AIAA-2013-1099

American Institute of Aeronautics and Astronautics

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Low Speed Powered Lift Testing of a Transonic Cruise Efficient STOL Military Transport

Fred J. Barberie1, Andrew T. Wick2, and John R. Hooker3 Lockheed Martin Aeronautics Company, Marietta, GA, 30063

Cale H. Zeune4 Air Force Research Laboratory, Wright-Patterson AFB, OH, 45433

he Speed Agile Concept Demonstrator (SACD) is an Air Force Research Laboratory (AFRL) sponsored contract program whose primary objectives were to raise the Technical Readiness Levels (TRL) of advanced mobility

component high lift technologies to a Level 5 or better through testing in a large scale relevant environment with jet engine propulsion simulation. The Lockheed Martin SACD air vehicle has a highly integrated propulsion system which significantly contributes to the high lift performance required to meet the short takeoff and landing (STOL) mission requirements. The SACD Task 2 effort which this paper documents was to validate the low speed high lift aerodynamic characteristics with a 23% scale Low Speed Powered Lift Model (LSPLM). The LSPLM was tested in the Air Force Arnold Engineering Development Center (AEDC) National Full-scale Aerodynamic Complex (NFAC) wind tunnel complex at NASA Ames in California from March through October of 2011. There was also a high speed compliment to this low speed test to validate the transonic performance which is documented in Reference (3) for completeness.

Nomenclature

AEDC = Arnold Engineering Development Center FG = Gross Thrust AFRL = Air Force Research Laboratories HPLS = Hybrid Powered Lift System ACMS = Actuator Control and Monitoring System IB = Inboard AIP = Aerodynamic Interface Plane LH = Left Hand AoA = Angle of Attack LSPLM = Low Speed Powered Lift Model AoS = Angle of Sideslip N1C = Corrected Fan Speed ATI = Advanced Technologies, Inc. NASA = National Aerospace and Science Administration CC = Circulation Control NFAC = National Full Scale Aerodynamic Complex CCW = Circulation Control Flaps OB = Outboard CFD = Computational Fluid Dynamics PSC = Preferred System Concept CL = Lift Coefficient REN = Reversing Ejector Nozzle = Ambient to Seal Level Pressure Ratio RH = Right Hand DAS = Data Acquisition System SACD = Speed Agile Concept Demonstrator DLPC = Core Distortion Limit Parameter STOL = Short Takeoff and Landing DLPF = Fan Tip Distortion Limit Parameter TRL = Technical Readiness Levels ECMS = Engine Control and Monitoring System WI = Williams International Company

1 Aeronautical Engineer, Senior Staff, Lockheed Martin Aeronautics Company, Marietta, Georgia

2 Aeronautical Engineer, Senior, Lockheed Martin Aeronautics Company, Marietta, Georgia

3 Aeronautical Engineer, Senior Staff, Lockheed Martin Aeronautics Company, Marietta, Georgia

4 Aerospace Engineer, Air Vehicles Directorate, AFRL/RQAA, 2130 8th Street

T

51st AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition07 - 10 January 2013, Grapevine (Dallas/Ft. Worth Region), Texas

AIAA 2013-1099

Copyright 2013 by Lockheed Martin Corporation. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.

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Figure 1, AEDC NFAC Wind Tunnel Complex

I. Introduction The SACD LSPLM program was begun in 2008 with the start of model design, and subcontracting to Advanced Technologies Inc. (ATI) for model fabrication and Williams International Co. (WI) to provide the FJ44-4A mixed flow turbofan engines, test hardware, and support. The testing was conducted in 2011 with an initial entry into the NFAC 80x120 wind tunnel in the spring, and a later entry into the 40x80 tunnel in the fall. The AEDC NFAC wind tunnel complex is shown in Figure 1 with the inlet to the 80x120 in the fore-ground. The LSPLM installed in the 80x120 test section is shown in Figure 2, and in the 40x80 test section in Figure 3. The priumary objective of the initial entry in the 80x120 was to perform static thrust calibrations and acquire data to assess blockage corrections between test sections. In the 80x120 the aft louvers were partially closed ensuring no natural or induced winds were incurred during the static calibrations. The model to test section span ratios are 34% and 51% in the two tunnels, and the test section cross sectional area ratio is 3.4. The majority of wind on data was obtained in the 40x80 where higher airspeeds were available above 100 knots consistent with aircraft approach operations (80x120 limited to 100 knots), in addition to cruise.

II. Test Background and Model Description

The SACD key enabling propulsion technology is the Lockheed Martin developed Hybrid Powered Lift System (HPLS)2. The principle components of the HPLS are the Reversible Ejector Nozzle (REN) and the Circulation Control Wing (CCW) which are highlighted in Figure 4. The REN and CCW have been designed and integrated to achieve an optimal balance between both the low speed high lift performance requirements which the LSPLM test demonstrated, and also high speed efficient transonic cruise performance. The REN is a multi function

Figure 2, LSPLM in NFAC 80x120 Figure 3, LSPLM in NFAC 40x80

Figure 4, Hybrid Powered Lift System Design

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nozzle which employs several control surfaces that articulate to align for cruise, stow flaps and deploy the bucket for reversing on the ground, and deflect the exhaust flow to create a jet flap effect over the inboard wing in addition to direct thrust lift. In this high lift mode a slot aperture is opened through the wing behind the nozzle to entrain flow providing thrust augmentation during takeoff. The outboard wing CCW system utilizes jet flap blowing through a thin, tapered, full span slot between the upper wing and flap. The source of CCW high pressure air is from the engine fan bypass stream. The CCW jet flap is created from the Coanda effect from attached high momentum flow turned over the deflected upper flap surface enabling very high airfoil circulation.

Changes to the Task 2 scale LSPLM were required relative to the full scale Task 1 airplane configuration to adapt for wind tunnel testing. The external outer mold-line (OML) differences are shown in the side-by-side CFD views in Figure 5, with the airplane shown on the starboard side and the LSPLM on the portside. The principle changes were to accommodate the engine installation with the airplane being a four engine design and the model a twin engine. In addition to the nacelle OML distortions, the inlet also required modification to a pitot type design to ensure flow quality to the engine compatible with its civil certification limits. Internally there were also differences with the fan bypass off-take for the CCW system again due to the integration limitations of the WI engines; however, these were slight and did not compromise the performance demonstration objectives.

There was a significant CFD3 optimization effort to develop both the model external moldline distortions and internal flow paths to ensure maximum performance and correlation with the Task 1 airplane, with some the details highlighted in Figure 6. The entire external moldline was modeled in CFD including the struts and even the test section to capture the wall effects. Internally the fan bypass and turbine exit streams were modeled including the daisy chute lobe mixer, with boundary conditions represented by engine cycle data. The primary nozzle, REN geometry, CCW off-take, plenum and slot geometry, and wing leading edge slats were all modeled in detail and optimized. Finally, there was significant development of the inlet including the boundary layer diverter details to ensure that clean flow required by the engine would be delivered at all planned test conditions.

Figure 5, Task 1 Airplane vs. Task 2 LSPLM Differences

Figure 6, LSPLM CFD Moldline Optimization

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The WI FJ44-4A engines were chosen for the model as they represented the best choice for matching fan pressure ratio (critical for CCW slot blowing momentum) and scaled thrust equivalency. A comparison of the Task 1 airplane gross thrust coefficient at the sizing condition vs. the Task 2 LSPLM in the tunnel is shown in Figure 7. The engines were in general limited to 93% corrected fan speed to maintain adequate margin to their inter-turbine temperature (ITT) control continuous operation limit. However, as can be seen in Figure 7 this maximum thrust was more than adequate to bound the range of the thrust required by the airplane. The thrust coefficient scaling, engine physical integration, and test section blockage were all factors in setting the model scale at 23%.

The LSPLM also had 12 actuated control surfaces as shown in Figure 8. They were designed to improve run time efficiency in acquiring more data minimizing tunnel stop/start cycles to make position changes. These include the elevators, rudders, inboard and outboard CCW flaps, and the REN mixing and aft flaps, symmetrical left and right. The control surfaces were independently controlled by a laptop based Actuator Control and Monitoring System (ACMS) located in the control room. The ACMS digital interface was designed to provide independent control of each surface including position feedback, error checking, and recording capability independent of the NFAC Data Acquisition System (DAS).

The range of control surfaces deflection were dependent upon air-loads due to airspeed but were designed to be unrestricted during low speed high lift testing. The CCW flaps were capable of deflecting from 0 to 90; the rudders 30; the elevators from -40 (TE up) to +15; the REN mixing flap from -135 (stowed for reverse) to 60, and the REN aft flap from 0 to 45. Only the rudder, elevator and CCW flaps were intended for deflections up to 10 above 150 knots.

In addition the LSPLM also had other fixed control surfaces which were installed or removed for long blocks of testing. These included wing LE slats, inboard and outboard (i.e. full span) spoilers, and a REN area relief flap which was deployed during reverse and high drag approach configuration testing.

The engines were also controlled remotely from the control room with an Engine Control and Monitoring System (ECMS) provided by WI as shown in Figure 9. The ECMS was adapted from the normal test cell configuration to independently control the two engines in the model and it included an interface switch panel (FADEC power, start, stop, etc.), throttles, and an engine health and monitoring laptop system. The FJ44-4A engines were FADEC controlled with a digital ECMS interface which was also independent of the NFAC DAS. Outside of the ECMS there were a few analog engine sensors incorporated that interfaced with the NFAC DAS for recording (engine face total pressure and temperature P2 and T2, and fan speed N1). Training was provided to A&P mechanics that operated the engines during the test.

Figure 7, Airplane vs. LSPLM Gross Thrust

Figure 8, LSPLM Actuated Control Surfaces

Figure 9, Engine Control and Monitoring System

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In addition to these systems and also instrumentation, the LSPLM also required a fuel delivery system for the engines, a zonal ventilation system to ensure both purging of potential flammable vapors as well as local component cooling as required, and a zonal fire extinguishing system in the event of a fire. NFAC provided and was principally responsible for the fuel system which included outside storage tanks, pumps and plumbing into the model. JP-8 was used due to its low flammability safe handling qualities. A high pressure/volume air compressor was also supplied and located outside the tunnel with plumbing into the model for ventilation. An intercooler was later also included downstream of the compressor for the 40x80 testing to lower the air delivery temperature and provide additional temperature margin. The sizing was consistent with aircraft design standards for bay changes per minute and the ventilation was networked internally to provide air distribution into the forward compartment of each engine bay and throughout the center fuselage. The flow rate was controlled by a fixed number of calibrated nozzles which vented to near ambient inside the model. Ventilation holes were also provisioned around the annulus of the primary engine nozzles and in the rear of the fuselage to ensure positive flow from front to rear. Finally, three fire bottles were installed inside the model sized appropriately for each of the engine bays and center fuselage. To avoid confusion over location in the advent of a fire the fire bottles were wired together to activate from one switch. Thermocouples were installed throughout the model and were monitored real time both for local overheating and fire detection. Emergency procedures were developed to shutdown fuel, ventilation, and power before discharging.

III. Test Approach

The primary objective of the LSPLM wind tunnel test was to verify the low speed high lift aerodynamic performance of the SACD configuration; however, there were several other key objectives as mentioned earlier. The 80x120 test block objectives are summarized in Figure 10, and this entry provided the static thrust calibrations for cruise, reverse and high lift, as well as wind on data. With the tunnel aft louver set closed the test section was ideally configured for static calibrations without natural wind or induced wind interference. Reverse was only planned for the 80x120 due to recirculation concerns and the wind on portion was not executed due to engine surging that was observed during the static testing at intermediate engine power. An identical set of cruise and high lift configuration wind on data was obtained in both tunnels to provide a basis for assessing the wall correction blockage effects. The 40x80 test block objectives are summarized in Figure 11. This entry was where the principal high lift configuration, wind on data was obtained. The 40x80 testing acquired basic aerodynamic performance as well as control surface perturbation data to assess incremental affects. Additionally control surface effectiveness and stability data was acquired to develop a data base for the SACD Task 4 flight simulation modeling including control in pitch, roll and yaw axis. Finally, similar cruise configuration control surface effectiveness and stability data was planned to be acquired but due to a model failure only an engine powered off series was executed with the control surfaces in neutral positions.

Blk Description1 STATIC LH ENGINE NO. 1 NOZZLE EXIT AREA SHIM RUN FOR OPERABILITY - BARE NOZZLE2 STATIC LH ENGINE NO. 1 NOZZLE EXIT AREA SHIM RUN FOR OPERABILITY (MODEL CHANGE)3 STATIC LH ENGINE NO. 1 NOZZLE EXIT AREA SHIM RUN FOR OPERABILITY (MODEL CHANGE)4 STATIC RH ENGINE NO. 2 NOZZLE EXIT AREA SHIM RUN FOR OPERABILITY (MODEL CHANGE)5 STATIC RH ENGINE NO. 2 NOZZLE EXIT AREA SHIM RUN FOR OPERABILITY (MODEL CHANGE)6 STATIC LH ENGINE NO. 1 BASELINE CRUISE THRUST CALIBRATION RUNS7 STATIC LH ENGINE NO. 1 REN MIXER FLAP DEFLECTION THRUST ISOLATION RUNS8 STATIC RH ENGINE NO. 2 BASELINE CRUISE THRUST CALIBRATION RUNS9 STATIC RH ENGINE NO. 2 REN MIXER FLAP DEFLECTION THRUST ISOLATION RUNS10 STATIC CHECK WITH BOTH ENGINES OF BASELINE CRUISE THRUST CALIBRATION RUNS11 STATIC CHECK WITH BOTH ENGINES OF REN MIXER FLAP DEFLECTION THRUST ISOLATION RUNS12 STATIC CHECK WITH BOTH ENGINES OF LOW CCW FLAP DEFLECTION THRUST ISOLATION RUNS13 WIND ON CRUISE CONFIGURATION14 STATIC LH ENGINE NO. 1 HIGH CCW FLAP DEFLECTION THRUST ISOLATION RUNS (MODEL CHANGE)15 STATIC RH ENGINE NO. 2 HIGH CCW FLAP DEFLECTION THRUST ISOLATION RUNS16 STATIC CHECK WITH BOTH ENGINES OF HIGH CCW FLAP DEFLECTION THRUST ISOLATION RUNS17 STATIC LH ENGINE NO. 1 THRUST REVERSE CALIBRATION RUNS (MODEL CHANGE)21 STATIC LH ENGINE NO. 1 REN MIXER FLAP DEFLECTION AT HIGH LIFT TAKEOFF THRUST CALIBRATION RUNS (MODEL CHANGE)22 STATIC RH ENGINE NO. 2 REN MIXER FLAP DEFLECTION AT HIGH LIFT TAKEOFF THRUST CALIBRATION RUNS23 STATIC CHECK WITH BOTH ENGINES REN MIXER FLAP DEFLECTION AT HIGH LIFT TAKEOFF THRUST CALIBRATION RUNS24 WIND ON HIGH LIFT CONFIGURATION - BASELINE ENGINE POWER SWEEPS

Figure 10, LSPLM 80x120 Test Block Objectives

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Blk Description0 80x120 INLET DISTORTION AND REN AFT TO MIXING FLAP PARAMETRIC OPTIMIZATION1 WIND ON HIGH LIFT CONFIGURATION - ENGINE OFF2 WIND ON HIGH LIFT CONFIGURATION - REN MIXER FLAP PARAMETRIC SWEEP3 WIND ON HIGH LIFT CONFIGURATION - CCW FLAP PARAMETRIC SWEEP4 WIND ON HIGH LIFT CONFIGURATION - BASELINE ENGINE POWER SWEEPS5 WIND ON HIGH LIFT CONFIGURATION - ELEVATOR PITCH TRIM SWEEPS6 WIND ON HIGH LIFT CONFIGURATION - RUDDER EFFECTIVENESS SWEEPS7 WIND ON HIGH LIFT CONFIGURATION - ROLL CONTROL PARAMETRIC SWEEPS (MODEL CHANGES)9 WIND ON HIGH LIFT, HIGH DRAG APPROACH CONFIGURATION - BASELINE ENGINE POWER SWEEPS (MODEL CHANGE)10 WIND ON HIGH LIFT, HIGH DRAG CONFIGURATION - SPOILER DRAG EFFECTIVNESS (MODEL CHANGES)11 STATIC CRUISE CONFIGURATION BASELINE THRUST CALIBRATION CHECK RUNS (MODEL CHANGES)12 WIND ON CRUISE CONFIGURATION - ENGINE OFF

Initial static runs in the 80x120 were necessary to set the FJ44 fan operating line to ensure safe engine operation during the test. This was achieved by starting off with the primary nozzle exit area more open and then judging how much to reduce the area with shims after assessing the installed engine effects and knowing the sensitivity. The shims adjusted the nozzle height in 1/8 increments or roughly by 5 in2. Initially both areas were set at 225 in2 to start the testing.

Within each block a series of points were tested representing different model parametric configurations, engine power settings or in some cases angle of sideslip (AoS). An angle of attack (AoA) sweep was executed for each wind on point. The resolution (coarse or fine schedules) and range of each AoA sweep varied by block objectives and were tailored to optimize with the occupancy hours allotted.

IV. Instrumentation

The LSPLM was fully metric including thrust and the forces and moments were resolved by the NFAC balance scale system into absolute values and coefficients. Static gross thrust measurements for each configuration were obtained and correlated accordingly, then applied to the wind on data in the body axis in accordance with the established thrust-drag bookkeeping system. In addition, inlet airflow measurements were made based on calibrated throat static pressures to derive ram drag in the wind axis which together with the computed gross thrust was subtracted from the baseline measurements to calculate the set of thrust removed aerodynamic coefficients. Analytical corrections were also made for the exposed metric portion of the strut tips based on the CFD computed drag adjusted for the as tested cross sectional area. A drag as well as pitching moment correction were made based on the drag acting through the strut tip centroid and offset from the model center. Following standard NFAC procedures weight tares were also applied as a function of angle of attack, and for each sideslip angle.

Pressures were primarily measured on 2.5 and 15.0 psid electronic scanning pressure (ESP) modules which were all connected to an Initium hub inside the model which multiplexed the signals to the control room for monitoring and recording. The lower range was used to monitor internal cavity (safety of test) and external pressures including the engine inlet AIP rakes, whereas the higher range was used to monitor the engine fan bypass, nozzle, CCW off-take and plenum. In addition an analog engine inlet PT2 total pressure sensor was installed in the inlet in front of the fan.

Temperatures were primarily measured on electronic scanning thermocouple modules which were connected to a VSI hub and multiplexed to the control room for monitoring and display. These were distributed throughout the internal cavity (safety of test) as well the engine inlet AIP rakes, engine nozzle surface, and CCW off-take. In addition an analog engine inlet TT2 total temperature sensor was installed in the inlet in front of the fan.

Strain gauges to measure the loads were installed on the actuator horns for each of the control surfaces which were ECMS controlled, i.e. rudders, elevators, IB and OB CCW flaps and REN mixing and aft flaps. The calibrations for these were developed by ATI through application of reference loads on each surface. The calibrations were incorporated by NFAC into the data reduction to compute engineering units. The loads were monitored during the test but never observed to approach any limits.

Figure 11, LSPLM 40x80 Test Block Objectives

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V. Internal Aerodynamic Results

The static nozzle area shimming to set the safe fan operating line was the first objective in the 80x120 test. The WI FJ44-4A fan operating line data taken for these initial cruise baseline configurations is shown in Figure 12. The key nominal engine conditions are the normal operating line, the 5% surge margin line, and the surge line. The goal for the baseline was to run near or slightly below the normal operating line, and allow excursions up to the 5% surge margin line for mixing flap deflection and reverser bucket operations. In the figure the Run 9 (green) data is the LH engine originally taken with the 225 in2 reference nozzle area. As seen, it was very close to the nominal op line which provided high confidence in the pre-test CFD analysis. It was decided then to open the LH nozzle area another 5 in2 to provide an additional bit of margin for REN mixing flap, reverser, inlet distortion, and other testing unknowns. That data is shown in Run 10 (red) and it can be seen that the fan operating line moved down only slightly, less than the engine cycle sensitivity predicted. Based on these LH engine results it was decided to skip the initial 225 in2 position for the RH engine and open the area to the same 230 in2. That data is shown in the figure as Run 11 (yellow) and its fan op line turned out to be lower which was OK, but showed a lack of symmetry suggesting that there was some unknown handedness between the two engine installations. Rather than go back to close the RH nozzle area to 225 in2 it was decided to leave the slight fan operating line asymmetry (same nozzle area) and continue on to avoid falling further behind on program schedule.

After the nominal nozzle areas were set, the REN mixing flap deflection testing was conducted and the fan operating line moved upward as expected until 45 was right at the 5% surge margin limit, and 50 was clearly over (not shown in figure). Based on that it was decided to set a REN mixing flap test limit of 45 even though the flap was capable of physically deflecting to 60. Later during the static thrust reversing calibration testing and at only intermediate power on the LH engine it was clear that the reverser mechanism and turning were throttling the effective flow area excessively and the fan operating line had exceeded the 5% surge margin line and in fact was approaching the surge line (not shown in figure). Based on those results it was decided to discontinue thrust reverser testing even though some useful data could have been obtained at low engine power.

Figure 12, LSPLM Fan Op Line Nozzle Area Shimming

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The static thrust calibrations were performed in accordance with the SACD thrust-drag Bookkeeping system and each engine was calibrated independently in the cruise configuration, and then together to confirm the combined thrust. The independent calibrations were then applied to the data reduction for calculation of the thrust removed aerodynamic coefficients. Note that these calibrations are for the computation of gross thrust which acts in the body axis and that inlet ram drag in the wind axis is also subtracted in the thrust removed wind on data. Ram drag is computed based on the measured inlet airflow and freestream velocity. The methodology assumes that thrust is fully recovered in high lift and that the differences are attributable to the aerodynamic drag. The corrected gross thrust (FG/) regression equations were 2nd order and correlated with fan bypass pressure ratio (PT16/P0). These engine static gross thrust calibrations are shown in Figure 13. The calibration shows a higher maximum fan bypass pressure ratio for the LH engine than the RH consistent with the fan operating line data results presented above. Additionally the RH engine shows approximately 5% higher thrust in the mid range pressure ratios which is not fully understood. The model was not designed for component build up to isolate the individual contributions to installed thrust and so the differences are further evidence of an unknown installation handedness. The individual IB and OB CCW stream thrusts were computed directly based on the CCW plenum total pressures, temperature and slot area. These were algebraically subtracted from the total thrust to also compute the individual primary nozzle stream thrust regression equations for analysis.

The REN mixing flap thrust turning effectiveness was assessed during the initial thrust calibration runs in the 80x120. The turning effectiveness as a function of the resultant thrust is shown in Figure 14. The effective thrust turning angle is the arctangent of the negative normal force divided by the axial force and the resultant thrust is the square root of the sum of the squares of the axial and normal forces. The data suggests that the effective thrust turning angle is about one-half of the actual REN mixing flap physical deflection or metal angle which is consistent with pre-test expectations.

The induced lift augmentation due to blowing effectiveness is shown in Figure 15 for the baseline power sweeps (including engine off windmilling) at 75 and 115 knots, the REN mixing flap at 30, and CCW flaps at both 60 and 90. The classic non-linear jet flap momentum coefficient induced lift augmentation that has been observed before is readily apparent in the LSPLM data as well. At 75 knots at both CCW 60 and 90 there is significant lift augmentation below a thrust coefficient of 0.4, and a reduced lift slope at higher thrusts even though still linear. Note that this is the thrust induced lift and that the thrust component due to turning has been removed. A similar trend is also observed at 115 knots for both CCW 60 and 90 where the non-linear thrust coefficient break is observed to be 0.2. The thrust coefficient break is consistent with the freestream dynamic pressure ratio between

Figure 13, Engine Static Thrust Calibrations

Figure 14, REN Turning Effectiveness

Figure 15, LSPLM CCW Lift Augmentation

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these two airspeeds which is approximately 2.3. The augmentation of lift at zero angle of attack suggests that most of the lift is augmented at the lower thrust momentum ratios. The data clearly demonstrates that the LSPLM exhibits a boundary between the Coanda effect (super-circulation) and the linear jet flap behaviour. Further it shows that the CC slot height, radius and flap chord are appropriate for SACD.

As mentioned earlier the reverse thrust testing was not complete due to fan surge margin issues, however, a limited amount of testing was accomplished on the LH engine before stoppage. The static thrust from these points is compared with the thrust from the equivalent runs in the high lift configuration with the CCW flaps all set to 90 so that the axial force differences are directly comparable. At the highest reverse power tested of 60% N1C the data suggests that a reverser thrust turning effectiveness of 65% was achieved as shown in Figure 16. It should be noted that 60% N1C is a very low power at less than 30% of max available forward thrust and reverser effectiveness trends arent linear with power, however, the data clearly supports that the reverser design is highly effective in turning the flow. Further studies are needed to understand the nature of flow turning separation reducing the effective engine nozzle exit area to improve the design.

A short series of runs were made early in the 40x80 to analyze the REN aft to mixing flap aerodynamic optimization. This was accomplished at 95 kcas with 2 engine power settings, 2 mixing flap settings and 3 aft flaps positions for each. The objective of this block was to either confirm pre test predictions or adjust the aft flap setting in the run matrix for the remainder of the test accordingly based on the results. The CFD predictions for SACD were that the aft flap should track the mixing flap angle one-to-one and this was confirmed at operationally relevant conditions as shown in Figures 17 and 18 at an AoA of 15. Figure 17 shows the CL result from the REN mixing flap of 30 and the aft flap swept from 15 to 30, whereas Figure 18 shows the result from the REN mixing flap of 45 and the aft flap swept from 25 to 45. Though the CL effect was weak the overall trend showed that the optimal CL was achieved when the aft flap and mixing flaps angles were set the same. Though CL was determined to more important with regard to optimization, the aft flap also has a significant impact on drag with lower aft flap deflections producing higher drag due to detached induced secondary flow underneath the aft flap, i.e. too large an expansion angle. This one-to-one flap angle relationship was set for the remainder of the test.

2.0

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AFT FLAP DEFLECTION ANGLE, DEG

SACD LSPLM NFAC 40x80 REN AFT FLAP SWEEP AT 30 MIX FLA(95 KCAS, 30.6 PSF, AoA 15, THRUST INCLUDED)

92% N1C

70% N1C

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SACD LSPLM NFAC 40x80 REN AFT FLAP SWEEP AT 45 MIX FLAP(95 KCAS, 30.6 PSF, AoA 15, THRUST INCLUDED)

92% N1C

70% N1C

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Speed Lines in %N1C2

FWD 810 LBS

REV 530 LBS

~65% REVERSER EFECTIVENESS

NOTE: BOTH FWD AND REV CONFIGURATIONS HAVE CCW FLAPS IN 90 POSITION

Figure 16, Reverser Effectiveness

Figure 17, REN 30 Aft Flap Optimization Figure 18, REN 45 Aft Flap Optimization

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The LH inlet AIP rakes were installed during the 80x120 entry and originally planned for removal during transition, however, not all inlet distortion survey objectives were accomplished and remaining points were carried over to the start of the 40x80 test. Validation of inlet design technology was not an objective for the LSPLM and the inlet distortion testing was only intended to clear the model AoA and AoS attitude envelope by verifying the inlet flow quality was within the FJ44-4A certification limits. This was accomplished early in the test because the AIP rakes were required to be removed due to fatigue life concerns and to mitigate the potential for ingestion damaging the engine. The AIP rake consisted of 40 total pressures which were located at the centroid of an equally area distribution based on 8 rake legs installed circumferentially and 5 probes per rake radially. Additionally there were 4 static pressures in the plane of the AIP rake and 8 total temperature probes installed to survey for exhaust reingestion during thrust reverse testing. The AIP pressure data were reduced with a WI inlet distortion analysis executable code which was integrated into the NFAC data reduction program. The principal parameter monitored during the test was DLPF which is the fan tip distortion limit parameter or surge margin consumed due to distortion, i.e. 100% indicative that the distortion had consumed the entire available stall margin and the engine were operating near surge. The core distortion limit parameter DLPC was also monitored but it was always effectively zero as expected from pre-test analysis. The DLPF trend vs. AoA during the cruise run at max power and zero AoS where the highest DLPF was recorded of 6.5% is shown in Figure 19. The distortion is essentially zero above zero AoA and slightly elevated and plateaued below. A further investigation into the rise revealed that the top-dead-center rake registered a pressure deficit of 5% on the probe 2nd from the outer duct wall which can be observed in the fan inlet pressure distortion isobars in Figure 20. This is indicative of ingestion of low pressure air from around the overhead boundary layer diverter separating the inlet from the underside of the forward chine. Subsequent 80x120 testing in the high lift configuration repeated the cruise results indicating that the circulation differences had no significant impact on the inlet flow field quality. Later testing in the 40x80 at 15 AoS (LH inlet leeward) showed even lower values of DLPF of around 2.5%. It should be noted that these distortion levels were extremely low and indicative of a well designed inlet. At the conclusion of the planned distortion testing the AIP inlet rakes were removed and appropriately no attitude or power restrictions were necessary.

The inlet rakes were also used to survey for exhaust gas reingestion during the thrust reverse testing. This potential for reingestion was a concern and the risk was that it could cause an instability and surge. Though very little reverser testing was accomplished as discussed earlier, the AIP rake thermocouple data showed no indication of reingestion.

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The CCW off-take and plenum performance were key to validating the HPLS performance for the test even though the off-take integration with the engine wasnt the PSC approach. The off-take mass flow was determined through integration by rakes in the diffusion duct midway between the inlet and first turn OB towards the plenum consisting of 3x3 equal area array of total pressures on 3 rake legs, 3 total temperatures (1 per rake) and 4 wall static pressures in the plane of the probes heads. Generally the flow uniformity was pretty good though there was some pressure deficit observed along the OB wall and in particular the lower corner. There was also a little bit of handedness suggesting that there could be some local separation beginning to form, likely due to overly rapid diffusion coupled with the initial turning away from the engine centerline. The flow measured during the initial thrust calibration runs

Figure 19, Max Fan Stall Margin Consumption with AoA Figure 20, Max Inlet Distortion at AoA -5

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is shown in Figure 21 for the LH and RH engines including a LH engine which shows negligible effect. As can be seen, the measured flows were very close to predicted based on the assumed pressures, losses and slot area.

The CCW pressure losses from the CCW inlet (engine station 16) through the off-take duct and plenum is shown in Figure 22. The instrumentation in the CCW plenum consisted of 2 rakes of 3 total pressures at the IB mid flap span and between the flaps, a single total pressure at the OB mid flap span, and static pressures. The CCW inlet pressure reference was adjusted from the measured fan bypass (engine station 13) accounting for the duct loss in accordance with the WI engine cycle model which is on the order of 3%. As can be seen the total pressure loss to the OB plenum is 10% at max flow as predicted, with 6% occurring in the short span from the inlet to the mid duct and 4% from the mid duct to the OB plenum, with negligible losses along the span of the plenum IB to OB.

Sixteen nozzle surface static pressures were installed to analyze the circumferential pressure distortion in the plane of the CCW off-take and nozzle inlet around the interior of the duct wall on both the LH and RH nozzles symmetrically. These were primarily intended to assess the mass flow matching of the CCW relative to fan bypass streams to ensure that the off-take did not create circumferential pressure distortion and potentially adversely affecting engine operation. There is no explicit engine installation design guidance for this based on standard commercial application, however, the CCW off-take system was extensively analyzed prior to test to ensure minimal disruption over the operating range. Those analyses were shared with WI and they concurred that the environment was satisfactory. The LSPLM test data confirmed the predictions showing that the circumferential pressure variation was indeed negligible. The resultant calculated duct Mach number from the initial static thrust calibration runs for two power settings are shown in Figure 23. As can be seen the overall duct Mach variation for any point is less than .05 which is excellent providing confirmation that the CCW off-take design performed as intended. The duct Mach is computed based on the upstream fan bypass total pressures installed immediately behind the fan (engine station 13) and the ring of static pressures in the nozzle (engine station 16). The fan bypass duct total pressure loss between the two engine stations is on the order of 3% at max power which if accounted for would result in an absolute Mach which was lower on the order of .04 less. The relative Mach in this instance is more important than the absolute Mach and again the difference is seen to be small. The distribution of pressures was such that 2 were located just ahead of the CCW off-take inlet with the other 6 more or less distributed uniformly

Figure 22, CCW Off-Take and Plenum Pressure Losses Figure 21, CCW Off-Take Airflow

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Figure 23, Fan Bypass Duct Mach at CCW Off-Take and Nozzle Inlet Plane

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around the circumference. The CCW off-take inlet is essentially at the top slightly offset from center and the two pressures ahead of the inlet can be seen on the far right of the Figure.

VI. External Aerodynamic Test Results

The primary objective of the test effort was to validate the trimmed high lift performance of the configuration. This included validation of the aircraft field length performance, maneuvering ability, and handling qualities. Primary high lift performance data was obtained in the NFAC 40x80 tunnel over a wide range of control surface and power settings, including REN, CC flap, and elevator setting variations. These variations were required to evaluate the performance over representative take-off and landing flight conditions, to determine the optimum control surface settings, and to establish the trim and handling quality characteristics of the aircraft. The required trimmed CLmax required to meet the as-sized aircraft field performance objectives is depicted in Figure 24, along with the test demonstrated lift characteristics plotted as a function of angle of attack. The test data is plotted with both CC flap 60 and 90 settings and at a maximum power conditions. Both the raw untrimmed data and trimmed corrected data is also plotted. The trim corrections were established using the mid-mission center-of-gravity location and the elevator control surface variations tested for each configuration. As depicted in Figure 24, the test demonstrated trimmed maximum lift coefficient meets the program required value verifying the aircraft meets required field length performance.

Required Trimmed CLmax

Angle of Attack

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Figure 24, Demonstrated Trimmed Maximum Lift Coefficient Meets Key Program Requirement

The test data was also used to assess the maneuvering ability and overall handling qualities of the SACD aircraft. A piloted simulation has been developed for use in the AFRL Large Amplitude Multimode Aerospace Research Simulator (LAMARS) facility for detailed evaluation by several pilots. Results from this evaluation have indicated that the aircraft meets handling quality metrics for both approach and take-off. In addition, a parametric study was

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performed using this simulation to establish the optimum control surface and power setting conditions to minimize take-off and landing field length performance.

VII. Summary

An overview was presented of the aerodynamic design, low speed wind tunnel testing, and validation of a revolutionary STOL, survivable, and efficient transonic cruise compatible military transport. The SACD LSPLM test accomplished all of the key program test objectives and validated the Lockheed Martin HPLS technology to a TRL of 5. The model aerodynamic design, scale, model construction and engine integration challenges were significant but all elements were successfully executed in allowing for essentially unrestricted WI FJ44 engine operability throughout the test. The REN and CCW systems were highly efficient in turning the flow with minimal losses, and the CCW provided the required slot blowing momentum coefficient and lift augmentation. The extensive use of CFD was also critical in the model design process optimizing the performance of both the internal flow paths and external aerodynamics. Overall the NFAC SACD LSPLM NFAC test met all expectations and demonstrating the critical STOL low speed high lift aerodynamic performance requirements.

VIII. Acknowledgements

The authors would like to acknowledge the generous help and support of fellow LM teammates Rich Peterson, Rich Orobitg and KC Martin for performance of the conceptual design effort, and Neil Hall for his expertise in evaluating the stability and control test data and development of the control surface effectiveness. The authors would also like to acknowledge the development efforts of Charlie Novak, who was the primary researcher of the HPLS technologies.

References

1. Hooker, R., Wick, A., Zeune, C., Jones, G., Milholen, W., Design and Transonic Wind Tunnel Testing of a Cruise Efficient STOL Military Transport, AIAA ASM 2013, January 2013.

2. Novak, C., Berry, O., Hardin. C., McAllister, B., Atlanta, GA, US Patent Application for an Airfoil Having A Movable Control Surface, Pub. No. US 2007/0290098 A1, 20 December 2007.

3. Barberie, F., Wick, A., Hooker, R., Zeune, C., Powered Lift CFD Predictions of a Transonic Cruising STOL Military Transport, AIAA ASM 2013, January 2013.

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