead wingtip ballast on the side opposite the prop-fan engine installation. The Gulfstream II has the structural capability for wingtip tanks, and since the testbed aircraft will not require these tanks and the wingtip ballast required for lateral balance weighs less than the tank and fuel, no additional structural changes should be required.

The wing tip ballast will also be of benefit to the longitudinal balance, since the ballast center-of-gravity will be considerably aft of the wing mean aerodynamic quarter-chord-point. This will tend to offset the effects of locating the prop-fan installation forward of the MAC quarter chord. The air- craft, although limited in payload capability, will still be able to accommodate the testbed propulsion system as well as the required ballast, within the zero fuel weight envelopes of the basic aircraft. The balance characteristics will, therefore, be maintained.

The flutter appraisal and balance characteristics and evaluations are shown on Table D-IX.

AIRCRAFT OPERATIONAL CHARACTERISTICS REQUIREMENTS The operational characteristics requirements for the aircraft must include compliance with design cruise conditions, a practical test mission duration, and acceptable aircraft stability and control and prop-fan installation effects.

Compliance With Design Requirements The testbed aircraft must comply with the required design cruise conditions of a cruise Mach No. of 0.8 at 9144m <30,000 ft) altitude and above and the proximity of the testbed aircraft cruise conditions to the high Mach number buffet limits which may impose constraints on the useable range of weight and lift coefficients at a Mach number of 0.8 must be determined.

EC B-1 Design Cruise Conditions Compliance Aircraft Speed/Altitude Capability High Speed Buffet Constraints Each testbed aircraft must be capable of performing the test mission at a Mach No. of 0.8 at altitudes of 9144m (30,000 ft) and above. Furthermore, the cruise capability should not be impaired by high-speed buffet constraints over the range of weights for the test-mission profile. A reduced buffet limit with the prop-fan installed has been determined for each aircraft. The combined data for cruise performance and buffet boundaries are shown in Figures D-8, D-9, D-10, and 0-11 for Lockheed C-141A, Boeing KC-135A, Convair 990 and Gulfstream These data show the speed/altitude capability at start and II, respectively.

FT I, /19 aUFFEr LIMIT :3 1I lc

,J

~ I" END TEST WEIGHT-' 6O,<J7.Kg (146.500 La)

:;..

1·') .., 3: :: JO )))7/;};;;7}7) 7 7 7.'1 7 )))); 'v))))) / ~ / ,..

I

~

II

_.~_' °.3 3[ IT·'" ,. W,,"",_ "," ~ "'MOO "'

'/

I , <Tnl ?,CP";">'N iNSTALJ....ITlCN

f

o ?ICP"'AN CPf~.ATING ':' --~--~'~--~'--~~--~'--~'.

'.73 .':0 .i2 .34 .30 .ia .;0

t

,: .\'Jt.CH NO.

\oI.jI.C~ NO.

Figure D-8. C-141A Speed/Atltitude Figure D-9. KC-l35A Speed/Altitude Performance and Buffet Boundaries Performance and Buffet Boundaries • I 9 aUffET LIMIT ;> ~RtC~UISE 'NT.

FT 4(l ~'\lIN PROP-FAN V: ruLl.

iAX ,8 rOWi;R I1N PlUlP-FAlIS MIN ~Net:1ES AT APPROX1MATZLY I:JLE 1\ ,6 :t END WEIGHT '\ ~.

:l4 '"", • 6lJ,769 Kg (1:l4,OOO L!)\ ~ I~ . ., 32 ~ ...

:>

lm:~~n"m")"

§

\ , -< JC ''';1;;''';~');7771);';)

STA~lJEST. WEIGHT \ /'

.82,337 Kg (182,000 LS~ i :6

i • XTiJl P~CP"'AN INSTALLATION

~ PRCP~AN OPERATING

1m

LIMIT .76 .dO .12 ~~~--~.3~0--~.3~2--~.~~--.~30~--.~3.~--j.l0 .'Meli NO.

Figure D-11. G II Speed/Altitude Figure D-10. Convair 990 Speed/Altitude Performance and Buffet Boundary Performance and Buffet Boundaries end cruise weights, together with the high-speed buffet limits and the boun- daries imposed by the design requirements.

The rating considerations are shown in Table D-X.

Test Mission Duration The test mission duration must be long enough to permit the acquisition of good test data economically.

EC B-2 Test Mission Duration Testbed aircraft will be ranked according to test mission duration.

The mission profile and the test duration for the Lockheed C-141A, Boeing KC-135A, Convair 990, and Gulfstream American Gulfstream II are shown in Figure D-12. All testbed configurations have acceptable test-mission duration.

TABLE D-X. EC B-1 DESIGN CRUISE CONDITIONS COMPLIANCE EVALUATION

REQUIREMENT - MACH 0.8 AT 9144 m (30,000 Fn AND ABOVE

WEIGHTING FACTOR = 4 0 .

MIN WT-Kg/AlT-m MAX WT -Kg/AlT-m RATING I CRI.

L\M TESTBED A/C VALUE SCORE MAX WT-lB/AlT-FT MIN WT-lB/AlT-F~ (PF OFF) 66,438/10,668 95,235/10,668 .006/.002 1 C-141A (146,500/35,000) (210,000/35,000) \ 4,852/10,688 ~ 75,734/10,058 .048/.01 (167,000/33,000) 1(107,200/35,000) KC-135A J 75,734/10,058 \ 4,852/11 ,277 .066/.057 ** 3 **, ** i(1 07,200/37,000) (167,000/33,000) 82,537/9,144 60,769/9,753 .08/.03 CONVAIR 990 3 12 (182,000/30,000) (134,000/32, 000) ~ 26,303/9, 144 ~ 19,500/9,144 .05/.04 (58,000/30 ,000) (43,000/30,000) GUlFSTREAM II 12 ~26,303/101668 ** ~ 19 I 500/9 , 144 .05* **3 ** (43,000/30,000) (58,000/35,000) HIGH SPEED BUFFET CONSTRAINTS "PLACARD LIMITED

- WEIGHTING FACTOR = 4

"P:~OP-FAN ON CRI- WT -Kg/AlT - m RATING TERION

TESTBED A/C L\~OT / L\MMARGIN

(WT -LB/AlT -FT) SCORE 66,438/10,668 .01/0 (146,500/35,000) C-141A 95,235/10,668 .002/0 (210,000/35,000) KC-J35A 48,615 (107,~00) .08/.01 3 12 60,769 (134, 00,) 3 CONVAIR 990 .08/.005 12 26,303 (58,000) GUlFSTREAM II 3* 12 .04/0 *GULFSTREAM II IS BUFFET LIMITED WITH PROP-FAN ON.

ALL OTHER AIRCRAFT ARE THRUST LIMITED EC B-1 DESIGN CRUISE COMPLIANCE OVERALL RATING AIRCRAFT CRUISE BUFFET OVERALL C-141A 4 4 KC-135A 12 12 12 CONVAIR 990 12 12 12 12 12 12 GULFSTREAM II RUNWAY CONDInONS

'\

ELfVATION - 701 In (2,:lQO FT.)

TEMPERATURE - 26.67"C (aOoF) ~'\ NO WIND NO GRADIENT PROP-fAN TEST '\.

MISSION PROFILE '\ , 1/ TA~EOFF .,..".

LANDING WEIGHnNG FACTOR. 3 START CRUISE"WT •.

MISSION DURATION

.J CRITERION I TESTBED

END CRUISE 'NT.

HRS RATiNGI SCORE I A/.'- 1<9 (LS) 95,235 {210,OOCl.

LOCKHEED I 4.33 3 9 C-141 66,-137 (l46,SOO) I

I I

75,734 (l67,OOO) 9 SOEING ~.69 3

I KC-135A

48,615 (lO7,200) !

CCNVAIR 32,337 (l82,aoo) 9 i 4.2 3 990 60,769 (l34,OOO) i

I

26,303 (sa,OOO) I GULFSTRfAM 1\ 3.3a 3 ? i 19,~ (42,964)

I I

Figure 0-12. EC B-2 Test Mission Duration Aircraft Stability and Control The testbed aircraft must be capable of operating as a stable platform to permit the acquisition of good test data.

EC B-3 Aircraft Stability and Control The prop-fan has a destablizing effect on both longitudinal and lateral-directional control, and this effect is more pronounced on the smaller aircraft.

Each testbed aircraft must exhibit good stability characteristics over the full range of prop-fan power settings and test conditions.

Each testbed aircraft must be able to achieve trimmed flight attitudes without large incidence and yaw angles on the prop-fan and "be able to trim at various angles of incidence when desired.

Estimates of the normal force caused by the installation of the prop-fan were used to determine the changes in stability derivatives Cn~ and C ' the ma yawing and pitching moment derivatives, respectively. These data are shown in Figures 0-13, 0-14, 0-15 and 0-16 for the aircraft with and without the prop-fan installation.

C-141A - The aircraft total C and C were obtained from C-141A data and nfJ rna are shown in Figure 0-13. These data indicate very little change in the levels of aircraft stability due to the prop-fan. The greatest reduction in the level of yawing moment derivative occurs at the low speed end of the Mach number band and amounts to a loss of 1.86 percent. The loss in pitch stability is almost constant over the entire speed range and amounts to 1 percent.

EFfECT OF PROP-FAN ON YAWING MOMENT EFFECT OF PROP.FAN ON YAWING MOMENT - WITHOUT PROP-FAN G' - WITHOUT PROP-FAN G' ...

--- WITH PROP.FAN ~ -- WITH PROP.FAN

... ~

... ~ 1.6 / 2 , I ..!S.

I "I.

= , u

u=<Q. , "

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i ~ 0 .2 .4 .a .a .~ ..

o • 2 MACH NUM8!R MACH NUMW EFFECT OF PROP.FAN ON PITCHING MOMENT EFfECT OP PROP-FAN ON PITCHING MOMENT -- WITHOUT PROP·FAN --WITH PROP·FAN - WITHO~I!.0~AN -- WITH PROP-FAN I

"

'I' 'I' ., ., ., 'I' ",' ..

....

,," ,..------,-~-~-"" .I.a~--~";;;;--~--~--~ o .2 .4 .6 .3 MACH NUMW .2 .4 .6 .a MACH NUMW Figure D-14. KC-l35A Effect of Figure D-13. C-141A Effect of Prop-Fan on Yawing and Pitching Moments Prop-Fan on Yawing and Pitching Moments EFFECT OF PROP-FAN ON YAWING MOMfNT EFFECT CF PROP-FAN ON YAWING MCMENT -- WITHOUT PROP-FAN 3.0 -- WITHOUT PROP-FAN --- WITH PROP-FAN is 'I ---- WITH PROP-FAN 'I

.. ~

'I' , 2.S C~ ~ U " Ol " C ...... Z 2.1 U "

----------

- --------_ .... '

--

~ -- .,-

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=<

" ;; 2.0

?;

"

~ ~ >- .a o .2 .4 .6 0 .2 .4 .6 .a 1.0 MACH NUMBER MACH NUMB£R EFFECT OF PRCP-FAN ON PITCHING MOMENT -- WITHOUT PROP-FAN EfFECT OF PROP-FAN ON PITCHING MOMfNT -- WITH PROP_FAN -- WITHOUT ?IOP-FAN --- WITH PROP-FAN s" u -3.4 Z ~ ::;:

"

-3.0 ~

~

.4 .6 .3 .2 MACH NUMBER , l..:l

.. .4 •• .3

o MACH NUMBER Figure D-16. Gil EffeCt of Prop-Fan Figure D-15. Convair 990 Effect of on Yawing and Pitching Moments Prop-Fan on Yawing and Pitching Moments KC-135A - The deri vati ves for the KC-135A were generated using DATCOM to determine the effect of the prop-fan installation on stability. The data, shown in Figure D-14 indicate a loss of C of 1 percent at low Mach numbers and less

than 1 percent at high speeds. Sim~~arlY, the loss in C is 1.6 percent at low

ma Mach numbers and 1 percent at high Mach numbers. No significant loss in stabil- ity, therefore, occurs with the installation of the prop-fan.

Convair 990 - The stability derivative data for the Convair 990 shown in Figur~ D-15 were also derived by means of DATCOM. The greatest loss in the yawing moment deri vati ve occurs at low Mach numbers and amounts to 1 percent with the prop-fan installed. At high Mach numbers, the loss in C is less than nfJ 1 percent. The losses in C amount to 1.4 percent at low Mach numbers, reduc- ma ing to 0.7 percent at high Mach numbers.

Gulfstream II - The deri vati ves for the Gulfstream II, also obtained by using DATCOM to determine the effect on stability with and without the prop-fan, are shown in Figure D-16. The data indicate that the prop-fan has a greater effect on the stability of the smaller aircraft than on the much larger candi- dates. The greatest change occurs to cn~ at low speeds for which the prop-fan changes the level by almost 8 percent. The decrease in stability, although not apparently dangerous. is signi ficant in that it does highlight areas having potential for problems such as engine-out characteristics and high-altitude dutch roll/dynamic stability.

The changes in the stability derivatives are shown in Figures 0-17. 0-18 • . 0-19. and 0-20 for the C-141A. KC-135A. Convair 990, and Gulfstream II, respec- tively. All the candidate testbed aircraft exhibit similar characteristics over the range of Mach numbers considered.

$ EFFECT OF P~OP-FAN ON YAWING MOMENt EFFECT OF PROP.FAN ON YAWING MOMENT .2.S ~~ $

.. ~ .2.6

~ ='Z:.

u ~ <l ·2.0 ~ U <l , .2.2

z

~ ~ ~ .1.S Z j < >- ii ~ .1.4 0 .2 ., .6 .2 .4 .~ .s .3 ~ ::: MACH NUMBER MACH NUMBER u EFFECT 01' PRCP-FAN ON PITCHING MOMENT is EffECT OF PROP-FAN ON PITCH IN G MOMENT

J

~ .

e ~ u ~ 1.5 ~ 1.4 u <l

-

~

~ 1.2 ~ Q 0: .4 ., .3 .2 .2 ... .~ .3 .'MCH NUMBER MACH NUM8ER Figure 0-17. C-141A Figure 0-18. KC-135A Change in Yawing and Pitching Change in Yawing and Pitching Moments Due to Prop-Fan Moments Due to Prop-Fan G" EFFECT OF PROP-I'AN ON YAWING MOMENT

J

2 -1.3 EFFECT OF PROP_FAN ON YAWING MOMENT ~ c<Z:.

u ~ I -1.1

i

g

.2 .4 .6 .s o .2 .4 .6 .a MACH NUMBER MACH NUMBER EFFECT OF PROP-I'AN ON PITCHING MOMENT G"

\

EFFECT OF PROP-FAN ON PITCHING MOMENT J 7.0

~ 2.0 o· S u ...

, 6.0 "

~

~ ~ S.O g ...

; .... 4.0 .2 .4 .0 .3 .2 .4 .0 .3 IMCH NUMBER MACH NUMBER Figure D-19. Convair 990 Figure D-20. Gil Change in Yawing and Pitching Change in Yawing and Pitching Moments Due to Prop-Fan Moments Due to Prop-Fan In general, no significant losses in stability and control have occurred as the result of the prop-fan installation.

The evaluation of the candidate testbed aircraft stability and control is shown on Table D-XI.

TABLE D-XI. EC B-3 AIRCRAFT STABILITY AND CONTROL EVALUATION WEIGHTING FACTOR .. 4 PROBLEM RATING TESTBED AIC % CHANGE C % CHANGE C AREAS VALUE CRITERION SCORE n m II f3 C-141A ::: -2 <-1 NONE 3 KC-135A :::::: -I <-2 NONE 3 CONVAIR 990 :l: -1 NONE 3 12 <-2 GULFSTREAM II -3 TO -5 -1.3 TO -7.8 DUTCH ROLL /DYNAMIC STABILITY HIGH ALT.

EC B-4 Prop-Fan Installation Effects The installation of the prop-fan propulsion system will affect the high-lift devices and flight controls systems to the extent that the operational characteristics of the testbed aircraft could be changed. The degree of interference caused by the prop-fan will be assessed and rated based on the magnitude of the problems.

The installation effects of the various propulsion systems are due to the interference of the prop-fan installation on essential devices such as high lift and flight control systems.

The principal effects on the C-141A, KC-135A, Convair 990 and Gulfstream II are shown on Figures 0-21, 0-22, 0-23 and 0-24, respectively, for the overwing installations.

r--_ ...... -. __ . ~.

-----

\ \ \

"

((tL, ~_ _____~_ ~ . \

" . "

, \ \

~,--- ..... c ~

~\ ----------~---- "

"

Figure D-21. C-141A Prop-Fan Installatioo Interference Figure D-23. Convair 990 Prop-Fan Installation Interference

h I

~. --::;C.c _.

_ ~., L .~:I '··1~4

II "\ I

. I I -r---.!- { I ,-t..-----.~ -------~ 1'$311.00 I

rfl.--- __

.c&= - -- - """i

~- -;:= - -- -=:l--

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Figure D-22. KC-l35A Figure D-24. Gil Prop-Fan Installation Interference Installation Interference In the case of the underwing installations, interference occurs only at the leading edge.

The ratings of the interference effects are based upon the magnitude of the problems caused, such as loss of maximum lift coefficient at takeoff and loss of control power for flight, and on the difficulty of rectifying such deficiencies where such appear mandatory. The ratings are shown in Table D-XII.

TABLE O-XII. EC B-4 INSTALLATION EFFECTS WEIGHTING FACTOR· 4 INTE~FHENCE ~DING ,ilAIUNG EDGE EDGE FLIGHT RATING CRITE~ION ruTllro A/C DEVICE DEVICE CONTROL VALUE SCORf NONE NONE 12 uw NONE 3 C-141A OW NON! FLAP S?OIUR 1.5 0

I

uw SLAT NONE NONE 2.0 KC-I35A OW SLAT NONE HIGH SPEED 1.0 AIURON

I I

uw SLAT OUTlID NONE NONE a 2.0 CONVAIR99Q cw SLAT OUTlID FLAP SPOIUR 1.0 <4

I

GULFSTlIEAM II O'N NONE FlAP SPOILER 1.0 <4 TESTBED PROGRAM OBJECTIVES ACHIEVEMENT The Testbed Program Objectives, established in Task I, form the basis for the evaluation criteria ratings that measure the suitability of each candidate The Task I order of priority for these objectives is as a testbed aircraft.

followed in the listing of the Evaluation Criteria.

Realistic Environment for Dynamic Loads Validation An important objective of the testbed program is the determination of the prop-fan cyclical loading to validate the structural integrity of the prop-fan structure.

EC C-1 Dynamic Loads Validiation Environment The assessment of each testbed installation will consider the degree to which each provides a representative environment for the validation of prop-fan structural characteristics and in- duced effects upon the aircraft structure. This will include consideration of the engine nacelle overhang, toe-in and the proximity of the wing leading edge, fuselage and other aircraft components.

The prop-fan blade dynamic response is a function of the blade aerodynamic and structural dynamic characteristics and of the aerodynamic flow field in which the prop-fan operates. One drive system was selected for the four can- didate testbed aircraft, the 2.89m (9.5 ft) diameter prop-fan/XT701/T56-A-14 combination, thus eliminating prop-fan blade aerodynamic and structural charac- teristics as design variables.

Prop-fan flow field variations, which induce blade dynamic loads, are caused primarily by configuration geometry, i.e., proximity of the wing and, to a lesser extent, the fuselage. Typically, a commercial passenger aircraft configuration would be a low-wing arrangement having approximately 0.523 rad (30 deg) of leading-edge sweep and two or four prop-fan propulsion units mounted over the wing to provide sufficient ground clearance for the large-diameter prop-fans.

On this basis, the testbed aircraft configuration for proper representation of the prop-fan environment should be a low wing. The prop-fan located for adequate demonstration of blade characteristics should be one prop-fan diameter from the wing aerodynamic center (assumed to be at the wing quarter-chord for this study). F~r proper representation of the equivalent fUll-power prop-fan propulsion installation, the ratio of wing chord, C ' to prop-fan diameter, D ' w p should be in the region of 1.0 for inboard engines and 1.5 for outboard engines.

The prop-fan tip/fuselage clearance, F, should be a minimum of 0.2D for p acceptable excitation and a.8D for acceptable acoustic environment p characteristics. The geometric parameters are shown in Figure D-25 , and the rating values, showing the degree to which each candidate simulates the dynamic loads environment, are given in Table D-XIII.

Acoustic Data Acquisition The ability of each testbed aircraft as an instrument for obtaining near- and far-field noise data will be evaluated.

C .. - CHORD LENGTH AT NACELLE CENTERLINE ALE - LEADING EDGE SWEEP Dp - nOP-FAN DIAMETER Figure D-25.

Geometric Parameters for Dynamic Loads Validation TABLE D-XIII.

EC C-l DYNAMIC LOADS ENVIRONMENT VALIDATION EVALUATION WEIGHTING FACTOR • 4 WING FUSITIP ENGINE

ALfl

Cw/Dp PARAMETER RATING CRITERION LOCATION RAD(DEG CLEARANCE F INSTAllATION DESIRED 1. 0 INBD 0.2 Op STRUCT VALUE SCORE .523 (300, LOW OVERWING VALUE 1.5 OUTBD 0.80 AC C-141A HIGH .488 ( a- 2.54 OVER/UNDER 28 1.35 Dp 1.5 6 u KC135A LOW .663 (38°) 2.15 OVER 1. 66 Dp 2 8

<

(707-l201 Q .....

CQ .....

CONVAIR 990 LOW 1.698 (40") 2.34 1.21 Dp OVER 2 8 VI .....

.....

.488 (28°) GULFSTREAM 11 LOW 1.36 0.44 DD OVER 3 C • WI NG CHORD w Dp = PROP-fAN DIAMETER EC C-2 Near-Field Noise Data Acquisition The effects of configuration geometry on near-field noise include prop-fan to fuselage clearance, propfan centerline to fuselage centerline relation, prop-fan/wing leading edge clearance, prop-fan/fuselage diameter ratio, prop-fan shielding by existing components and proximity of other powerplants.

Other considerations include flap and control surface immersion in prop-fan wash and the effect of testbed attitude character- istics on near-field noise measurement.

Since one of the major objectives of the testbed aircraft is to investigate near-field acoustic characteristics, it is important that the prop-fan be properly located so that clear noise signals can be obtained inside and outside of the fuselage. In addition, the fuselage structure and interior trim and furnishings should be representative of the commercial aircraft environment.

Furthermore, the fuselage structure in the region of the prop-fan plane should be capable of modification to test various noise-attenuation concepts.

This evaluation criterion includes all these considerations as shown in The interiors of fuselage for the C-141A and KC-135A are con- Table D-XIV.

figured for military use and are, therefore, not representative of commercial configurations. Some modification of the basic aircraft would be necessary in the prop-fan plane region to simulate a passenger aircraft configuration. In the case of the KC-135A, this deficiency could be overcome by using the 707-120 series aircraft.

TABLE D-XIV. EC C-2 NEAR-FIELD NOISE DATA ACQUISITION EVALUATION WEIGHTING FACTOR = 4 TESTBED AIC FUSELAGE INSULATION INTERIOR AC & PRESS. CRITERION COMMERCIAL ACOUSTIC RATING WINDOWS TRIM DUCTING REPRESENT MODS SCORE VALUE NONE 1 4 C-14lA UW NONE INTERNAL NONE REP V V OW BLANKETS MILITARY 2 8 INTERIOR INTERNAL NONE NONE REP 2 KC135A UW NONE V V BLANKETS MILIT. INT.

(v)

( 707-1201 OW (.J) (.J) (REPI 3 (.J) 6/) UW PASSENGER 2 CONVAIR REP

v .J .J .J

OW CONFIG 3 PASSENGER GliLFSTREAM 2 OW REP .J .J .J .J CONFIG II EC C-3 Far-field Noise Oata Acquisition The capability of the testbed aircraft to provide prop-fan noise detect ability above the levels of the basic aircraft and engines will be evaluated by comparing the noise signature of the basic aircraft and engines with the predicted noise of the testbed prop-fan over a range of frequencies wide enough to provide useful data.

The acquisition of a good, clean, prop-fan noise signal depends on the ability to reduce background noise, generated by other noise sources, on the candidate testbed aircraft. This can be accomplished by: o Operating the prop-fan at the highest power setting (loudest) o Providing noise suppression for the prop-fan driver o Operating the primary engines at the lowest possible power setting (flight idle) o Operating the airframe in a "clean" configuration It is considered that prop-fan noise of good quality can be obtained by the above means, which could be used to validate prop-fan noise prediction methodologies.

The predicted noise characteristics for the prop-fan and driver are shown in Figure 0-26. These data clearly indicate the need for suppressing the XT701 driver noise to allow the prop-fan signal to dominate the noise spectrum. The aircraft component noise spectra for the Lockheed C-141A, Boeing KC-135A, Convair 990. and the GII are shown in Figures 0-27, 0-28, 0-29 and 0-30.

The ranking of the candidate testbed aircraft for far-field noise prediction methodology validation is shown in Table O-XV. The ranking is based on how well the prop-fan signal, S, is separated from the background noise, N.

in one-third octave band level decibels. The SIN factor is presented for the prop-fan fundamental tone and high frequency noise e.g. )1,000 Hz.

• 304.8 m (1000 FT)ALTrTUDE • 5964.8 KW (8000 SHP)

• DIAMETER = 2.89 m (9.5 FT)

XT701 DRIVE SYSTEM • NO. OF BLADES = 8 co ': 100 • TI P SPEED = 244 m/s (800 FPS) • GRD FLUSH MICROPHONE • ~ = 1.57 RAD (90 DEG) III > - ..........

~ 60

o

' ......

-..

Driver With ~ """ ~ 15d8 Exhaust "- 'i" 40 III Suppression '\

o

3451>7891 3.

20000 FREQUENCY- J HERTZ Figure D-26. Prop-Fan Driver Suppression ...

...

I 100- • 304.8 m (1000 FT) ALTITUDE

•

j • GRD FLUSH MICROPHONE t .a • NO. OF ENGINES a 3

i

EST. '.ren, ;1. = I.ST RAD (90 DEG) .

~ Engine Nol ..

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......... ~,Driver "", (8000 SHp) , 2 3 S 3 •

•

2\} 100 20000 FREQUENCY, HERTZ Figure D-27. C-141A Component Noise Levels • 304.8 m (1000 FTlALTITUOE .

• GRD FLUSH MICROPHONE " ~ 100 .

• NO. OF ENGINES = 2

.s

1 • ;1- 1.57 RAD (90 DEG)

• l EST. Poren' Engine Nols.

~ ~ @FI,ldl.

~ lo ~ Supp.ouod Driver (8000 SHp) 23.567891 3. S67891 2 34567891 1000 ~ 100 20000 10000 FREQUENCY, HERTZ Figure D-28. KC-l35A Component Noise Levels • 304.8 m (1000 Fn ALTITUDE CO2 ""0 • GRD FLUSH MICROPHONE I "i • NO. OF ENGINES = 3 > GI .....

• ;1 = 1.57 RAD (90 DEG) ""0 C

.s

GI > .2 u EST Parent Engine No_ "'E @Fltldl.

~ 40 ....

I GI C 2 4567.91 234567891 2 :000 10000 20000 FREQUENCY' I HERTZ Figure 0-29. Convair 990 Component Noise Levels • 304.8 m (1000 FnALTITUDE co 100 ""0 • GRD FLUSH MICROPHONE

• NO. OF ENGINES = 2

• )1 = 1 .57 RAD (90 DEG)

4 5 100 1000 20000 FREQUENCY I HERTZ Figure 0-30. Gil Component Noise Levels TABLE D-XV. EC C-3 FAR-FIELD NOISE DATA ACQUISITION EVALUATION WEIGHT FACTOR • 2 TESTBED AIC FUNDAMENTAL HIGH FREQUENCY RATING VALUE CRITERION SCORE: S/N dB S/N dB C-141A 0 25 1 2 KC135A 25 0 CONVAIR 990 25 0 1 GULFSTREAM II 30 10-15 3 S/~ Signal to Noise Ratio Prop-Fan Scale EC C-4 Prop-Fan Scale Effects The diameter of the prop-fan is important in confirming manu- facturing and fabrication feasibility and for scaling laws validation.

Hamilton Standard recommends that the minimum diameter for the prop-fan, to ensure a representative structural configuration and characteristics, should be in the range of 2~44 to 3.05m (8 to 10 ft). The XT701 propulsion system is capable of driving a prop-fan having a diameter of 2.89m (9.5 ft) at current power levels and disk loadings.

Prop-fan scale should be consistent with the Hamilton Standard recommenda- tions for demonstrating the manufacturing and fabrication feasibility of the prop-fan. The recommended minimum value is 2.44m (8 ft). The XT701 drive system is capable of powering a prop-fan diameter of 2.89m (9.5 ft) and satis- fies this requirement. The rating values and criterion score are shown in Table D-XVI.

TABLE D-XVI. EC C-4 PROP-FAN SCALE EVALUATION WEIGHTING FACTOR • 4 RATING CRITERION PROPFAN SATISFY SCORE VALUE DIAM -FT REOMT TESTBED AlC UW 3 12 YES 9.5 C-14lA OW UW 3 12 YES 9.5 KC135A OW UW 3 12 9.5 YES CONVAIR 990 OW 3 12 YES OW 9.5 GULFSTREAM II Installed Propulsive Efficiency and Interaction Effects An objective of the testbed program is to demonstrate that the prop-fan net efficiency of 80 percent can be achieved at a cruise Mach number of 0.8. The evaluation will consider the degree to which this may be accomplished on the various configurations.

EC C-5 Installed Propulsive Efficiency Validation Each testbed configuration will be evaluated for suitability to obtain the data necessary for propulsive efficiency validation.

This will include consideration of the test equipment required and the accuracy of the data obtained.

The aerodynamic/propulsion system integration will be concerned with the need to maintain the aerodynamic efficiency of a modern, high speed airfoil immersed in the slipstream of a prop-fan. Also of concern, will be the need to recover some of the propulsive thrust from the slipstream swirl.

Two methods of establishing installed propulsive efficiency are by: 1. Evaluation of flight test performance data 2. Conducting wake surveys by means of pressure rakes The first method is indirect. Unless the candidate testbed aircraft can fly on prop-fan power alone at the cruise point, at which measurement of the propulsive efficiency is to be performed, the speed-altitude conditions must be attained through a combination of prop-fan and primary engine thrust. The thrust of the primary engines must, therefore, be separated from the total thrust to obtain the performance of the prop-fan. This procedure necessitates accurate determination of the thrust of all of the propulsive units contributing to the total thrust and may suffer somewhat in accuracy.

The second method, which is also the preferred method, determines propul- si ve efficiency by conducting wake surveys with pressure rakes located behind the prop-fan and behind the wing to measure the average momentum of the prop-fan wash and wing wake. Instrumentation of the wings of the candidate testbed aircraft can be accomplished with varying degrees of difficulty. The C-141A and KC-135A do not present problems, the Convair 990 requires removal of the anti- shock bodies behind the prop-fan unit. Finally, the GIl instrumentation instal- lation requires care to minimize flow distortion to the aft mounted primary engines. Although the candidate testbed aircraft differ, the ability to obtain accurate test data by this method is not significantly affected by the varia- tions. The configurations are ranked equally for the evaluation, but no account of this feature will appear in the evaluation score.

It is important that the configurations on which installed propulsive efficiency is measured represent realistic geometric conditions, to remove uncertainties due to the size or scale of the prop-fan propulsion unit.

Representative values of geometric relationships, such as prop-fan diameter/ fuselage diameter, prop-fan diameter/wing chord, effective disc area, and slipstream affected area/total wing area, are important if a realistic environment for propulsive efficiency validation is to be provided. Wind tunnel tests have shown that the propulsive efficiency of the uninstalled prop-fan may be increased as much as 5 to 6 percent if the energy locked in the swirling com- ponent of the propwash can be extracted.

It is possible that, by properly contouring the wing and nacelle in the propwash, some of this thrust loss may be recovered. Therefore, if the prop-fan size relative to the wing size is not realistic, the possibility of verifying thrust recovery techniques diminishes.

Realism in the proportionate scale of the prop-fan and aircraft is thus of primary importance.

It has been analytically determined, Reference 3, that optimizing aircraft, on the basis of takeoff noise footprint, results in a ratio of prop-fan diameter/fuselage diameter of 1.5. It has, therefore, been concluded that the ratio of Dp/Df for the testbed aircraft should be as close to 1.5 as practicable to yield representative data.

The data fr(lI'.l Reference 3 also showed the ratio of prop-fan diameter/wing

chord, D Ie , to be in the range of 0.95 to 1.5, depending upon the locetion and

p w number of propulsion units. Furthermore, the scale effect of slipstream-washed wing area/total wing area and power loading ratio will be realistic if the ratio for S IS is 0.17, and the power loading ratio is 0.4.

w slip w total The effective disc area, which is a measure of the nacelle blockage, is based on a ratio of Dn/Dp = 0.35, where Dn is the equivalent diameter of the nacelle. In the case of the XT701 nacelle, this value is 0.39, which although slightly larger than optimum, can be reduced by further refinement of the nacelle.

The values of the various ratios and the evaluation to determine the suit- ability of each candidate testbed aircraft as a vehicle for installed propulsive efficiency validation are shown in Table D-XVII.

TABLE D-XVII. EC C-5 INSTALLED PROPULSIVE EFFICIENCY VALIDATION EVALUATION WEIGHTING FACTOR = 2 o 10 C /0 0/0 DISC LOADING/ ITEM DFus/C SWSli/Sw w RATING CRITERION p' Fuse w p n p WI NG LOADI NG SCORE VALUF DESIRED VALUE 1.0101.5 I TO 1.5 0.35 1.0 0.4 0.17 C-141A UW 0.39 0.59 0.423 0.082 I 0.67 2.54 OW Y.

KC-135A UW I 2 0.80 2.15 0.39 0.59 0.432 0.0896 « OW CONVAIR UW w 0.80 2.34 0.39 0.54 0.370 0.1116 4 .., 2 OW l- V> W I- GULF- 0.60 OW 1.21 1.36 0.39 0.493 0.1553 3 6 STREAM II D = PROP-FAN DIAM = SLIPSTREAM WASHED AREA S p w S1ip = FUSELAGE DIAM D Fuse S = WING TOTAL AREA w D = NACELLE EaUIV. DIAM n C = WING CHORD LENGTH w EC C-6 Interaction Effects Validation The effects of slipstream superveloci ty and swirl can be re- duced by local tailoring and contouring of the wing. Each con- figuration will be evaluated by considering testbed installa- tions with regard to their relative sizing of the prop-fan, nacelle, and wing reflecting the ability of each to render representative aerodynamic data.

EC C-6 Interaction Effects Validation - The ability of the candidate test- bed aircraft to yield valuable data on interaction effects is strongly reflected by the evaluation for propulsive efficiency validation. In the case of inter- action effects, the principal considerations for a realistic environment were dependent upon geometric relationships.

In the case of installed propulsive efficiency, however, the important considerations for a realistic environment depend upon the position of the nacelle relative to the wing and on the wing section sensitivity to swirl and supervelocity effects. For the first of these considerations - the position of the nacelle on the wing - only two configura- tions are of interest: (a) the nacelle placed on top of the wing, and (b) the nacelle placed under the wing. In the general application, both locations are likely to be encountered and will depend on aircraft type.

It has been speculated that the underwing location, which is typical for high-wing cargo aircraft configurations, may have less adverse effect on wing flow. This, however, has yet to be proved. Commercial passenger aircraft are generally 10w-loItng configurations, which for prop-fan dedicated aircraft, would require an overwing nacelle installation. Because of the lack of data, compari- son of the two nacelle locations cannot be featured in the eval uation from the point-of-view of establishing interaction effects.

Proper evaluation of the interaction effects requires that the testbed air- craft wings should be representative of the wings of future prop-fan-powered aircraft. This means that the wings should have an advanced, transonic airfoil section with the thickness and sweep associated with cruise at the appropriate Mach number. Because all of the candidate testbed aircraft are configured from existing aircraft of varying age, this criteria cannot be met so that some compromise is necessary.

This suggests that the nacelle/wing relationship be such that some local reshaping to approximate a realistic aerodynamic environment would be desirable.

The geometric characteristics for propulsive efficiency validation apply to the evaluation of the interaction effects.

However, two additional geometric parameters are considered in the interaction effects validation: nacelle overhang and leading-edge sweep. The position of the nacelle on each of the testbed aircraft configurations is arranged so that the prop-fan plane is one prop-fan diameter from the wing leading edge at the center line of the nacelle. Al though this arrangement produces low excitation factors for the prop-fan, sufficient clearance is provided to enable filleting and contouring of the nacelle/leading edge to be performed in order to investigate interaction effects.

Where propulsion system substitution has been performed, the amount of modification permissible is somewhat limited. For the case of propulsion system addition, where more extensive modification to the wing is required, the oppor- tunity to extensively contour the wing/nacelle intersection is much greater.

This evaluation is, therefore, based on the degree to which a realistic environ- ment for interaction effects can be simulated on each testbed aircraft.

The data for the evaluation are shown in Table D-XVIII.

TABLE D-XVIlI. EC C-6 INTERACTION EFFECTS VALIDATION EVALUATION WEIGHTING FACTOR = 2 LE SWEEP A/C TYPE BASIC A/C NACELLE IP EFF WING RATING CRITERION TESTBED A/C A LE SIMULATION OVERHANG TYPE EXC. FACTOR CONTOURING VALUE SCORE RAD {OEG UW MIL/CARGO MllI~ARY 1.00 .489 (28'1 - LIMITED FWD C-141A I P OW COMM/PASS - OF F/S UW MILITARY 1.00 .663 (38'1 CLIMB 2.81 LIMITED FWD KC-135A 2 P OW COMMjPASS CRUISE 1.89 OF F/S CONVAIR UW COMM 1.00 .680 (39°)

- LIMITED FWD

P 4 OW 990 COMMjPASS OF F/S GULF COMM 1.00 .506 (29'1 OW COMM/PASS CLIMB 2.75 EXTENSIVE IN 3 6 P STREAM II CRUISE 2.47 REBUILT WING AIRCRAFT, HARDWARE AND DATA AVAILABILITY AND MODIFICATION POTENTIAL This category of evaluation criteria relates to the ability to assemble the components for the testbed aircraft in the early to mid-1980 time frame.

Aircraft Survey A survey of aircraft in the NASA inventory was made in conjunction with the Aircraft Office at NASA Headquarters, Washington, D.C. Out of a total of 110 aircraft, either belonging to, or on loan to the NASA, only 7 were found to be compatible with the design requirements for the Advanced Turboprop Testbed Aircraft. The 7 aircraft are: 0 Lockheed C-141A 0 Lockheed -6 JetStar 0 Boeing KC-135A 0 Boeing 737 Boeing B-52B Convair 990 Gulfstream American Corporation Gulfstream II These aircraft were subjected to an initial screening to establish testbed suit- ability. As the result of this screening, the Lockheed -6 JetStar and the Boeing 737 were eliminated. The data relating to the survey are shown in Table C-I and include the the location of each aircraft, the current or planned configuration, and availability.

Aircraft Availability The availability of each aircraft for testbed service in the mid 1980's was examined as part of the survey. Except for the Boeing B-52B, which has limited application for testbed use, all of the aircraft considered are either engaged in long-term programs or are returning to their parent organizations on comple- tion of the current activities.

Aircraft in the NASA inventory are, therefore, not likely to be available for this program. The Boeing KC-135A and the Gulfstream American Gulfstream II, could be obtained from the USAF and on the used-aircraft market, respectively.

Alternatively, a Boeing 707-120 could be substituted for the KC-135A.

Checks of the used aircraft market indicate that early models of the Boeing 6 6 707 aircraft are available in the price range of $1.0 x 10 to $1.4 x 10 . It is clear from the survey that, unless NASA priori ties change, none of the desired aircraft will be available for the Advanced Turboprop Testbed.

Data Availability The modification of the base aircraft to the testbed configuration will require detailed knowledge of the structural and systems design of the selected testbed aircraft. There is concern that, because of the age of many of the aircraft designs, the data to perform the required modification may be difficult to acquire. During the lifetime of some of the aircraft, changes have occurred which further complicate data acquisition. These changes include change of manufacturing organization, termination of manufacture, extensive modification to later models and type serialization.

EC D-1 Aircraft Data Availability The Candidate Testbed Aircraft will be evaluated for data availability by establishing the degree to which the data are available, degree of cooperation extended to contractor by the appropriate manufacturer, arrangement by which data may be acquired and, in the case that the necessary data are not currently available, the ease with which the data required may be reconstructed.

Data such as basic aerodynamic, propulsion, structural, and aircraft performance as well as control system characteristics and aircraft subsystems information would be required to perform the aircraft modification.

The position, as far as data availability is concerned, ranges from the immediately accessible contractor data for the C-141A to doubtful acquisition of such information in the case of the oldest of the aircraft under consideration, the Boeing KC-135A. The prototype for this aircraft, the 360-80, first flew in May 1954.

Airframe manufacturers are reluctant to share proprietary information with competitors, however, avenues such as U.S. Air Force channels may provide access to the necessary data for the Boeing KC-135A.

In the case of the Convair 990, General Dynamics and the Lockheed-Georgia Company have an agreement of mutual assistance for providing information required for the aircraft modification. General Dynamics has already supplied data for Task III of this study, and further assistance either by data purchase or subcontract participation has been pledged.

The Gulfstream II, originally manufactured by Grumman, is now a product of Gulfstream American Corporation, Savannah, Georgia. Some contact has been made with Gulfstream American and information obtained. There is every indication that further information may be obtained by subcontract or through data purchase.

The availability of data may influence the selection of the testbed aircraft for further study. However, this factor, although important, will not be an overriding element in the evaluation process.

The weighting factor for this evaluation criterion has been set at 2 to prevent the criterion from unduly influencing the final choice of testbed aircraft. This evaluation is shown in Table D-XIX.

TABLE D-XIX. EC D-l CONTRACTOR ACCESS TO AIRCRAFT DATA EVALUATION WEIGHTING VALUE. 2 CONTRACTOR SUBCONTRACT U.S. AIR FORCE TESTIED DATA OR OR RATING CRITERION SCORf DATA PURCHASE OTHER CHANNELS VALUE AIC ACCESS C-141A NA NA J 6 ../ I KC-IJSA x V V ../ CONVAII990 X NA J 6 GULFSTREAM II X NA 2 4 v' The evaluation assumes that the Lockheed-Georgia Company performs the testbed aircraft modification.

Potential for Modification to Research Aircraft Configuration The performance of each testbed aircraft will be evaluated with the available and projected drive systems for potential for modification of the testbed to a research aircraft configuration where all or most of the propulsive thrust is obtained from prop-fan propulsion, since this is an important long- range consideration.

EC E-1 Potential for Modification to Research Aircraft This evaluation is based on the possibility of achieving research aircraft status, with the selected drive system.

The possibility of the candidate testbed aircraft undergoing further modification to a research aircraft configuration, where two or more prop-fan propulsion units provide all or most of the propulsive thrust, is limited by the power of the XT701 drive system. The thrust available from two and four XT701 uni ts and the thrusts required by the candidate testbed aircraft are shown in Figure D-31. These data indicate that the choice of a twin engined research aircraft is limited to the GIl. The GIl also has the advantage of having the prop-fan units as additions so that the aircraft could meet the design speed/altitude requirement with power from the primary engines at the maximum takeoff gross weight of 27210 kgs (60,000 lb). Alternatively, the primary engines could be removed and the speed/altitude requirement could be satisfied at a maximum weight of 25396 kgs (56,000 lb). At 27210 Kgs (60,000 lb), the prop-fan-dedicated GIl could achieve a Mach number of 0.783. The configuration is shown in Figures D-32.

The C-141A, KC-135A and the Convair 990 would fall short of Mach 0.8 at 10668m <35,000 ft) if the two inboard engines were replaced by XT701 prop-fan drive systems. Conversion to a dedicated prop-fan for these candidate testbed aircraft is al so out of the question, as the data of Figul'e D- 31 show. Four XT701 drive systems would produce 39,142N (8800 lb) of thrust, and at the lowest flight weights for the test mission, the C-141A, KC-135A, and Convair 990 all require greater thrust to satisfy the design requirement.

La lI. lO·J klI to SO ~C.'" ,.

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I

r .s.

~77 .00 Figure D-32. Gulfstream II Twin-Engine Testbed Overwing Configuration To achieve the performance level required for a four prop-fan propulsion system arrangement, the power level required would be as shown in Table O-XX.

An improvement in the modification potential of the three aircraft listed in Table O-XX could result if the design philosophy is changed from that of propulsion system substitution to one of addition as in the case of the GIl.

This has the advantage that the primary propulsion is retained and therefore the aircraft, when modified for the addition of one or two prop-fan units located on TABLE D-XX. POWER LEVEL REQUIRED r XT701 GROWTH XT7XX PROP-FAN DIA.

TESTBED A/c

I

kW (SHP) FACTOR (SHP) kW m (FT)

I

--

,

~

6019 (8071) 1.8,0 10,835 (14,530) 4.08 (13.4) C-141A ;

I

i (12.1) 3.69 KC-135A 6019 (8071) 1.47 8,844 (11,860) \ 8,426 (11,300) 3.60 (11. 8) 6019 (8071) CONVAIR 990 1.40

I

the wing between the fuselage and the inboard engine, do not suffer significant performance degradation.

The twin-engined testbed configurations for the C-141A, KC-135A, and Convair 990 are shown on Figures D-33, 0-34, and 0-35. All are overwing installations so that clearances are maximum. Of the three arrangements, the KC-135A appears to be the best since the inboard engine is so far out on the wing, 11 = 0.41, that ample clearance between the prop-fan and the fuselage and engine nacelles exists. These twin-engine testbed configurations do not fulfill the previously defined "research aircraft configuration" role in that the prop- fan units do not provide a significant portion of the total required propulsive thrust. However, they would provide additional valuable accoustic data relating to multiple sources and their interactions. Therefore, a twin-engine testbed of this form might prove highly desirable.

The potential for modification evaluation takes into account the change in design philosophy which is reflected in the rating values of Table D-XXI.

------~, ',~,~~~--~------- -af--~,~~"l-' II ~---~ S.L.

-'E~~m- 182.00 W.L.

l13.m I j Figure D-33. C-141A Twin-Engine Testbed Overwing Configuration Figure D-34. KC-l35A Twin-Engine Testbed Overwing Configuration

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S.L. -iE:~~1=- 266.00 X"I701 ORIVE S!ST!lI Figure D-35. Convair 990 Twin-Engine Testbed Overwing Configuration TABLE D-XXI.

EC E-l POTENTIAL FOR MODIFICATION TO RESEARCH AIRCRAFT EVALUATION WEIGHTING FACTOR = 2 RATING TESTBED AIC VALUE CRITERION SCORE C-141A 1 KC-135A CONVAIR 990 1 GULFSTREAM II 3 6 RELATIVE COST OF TESTBED SYSTEMS The ROM cost data for modification to the testbed configuration derived for Task III will be reviewed and will form the basis for cost comparison of thp.

candidate testbed aircraft.

EC F-l Modification Cost Data Ranking This evaluation will be based on the comparison of the ROM costs to modify each testbed aircraft and will include identi- fication of the cost drivers.

ROM cost data estimates have been prepared for each candidate testbed aircraft. These data include the following: Structure and Systems o New nacelle structure o Engine/prop-fan controls modification o Surface controls modification o Fuel system changes o Flap and spoiler modification for prop-fan loads and temperature effects o Wing Structure changes for engine QEC pick-up and resulting spar, cover and rib changes o System changes for hydraulic electric and aircraft systems affected by the deletion of a primary engine Engineering and Test o Design of structure and systems o Design support, i.e., structures, aerodynamics, propulsion, flutter, and vibration o Ground test of components and installation on aircraft o Modification of the aircraft to the testbed aircraft configuration The ranking of the cost data is shown in Table D-XXII. The data for the C-141A, KC-135A, and Convair 990 are all of the same order. Aircraft size and amount of modification required are similar. Extensive structural modification for the wings is not required, since the prop-fan installation is located. in the same place as the primary engine. The modification to the GIl is, however, much This greater, since the prop-fan installation is added to the wing.

necessi tates extensive rework of the structure of the wing inboard of the prop-fan installation. This fact is reflected in the ROM cost for the GIl, which has the highest cost of the four testbed configurations.

The cost data of Table D-XXIl do not include the cost of modifying the drive system gearbox. DDA estimates this cost to be in excess of $400,000.

The cost data have been estimated on the basis that the power section and unmodified gearbox are government-furnished equipment.

All dollar values are in 1980 dollars.

TABLE D-XXII. EC F-l MODIFICATION COST DATA RANKING WEIGHTING FACTOR = J TESTBED ROM COST RATING CRITERIA SCORE AIRCRAFT ESTIMATE $ X 10- VALUE C-141A II .7 9 J KC-IJ5A II .7 CONVAIR 990 11.8 GULFSTREAM II 12.5 2 6 CANDIDATE TESTBED AIRCRAFT EVALUATION AND RECOMMENDATIONS The data from each of the evaluation criteria have been consolidated in Table D-XXIII and the total weighted score computed. This evaluation shows that the C-141A is not a suitable candidate for the advanced turboprop system, mainly because of the marginal performance at the design conditions. The speed and altitude increments beyond Mach 0.8 and at 9144m (30,000 ft) do not provide sufficient flexibility for test purposes or to accommodate increases in aircraft drag should more refined analyses show this to be the case.

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The possible candidate testbed aircraft are. therefore, as follows: Weighted Score 122 Gul fstream II Weighted Score OW 118 Boeing KC-135A UW 118 Weighted Score OW 118 Convair 990 UW 118 Estimation of the score subtotals for these aircraft in the categories of Aircraft Safety. Operational Characteristics, and Testbed Program Objectives Achievements show some interesting results, shown on Table D-XXIV. The scores shown are for overwing installations, since these are considered to be most representative of commercial aircraft application.

From the testbed aircraft safety standpoint, there is a little difference between the four candidates; all can be operated safely as testbed aircraft. On the basis of operational characteristics as testbed vehicles, the KC-135A and Convair 990 are best, offering the most stable platforms with greatest per- formance margins. The GIl is not quite as good, primarily because of the re- quirement for ballasting to maintain balance.

In the area of meeting testbed objectives, however, the GIl clearly emerges as the best candidate aircraft with a score in that category high enough to make it the best overall candidate.

The second choice is between the KC-135A and the Convair 990 for which the evaluation total scores are identical.

The weighted score subtotals for the evaluation criteria of categories A, Band C of the evaluation are also identical as shown on Table D-XXIV, Testbed Final Selection. These scores re- late to single prop-fan testbed configurations only. If, however, the potential for modification to a multi-prop fan arrangement is a consideration, any in- herent advantage in one configuration may be of significance. This is shown to be the case, since scrutiny of the configurations, Figures D-34 and D-35, shows that conversion can be accomplished on the KC-135A without infringing on the important clearance parameters, since the inboard engine is located at 41 percent of the wing semi-span against 36 percent for the Convair 990. When the weighted score for this criteria is considered as shown on Table D-XXIV a clear choice is possible for the second testbAd aircraft, since the KC-135A has a total of 107 against the 103 for the Convair 990. It should be noted that the GIl continues to have the highest score.

TABLE D-XXIV. TESTBED FINAL SELECTION SUBTOTAL SCORES SUB- MOD.

OPERATIONAL PROGRAM A/C TESTBED A/C TOTAL POTENTIAL TOTAL SAFETY CHARACTERISTICS OBJECTIVES C-141A 22 31 28 81 2 83 KC-13SA 22 37 42 101 6

e

CONVAIR 990 22 37 42 101 2 103 GIl 23 33 50 105 6 ~ TESTBED SELECTED A!C The following are the recommendations for the testbed aircraft and for the study Task V activities: (1) Recommended Drive System o Detroit Diesel Allison XT101/T56-A-14 Gearbox (2) Propulsion System Configuration o Overwing Installation o Universal QEC Design (3) Recommended Testbed Aircraft o Boeing KC-135A o Gulfstream American Gulfstream II TEST PLAN AND INSTRUMENTATION RECOMMENDATIONS This recommended test plan is submitted in response to Task IV of the Statement of Work for the "Advanced Turboprop Testbed Systems Study." The test plan will be expanded and/or modified as required in Task VII.

Test Article The test article will be fitted with the recommended drive system: Detroit Diesel Allison XT701/T56-A-14 gearbox with a universal, QEC-type, overwing installation. The recommended testbed aircraft, either the Gulfstream American Gul fstream II or the Boeing KC-135A, is considered as the vehicle for the test program.

Approach Contractor flight test personnel will conduct the instrumentation system installation during the modification span for powerplant installation. The instrumentation system design and installation will be supplied by the Lockheed-Georgia Company to include the recording system, wiring, transducers, and support equipment for all measurements except those required by Hamil ton Standard for the test propeller and DDA for the propeller drive system. The Hamil ton Standard recording system will be installed by the Contractor and necessary wiring incorporated from the propeller to the recording system. The associated slip rings and engine wiring for the propeller and propeller shaft instrumentation will be supplied by Hamilton Standard/DDA.

All instrumentation recording systems will be installed in the passenger/cargo compartment of the test vehicle. The test instrumentation will be maintained by Contractor personnel, except for the propeller and propeller shaft instrumentation systems.

It is assumed that the test engine{s)/gearbox will be fully qualified and will be received as calibrated units for test purposes.

Objectives The objectives of the flight test program are to assure the airworthiness of the installed prop-fan test system and to provide data to verify the goals of the program. Tests will be conducted for data acquisition to: o Verify the test system/airframe airworthiness o Evaluate prop-fan control system function o Evaluate propeller and propeller shaft structural integrity and dynamics o Determine airframe dynamic and vibratory characteristics induced by the propeller o Evaluate cabin noise levels and the benefit of additional cabin acoustic treatments o Evaluate near- and far-field noise levels o Evaluate engine inlet performance o Substantiate scale effects Test Program After instrumenting the aircraft and the test drive system, the testbed will be prepared for flight and enter a ground test phase.

Ground Tests The ground tests to be conducted will include the ground vibration tests of the airframe structure, propeller shaft and blade stress testing, noise evalua- tion, propeller control tests, and aircraft ground control tests.

Flight Tests The flight test phase will involve evaluation of propeller shaft and blade structural dynamics characteristics, airframe flutter characteristics, airframe/ test system airworthiness, near- and far-field noise, engine inlet performance, and propeller control system operation.

Instrumentation Requirements The recommended instrumentation requirements will be developed to support the program objectives. The following instrumentation groupings are estimated to provide the data required to support the test program.

Parameter/Instrumentation Quantity (est) Location 0 Noise (Microphones) Fuselage, Wing, Empennage 0 Accelerometers Wing, Empennage, Nacelle 0 Surface Pressures 40 Wing 0 Wing Pressure Rake 2 0 Wake Rake Wing Engine Inlet 0 Engine Inlet Rake 0 Basic Engine Parameters 6 Engine o Basic Aircraft Parameters 16 Motion, control position, airspeed/altitude, accelerations Propeller o Propeller Strains o Propeller Shaft Strains Propeller Shaft o Engine Acceleration Engine o Engine Pressures and Temperatures Engine Typical instrumentation is shown in Figure D-36.

UPPEl{ SURFACE ROW LOCATIONS 145 250290 120 ' 170 210 I 1 1001 ' I

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I I I I I .305 m (12 IN) 0.305 m (12

IN )-11- AFT OF WING] r:-

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ROWS OF STATIC T RAKES

PRESSURE ORIFICES

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WAKE RAKE Figure D-36. Instrumentation for Gil Testbed Aircraft

APPENDIX E - CONCEPTUAL DESIGN OF TESTBED SYSTEMS - TASK V· At the conclusion of Task IV, the evaluation showed the candidate testbed aircraft configurations to be the Boeing KC-135A, the GAC GII, and the Convair 990. The two testbed systems recommended for further study were selected from the three candidate systems and identified as the Boeing KC-135A and the GAC GII. Although the KC-135A is a military aircraft configuration, the commercial counterpart is the Boeing 707-100 series, which was derived from the KC-135A.

The Boeing 707-100 can, therefore, be substituted without changing any of the findings of the evaluation and without change to the recommendations. These testbed aircraft systems were recommended in conjunction with the DDA XT701 drive system powering a 2.89 m (9.5 ft) diameter prop-fan. Following the Task IV "Evaluation and Recommendations," the NASA Lewis Research Center directed that single prop-fan aircraft designs should be discontinued and that the Task V "Conceptual Design of Testbed Systems" should proceed with twin prop-fan configurations.

The overwing "Pinion-High" installation was chosen as the drive system most representative of future aircraft applications, and the drive system/airframe integration was performed without attempting to optimize the arrangement aerodynamically.

Gloves and fillets at the wing/nacelle intersections are contemplated, however, to obtain an efficient installation.

The change to the twin prop-fan design did not affect the choice of candidate testbed aircraft, as the potential for modification to multi-prop-fan arrangements was an evaluation criterion in Task IV. Selection of the candidates, therefore, included this consideration. The effect on the GII was to add a second prop-fan QEC unit to the left-hand wing of the aircraft. In the case of the KC-135A, however, the effect was to change the design approach from a prop-fan substitution to a prop-fan addition. Since the inboard pr imary engine is located at 43 percent of the semi-span, no difficulty was encountered in positioning the prop-fan units on the wings, between the inboard engines and the fuselage, to give the recommended 0.8 Dp from prop-fan tip to fuselage wall, and 0.2 D from prop-fan tip to engine nacelle clearances.

p Detailed conceptual designs were completed for each of the recommended testbed candidates to further confirm the suitability and adaptability of each system to the flight test program. This design effort was aided by the loan of design and technical data to the Lockheed-Georgia Company by GAC and by a review of the Lockheed design by GAC for feasibility and practicality. Data for the KC-135A were obtained from the public domain through Wright-Patterson Air Force Base.

The following text describes this design process which is divided into three distinct sections. First, a description of the quick engine change (QEC) nacelle design, including rationale for the selection of the basic nacelle contours, engine air inlet design, nacelle structural design, and drive system installation is given. Second, the KC-135A testbed system design is reviewed covering the drive system location and geometry, the aft nacelle structure, a flutter analysis, the testbed operating envelope, the testbed performance, and a summary of the KC-135A testbed weights and balance. Finally, the testbed system design utilizing the GAC GIl is covered including the same design details as for the KC-135A system, and additional details concerning trim capability, required wing modifications, estimates of the prop-fan slipstream characteristics, and estimates on near-field noise characteristics.

QEC NACELLE DESIGN CONSIDERATIONS The QEC nacelle envisioned for the prop-fan testbed was designed to contain the drive system and its associated support systems and structures, and to duplicate, as nearly as possible, the experimentally derived flow field through the prop-fan, in an attempt to validate propulsive efficiency gains theoretically possible from this propulsion system.

Nacelle Contours The XT701 nacelle contours were designed to provide the same envelope for "pinion-high" and "pinion-low" drive system configurations. Since the overwing installation was chosen for both conceptual designs, the development of the nacelle envelope for the "pinion-high" arrangement only will be addressed.

The nacelle contours were based on the NASA spinner/hub area distribution, Figure B-6, and are arranged to permit the use of the main forged support frames and supporting V-frames from the Lockheed P-3C T56 engine installation, modified for the DDA XT701 drive system installation.

The nacelle contours, Figure E-1 are arranged to provide an envelope for the drive system with air induction systems for the engine and oil cooler arranged on the upper portion of the nacelle in a stacked and staggered configuration. The oil cooler inlet and ducting are designed to house the C-130 - T56 oil cooler.

Engine Air Inlet Design A scoop type inlet was selected for the XT701/T56A-14 engine/gearbox arrangement, as shown in Figure E-1. The general design philosophy adopted was that the engine should perform reliably and efficiently over the range of test condi tions at the expense of drag minimization. Thus, in choosing between efficient internal or external flow performance, internal performance was considered more important.

Turboprop-powered aircraft have not heretofore been designed to cruise at Mach numbers higher than 0.6. Consequently, large inlet areas have been used, resulting in a contraction in the duct between inlet and engine compressor face.

At Mach 0.5 or 0.6, this results in moderate flow spillage around the inlet lips, and insignificant spillage drag. However, at Mach 0.8, spillage drag may be significant. From the drag standpoint, therefore, it is desirable to keep the inlet area as small as possible. A small inlet, however, may result in excessive internal flow pressure losses and flow distortion at the engine compressor face. For the testbed nacelle, the inlet area was selected to be equal to the compressor face area - a compromise intended to provide good internal flow without excessive drag.

The shape of the engine air inlet, as shown in the front view of Figure E-1, was selected to minimize departures from symmetry about the nacelle axis.

This results in a high-aspect-ratio inlet shape or one in which the inlet encircles a large portion of the upper half of the nacelle. Use of this arrangement was based on: (1) knowledge that large obstructions behind the prop-fan would induce 1-P dynamic loads on the prop-fan, and (2) the premise that since the forebody design criteria were based on tests of bodies-of- revolution, the design risk is minimized by using nacelle shapes which approach symmetry as closely as possible.

A third inlet design consideration concerned the fore-and-aft location of the inlet. In the case of the testbed, the inlet was located farther aft than .

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II Figure E-l. XTlOl QEC Nacelle Contours (Cont'd) on existing turboprop installations in order to maintain adequate airflow over a wide range of prop-fan blade angles.

As shown in Figure E-1 the oil cooler for the testbed is located in a stacked and staggered position relative to the engine inlet S-duct. The con- figuration selected uses a fixed lip inlet and variable geometry (flapped) exhaust for the cooling air flow. If necessary, the air flow can be augmented at low forward speeds, ground idle, and other critical conditions by an engine- bleed-powered ejector pump downstream of the heat exchanger.

NACELLE STRUCTURAL DESIGN The externally applied loads for the drive system nacelle design were derived from flight envelope data for the KC-135A and for the GII. The external limit loads data, Table E-I, include positive and negative vertical accelera- tions, positive and negative lateral accelerations, torque, and shear loadings.

Since the external loads, Table E-I, are common to both testbed configurations, the internal loads and sizes of structural members are independent of the recei ving airframe. This results in a common structure in the QEC up to the mating plane. The structure on the receiving airframe, from the mating plane aft, is designed to be compatible with the QEC structure. The nacelle shapes are arranged to facilitate manufacture and consist of a body-of-revolution for the spinner/hub region, changing to an upper and lower radius joined by straight sides over the remaining portion of the nacelle. The nacelle is 3.56m (140.0 in) long, 1.43m (56.4 in) deep, and 1.01m (40.0 in.) wide. The cross-sectional area is 1.32 sq m (14.2 sq ft) and the wetted area is 11 sq m (122 sq ft).

A finite-element analysis was performed to establish the sizes of the structural members of the nacelle, to check the capability of the P-3C members to be used in the design and to provide data for weight estimates of the nacelle and testbed aircraft. The respresentations of the nacelle structure are given in Figure E-2, which is the engine facsimile; Figure E-3 the representation of the longerons and truss members; Figure E-4, which shows the nacelle structure shear panels; and Figures E-5 and E-6, the representations of the main support and mating plane frames, respectively. The results of the analysis are shown in Tables E-II and E-III for the axial elements and shear panels, respectively.

TABLE E-I. NACELLE - EXTERNAL LIMIT LOADS • lOAC CASE -5.0g (LIMIT) .

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91 29 1109 I 29 59 .3765 31 59 .1933 Figure E-3. FEA Longerons and Truss Members FORWARD MAIN FRAME THICKNESS SHEAR PANEL 29 .0400 33 .0400 Figure E-4. FEA Shear Panels " t.

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(CONT'D) FEA AXIAL ELEMENTS TABLE E-II.

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Be 1. C:lO .1~:JC .ula

1'+2 '4~ 7:; 117 ~.'ion • -,] C • I"!"''!

1~4 .1::1CO ae So 1.00D .C1a 7q lite: 81 '.rOD .C10

·cno

1 '+6 a: sa .1JCa

1.000 .::10 1147 l.OOO 19 89 .1000 .ClC He 79 91 1.r'!OJ .1C::~ .010

-,-

03 l,l"'Jro .'~I'1[:; 149 fd ebC.L d 1St Sl 92 1.000 .1:1:0 .CI0 1 :: , 81 .1~"'C 94 1.000 .01 C • c:- S9 .247 61J 24.COO .1C:~

~~ "

I:: .., Q .246 • ;j8 c: f,e .l-~O .1COa .'+3!

15" 59 95 42.:48 ]55 9f= 97 1.qCO .1cr;O .010 .1OuG ISf- 95 98 1.COn .tHO

_.-

1 • ,., no .1;"!C~ r!· I ' 15 ; 96 '01 lSE l.O:!!] .1~a(j 9~ 1('2 .01: '59 96 Leco .01 a • 10,,0 9~ l~Q 1.CCJ .100 !b~ u .CIO 1 ~ 1 (;4 O , ~88 f;0 37.6 "f ,1000 : 0, .388 :59 63 37.693 .1:aO f, 1 0 .1,. .... ::1 F 3 ."'00 ."02

iCTAl .. Ertan or nIH. r:lt .. t NTS 1':"'.96: POUNCS

••••••••••• ----------- TABLE E-III. FEA SHEAR PANELS --------- ~'ODE ,.

HE~8E2 ARe::A NODEl NOOEZ NODE'I ., Z6 1 7

25 28.625 .;:~ aa :fla

a 25

...

3 .r , c: 6·"'cn .0 aCi"' Z 4 9 6 e.:CO .c~O;] .C2: I: . 1 11 9 3J.J4~ ·1,11

6 a 12 10 3J.~CO .12"

.'~"C: () 11 15 13 27.230 • 11 ,

7 .:"GJ

.3 lZ 16 l~ 27.230 .112 t; 11 11 26.596 9 .1 l' Ii:: 16 2J 1a 26.596 .Co3~ .17" 1" 19 2, 21 b.~S~ .1733 .114 1: 18 20 24 22 6.:5~ .11Z :9 17 29 23 42.~5~ .-40, .ll!

14 ,1 2~ 24 42.:50 .:4CJ .17!

1~ 53 5u 3S 24.253 .:"OJ

.lce

~3 54 36 24.2~e .J4QJ If: 32 4 3 17 31 35 33 4.:uO .10

.,-

J • 32 35 34 4.:~C .1029 .a~2 33 ;5 39 37 Z4.aery .:4:~ .li'

. ~

34 J6 ~a 38 24.8~J .e"ca ......

. .

3 ; ~ 9 It 'P 9 1 2 ~ • -:; ~ , a 4.C...,O"-- ____ ......... _- .... ::I_~ __

4+ 38 4J 44 ~2 2G.:~J .~"a~ .aaZ 41 43 9 7 95 21.<;-, ,"4C; .Li 9 7 2~ 42 44 4~ 46 23.5u: .~4CJ .C97 ________ ~2~5 _________ 4~S ____ ~4~1~ ____ ~5~1 ____ ~4~9~ ______ ~4~.~;~G~J ________ ~.~ca .016 .~l~ 26 46 48 52 5: 4.:0~ .a401 ,7 4~ 1:5 56 :1 34.712 .:40J .14!

:~ 48 55 56 52 34.112 .C40J 21 49 29 l~aG.!,! .~4C::: - ,.

:. .. 2 22 5~ 3~ 153D.!Zl .:4C~ 6.511

31 49 61 59 1358.079 .~4C~ :.5;5 5C 6~ 6: IJ58.:79 .:~CJ 5.595 Z2 5~ ~9 4599.:C~ .C4CJ 6.558 .

. 3~ 29 114S.:CJ .:4CO

35 49 67 61 144T.~cr .ruGa

.. :.ooc~ 536 • .3C6

59 59 521.111 9 • :"j!t 18 Z .4 2:: 6? 6 562·531 3<3 l6 S6 ,:42::1 68 7: 562.531 39 67 69 71 31 189.908 .C4bY .:::465 .9 :C 68 70 72 32 189.9~8 15 73 7~ 69 453.~96 1.B b 7 .Clte" '+2 68· 67 614.4u~ .:4C::: 2.779

43 75 71 194.274 .C 00 .3 ce

1.2 U: f:7 32 31 !09.601 oS • e ;.8 ;4 19 b§ 65. 689.559

. -

72 71 389.664 .alto;:] 1.6:'5 • -44') 7 74 73 19U.B32 .s a ...

74 73 194.332 .:883 1.771 48 68 7C ,+0 .884 67 69 74 73 194~a3Z ."'44'1 ~ ..

76 14 194.214 .8 ::C

- -

T 7Q 70 4S .-95 l·a~"7 7: 7!: 12 194.214 .B C t 6C5.923 POUNDS TOTAL -EIGHT OF SHEAR PANEL ELEMENTS ••••• Nacelle Structural Details The nacelle structure, Figure E-7, consists of a drive system suspension and mounting, and an aluminum-alloy envelope supported by frames and longerons.

rps 0.00

PPS 24.50 PPS 77.70 PPS 101.15

I

PPS 91.40 I 'i

I : ----- 1---- -tl~----::l : : I

'I I i / I I I I

I I I I I I I

: :1" : : " I I I I

---i1~----ilL ----ll-~~-- L---I -j--- I --- -J- -~'~ 1~~:;::':~J

- PROP- FAN HUB

--r' ,---oil r-:=~"::'l-t"~:S~~~: ~~;:71

II I .- __ "- I I II / / I

\,~.-'~ :

I ~' I-----~~, + --' ,-- l WELDED JOINT: .ir:/ I

r --.. I I I -I?, I

~- I

-'--- 1-.. \., I I /1}1 ! i

( I : :---- ,I : ,,~.~ # ENGINE ACCESS PANEL -0.04?, AI;~ SHEET I

I

I I "- .. " ...... ~I I' // :111 I

'-. ....... I ~ // I _:._ I ---'''''., . II - - I --/~r- PPWL 120.40 I

+. i

---- : l' "'-. I /~/ : ,:: I

'II I .'-..... '-." I </ : : I

! : !

: II A ~ I ,.'" I // I ': I

r

I I' I ________ !__ ~-------~ -:--"~IJ :.'~--::.::-=--~==--:-.:... ....::-::.~~

1"- ~.::.=.:- --.::=:.-- _==~..,r.~ .. ~~JO-.r_:.::-..: - - ----fir :

L__ ,\

I r t ,------ ------ F----- ----,-:':;1)/1-- I: ; . t i

I

--.........Jl

:' , : I '''''\if III I I I I Ijl • • : A FRAMES MODIFIED P3C MODIFIED P3C V-FRAME ASSEMBLY \ \

\

SECTION 'AA I \.

)

\ I ..

ENGINE ACCESS PANEL / \ LOWER LONGERON·

-'. ( , ( )\

.,~ __ ,./ .:1,. /1

LOWER PANEL FRAME r-----Of~

LOWER PANEL SKIN / I

.,1

CENTER STIFFENER' PPBL 0.00 Figure E-7. Nacelle Structure XT701 Mounting System - The mounting system for the XT701 is similar to that of the T56 installation in the Lockheed P-3C and consists of a suspension system and the supporting truss, longeron, and frame members.

The suspension system consists of seven mountings located and identified as follows: Two (2) side front mountings, Lord PIN LM-204-5A19 Two (2) top front mountings, Lord PIN LM-204-5A28 One (1) bottom front mounting, Lord PIN LM-204-5A21 One (1) top rear mounting, Lord PIN LM-204-5A8 One (1) side rear mounting, Lord PIN LM-204-5A30 These mountings provide restraint in pitch, yaw, and torque and have been analyzed using the XT701/T56-A-14 limitations of 4175 kW (5600 shp) at 1600 rpm and a maximum torque of 2540 m-kg (220,500 in.-lb).

The analysis shows the P-3C suspension system to be acceptable for testbed aircraft application with a limit of 300 flight hours. A flight program beyond 300 hours will require further analysis to establish mounting suitability. The suspension system locations are shown on Figure E-8. The main mounts on each TOP FRONT MOUNTS - 2 - LORD LM-204-SA28 TOP REAR MOUNT - 1 - LORD LM-204-SAB SIDE REAR MOUNT - 1 - LORD LM-204-SA30 SIDE MOUNTS - 2 - LORD LM-204-SA19 BOTTOM FRONT MOUNT - 1 - LORD LM-204-SA21 Figure E-8. DDA XT701 Suspension System Mount Location side of the gearbox react loads in all three directions. The top and bottom front mountings react fore and aft loads. The bottom mount is also designed to react vertical load in the event of a main mount failure. The aft upper mount and the side mount on the rear casing of the engine are designed for vertical and lateral loads respectively. The mounting structure consists of two forged frame members adjacent to the gear-box mounts, fore and aft V-frames, aft diagonals, and upper and lower longerons as shown in Figure E-1.

The V-frame members are fabricated from the P-3C nacelle V-frame part number 918829-1, parts as follows: The upper tube Part No. 839411, made from 8630 steel tubing heat treated to 862,OOOkPa - 655,000 kPa at the flashwelds (125,000 psi - 95,000 psi at the flashwelds), is extended at the aft end by flash welding a similar piece of 8630 steel to form the lengthened upper tube of the truss. The aft upper fitting is a new part fabricated from 4340 steel welded to the upper tube. The new part is required because of the change in the angle of the faying surfaces at the mating plane. The forward fittings, also flashwelded to the upper tube, require new parts for the same reason.

The lower diagonal of the V-Frame, part number 839428, is used directly from the P-3C structure. This part is fabricated from 8630 steel tubing, heat-treated to 862,000 kPa - 655,000 kPa at the flashwelds (125,000 psi - 95,000 psi at the flashwelds) and is tapered from each end toward the center, with areas of 3.2 sq cm (0.495 sq in) at the ends and 3.32 sq cm (0.515 sq in) at the center.

The lower attachment is a new part fabricated from 4340 steel.

Al uminum alloy, built-up aft diagonal members are connected to the V-Frame at the upper and lower ends. The outer flanges of these members are also attached to the nacelle outer skin. The lower longeron extension is also connected to the rear fitting of the lower diagonal and to the nacelle skin.

The forward support frames are manufactured from the P-3C forgings used for part numbers 918459-1 and 918468-1. The upper portion of each frame is modified to accommodate the engine air inlet for the XT701 engine, and the lower portion requires changes in the flange bevel angles for compatibility with the "area ruled" spinner/hub.

The nacelle structure extends forward from PPS 0.0, the mating plane location, to PPS 101, the hub plane. The nacelle forebody consists of a formed, aluminum alloy upper portion, 1.0 mm (0.04 in) thick between PPS 77.7 and PPS 101.15 and includes the engine inlet lip. The lower portion is also a formed aluminum alloy structure 1.0 mm (0.04 in) thick. Channel-section frames are located at intermediate stations to support the forebody structure. Between PPS 0.0 and PPS 77.70, the nacelle structure is an arrangement of shear panels, longerons and frames. Frames are located at approximately 25.4 cm (10.0 in) spacing and consist of aluminum alloy channel sections which, with the upper and lower longerons, form the basic skeletal structure of the nacelle. The upper and lower shear panel skins are aluminum alloy 1. 0 mm (0.04 in) and 2.54 mm (0.10 in) thick. The longerons are built-up from aluminum alloy extrusions and sheet and have cross-sectional areas of 0.65 sq cm (0.10 sq in) for the upper and 0.8 sq cm (0.12 sq in) for the lower. The lower longeron extension, from the aft portion of the V-Frame lower diagonal, is reduced to an area of 3.7 sq cm. (0.57 sq in) and is also fabricated from aluminum alloy. The aluminum alloy diagonal member, extending upward from the lower fitting of the V-frame lower diagonal to the V-frame upper tube joint, at the mating plane bulkhead, is a buil t-up structure having a total cross-sectional area of 3.22 sq cm (0.5 sq in).

The side panels between PPS 0.0 and PPS 77.70 and the upper and lower longerons are aluminum alloy 1.0 mm (0.04 in) thick. Portions of the side panels are removable for access to the engine. An S-duct, located between PPS 77.7 and the engine compressor casing, is fabricated from stainless steel sheet supported by external rings.

The upper shear panel is configured to accommodate the ducting for the oil cooler and its inlet and exhaust. A controllable flap is provided on the exit duct to control the cooling air mass flow as required.

The mating plane bulkhead located at PPS 0.0 also serves as the fire barrier and is fabricated from 1.0 mm (0.04 in) thick titanium sheet, for the web, and aluminum alloy extruded sections at the inner and outer boundaries.

The main attachment points for the engine nacelle to the airframe aft nacelle are located on the bulkhead at the ends of the upper tubes and lower longeron extensions.

DRIVE SYSTEM INSTALLATION The drive system is installed in the nacelle. together with those accessories and systems necessary to operate the prop-fan unit. as shown on Figure E-9. A modified 54H60 control unit is used for prop-fan control and is located at the rear of the prop-fan hub. The engine fuel control is a hydro- mechanical device having an electronic supervisory system. The engine starting system uses air bled from the primary engines. conducted to an Ai Research Starter No. ATS100-397. located on the underside of the XT701 compressor case.

Fuel and air line disconnects are provided on the mating bulkhead for the QEC.

The oil cooling system uses the heat-exchanger from the C-130 T56 installation.

A new oil tank is located below the torquemeter immediately behind the gearbox.

I{EAR MOUNT

~]-::--=--/ .---

LLl----

-li--_t__--+f PI' S'I'A n.40 V-~'I{AMf: ASSY. LOWEI{ PANEL ASSY PI' !iTA 140 LOWER LONGERON I'P S'J'A 80.00 1'1' S'l'A 0.00 1'1' !iTA IOL. L50 Figure E-9. Drive System Installation BOEING KC-135A TESTBED SYSTEM CONCEPTUAL DESIGN The USAF series KC-135A aircraft is a high-performance, jet-propelled, tanker-transport, low-wing aircraft. The C-135A and the Boeing 707-100 series are outwardly identical to the KC-135A, except for the removal of the aerial refueling boom. The KC-135A can, therefore, be regarded as a reasonable representation of a commercial aircraft configuration for the purpose of the conceptual design. The main differences between the military tanker-transport and commercial versions are the lack of windows, cargo stressed floor, side cargo door, and the lack of commercial cabin furnishings and trim. These are not, however, considered to significantly affect the KC-135A suitability as a vehicle for the test of acoustic attenuation concepts for near-field acoustic tests. The principal dimensions and characteristics for the C-135A, KC-135A and 707 -100 series airplanes are given on Table E-IV. The KC-135A configured as a twin prop-fan testbed aircraft is shown on Figure E-10 for the "Pinion-high" overwing installation.

Figure E-lO. KC-J35A Twin Prop-Fan Testbed Configuration Drive System Location and Geometry The inboard wing scantlings and sections of Figure E-11 were used to locate the prop-fan drive system installation and for the development of the wing sections adjacent to the installation. The drive systems are "pinion-high" overwing installations located at WBL 217 LH and RH, with each installation vertical plane normal to the wing chord plane. Each nacelle is placed with the prop-fan centerline located at WL 217.995, which provides adequate clearance between the wing upper cover and the jet exhaust pipe. The nacelle installation geometry is shown on Figure E-12. Overall, the length of the installation from the spinner tip to the end of the jet pipe is 9.2 m (362 in) and is a maximum of 1.6 m (62 in) wide. The height of the nacelle above the wing chord plane is 1.04 m (40.8 in).

" I \. no?·rA..": DRI\'£ sysre: .'t~ ':17 Figure E-ll. KC-l35A Inboard Wing Scantlings KC-135A Aft Nacelle Structure The aft nacelle structure consists of a skin-frame-Iongeron structure extending aft from the nacelle mating plane at FS 707.56 to approxim<ltely FS TABLE E-IV. KC-l35A PRII'CIPAL DIMENSIONS AND CHARACTERISTICS KC-13SA C-13SA 707-100 RAMP :./EIGHT Kg (IJI) 136,926 (301,600) 126,893 (279,500) 117,132 (258,000) OPERATING :./EIGHT Kg (LB) 117,132 (258,000) 65,494 ( 99,359) 49,801 (109,695) ( 90,300) FUEL WEIGHT Kg (LB) 92,071 (202,800) 85,656 (188,670) 40,996 WING SPAN M (FT) 39.87 (130.83) 39.87 (130.83) 39.87 (130.83) OVERALL LENGTH lo( (FT) 4L 56. (136.25) 4L02 (134.5) 41.20 (135.083) !lEIGH! lo( (FT) 12.7 (4L67) 12.7 (4167) 12.7 (4L67) wnlG AREA SQ lo( (SQ FT) 226 (:!433) 226 (2433) 226 (2433) ROOT CHORD lo( (INS) 3.58 (337.98) 8.58 (337.98) (337.98) '!:I!' CHORD M (INS) 2.84 (112) 2.B4 (112) 2.84 (112) MAC ~ (INS) 6.14 (241.88) 6.14 (24L85) 6.14 (24L88) tic ,~ WBL 70.5 15.6% 15.6% 15.6% tic @ WBL 360 9% 9% 9% tIc @ WBL 780 9% 9% 9% r::CIDENCE RAnS (DEGS) 0.035 (2) 0.035 (2) 0.035 (2) DIHEDRAL RAnS (DEGS) (7) (7) 0.122 0.122 0.122 (7) SWEEP ;~ C RAnS (gEGS) 0.61 (35) 0.61 (35) 0.61 (35) ASPECT RATIO 7.065 7.065 7.065 HORIZONTAL STABILIZER 46.51 AREA SQ M (SQ FT) 46.51 (500) (500) 46.51 (500) SPAN lo( (FT) (39.7) 12.11 (39.7) 12.11 11.50 (37.7) ROOT CHORD M (INS) 5.28 (20B) 5.28 (208) 5.28 (208) TIP CHORD M (INS) 2.41 (95.05) 2.41 (95.05) 2.41 (95.05) MAC M (INS) 3.99 (157) 3.99 (157) 3.99 (157) ASPECT RATIO 3.2 3.2 3.2 VOLllHE COErP V .62 .62 .62 H VERTICAL STABILIZER AREA SQ M (SQ FT) 30.5 (328.3) 30.5 (328.3} 30.5 (328.3} SPA.,{ M (FT) 7.53 (24.7) 7.53 (24.7) 7.53 (24.71 ROOT CHORD M (L.'1S) 6.15 (,242) (242) 6.15 6.15 (242) TIP CHORD M (L.'~S) 2.21 (86.92) (86.n) 2.21 2.21 (86.92) MAC ~ (L.'lS) 4.44 (174.6) 4.44 (174.6) 4.44 (174.6) ASPECT RATIO (1.8) VOLllHE COEF'F Vv .064 .064 .064 FUSELAGE MAX WIDTH l{ (FT) 3.66 (12) 3.66 (12) 3.66 (12) MAX !lEIGH! M (FT) 3.44 (17.83) 5.44 (17.83) 5.':'4 (17.83) .. 2.3 (138.83) OVERALL LENGTH M (FT) 39.27 (128.83) 39.27 (128.83) PROPULSION SYSTEM P&W JT3C-6 P!.W .J57 TIPE P&W .]57 ..., a 0.402 WBL 315 INBOARD LOCATION ;zm. 315 'I" 0.402 I.1IL 315 '7 - 0.402 OUTBOARD LOCATION WBL 545 'I,. 0.694 I.1IL 545 'I a 0.694 ;rnL 545 '7 - 0.694 ~ TAKEOFF THRUST (NET) 60,075 (13,500) 49,840 (11,200) 49,840 (11,200) N (LE) \... C PaoP-FAN L I .

t......! 3.56 m

I : Ul.67 ~) ~'<ACEl.LE LTI!lIFACE

II.B.I..Zl~ )

, ' I' .

~ I

VIElI LOOKING An , F.S. 597.66 F.S. 707.56 Figure E-12. Nacelle Installation Geometry 928, as shown on Figure E-13. This portion of the nacelle is 5.6 m (221 in) long and varies in height above the wing from 1.14 m (45 in) to 0.9 m (35 in) at the center line. As far as possible, the aft nacelle contours are designed with single curvature panels and consist of a semi-circular upper section and straight sides from the maximum beam to the intersection of the nacelle side wall with the wing upper contour. An aluminum alloy "skate" angle is attached to the wing upper surface, providing attachment for the nacelle side walls and for lower pick-up points on the engine nacelle. Upper diagonal ties from the upper attachment are secured to the front spar, adjacent to the skate angles, on the wing upper surface. Lower diagonal truss members are attached to the lower QEC unit pick-up points and extend downward and aft to an attachment located on the nacelle centerline at the front spar flower cover junction. These member s form a V-truss and transfer load into the lower skin cover by means of an external "tee" support. Reinforcement of the covers, except for local increases in thickness to provide bearing material for nacelle structure attachment, is not required. The aft nacelle structure frames are spaced approximately at 30.5 cm (12.0 in) intervals. The frames are formed aluminum alloy channel sections 1.0 mm (0.04 in) thick. Extruded aluminum-alloy longerons, at the maximum beam of the nacelle, form the boundary between the straight-sided walls of the nacelle and the semi-circular upper covers. Access to the jet pipe is provided by three removable panels. In general, the nacelle skins are aluminum alloy 1.0 mm (0.04 in) thick supported by longitudinal "tee"-section aluminum alloy extruded stiffeners. The aft portion of the nacelle terminates slightly forward of the trailing edge of the inboard spoilers, and a fairing is added to protect OL\CONAL TIE AL. AL. 0.1 SQ [NS ~---r--Rf!tOVABLE PANELS

/

fUKINC ~TTnT.~~Tr~~--~_,/ ... \ .-/~ .. .,~""", \ \\

_ .. ,,-/.-- ./

/ /-/~\ .

\.----- ... .-----

Figure E-13. KC-l35A Aft Nacelle Structure the upper surface of the flap from the jet blast. Because the nacelle covers the inboard spoilers, it is necessary to lock-down both spoilers and disconnect both from the spoiler system. The KC-135A wing is a two-spar distributed struc- ture consisting of constant-thickness, roll-tapered or machined, aluminum-alloy skins and of extruded "Z"-section stiffeners. In general, the stiffeners are 6.35 cm (2.5 in) deep and vary in thickness from 2.39 mm to 7.62 mm (0.094 in to 0.30 in). The upper cover thicknesses in the area of the prop-fan installation vary from 3.18 mm to 6.35 mm (0.125 in to 0.25 in). Cover and stiffener material is aluminum alloy 7178-T6. The front spar in the region of the prop- fan installation has a web 2.29 mm (0.09 in) thick and extruded aluminum-alloy "tee" section caps. Where possible, attachment of the nacelle structure will be accomplished by picking up existing fastener locations in the upper cover. The addition of fasteners in excess of those already in the structure will be per- formed without degradation of the strength or stiffness of the wing primary structure.

KC-135A Testbed Flutter Analysis A preliminary wing flutter analysis was performed for the KC-135A testbed configuration to determine the effects of the prop-fan power plant installation on the wing flutter stability. A semi-span (half-airplane) mathematical model •.

which implies structural and aerodynamic symmetry about BL 0.0. was used. The resul ts of the analysis are. therefore. directly applicable to a symmetrical 2-engine testbed configuration.

Structural Representation - The structural representation used in the analysis consisted of a flexible wing with flexibly mounted turbojet and prop-fan powerplants and a rigid fuselage-empennage. The flexible wing was represented as a 10-lumped mass system with freedom in vertical and fore-and-aft bending. and torsion. The pylon flexibilities for the primary turbojet engines were each represented by one vertical and one side bending mode. All mass and stiffness data for the basic airplane were taken from the Reference 4.

The prop-fan powerplants were represented as additional sprung and unsprung mass lumps located at BL 217. The sprung mass of 1378 kg (3040 Ib) represented the propeller. gearbox. and engine and had uncoupled mode frequencies of 5.72 Hz (lateral). 7.51 Hz (vertical). 8.68 Hz (yaw), and 9.93 Hz (pitch). The unsprung mass ~f 703kg (1550 Ib) represented the fixed nacelle structure and wing local strengthening.

Aerodynamic Representation - The unsteady aerodynamic forces on the wing were computed by the Theodorsen strip theory. Finite span and compressibility effects were accounted for approximately by local lift-curve slope and aero- dynamic-center modi fications. which were based on vortex lattice calculations

for a speed of M = 0.85. The wing aerodynamic forces were computed for '0)

strips which coincided with the 10 mass panels. No unsteady aerodynamic forces were applied to the fuselage or empennage surfaces.

Aerodynamic forces on the prop-fan were computed by quasi-steady strip theory. modified for lift lag due to unsteady flow. The prop-fan blade lift-curve slope distribution data were calculated by Hamil ton Standard for the S-blade SR-3 prop-fan operating at M = 0.8.

Flutter Analysis Results - The results of the fl utter anal ysis are sum- marized in Figure E-14. The unmodified KC-135A wing was analyzed first, since Boeing data were not available to form a basis for comparison. A single weight (:ondition of 36,432 kg (190.590 Ib), condition "C" of Reference 4, was analyzed.

This weight condition includeS structural r~serve wing fuel of 1405 kg (3100 Ib) and 37,786 kg (83,323 Ib) of fuselage fuel. The critical flutter mode for the symmetr ic and unsymmetric cond itions was characteri zed by wing outer panel bending-torsion at a frequency of about 11 to 12 Hz. The symmetric flutter HINlHUH RESERVE VDC FUEL ALTITUDE FT m SYMMETRIC ANTISYMKETRIC 40000 12000 D BASELINE CONFIGURATION • TESTBED CONFIGURATION 101000

\

30000

\

8000 \

\

20000

~

6000·

\

\

\

10000

\

\

mls 100 200 250 300 0 50 350

________ ~'~----~~'~------~'~------~~----~~'~----~~'~----~n KNOTS

0 100 200 300 400 500 600 700 EQUIVALENT AIRSPEED Comparison of Predicted and Measured Wing Vibration Modes (No Prop-fans) SVlIDetric Predicted Measured Mode Freg, (Hz) Freg, (Hz) fp/fm Wing 1st bending 4,31 4,40 ,9S0 Wing 2nd bending 16,42 13.56 1.21 Wing 1st torsion 19.27 19,42 ,992 Wing 2nd torsion 3S,39 3S.76 .990 AntisI!!!!!etric Wing 1st bending 5.15 5.12 L06 Wing 2nd bending 12.30 9.66 1.27 1st IS,18 ,992 Wing torsion 18.32 Wing 3rd bending 35.70 36.59 .976 Figure E-14. KC-l35A Testbed Flutter Boundaries speed was lower, as shown in Figure E-14, but was outside the required 1. 15 VD envelope of the unmodified KC-135A.

The addition of the prop-fan powerplant, with nominal attachment flexibili- ties and propeller aerodynamic and gyroscopic effects, caused the flutter speeds to change slightly, as the flutter mode involved mainly outer wing motion. The unsymmetric fl utter speed decreased slightly, and the symmetric fl utter speed increased slightly as indicated by the solid square and circle symbols, respectively. Elimination of the propeller aerodynamic and gyroscopic effects and changes in the prop-fan power plant attachment flex ibil i ties caused negligible changes in the flutter speeds. It was concluded from these results that the prop-fan installation will have negligible effect on the wing flutter characteristics of the KC-135A aircraft and that no changes to the wing structure will be required for flutter prevention.

KC-135A Testbed Operating Envelope The KC-135A testbed flight envelope, Figure E-15, was derived from KC-135A data. The design dive Mach number, 0.88, is sufficiently beyond the testbed design requirements of Mach 0.8 at 9118 and 10,668 m (30,000 and 35,000 ft) to obviate the need for speed restrictions on the testbed aircraft over the full range of flight conditions.

ALTITUDE FT M ·O.88 PROP-FAN OPERATING D 4O,COO 12,000 10,000 30,000 20,000 6COO 10,000 .. /.

o ISO 0 !O 100 200 L-_---',"___---',"___--" __ --" KNOTS I 0 100 :00 300 400 EQU~/4LENT 41RSPEED Figure E-15. KC-l35A Testbed Operating Envelope KC-135A Testbed Performance The performance of the KC-135A testbed aircraft with twin prop-fans is shown on Fi gure E-16.

Data are given for two weight conditions representing start and end cruise test weights of 81,630 kg (180,000 lb) and 54,420 kg (120,000 lb), respectively. The capability of the KC-135A is given for the unmodified aircraft for two conditions: three engines at normal rated thrust (NRT) and one windmilling and with four engines at NRT. These data show the capability of the basic aircraft to satisfy the testbed aircraft design requirements. The remaining data demonstrates the capability of the testbed configuration to meet the design requirements. Operating three primary engines at NRT and the two prop-fans at a power equivalent to the thrust output of a primary engine at NRT provides a speed margin of L1 M = 0.05 to L1 M = 0.07 over the altitude range 9144 m to 12,192 m (30,000 to 40,000 ft). The testbed aircraft, at the true start and end cruise weights of 84,673 kg (186,710 lb) and 55,476 kg (122,330 lb), provides a test mission duration of 4.7 hours. This test mission duration will provide adequate time to set up test conditions and to accumulate test data.

STTl lJAY GUNlJlnuNS, .1-57-['-59:; MIN P\oo1{ I'LT FT m

\ ~ \

11,000 36.000

:\ '~; ENGINES ,~ NRT PLUS

'Ii ~ 1 PROP-FAN

1 : ,l/

34,000

~~ '" '" ~) 5,~42 i(g, (120,000 !..3S)

"'0'\ :v<? :;

I

.DO 10,000 '" .... 0 '7' --.J '::x:J'" ."" ." .

32,000

~":.c!"" "/

II ~ ,,/ 0" ~ "" 3~ 4:'V

.::: 1

30,000

i ,,~1t'<'~~~ki'~"""""~~~~'"

9,800 3 ENCINI::S ~~ NRT "II EW:lNES ,·1 :-<&1'

~~~---~~~------------ .70 .~O .84 .88 .n MAClI NO.

~ SPEEO/ALTITUDE BI::GIN TEST Xl' ~~D ['EST AT 10.668 m :0,668 " (35,000 FTlO.3M RA.'1P ',EIGHT = 91,~75 Kg (201,710 LB) START CRUISE .T = 84.672 Kg (186,710 LB) '3S'OOO FT)O.8M /; RUNWAY CONDll'IlJNS END CRUISE WI ~ 55,~i6 Kg (J22,330 LB) I,LEVAl'ION - 701 m FUEL REQUIRED ~ 29,196 Kg (t~, 380 LE) (2300 FT)

!: TEMPERATURE - ~ouF

rEST TUIE ",0- NO '.-nND ~O I~RADIENT 2 4. J HRS s:.

LANIlINt; TAKEIlFF Figure E-16. KC-l35A Performance KC-135A Testbed Weight and Balance Weight and balance data for the testbed aircraft are shown on Table E-V.

The wing fuel capacity of the unmodified airplane is 49,431 kg (109,500 lb).

Since the mission fuel required is less than the wing-tank capacity, all mission fuel can be carried in the wing tanks so that the center-of-gravity will move The center-of-gravity at aft as fuel is loaded and forward as it is consumed.

operating weight can be maintained in any position for the modified aircraft by proper location of the test equipment. The normal range of center-of-gravity movement is from 12.5 percent MAC to 35 percent MAC. At ramp gross weight the testbed aircraft center-of-gravity at 26.4 percent MAC can operate within this range as shown in Figure E-17.

'I I I ( TABLE E-V.

I !:: ~I KC-l35A WEIGHT AND BALANCE 220 100 g ",I c; ~I WICHT LOCAIIOH 200 WICII'I COMPONE:r.: LB Aa.'f fS ': MAC KG ..

...

(90,666) 850,6 OPEllTL'I: WICK! 30 .ll17 ... ~ ,

·

UlIIIOOIFIEIl 180 ;: I (3614) 66Z.9 2-X'I70l PROP-Fo\.."1 3906 ~ PACKAGES :; ~ '" (654) 798.0 .96 '" OVERIIL'IG "ACELLE ~ SnwCU'RE 1.:.0 (3000) 964.1 TEST E~I~ 838.5 1.:0 25 (l08,334) ZnQ FeEL ~'I.tGH'r :.9129

·

so (93,376) reEl. 42J46 (~01. ;:'0) 10 :0 )0 ,,,"11' GROSS WIGl!T 91:'75

·

Cz..'tI£R JP' G'AAVr.:y - ':: ~~ Figure E-17.

KC-l35A Center-of-Gravity Range GAC GIl TESTBED CONCEPTUAL DESIGN The GAC GIl is a high-performance, jet propelled, low-wing, business/execu- ti ve aircraft. The GIl has sufficient volume in the passenger area to seat a

maximum of 19 passengers and can fly cruise missions at M = 0.80 and M = 0.85 of

4723 to 5648 km (2550 to 3050 nm) and 3389 to 4000 km (1830 to 2160 nm) respec- tively. The unmodified aircraft performance, i.e., speed/altitude, has been shown to be in excess of the design requirement for the testbed aircraft, and previous analyses have shown that the XT701 power/aircraft/prop-fan scale to be compatible with a GIl powered by the DDA XT701/T56-A-14 drive system. The principal dimensions and characteristics for the GIl are given in Table E-VI.

Two testbed configurations were investigated; that shown in Figure E-18 is the testbed with the drive systems located at BL 145 left and right, TJ = 0.35, for which Figure E-19 shows the corresponding nacelle geometry. Figure E-20 illus- trates the drive systems located at BL 185, TJ = 0.45, the limiting outboard location.

GIl Drive System Location and Geometry The GIl uing scantlings and sections, Figure E-21 were obtained from data supplied by GAC and were used to locate the drive systems on the wings. Two locations were investigated, as indicated above. The first location chosen, shown on Figure E-18, was at BL 145.0 TJ = 0.35, because a change in wing thick- ness occurs from this location inboard and adequate back-up structure for the prop-fan installation exists in the wing. This location is the limiting posi- tion inboard for the 2.89 m (9.5 ft) diameter prop-fan as far as clearance and interference with the airflow to the primary engines is concerned. Mounting the drive system at BL 145.0 places the mating plane for the nacelle/airframe at FS 385.98, the location of the wing leading edge at BL 145.0. In keeping with the recommendations of Hamilton Standard for minimizing excitation factors, the mating plane is inclined forward 0.0174 rad (1 deg) from the vertical plane through FS 385.98. In the normal ground attitude, the ground/prop tip clearance is 0.513 m (20.2 in). The prop-fan shaft centerline, located at WL 73.52 and FS 386.98 at the mating plane, positions the power section centerline so that TABLE E-VI. GULFSTREAM II PRINCIPAL DIMENSIONS AND CHARACTERISTICS RAMP tJEIGR'I Kg (LB) 28,375 (62,500) OPERATING WEIGR'I Kg (LB) '15,481 (34,100) ruEL WEIGR'I MAX Kg (LB) 10,578 (23,300) YING SPAN M (n) 21.0(68.83) OVERALL LENGTH M (n) 24.4 (79.92) OVERALL HEIGR'I M (n) 7.5 (24.5) WING AREA SQ M (SQ n) 73.7 (793.5) ROOT CHORD M (F'r) 5.08 (16.67) (6.34) TIP CHORD M (n) 1.93 MAC M (F'r) 3.75 (12.28) tIc ROOT 12.05% tIc TIP 8.42% (.10) DIHEDRAL RADS (DEGS) 0.052 (25°) SWEEP C/4 RADS (DEGS) 0.436 ASPECT RATIO 6.0 HORIZONTAL STABILIZER AREA SQ M (SQ n) 16.93 (182.25) SPAN M (FT) 8.23 (27) ROOT CHORD M (rT) 2.74 (9) TIP CHORD M (n) 1.37 (4.5) MAC M (FT) 2.13 (7) ASPECT RATIO 4.0 VOLUME COEFF V H 0.677 DIHEDRAL o VERTICAL STABILIZER AREA SQ M (SQ FT) 14.38 (154.7) SPAN M (F'r) 3.75 (12.3) ROOT CHORD M (FT) 4.65 (15.25) TIP CHORD M (FT) 3.022 (9.92) MAC M (FT) 3.89 (12.77) 1.0 ASPECT RATIO VOLUME COEFF 0.073 ruSELAGE 2.39 (7.84) MAX WIDTH Men) 2.39 (7.84) MAX HEIGR'I M (n) LENGTH M (FT) 21.74 (71.33) PROPULSION SYSTEM RR "SPEY" KKSl1-8 (RB 163-25) TYPE AFT FUSELAGE LOCATION MAX TAKEOFF THRUST N (LB) 50707 (11400) Figure E-18. Gil Twin Prop-Fan Testbed Configuratioo WI 100 (Rtf) . WL 73.5~L

.I1:J'1

FS 385.62 ~/ Af ItHERSECI fS 385.98

/.-.-,\

.- - ~.- "12~~j --- .. -._--

-- .. -,,~.-.--- ..... --.-.- , , FAlklNG ~---------·--~\l-~~~··~~--·~r~\-l~····~-·-~-~·-·~-~\~--~·--~- .L.l-J ~, -.-.- --\,--,- --- ; ::!; \ ._ ..•.• I ___ .. ____ \' ..

c

·- __________ ~~~~-:....=..~~-n,T-'"T

\ WING REAK BEAM ---\ . \ - \ (TRACE Al WRfj .'- .\- \.

\ ___ ." II WING HONI BEAMV"'\· .- ..... ..\ \ .

(TRACE Al WRf) '\ \ • \ '; , \.. \ Figure E-19. Gil XT701 Nacelle Geometry WlIOO.oo· . WL 53.04 AT ______ --~~\7.~::~::--~--------~~.-.EA~NGEDGE Wl47.5O-" f.S. JOB.HH MAX CONTOUR BL 145.00 lWR WING SURf WING REF PLANE AT .I:~!~~':~'O> ~'~_______ +;~ -~ _, ___ .

_. [GROUND LINE ATPV:NE OF pWP Figure E-19. Gil Testbed Nacelle Geometry B.L. 185 (Cont'd) US.OO '''5 Figure E-20. G II Testbed Nacelle Geometry B. L. 185 -=-- .s=-:_, .. --............. .. _ --- ---~.

.------- ~-- --_._------ --- ~ -------

-

---------

-------- =-~... - --'~-==-=-

. -._-=_ .. _-_ .. ........... - .

Figure E-21. G II Wing Scantlings sufficient clearance is provided between the upper surface of the wing and jet pipe. This geometry is a compromise location to minimize the torque effects of the prop-fan thrust on the wing box structure and to maximize the prop-fan tip/ground clearance in the normal ground attitude.

A second location at BL 185. 11 = 0.45 Figure E-19. was also investigated.

since this position represented the limiting position outboard on the wing. at which engine-out conditions could be controlled. At thi s location the prop-plane required 0.914 m (36 in) of movement aft to partially satisfy flutter requirements.

The nacelle/airframe interface plane is located at FS 407.25 and WL 54.62, the wing leading edge at BL 185, and is inclined 0.017 rads (1 deg) forward from the vertical plane through this location. The intersection of the prop-fan plane and shaft center line is at FS 332.56 and WL 73.72 and the prop-fan shaft center line through this point is normal to the interface plane.

The data for the two locations, which represent the limits of inboard and outboard movement for the drive system, were generated to establish the clear- ances between prop-fan tip and fuselage external surface. Near-field acoustic considerations, relative to cabin noise attenuation, recommended a minimum clearance of 0.8 D (where D is the prop-fan diameter). At the BL 145 location p p the clearance is 1.087 m (3.57 ft) or 0.375 D. At BL 185 the clearance p increases to 2.062 m (6.77 ft) or 0.71 D. Although less than the recommended p clearance these clearances should not prevent the use of the GIl a vehicle for acoustic tests.

GIl Aft Nacelle Structure, Drive System at BL 145 The aft nacelle structure mounted on the wing upper surface at BL 145.0 is The nacelle structure consists of two vertical side shown on Fi gure E-22.

panels capped by a semi-circular removable cowl structure. The aft nacelle extends from the mating plane to the end of the jet pipe located approximately at the trailing edge of the spoilers. The structure consists of an assembly of skins, frames, longerons, and stiffeners of aluminum alloy. The engine QEC pick-up points match similar attachment points on the aft nacelle at the mating plane, and the structure is arranged so that the upper attachments coincide with the main diagonals which are connected to the rear spars of the wings at the lower end. The nacelle attachment angles on the upper surface of the wing pick-up the QEC lower attachments and the diagonal members at the aft ends. The skins, which are 1.0 mm (0.04 in) thick are supported every 25.4 cm (10 in) by 10 cm (4.0 in) deep channel section frames. Longerons on each of the nacelle walls located on the maximum beam of the envelope also serve as boundary members for the removable semi-circular panels. Frame thickness is 1.0 mm (0.04 in) and the areas of the longerons, main diagonals and attachment angles are approxi- mately 4.0 sq cm (0.6 sq in).

DIAGONAL MEMBER AL. AL. 0.10 SQ INS SKIN AL. AL. 0.05 INS TYPICAL SECTION AA Figure E-22. Gil Aft Nacelle Structure B. L. 145 GIl Aft Nacelle Structure. Drive System at BL 185 The flutter analysis with the QEC mounted at BL 185 indicated an unaccept- able condition, as far as damping modes were concerned, so that mounting the QEC, Figure E-9, at this location was not practical. Further analysis with the propeller plane moved aft, toward the leading edge, 9.15 em (36 in) produced a marginal condition. The geometry for this location is shown in Figure E-19.

The lines and contours for the shortened nacelle were developed from those of the QEC unit by removing 35.6 cm (14.0 in) from the aft portion of the nacelle.

Due to the relocation of the prop-plane, the mating plane is reconfigured to a sloping bulkhead between the nacelle attachment points. Above and below the attachments points, the mating plane bulkhead segments are normal to the nacelle center line. The nacelle mounting consists of the Lockheed P-3C V-frame mem- bers, with the appropriate changes in the attachment fitting angles at the mating plane. The aft nacelle structure, Figure E-23, consists of that portion of the structure from the sloped mating plane to the trailing edge. The aft nacelle supports the drive system installation by means of diagonal members from the upper attachment points downward and aft to the rear spar and by "skate" angle and nacelle lower longeron extensions which are attached to the front spar. The nacelle structure is fabricated from aluminum-alloy skin, frames, and longerons/ stiffeners. Skin thickness is 1.0 mm (0.04 in) and frames are 10 crn (4.0 in)-deep channel sections. The upper, semi-circular portion of the nacelle FAIRING \ \; "'-TRA[LING EDGE FRAME TYPICAL ~GINE MOUNT I DRAG ANGLE , ['AIRING Figure E-23. G II Aft Nacelle Structure B. L. 185 is removable to provide access to the jet pipe installation. The main attachment of the nacelle to the wing upper surface is by means of chordwise "skate" angles. A fairing is provided at the tail pipe to protect the upper surface of the flap from the jet efflux. Because the turbine section of the power unit has moved aft to a position above the wing upper surface primary structure, provision for blade turbine containment is required in this area.

GIl Wing Modification The GIl wing structure consists of integrally stiffened machined upper and lower skin panels and front and rear spar structures, which together form the wing box beam structure. The upper surface of the wing has approximately 26 circular or elliptical access panels on each side, and changes of curvature occur at BL 145 inboard to the center lines. Increasing the torsional stiffness 60 percent, for either of the drive system locations investigated, requires the addition of doublers to the upper and lower surfaces of the wing and to the front and rear spars. Aluminum-alloy doublers, Figure E-24, 1.9 to 2.1 mm (0.075 to 0.084 in) thick are required for the upper and lower surfaces, respectively. These doublers would be attached to the existing skins with mechanical fasteners and would be arranged to accommodate new covers at each access panel location.

Because double curvature exists on the wing from BL 145 inboard, perfect matching of the doublers and skin is not possible and liquid shim would be applied to the faying surfaces. Machined plate, aluminum alloy doublers approximately 2.54 mm (0.1 in) thick, would be added to the forward face of the front spar and to the rear face of the rear spar. Addi tion of the doublers would require removal of the leading and trailing edge structures. No problems are anticipated with the front spar reinforcement, but the doubler applied to the rear spar presents a major undertaking since the removal of the landing gear support is involved. The wing reinforcement would extend from BL 172 L to BL 172 R for the drive system located at BL 145 and from BL 220 L to BL 220 R for the BL 185 location. Finally, modifications to the spoiler system are necessary. These would consist of eliminating the ground spoiler for the inboard location or deactivating the inboard flight spoiler for the outboard drive system location.

UPPER SKIN (0.084 in) .

EXT. DOUBLER 0.21 em I UPPER AND LOloJER - I

/

DOUBLERS \ ,

\

,

(

) i ,

)

W~i 172.n \-JS 0.0 EXT. DOUBLER 0.19 em (0.075 in) NEW MTG.

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= ...... '-==1

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-,<--"-~ '.:-

.

EXT. DOUBLER 0.19 em: (0.075 in) EXT. DOUBLER 0.21 em (0.084 in) ........ ===-.

~~-------- REAR BEAt-!

Figure E-24. G II Wing Modification GIl Flutter Analysis As previously noted, preliminary wing flutter analyses were performed for the GIl testbed configuration to determine the effects of the prop-fan powerplant installation at BL 145 and BL 185. A semi-span (half aircraft) mathematical model was used, implying structural and aerodynamic symmetry about

BL o. o. The results are, therefore, directly applicable to a symmetrical

2-engine testbed configuration.

Structural Representation - The structural representation used in the ana- lysis consisted of a flexible wing with a flexibly mounted prop-fan power plant and a rigid fuselage-empennage. The flexible wing was represented as a 10-lumped mass system with beam vertical bending and torsion degrees of freedom.

The wing mass data were taken from Table IV of Reference 5, and stiffness data from Figures VIIl-3 and VIIl-4 of Reference 6. The torsional stiffness was increased by 12 percent to obtain better correlation with the vibration test resul ts of Reference 7. For the unsymmetric case only, a wing-to-fuselage flexibility in roll was added for the same reason.

A comparison of the predicted and measured wing vibration mode frequencies for the unmodified GIl, without prop-fan powerplants, is shown in Figure E-25.

The comparison is very close for both the symmetric and unsymmetric cases, with the exception of the second bending mode frequencies, which were overpredicted by 21 and 27 percent for the symmetric and antisymmetric cases, respectively.

It was concluded that these differences were caused primarily by engine pod flexible mode coupling, which was not represented in this analysis.

~U"'IOUlmWltIC U""'''IQ Gil 1110 '1IOf'-',"" 1000 " ..

;.

1.

11.1S ~~,~~~-.~,!~---+',W~~ ... r-~J~~~,~~--~:1 ~s ~IVALDn' AIl:irUD

Figure E-25. G II Flutter Boundaries - Prop-Fan at W. s. 145

The prop-fan powerplants were represented structurally as additional sprung and unsprung mass lumps located at BL 145 and BL 185. The sprung mass of 1378 kg (3040 lb) represented the prop-fan, gearbox and engine and had uncoupled mode frequencies of 5.72 Hz (lateral), 7.51 Hz (vertical), 8.68 Hz (yaw) and 9.93 Hz (pi tch) , which were derived from Lockheed "Electra" nacelle and engine mount stiffness data. The unsprung mass of 474.8 kg (1047 lb) represented the fixed nacelle structure and wing local reinforcement.

Aerodynamic Representation - The unsteady aerodynamic forces on the wing were computed by Theodorsen strip theory. Finite span and compressibility effects were approximately accounted for by local lift-curve-slope and aero- dynamic-center modifications, which were based on data from Table VI of Refer- ence 5. The wing aerodynamic forces were computed for 10 strips, which coin- cided with the 10 mass panels, as shown in Figure 2 of Reference 5. No unsteady aerodynamic forces were applied to the fuselage or empennage surfaces, since previous experience has shown that these are relatively unimportant for wing flutter predictions.

Aerodynamic forces on the prop-fan were computed by quasi-steady strip theory, modified for lift lag due to unsteady flow. The prop-fan blade lift- curve-slope distribution data were supplied by Hamilton Standard and were for the 8-blade SR-3 prop-fan operating at M = 0.8.

Flutter Analysis Results, Drive System at BL 145.0 - The results of the wing flutter analysis are summarized in Figure E-25. The unmodified GIl wing was analyzed first and compared with the results of the Grumman analysis to validate the mathematical model. The flutter boundaries agreed within 2 percent, as indicated by the circle symbols in Figure E-25, even though the Grumman mathematical model included flexible fuselage and empennage effects, which were not included in the Lockheed analysis. The flutter mode involved is a 7 to 10-Hz anti symmetric wing bending-torsion mode.

The addition of the prop-fan powerplants at BL 145 caused a 5-Hz symmetric flutter instability inside the testbed dive speed envelope, as indicated by the solid square symbol. When rotating prop-fan aerodynamic and gyroscopic couplings effects were added, the speed of this instability increased by about 23 mls (45 knots), but was still unsatisfactoriily low, as shown by the open square symbol.

To increase the flutter speed to a satiisfactory level, a substantial in- crease in the wing torsional stiffness inboard of BL 145 is required. The effect of a 60 percent increase is shown by the solid and open triangle symbols for the feathered and rotating prop-fan conditions, respectively. Although a somewhat smaller stiffness increase might be satisfactory, a more elaborate and comprehensive flutter analysis will be required to determine a precise figure.

Flutter Analysis Results, Drive System at BL 185 - Relocating the QEC to BL 185, and with the prop-fan plane one diameter ahead of the wing leading edge, increases the flutter speed above that for the powerplant at BL 145. The damping of the fundamental wing torsion modes (both symmetric and unsymmetric) is, however, unsatisfactorily low at airspeeds well wi thin the limit-speed envelope. Attempts to stabilize the mode by increasing the wing torsional stiffness actually reduced the damping, so that it became obvious that no reasonable amount of wing stiffening would solve the problem.

It was found, however, that moving the prop-fan plane aft 91.4 cm (36 in) improved the damping of these modes, which when combined with a 60 percent increase in wing stiffness out to BL 200, provided satisfactory damping, within the limit speed envelope. It should be noted that the damping is only marginal, as indicated in the Figure E-26, and is sensitive to changes in altitude, power plant mounting stiffness, prop-fan aerodynamic characteristics, and other parameters not investigated. Additional investigation is required to verify the damping characteristics at this power plant location and to determine, more precisely, the wing and powerplant design requirements for flutter prevention.

-1000 rT 1000_ ..

.0 lO '0 .0 0 .00 ,GO 'GO :'0 ".

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... ,OQ

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'00 '00 >GO '00 ......

~I'IALDt .UIPUD Figure E-26. Gil Flutter Boundaries - Prop-Fan at W.S. 185 GIl Operating Envelope The operating envelope for the GIl, Figure E-27, was established by analyzing the 0.13 rad (7.5 deg) upset condition for 20 seconds to determine the dive speed. The points analyzed were those at altitudes of 9118 m (30,000 ft) and 10,668 m (35,000 ft), starting the upset at a Mach number of 0.8. The upset condition onset at 9144 m (30,000 ft) results in a Mach number increase to 0.89 at the end of 20 seconds and an end altitude of 8534 m (28,000 ft). Below this altitude the testbed aircraft speed is restricted to 172 mls EAS (335 KEAS) in order to minimize weight penalties arising from wing torsional stiffness increases.

.IL!1'Ml!

n: soooo ~DlUM OPEllA'tt!iG :'! - 0.35 1.0000 30000 • :-!AXnfL"M om spEEJ) \rITI'H PROP-FAN OPDAT~ 20000 10000 -:.15 0 SO 100 150 250 I ( I r '~TS C 100 200 300 400 300 Figure E-27. GlI Operating Envelope GIl Prop-Fan Testbed Trim Capability The outboard limit for locating the drive system is BL 185, TJ = 0.45, which is dictated by the aircraft trim capability. The data of Figure E-28 show that, with one prop-fan windmilling and 100 percent power on the other, the testbed aircraft can be trimmed for engine-out conditions at 77.7 mls EAS (151 KEAS) in free air with a 0.087 rad (5 deg) angle-of-bank or at 90 mls EAS (175 KEAS) on the ground. The use of the T56-A-14 gearbox restricts the power input to 4101 kW (5500 shp) so that, when this constraint is applied to the data of Figure E-28, the power setting of the prop-fans, at the conditions indicated, is limited to approximately 75 percent of takeoff power. At this power setting.

puaNt PaWl.

/ / GlAUOX CAPAliUty- / ..

CUT"'" / / od PlOP"''''' WtMMILL .

.- ;' +0 ONG~// -' '" ! ! !

;zo L~O L..o :.&0 LlO ,'" !'QDt'fALm AIUPUD PROP FAN LOCATED AT WS 185 Figure E-28. Gil Trim Capability - Prop-Fan at W.S. 185 the engine-out, free-air trim capability can be achieved at a speed of 66.3 mls EAS (129 KEAS) and 77.0 mls EAS (150 KEAS) on the ground.

The data of Figure E-28 apply to takeoff conditions at sea level standard day, with a flap setting of 0.35 rads (20 degs), at a weight of 25396 Kg (56,000 lb), with center-of-gravity located at 32 percent of the MAC. The trim capability is shown for 95 percent rudder movement.

GIl Testbed Weight and Balance Weight data are presented for both drive system locations in Table E-VII.

The essential difference between the weights is due to the increased doubler weight for the BL 185 drive system location. The operating weight of the unmodified aircraft is 15,464 kg (34,100 lb), which increases to 21,508 kg (47,428 lb) and 21,622 kg (47,678 lb) for BL 145 and BL 185, respectively. The difference in fuel weight for the two configurations is about 113 kg (250 lb).

The take-off weight of 28,344 kg (62,500 lb) is based on the maximum ramp weight of aircraft serialized from 101 to 216, inclusive. The small size of the testbed configuration, and the large weight increment percentage needed to convert the basic aircraft to the testbed vehicle, requires careful control of the center-of-gravity. Balance checks of the testbed configuration show that, for either of the drive system locations, the aircraft center-of-gravity can be maintained wi thin the envelope for the existing aircraft at all weights, by TABLE E-VII. GIl TESTBED WEIGHT AND BALANCE WS 145/FS 385.98* WS 185/FS 332* DRIVE SYSTEM LOCATION i WEIGHT WEIGHT WEIGHT COMPONENT % MAC ARM FS ARM FS % HAC Kg Kg LB LB 0 OPERATING WEIGHT-UNMODIFIED 39.3 15,464 (34,100) 462.0 15,464 (34,100) 462.0 39.3 XT-701 PROP-FAN PACKAGES 3,907 ( 8,614) 342.2 3,907 ( 8,614) 395.1 OVERWING NACELLE STRUCTURE 233 ( 514) ( 424.9 233 514) 441.8 WING DOUBLERS 544 ( 1,200) 408.0 657 ( 1,450) 410.9 TEST E~IPMENT 1,360 ( 3,000) 538.0 1,360 ( 3,000) 530.0 0 ZERO FUEL WEIGHT 26.6 21,513 (47,428) 443.3 32.8 21,626 (47,678) 452.4 FUEL 6,837 (15,072) 418.5 6,723 (14,822) 418.3 0 RAMP GROSS WEIGIIT 22.5 28,350 (62,500) 437.5 27.3 28,350 (62,500) 444.3 *PROP-PJ.ANE LOCA'CION placing the test equipment in the passenger compartment, the center-of-gravity problems are eliminated.

GIl Testbed Performance The mission performance of the GIl twin prop-fan testbed is shown in Figure E-29. At a ramp weight of 28,344 kg (62,500 Ib), the start cruise weight at 10,668 m (35,000 ft) is 27,317 kg (60,235 Ib), and the end cruise weight is 22,109 kg (48,752 lb). Cruising at Mach 0.8 gives a test mission duration of 2.68 hours. The speed/altitude performance, also shown in Figure E-29, shows that a Mach number margin of I!J. M = 0.04 to 0.05 exists over the design condi- tions for the twin prop-fans operating at full power with the primary "Spey" propulsion slightly above idle power setting. Because of the gearbox power absorption limitation and the design of the basic aircraft flap and spoiler systems for operation without propeller Slipstream effects, the use of the prop-fans on takeoff and landing would be restricted to a condition of zero net thrust.

1000 FT 1000 m TWIN PROP-FAN V FULL MAX POWER ON PROP-FANS MAIN ENGINES AT APPROXIMATELY IDLE PLACARD LIMIT . 76 .80 .84 .88 .92 MACH NO.

SPEEn/ALTITUUE BEGIN TEST AT 10,668 m (35,000 FT) O.8M RUNWAY CONDITIONS ELEVATION - 700 m (2300 FT) o TEMPERATURE - 26.6 C

()o °

~ TAKEOFF (80 F) LANDING NO WIND NO GRADIENT RAMP WEIGHT 28,350 kg (62,500 LB) START CRUISE WEIGHT 27,322 kg (60,235 LB) END CRUISE WEIGHT 22,114 kg (48,752 LB) FUEL WEIGHT 6,837 kg (15,072 LB) TEST TIME 2.68 HRS Figure E-29. Gil Testbed Mission Performance GIl Estimates of the Prop-Fan Slipstream Characteristics Slipstream characteristics were calculated for the GIl, in terms of veloci ty and swirl angle, using an SR3, 10-bladed prop-fan configuration. The data shown in Table E-VIII are consistent with values used to determine slipstream effects for the estimation of the GIl testbed mission performance.

TABLE E-VIII. Gil SLIPSTREAM CHARACTERISTICS CONDlTlON % RADIUS ALT kW/m2 VrlD/s CASE .249 MN C J .291 .361 .455 .563 .676 .785 .878 .949 .990 m (FT) (fps) p (SIlP/D .006 .0079 .0107 .0162 t::.v/v .0139 .0179 .0195 .0234 .0186 .0191 183 10668 1 0.8 2.789 4.08 (600) (35000) (26) 4.58 5.19 5.50 4.98 tit 5.47 4.53 4.20 4.60 3.29 3.27 .0077 .0102 .0138 t::.v/v .0178 .021 .0239 .0270 .0298 .0216 .0247 2 0.8 2.027 3.497 (30) (700) 4.37 5.02 5.49 5.61 5.34 5.07 tit 4.94 4.91 3.20 3.57 6v/v .0092 .0123 .0170 .0223 .0273 .0326 .0383 .0330 .0281 .0338 301 244 3 0.8 1.697 3.06 (37.5) (800) 4.23 4.96 5.67 5.98 6.03 6.06 6.18 4.67 3.65 4.29 tit .0085 .0113 .0155 .0198 .0237 .0279 .0325 .0254 .0206 .0246 6.V/v 241 244 0.8 1.3579 3.06 (30) (800) 3.76 4.39 4.99 5.16 5.06 5.05 5.11 3.51 2.61 3.0~ tit 6.v/v .008 .0106 .0146 .0215 .0184 .0252 .029 .0211 .0163 .01\14 5 0.8 1.1769 3.06 (26) (800) 3.50 4.08 tit 4.61 4.69 4.52 4.49 4.51 2.87 2.04 2.37 6.v/v .0118 .0157 .0219 .0293 .0368 .0449 .0534 .0629 .0519 .0593 301 244 6 0.7 1.697 2.677 (800) (37.5) 4.67 5.45 6.27 6.85 7.10 7.34 7.60 8.15 6.07 6.78

'"

6.v/v .0101 .0132 .0179 .0234 .0286 .0341 .0483 .0396 .0497 .0471 7 0.7 2.027 3.06 (30) (700) 4.85 5.50 6.10 6.38 6.37 6.40 6.43 7.42 6.45 6.37 tit t::.v/v .0082 .0106 .0144 .0188 .0228 .0267 .0428 .0459 .0303 .0364 209 183 10668 8 0.7 2.789 3.57 (26) (600) (35000) 5.26 5.83 6.28 6.36 6.16 6.00 5.84 7.15 tit 6.40 7.43 GIl Prop-Fan Near-Field Noise Characteristics Free-field peak sound pressure levels and noise contours were generated for the GIl fuselage at the flight conditions shown in Table E-IX with the BL 145

drive system location. A peak noise level of 147.7 dB occurs at M = 0.8 with a

2 2 tip speed of 249 mls (800 fps) and a disc loading of 301 kW/m (37.5 shp/ft ).

The noise levels decrease as Mach number, tip speed and disc loading decrease.

Relati ve sound pressure levels estimated for conditions up to the tenth blade passage frequency harmonic. for tip speeds of 183. 213. and 244 mls (600. 700 and 800 fps). respectively, are shown on Figures E-30, E-31, and E-32. These data represent the explicit cruise conditions of Table E-IX and should not be extrapolated to other conditions. The noise contours on the fuselage are shown on Figure E-33 for the XT701 and SR3, 10-bladed prop-fan drive system at the cruise conditions of Table E-IX. At these conditions, the sound pressure level of blade passage frequency harmonics on the noise contour may be determined by algebraically adding the data for each tip speed from Figures E-30, E-31 and E-32 to the OASPL of Table E-IX.

\

"'

-LO \ Sal PROP-FA.'

\ \ \ \ -20 \ \ \ ....

\ -)0 ....

-)0 ....

\ ....

\ \ , \ 0.3 ":i CIWtSE -'00 -'00 \ \ \ \ -50 -50 \ \ \ "- -60 "- "- "-

........... ::i ~ ';R::rst

'0 Figure E-30.

Figure E-31.

Gil Blade Frequency Harmonics G II Blade Frequency Harmonics

@V = 183 m/s (600 fps)

TIP

@ V = 217 m/s (700 fps)

TIP -10 ',~,~---------------- ' ... 0.3 :~ CRUISE .......... ", -10 , , , \.

, -JO \ , , , , , , -40 O. 7 ~I CaUIS!

> '" ~ -50 ::l " :;j l: ...

-,>0 Figure E-32. Gil Blade Frequency Harmonics @ V = 244 m/s (800 fps) T1P '"t.

-l- \.. \ \ \ I I

'\ " \ \

!

\ \ i \ ' i 1\ 1\ -d- \ \ \ \ \ \ : i·' .., :·,1 ..... ,-

"1. :';I~

~ .-j i-T' ~ III I ,I : / I /1

/ /,1

I j I I If

\\

/ I I' Figure E-33. G II Fuselage Noise Contours TABLE E-IX. FREE FIELD PEAK OVERALL SOUND PRESSURE LEVELS SR-3 CONFIGURATION ON. GULF STREAM II TESTBED @ 10668 m (35,000 FT.) CRUISE ALTITUDE TIP SPEED (SHP/D2) kW/m2 CASE OASPL (dB) CRUISE M mls (f~_s) 1 (,26.0) 183 (600) 142.0 0.8 209 2 0.8 241 (30.0) 217 (700) 146.7 3 0.8 301 (37.5) 244 (800) 147.7 244 (800) 146.8 4 0.8 241 (30.0) 5 0.8 209 (26.0) 244 (800) 147.2 6 244 (800) 145.4 0.7 301 (37.5) 7 0.7 241 (30.0) 217 (700) 137.3 8 0.7 209 (,26.0) 183 (600) 129.1 GIl Twin Prop-Fans at BL 185 Stability Derivatives The stability derivatives; yawing moment coefficient due to sideslip, C ,

rolling moment coefficient due to sideslip, C and the sideforce coefficientn~ue

l

to sideslip, C were estimated for the GIl tJ Wi th the prop-fans located at BL

ytJ

185. These coefficients were then compared with those for the unmodified aircraft at takeoff and cruise conditions. As indicated in Table E-X, the coefficients change very little as a result of adding the prop-fan. There is, however, a loss in roll control effectiveness due to the elimination of the inboard flight spoiler. It is estimated that the roll control effectiveness would be reduced to 69 percent of that of the unmodified aircraft.

TABLE E-X. STABILITY DERIVATIVES FOR Gil - PROP-FAN LOCATED AT W.S. 185 TWIN PROP-FAN Gil CONDITION DERIVATIVE UNMOD IFIED Gil C n (yaw) 0.00202 0.00l4 TAKEOFF C (roll) -0.00295 -0.00295 . 1{3 (side- -0.0164 -0.0150 C y force) f3 CRUI SE C (yaw) 0.0023 0.00206

'1l

C (roll) -0.00170 -0.00170

7l

C (side- -0.0144 -0.0153 ____________ ~ __ ~~f_or_c~e)~~ __________ ~ ____________ _ ROLL CONTROL EFFECTI VENESS WITHOUT INBOARD SPOI LER REDUCED TO 69% OF UNMOD IFIED G II VALUE

APPENDIX F - WIND TUNNEL TEST PLAN - TASK VII The Wind Tunnel Test Plan developed in TasK VII is directly related to the Testbed Program Objectives outlined in TasK I, Appendix A, where four technological areas are identified as follows: o Integrity of the Structure o Acoustic Environment o Aircraft Performance o Systems Operation The objectives, wi thin each technology area, were also identified and assigned priority, as shown on Table F-I which is essentially a repeat of Table A-I, but with the addition of NASA-sponsored programs providing useful data concerning the specific technology objectives. Several methods of solution to satisfy the objectives are presented, and the preferred methods identified. The preferred methods of solution, identified by the circles, are not absolute or singular methods, but are merely recommendations. It is, therefore, recognized that the same or similar data for specific objectives can be obtained by al- ternative approaches.

Wind tunnel investigation was identified as the preferred method of solution for three objectives: o Propeller Aerodynamic Data for Flutter Analysis o Verification of Propulsive Efficiency o Effect of Propeller Flow Field on the Wing As a result of this preferrential selection for wind tunnel testing, a Wind Tunnel Test Plan has been developed to augment the Testbed Objectives and to provide answers to technology questions that are uniquely testbed configuration dependent.

The Wind Tunnel Te st considerations, structured around the Program Objectives and Priorities, preferred methods of solution fall into three areas as follows: TABLE F-I. PROGRAM OBJECTIVES AND PRIORITIES PROUI,EM SOUl'l'lUN HE'fIIOD l'U;UHUL()(;¥ ~u.

l'''loWi'l'Y' 'JI':~ TICST STA11lS UUJU;nVE HI'~ U~ PRJORI1'Y AREA USTBW WIND '1'UNNt:L S'fAnC AIKCRAF'I' n:ST ANALYSIS liS I.S .'

1 IHTU:WIH Of I Blade dynwdc reMpOIUI~ 1 ® 'WE STRUC'llJRE vaUdattoll 0 frollcUt=:[ 2 Bladu elasldeal t luttcr 2 X II ® ~tl'"uctural validation lntl!¥f Jty , 3 aJud~ »tall t1ULlcC valhlatton l X X dynuIIIJc .. 4 Critical tIlpt:t!d & hub tittffntHU' 4 X

ffi

vctlfdatluu ------ 0 Prupel1er 5 Deter.iau! aerodyn<Jlillc data X ® ® induced tor flutter analytiulil vlbrttt1un 6. 6 Ikl~nw.hu: tltl'"uctul- .. l VJ.b"MtJun 2 X II II X ® dynWltJcM lipt=ctra _atilt tude 7 DrJ VI! l:iy&t~. dyna.lc loa"'. 3 II X X ® 6. induced cfft!cts Scale 8 VaUJatd or develop deal to" 1 X II ® effect ..

JdWS Blitde ... " .. , tiLtffuetili 2 X II X ® dlutrlbuUOll d~h:r""DaUoD 10 HeMKJlUltrate full .. 1zf;l: vrup-fan J X tOlbl'lcatJon fe .... lbillty 11 lUabl ll1h ..!riv~ tiy8L~ feasi- 4 X bl lilY for 15,000 SHP , abu •• 2 ACOUS'flC 12 Sound ~r-~tiIliUr-d dlr-dct1vity 1 X X ® t:NVuon.It:N'r and 8lJectra var-iat1on u Pro~Jlt:r- lJ Suund I)r-~fu.lurt: levels on 2 II ® gl.!uerdted Pl'(!lhiur-lzeJ sur-faced n~ar-flelll 14 Hohu. ItItrt:ujl,th , dtr.:ctivlt)' J X X ® nulM~ dt:tun,Jnat1ou I) Fluctu_tin, prt:t:t8un: aip~cua 4 1I X 16 £f'~ct .. of tUti~la't: curavtur-e 5 II 17 Cco.~tr)' of corrt:lated sound 6 X

~

prt:ssur-e .ne_ 0 PrapelJer 18 Vertfy prop-fan co .... Uance 1 ® ~t!lIl(:rHted with "AR Pari lb far-field noiait!

- Pa¥sengel" 19 HJDJlIIolzation of sound Uans- 1 X ® e.lIio noltus .1.111101\ , vlbrollotion 20 Rt:tlon~ut he'lut:llcy ..,Jal 2 X ® aur-vey II FU8~hge awdea aud t:xterual J ® n016e relatiOn H Not,t"'! reduction & Mtructural 4 II ® rt:Mpollse .101111"&tJoo by synchro.,ho:ts1ug lJ JPlprove.cnt thru optlI11" .. tlon 5 X X ® c.f ahell .uduil 24 Nultie reduction Lliru cabin 6 l( II ® dll1t:n::.iou changt:ti *Sce Holt: lhduw J AIRCRAfT 25 Vl.:cUy pcopul.tv" effteleuc, 1 X X li'wu.lallh:::lIla I PtlU'OI<llANCt 26 Det"rll!Dt: flow field effect 2 X X PrulJ-t'un

fB

on wJng 1't:I,cl.uul ll SY 27 Vt;:[Uy ~ugJue 1111~t pt:rforlUlncc J I X X ® 4 sys'rms 28 Vt::rify drive "Ydte.. coutrol 1 X ® OPERATION "y"t~ 29 Yt::rUy r6=yerai~r effectlv~ut:»" 2 II <V ]0 Dcter.tne prop-tan vuu~r- ] I ® _bUlly lu .'00 PREFERRED METHOD OF SOLUTION ~,'!

Swirl Simulator ARC 14' Completed Semi-Span, Powered ARC 14' Completed Semi-Span, Unpowered ARC 11' x 11' Completed Semi-Span, Powered ARC 11' x 11' To be tested 2D Wing & Unpowered Nacelle LRC To be tested 2D Wing & Powered Nacelle LRC To be tested Prop-Fan & Non-Symmetric AAC LeRC To be tested Prop-Fan & Flow-thru Nacelle To tested be o Wind tunnel tests that demonstrate the operational readiness of the prop-fan drive system through proof testing procedures.

o Wind tunnel tests that validate and/or advance the fundamental prop-fan design state-of-the-art.

o Wind tunnel tests that validate the airworthiness and the predicted performance levels of the selected testbed configuration.

The first of these technological areas is not aircraft-dependent, but re- lates to the selected drive system for the testbed aircraft. The second area of concern relates to the development of fundamental data that contribute to a better understanding of the technologies associated with the prop-fan concept.

Present NASA programs are also directed toward providing answers to prop-fan concept problems such as prop-fan/nacelle/wing interaction and improved pro- pulsi ve efficiency through propeller swirl recovery. The third area of tech- nological concern is testbed aircraft-oriented, such that tests in this area will relate directly to the airworthiness of the testbed vehicle, to the pre- flight prediction and determination of the testbed aerodynamic characteristics, and to the development of a wing/nacelle installation final design. Further- more, these tests will provide technology-related data uniquely associated with the testbed aircraft.

Recommendations concerning the Wind Tunnel Test Plan are arranged to satisfy the Testbed Program Objectives and to demonstrate airworthiness of the Flight Research Vehicle, and are outlined in the following text. The Schedule of Testing presented in Figure F-1 should provide answers to the first and third areas of technological concern listed above. Because the testbed' concept is intended to augment test data in the second area of concern, no wind tunnel tests are recommended for this area.

KC-135A GULF STREAM II MONTHS FRctI GO-AHEAD MONTHS FROM GO-AHEAD

I

o 12 18 24 30 36 o 6 12 18 24 30 36 6 !

I : I !

I : I l I I

• STATIC TEST STAND ~",zt cbI

I , ~ I

20,540 Mil I

I I 20,540 MIl SUPPORT I (FULL SCALE PROPFAN/NAC) $100,000 ETS I I : $l~f:~fgJ1f?gr~p $22,560 MATL I I I • LOW SPEED W.T.

I~i I (NOT RF.QIJlIir.o) I LOCKliEED-GA. L. S. W. T.

I 5640 KII I II I I I ' I I $ 20'4,000 S'f'POIT I I I I

I i I I

I ' I • HIGH SPEED W. T.

I I I I I I I

I I I I I I I

AEDC I6T ~II

~!!II

(SEMI-SPAN - SCALE - 0.21) 17 , 340 MIl + I I 24,840 MIl MODEL & SUPPORT I $36,000 MATL & MODEL COSTS $21;OOOlMODEL & SUPPORT 'I , , I I I I I I I I .

I I (NOT REQUIRED) • HIGH SPEED FLUTTER TEST I~' I I .

I I LRC 16' TOT , 12,240 MH : I I I I I (SCALE 0.13) I SUPPORT & MODELl I I I I I I I I I

l I I I I

I I I I i I Figure F-l. Wind Tunnel Test Program Schedule DRIVE SYSTEM OPERATIONAL READINESS DEMONSTRATION Operational Readiness Test Plan Recommendation The recommended Test Plan to demonstrate operational readiness, preceding actual flight test of the Advanced Turboprop Testbed System. involves static test stand experimentation only. These tests will provide proof of operation of the prop-fan and the drive system, as well as some near-field acoustic environ- mental data for the nacelle and adjoining structure. These tests would be applicable to either of the selected testbed systems, i.e., the GIl or the It- KC-135A. A suggested test program, including test site and manhour/cost estimates, will be presented subsequently in this section.

Lockheed-Georgia decided not to recommend wind tunnel testing of the propeller drive system prior to actual flight test for the following reasons: o Most of the available wind tunnels are not capable of Simulating the prop-fan design flight environment in terms of dynamic pressure, Mach number, and temperature.

o Most of the available wind-tunnel-flow, solid-wall blockage limits are exceeded with the selected full-size nacelle, prop-fan, and wing- section installation.

o Low-speed testing does not directly address the design point of the prop-fan.

o Costs of wind-tunnel testing are high relative to the Usefulness of the data obtained.

These reasons are justified in the following discussions.

Wind Tunnel Simulation of Prop-Fan Flight Conditions Inadequate With the selection of the 2.89 m (9.5 ft) diameter prop-fan, several wind- tunnel facilities are eliminated from consideration, leaving the Ames 14-foot Transonic Wind Tunnel, the AECD 16T wind tunnel, and the Modane, France, S-1 wind tunnel as candidate facilities. The Ames 14-foot and the Modane S-1 wind tunnels are both atmospheric and are not capable of simulating the Mach

number/altitude (M = 0.8/10688 m (35,000 ft)) environment of the prop-fan design

point. To test at Mcdane, the largest of the available wind tunnels, the Mach number would have to be reduced to about 0.41, so that the operational design point dynamic pressure of the prop-fan would not be exceeded. Testing at the Modane facility would require a strengthened test article, which would not be representative of the flight article. The AEDC 16T, although capable of testing operational gas turbine powerplants in the test section, has not been used for a propeller test in over 20 years. The AECD 16T is, however, a pressure tunnel and is capable of simulating near-design flight conditions of Mach number and 0 0 0 0 pressure altitude with a slight n:tismatch of 33 C - 39 C (60 - 10 F) in stag- 0 o nation temperature. This mismatch could be reduced to approximately 11 C (30 F) with additional tunnel cooling. Other test constraints, such as wind-tunnel blockage and model size, limit the use of the 16T for full-size nacelle testing.

Wind Tunnel Blockage Limits A criterion to be considered in selecting a suitable wind tunnel is the

model blockage in the test section, as measured by the ratio of the test article maximum cross-sectional area to wind-tunnel test-section area, ~/ AwT. The proposed prop-fan test article cross-sectional area would include a wing section. No contribution due to the prop-fan blades would be included. Table F-II identifies the estimated area ratio of the test article for each wind tunnel/ test candidate combination and compares it with the chocking limit of a solid-wall wind tunnel at M = 0.8.

TABLE F-II. WIND TUNNEL BLOCKAGE WIND TUNNEL BLOCKAGE Solid Wall Area Ratio Choking AM/AwT Tunnel Limit GII Nacelle Only

@ M = 0.8 KC-135

. Modane 3-1 .037 0.122 0.089 0.026 AEDC .037 0.180 0.136 0.055 Ames 14-FT .037 0.215 0.165 0.072 The lowest nacelle cross-sectional area ratio, 0.026, shows that the only tunnel to meet the solid-wall choking limit is the Modane facility. This ratio, based on a nacelle cross-sectional area of 1.32 m (14.2 sq ft), does not include model support contributions, which would further increase the area ratio toward the choking limit for a solid-wall tunnel. Recommendations of 2/3 or, prefer- ably, 1/2 of the choking limit are usual for solid-wall tunnels such as Modane.

Testing of the nacelle/prop-fan on a vertical type of support, therefore, be- comes critical because the area ratio approaches the choking limit. Wall effects and blockage interference corrections would be significant for these tests, thereby introducing further que'stions concerning proper data reduction and analysis.

The Ames 14-foot wind tunnel, which ·as a slotted-wall test section, would alleviate the solid-wall blockage limit; however, the blockage ratio is excessive for the nacelle alone. The AEDC 16T wind tunnel, which has a porous- wall test section, may be able to accommodate the nacelle-alone test article.

Wall effect and blockage interference in a porous wall test section are presently unknown quanti ties. Currently, three-dimensional blockage analysis procedures are being developed, but are, as yet, uncorrelated. A significant amount of additional, time-consuming, costly testing would be required to fully develop and understand the blockage effects. The blockage interference problem is neither unique to the testbed development, which could be completed without this information, nor is it deemed critical enough to encumber the program with the additional costs necessary to provide its solution.

The chief purpose of wind tunnel testing of the prop-fan installation would be to obtain valid data for correlation with flight-test data. The use of the testbed full-size nacelle, with its associated blockage problems, would place such serious doubt on the validity of data obtained from wind tunnel testing, this type of test is not justifiable.

Low-Speed Wind-Tunnel Testing Validity A full-scale, low-speed wind tunnel test of the operational prop-fan installation is not recommended. A limited amount of data only can be obtained from such tests due to the physical size of the testbed drive system nacelle. A possible installation of the GIl testbed in the AMES 40x80 foot wind tunnel is shown in Figure F-2. The 18.3 m (60 ft) span of the GIl fits snugly into the test section with no apparent problems in adapting the aircraft to the Ames strut-type mounting system. Based on NASA experience with other full-scale aircraft installations in the 40x80 tunnel, ARC personnel suggest that component data only can be expected to be valld. Total balance data would not be valid, in part, due to wind tunnel wall interference with the aircraft. The KC-135A testbed article would present an even more serious problem. Because of its size, only a segment of the fuselage and wing span can be accommodated in the tunnel as shown in Figure F-3. This type of arrangement negates the possibility of obtaining data directly applicable to the total aircraft configuration.

Although the acquisition' of basic aerodynamic data on the testbed con- figurations appears unlikely from these types of installations, the acquisition of some valid component information, such as pressure data over the wing and nacelle, should be possible. Total pressure data in the wake of the Wing/nacelle would also be valid. Strain-gage information on the propeller blades, propeller shaft, wing/nacelle structure, flaps, and other components should also be valid. Angle-of-attack excursions of the installation will provide useful information for flutter analysis of the prop-fan blades.

11lNNEL TRIPOD MOUNT Figure F-2. Gil Installation in Ames 4OXOO Tunnel RIGUT WING B. L. 217.000 /\~-----PR-0-P-E-LL-E-R--RO-T-A-TI00 1 15.3 m (50.16 FT) )( LEFT lUNG

/ DUMMY '\BL 315.00

/ FUSELAGE

I

\\ )

PITCH ACTUATOR

21.5 m (70.75 FT) ----I I

DUMMY FUSELAGE AFT FAIRING I I

I i

i

I STABILIZER ------- 24.4 m (80 FT) Figure F-3. KC-l35A Installation in Ames 4OxOO Tunnel The necessity for low-speed testing in the development of any aircraft.

particularly for flight safety. and to identify the aircraft takeoff. landing.

and low-speed flight characteristics. is recognized. This type of information is more readily available through testing of scale models of the testbed con- figurations and will be addressed later in this section.

Testing Costs Relative to Necessary Test Requirements Because costs for wind-tunnel model design. development. fabrication. and operation have escalated in recent years. the requirements for testing have be- come more stringent. Wind tunnel testing requirements for the testbed develop- ment should. therefore. be based on absolute need. rather than on a desire to increase the prop-fan technology data base.

OPERATIONAL READINESS STATIC TEST STAND TEST PLAN The Testbed Program Objectives that can benefit by the test stand static tests are: (1) the investigation of propeller generated near and far-field noise. and (2) the proof testing of the operational aspects of the prbp-fan/ gearbox/drive train assembly and systems operation.

The static test. as presented in Figure F-1. would be the initial test of the complete prop-fan/powerplant/nacelle assembly in the QEC configuration. At the conclusion of this testing. the QEC assembly would be removed from the test stand and installed on the flight-research aircraft.

A recommended facility for the drive system static test is the Lockheed- Georgia non-metric engine test stand shown in Figure F -4. Assuming that Lockheed-Georgia would modify the testbed aircraft and install the prop-fan assembly. it would be advantageous to conduct the static test with the same personnel as would be involved in·the flight testing of the testbed aircraft.

Instrumentation for the static test stand would include. but not be limited to. equipment to measure nacelle surface· pressures •. wake pressures at several longitudinal pOSitions aft of the prop-fan plane. pressure rakes for inlet in- vestigations. blade strain gage. acoustic transducers for near- and far-field noise measurements. prop-fan blade position. and power plant parameters.

Measured data requirements would include information to verify the structural integrity of the prop-fan installation. i.e •• prop-fan blade, gear- box, drive train and attachments, and acoustic environment parameters to iden- tify prop-fan sonic pressure intensity and direction. Also, propulsive data that would include pressure profiles for compressor and oil cooler inlets and the power plant exhaust, as well as conventional power-plant parameters, would be required.

Figure F-4. Static Test Stand The manhour/cost estimates for the static test, given in Table F-III, in- clude the cost of the engine test stand start-up phase.

TESTBED AIRWORTHINESS AND TECHNOLOGY VALIDATION TESTS The prop-fan operational environment encompasses both high- and low-speed flight regimes. It is possible that some testbed design features will be com- promises between the high- and low-speed design point performance considerations and high- and low-speed safety-of-flight considerations. To investigate these areas of concern, and to provide supportive and validation data to the testbed aircraft prior to flight, low- and high-speed model wind-tunnel tests are recom- mended. These tests are summarized in Figure F-1.

t-3 ~ I-' CO t'%j GULF STREAM II I KC-135A I H GELAC ETS AEDC 16T-IISIH NASA tRe 161'U'1 GELAC LSWT H GELAC ETS AEDC 16T-IISWT H ELAPSED TlME, TEST PREPARATION AND SET-UP 50 WEEKS 26 WEEKS 18 WEEKS : 18 WEEKS 50 WEEKS 26 W,EEKS ~ t-'.

I tj TEST DURATION 4 WEEKS 4 WEEKS 3 WEEKS 4 WEEKS 4 WEEKS 4 WEEKS i .

P.

26 WEEKS 26 WEEKS 4 WEEKS 8 WEEKS DATA ANALYSIS 8 WEEKS 26 WEEKS t-3

I

1 §

$ CO TEST SUPPORT - MANIIOURS AND CHARGES: Il/H $ MIH MIH $ HIli $ 11/11 $ MIll $ -I-' AERODYNAMICS 400 2,000 480 440 400 2,000 Pl tj 1,600 200 1,600 p.

WIND TUNNEL SUPPORT I en 2,520 200 PROPULSION 1,000 1,000 2,520 I rt Pl 2,800 1,680 720 ADVANCEn STRUCTURES 1,680 720 rt J t-'.

STAB II. tTy A~1l CONTROL 360 880 360 n 80 360 80 I t-3 MOIlEL DESIGN 1,900 2,320 2,320 I CO , en ICXI'ERIMENTAL SHOP (FABRICATION) 7,820 6,700 3,120 15,320 rt EN(;IN~:ERING FLIGHT TEST 1,600 1,600 t%l I en rt t-'.

START-lIl' ENGINE TEST STAND 4,500 4,500 ~ rt III R ECT CIIAR(; ES $100,000 CO $100,000 P.

EN(; INt: INSTALLATION ON ETS 11,280 $22,560 11,280 $22,560

r

a c: Ii WIND TliNNEL DIRECT CHARGES en MATER IAL COSTS SI5,OOO $4,000 $30,000 Pl tj p.

III RECT CHAR(;ES $0,000 S200,OOO $6,000 t;j t-'.

I Ii $122,560 12,2!.O .... ),h40 TOTALS: 20,540 17, -!40 521,IlOO ~2{)!~ ,00~D 20,540 $122,560 24,840 $36,000 CO n • rt (') TPTAL MANIIOUI{S ')'),7hO 41),1811 ::r'

J

Pl IJIRECT C1lAR(;ES ~ 31.7, r;f)() $1,)8,560 I Ii ()Q - -- -- -- CO en w W \0 High-Speed Model Wind-Tunnel Testing The high-speed wind tunnel test plan will contribute to the satisfaction of the Objectives of the Advanced Turboprop Testbed Program Plan by the following: (1) Validation of prop-fan blade classical and stall flutter characteristics.

(2) Provision of the necessary aerodynamic and structural data required for a testbed airplane flutter analysis.

(3) Verification of the propulsive efficiency of the prop-fan installation for the individual testbed candidates.

(4) Investigation of the wing/nacelle/prop-fan interactions through flow- field studies.

Semi-Span High Speed Wind Tunnel Test - NASA has previously conducted wind- tunnel tests to investigate the uninstalled performance of several prop-fan con- figurations. Prop-fan swirl effects and interactions on a supercri tical wing were investigated at the Ames 14-foot wind tunnel with the use of a slipstream simulator. Recently, a powered semi-span model using a 0.62 m (2.0 ft) prop-fan and a supercritical wing was tested at the Ames 11 x 14-foot wind tunnel to in- vestigate the installed effects and interactions of the nacelle and wing. These tests, and others already planned. will provide a sUbstantial data base for analytical studies of the prop-fan testbed interaction question.

A possible alternative to the recommended high-speed wind-tunnel model would be the adaptation of an existing 1/8. 8-scale semi-span GIl model. To scale the prop-fan correctly a 0.33 m (1.08 ft) diameter propeller would be re- quired.

The selection of model scale and test site are. inseparable. The critical dimension for establishing model scale is the prop-fan diameter. The use of existing prop-fans with a 0.62 m (2.0 ft) diameter would result in a scale, relative to the 2.89 m (9.5 ft) prop-fan of approximately 0.21. NASA already has a drive system for the 0.62 m (2.0 ft) prop-fan. however, the prop-fan diameter could be increased to 1.24 m (4.0 ft) with this drive system. and the model scale doubled; i.e., the scale would be 0.42. Table F-IV shows the esti- mated characteristic dimensions and the estimated cross-sectional area of a semi-span model for each testbed candidate, and Figures F-5 and F-6 illustrate a possible installation of a model for each candidate testbed aircraft.

TABLE F-IV. WIND TUNNEL MODEL CHARACTERISTICS Wind Tunnel Model Characteristics Characteristic Dimension Gulfstream II KC-135A Prop-Fan Diameter - m (ft) 0.62 (2.0) 0.62 (2.0) Model Length - m (ft) 5.1 (16.78) 11.55 (28.6) Wing Semi-Span - m (ft) 2.2 (7.22) 4.19 (13.74) Cross-sectional Area (est.) - 0.309 (3.32) 0.552 (5.94) 2 2 m (ft ) Prop-Fan Diameter - (ft) 1.24 (4.0) 1.24 (4.0) m Model Length - m (ft) 10.23 <33.56) 17.43 (57.2) Wing Semi-Span - m (ft) 4.4. (14.44) (27.48) 8.37 Cross-Sectional Area (est.) - (13.26) 1.23 2.21 (23.76 ) m (ft2) t----4.9 a U6.0 l'Tl .~ i ./ SCALI • 0.:'1 Figure F-5. Gil in AEDC 16T Tunnel 34-1 (16.0 FT)~ TUNNEL SIDEWALL\ , JI ~ 0.62 m (2.0 FT) SCALE = 0.21 Figure F-6. KC-l35A in AEDC 16T Tunnel Figure F-7 compares the model-to-test-sectional area ratio for three different wind tunnels using the solid-wall choking as the limit. The NASA-Ames 14-Foot and 11-Foot Tunnels have slotted test sections, while the AFDC 16T has a porous wall test section. The model-to-test-section area ratio for both tunnel types may be permitted to exceed the solid-wall choking limit of Figure F-7. It is generally accepted that an area ratio of 0.05 is the limit for models in vented test sections, however, this depends on the tunnel and its porosity.

The data of Figure F-7 show that a model at a scale of 0.21 could meet the area ratio requirement of the Ames 11-Foot Tunnel, but would be marginal for the KC-135A. The limiting factor is model length and/or wing span. The 0.42 scale can be used only for the GIl and only in the AEDC 16T. The KC-135A can be tested only at the AEDC 16T at a scale of 0.21. If the GIl is the selected testbed configuration, a O.42-scale model would require the development of new prop-fan models and some adaptation of the existing drive system. Although this would provide new data for use in determining scale effects, development of this model would incur high costs in model construction, testing, and correlation with previous testing. Based on these considerations, Lockheed recommends the O~21-scale~ O~62 m (2~O ft) prop-fan diameter for the testbed wind tunnel model wi th testing performed in the AEDC 16T tunnel, where wind-tunnel wall interference would be at a minimum level.

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HACK NUHBER Figure F-7. Tunnel Blockage An additional consideration in the selection of the AEDC 16T is the lower Mach number capability. This would provide data in the low-speed flight regime that is not currently planned.

Instrumentation requirements would include force balance for prop-fan, surface pressure taps for wing and nacelle, flow-field rakes for measuring surface pressures and flow angularity, strain-gages for blades, and wind-tunnel wall pressures.

Force and pressure data at several Mach number and Reynolds number combina- tions are needed; prop-fan blade angle is of prime importance so that a feathered flight position and thrusting positions from flight idle to maximum power can be tested. Since the proposed model would be a semi-span configura- tion, only longitudinal data will be measured. Although the recommended drive system for the prop-fan testbed would be the DDA XT101, a free turbine design which operates at near constant RPM, the wind tunnel test can be structured to test with RPM as a variable in order to investigate excursions from the desired value.

The manhour/cost estimates for the high-speed wind tunnel test of the Test Program Plan are presented in Table F-III.

High-Speed Flutter Test - One problem area identified with the selection of the GIl as the test-bed vehicle is a possible reduction in the testbed aircraft flutter speed. Preliminary flutter analyses have shown that this problem would not affect the prop-fan testbed design point, but has been identified at off- design conditions (high speed at low altitude), and modifications to the wing structure have been suggested to overcome this problem.

The high-speed flutter considerations apply only to the GIl and are not anticipated for the KC-135A testbed.

A preliminary flutter analysis of the GIl indicates that the installation of prop-fan powerplants at WS 145 reduces the wing flutter speed to an un- satisfactory level, and that a 60-percent increase in torsional stiffness from WS 145 inboard is required to restore the flutter speed to a satisfactory level.

In addition to the flutter analysis, a minimum-risk testbed program would also include a high-speed flutter model test to verify the transonic wing and whirl flutter stability of the aircraft. A dynamically scaled model of the complete aircraft, including windmilling prop-fans, would be tested in Freon in the NASA Langley 16-foot Transonic Dynamic Tunnel (TOT). The test would reduce the uncertainty in the analytical results, due primarily to inaccurate repre- sentation of the transonic, unsteady, aerodynamic forces in the wing and prop- fan.

The flutter model scale would be approximately 0.13 and would be compatible wi th the recommended test site: Langely Research Center's 16-foot TOT. The model would require a scaled prop-fan capable of windmilling operation.

Instrumentation for this model would include wing, nacelle and empennage accelerometers, wing spar strain gages, and high-speed motion pictures.

Test parameters would include wing variable fuel distributions, atmospheric densi ty and dynamic pressure, and Mach number. Data output would provide flutter speed and frequency and sub-critical damping.

The manhours/cost estimates for the flutter test and model development are presented in Table F-III.

Low Speed Wind Tunnel Testing The purpose of the recommended low-speed wind-tunnel test is the airworth- iness verification of the testbed aircraft. Preliminary examination of critical engine failure during takeoff on the KC-135A indicates the availability of adequate control power to overcome the addition of the prop-fan thrust so that low-speed wind tunnel test of the KC-135A is not required. As a more definitive prop-fan design is developed, control power for the KC-135A would be continually monitored to ensure the airworthiness of the testbed aircraft. The prop-fan tip clearances between the fuselage and inboard primary engines do not significantly affect the operation of the high-lift system. Longitudinal control effective- ness could be determined at the low Mach number spectrum of the recommended semi-span test. While some change in effectiveness of flaps and horizontal tail is expected, estimates have not indicated a critical situation.

The GIl thrust requirements are more nearly matched to the thrust available from the prop-fan than is the case for the KC-135A, and to ensure that the GIl testbed vehicle will be completely airworthy, Lockheed recommends a low-speed wind-tunnel test to examine the low-speed longitudinal, lateral, and directional aerodynamic characteristics. GAC advises that a 1/10-scale, low-speed model of the GIl is available that could be tested in the Lockheed-Georgia Low Speed Wind Tunnel, which has a test section 4.9 m x 1.0 m (16 ft x 23 ft). Lockheed- Georgia also has 50 hp electric motors available that could be used to drive the prop-fan to simulate thrust and slipstream effects on the testbed aircraft.

Testing requirements would include: o Basic longitudinal stability and control characteristics with and without prop-fan power.

o Basic lateral-directional stability levels with and without prop-fan power.

o Rudder effectiveness with and without prop-fan power.

o Lateral control effectiveness with and without prop-fan power.

Model requirements would include multiple elevator, aileron, rudder, spoiler and flap positions, static pressure measurement capability, and the operating prop-fans capable of simulating variable thrust levels.

Instrumentation requirements would include the basic tunnel balance system, a force balance for the prop-fan, surface pressure taps on the wing and nacelles, and flow-field rakes for measuring wake pressures and flow angularity.

Manhour/cost estimates for this low-speed test are presented in Table F-III. These costs include the estimate for fabrication of a new wing designed to accommodate the prop-fan drive train adapters required on the prop-fan in- stallations.

APPENDIX G - LIST OF SYMBOLS AND ABBREVIATIONS Cross-sectional Area-to-Maximum Cross-sectional Area Ratio AlA MAX - Reference Area AREF - Wind Tunnel Test Section Area AWT - Aircraft AlC - Arnold Engineering Development Center AEDC - Ames Research Center ARC - British Aerospace Corporation BAC Buttock Line BL - Degrees Centigrade (Temperature) °c Rolling Moment Coefficient Due to Sideslip Pitching Moment Coefficient Due To Angle-of-Attack - Yawing Moment Coefficient Due to Sideslip - Wing Chord Side-force Coefficient Due to Sideslip Propeller Normal Force - Center-of-Gravity Centimeter cm - Fuselage Diameter DF - Nacelle Diameter DN Prop-Fan Diameter Dp - Nacelle Diameter-to-Prop-fan Diameter Ratio Dn/Dp dB - Decibels - Detroit Diesel Allison DDA - Degrees (Angle) deg - Direct Operating Cost DOC - Equivalent Air Speed EAS Evaluation Criteria EC - Evaluation Criteria Ranking ECR Prop-fan Tip-to-Fuselage Clearance F of - Degrees Fahrenheit (Temperature) - Federal Aviation Administration FAA FAR - Federal Air Regulations - Foreign Object Damage FOD fps - Feet per Second FS - Fuselage Station FT - Feet GAC - Gulfstream American Corporation gal - Gallon GFE - Government Furnished Equipment GW - Gross Weight Hz - Hertz in Inch ISA. International Standard Atmosphere oK - Degrees Kelvin (Temperature) kg - Kilograms kN Kilonewton KPa Kilopascal KTAS - Knots True air Speed KTS Knots kW - Kilowatts LB Pound LID - Lift-to-Drag Ratio L/D - Maximum Lift-to-Drag Ratio MAX LeRC - Lewis Research Center LH - Left Hand LRC - Langley Research Center LSWT - Low Speed wind Tunnel m - Meter M - Mach Number Design Dive Mach Number MD MAC - Mean Aerodynamic Chord M&DC - Material and Direct Charges mm - Millimeter I-1RT - Maximum Rated Thrust m/s - Meters per Second N Newtons or Background Noise NRT - Normal Rated Thrust NTS - Negative Torque Sensing OSPL Overall Sound Pressure Level

-

OW Overwing

-

PPS Power Plant Station P&W Pratt & Whitney QEC Quick Engine Change Radians rad RH Right Hand

-

ROM Rough Order of Magnitude RPM Revolutions per Minute

-

S Prop-fan Signal Ratio SIN Signal-to-Noise

-

- Slipstream-wetted Wing Area-to-Total Wing Area Ratio SWSLIp/SWTOTAL SHP - Shaft Horsepower SHP/D Power (Disk) Loading SL - Sea Level SLS - Sea Level Static SPL - Sound Pressure Level TDT - Transonic Dynamic Tunnel UW Underwing - Design Dive Speed V D - Stall Speed Vs V Propeller Tip Speed T WL - Water Line WRP - Wing Reference Plane WS - Wing Station X/L - Location as a Fraction of Total Length ZFW - Zero Fuel Weight Il Increment TI - Location as a Fraction of Wing Semi-span Propeller Efficiency TIp - Wing Leading Edge Sweep ALE

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--T- REFERENCES 1. M. L. Boctor, C. W. Clay, and C. F. Watson, "An Analysis of Prop-Fan/ Airframe Aerodynamic Integration ," NASA CR-152186, Boeing Commercial Airplane Co., October 1978.

2. J. A. Baum, P. J. Dumais, M. G. Mayo, F. B. Metzger, A. M. Shenkman, and G. G. Walker, "Prop-Fan Data Support Study," NASA CR-152141, Hamilton Standard, February 1978.

3. J. C. Muehlbauer, J. G. Hewell, Jr., S. P. Lindenbaum, C. C. Randall, N. Searle, and F. R. Stone, Jr., "Turboprop Cargo Aircraft Systems Study Phase I," NASA CR-159355, Lockheed-Georgia Co., November 1980.

4. R. N. Latz, "KC-135 Power Spectral Vertical Gust Load Analysis ," AFFDL- TR-66-51, Vol. II, July 1966.

5. K. Wilkinson and J. Zambito, "Final Flutter Analysis on the Gulfstream II Airplane ," Grumman Aircraft Engineering Report No. LD 1159-1831, May 1961.

6. R. F. Mohrman, "Aeroelasticity Grumman Design G-1159," Grumman Aircraft No. LD 11S9-102.2, Section VIII, October 1965.

7. D. Wist and J. K. Zentgraf, "Final Results Ground Vibration Survey of the Grumman Gul fstream II," Grumman Aircraft Engineering Report LD 1159- 203.1, January 20, 1961.

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