Overview of Experiment Flight Conditions M=0.56.25000feet (Subsonic condition M=0.86.36000feef (Transonic condition) Nose-To-Tail (N2T) Distances 20,55,110and190feet Nomenclature X-direction(longitudinal) Y-direction(lateral) Z-direction(vertical) The subsonic flight condition(M=0.56, h=25000 feet) was selected to match pre-existing data from vortex-effect prediction codes. These codes needed to be validated to determine their utility ure applications of AFF. Since a possible future application of A FF is for transport airplane, flight data were also acquired at M=0.86 and h =36000 feet. This transonic flight condition is representative for that class of vehicle The vortex effects were also mapped at different longitudinal distances behind the leader airplane. These Nose-To-Tail (N2T) distances were monitored by the control room and maintained by the pilots through periodic radio calls. Only the results from the subsonic condition, 55 N2T will be presented here. 55 is equal to the length of the FlA-18. The reference axis system was as shown above. It should be noted that although Z is positive down, this presentation will refer to positions above the lead airplane(or high)as positive and positio below the lead airplane(or low)as negative BELT Autonomous Formation Flight Page 1
Overview of Experiment • Flight Conditions – M = 0.56, 25000 feet (Subsonic condition) – M = 0.86, 36000 feet (Transonic condition) • Nose-To-Tail (N2T) Distances – 20, 55, 110 and 190 feet • Nomenclature – X-direction (longitudinal) – Y-direction (lateral) – Z-direction (vertical) The subsonic flight condition (M=0.56, h=25000 feet) was selected to match pre-existing data from vortex-effect prediction codes. These codes needed to be validated to determine their utility on future applications of AFF. Since a possible future application of AFF is for transport airplane, flight data were also acquired at M=0.86 and h =36000 feet. This transonic flight condition is representative for that class of vehicle. The vortex effects were also mapped at different longitudinal distances behind the leader airplane. These Nose-To-Tail (N2T) distances were monitored by the control room and maintained by the pilots through periodic radio calls. Only the results from the subsonic condition, 55’ N2T will be presented here. 55’ is equal to the length of the F/A-18. The reference axis system was as shown above. It should be noted that although Z is positive down, this presentation will refer to positions above the lead airplane (or high) as positive and positions below the lead airplane (or low) as negative. Page 1 Autonomous Formation Flight Program NAS4-00041 TO-104
Lift and Drag Analysis Flight Test database Air data Engine Data JINS IS Data Air Data Computations Gross Weight, Vint P In-Flight Thrust Model Wind Axis Accelerations FE A = G RAM.DRAG XW,YW2ZW Fry=GW*A Predicted Performance Performance Model (outside vortex influence) D=cos(aestFG-FRAM-FEr DRAGEX LD,D Vortex effect= Vortex-Baseline %△Cn,%△Cn,%△WFT Performance data was determined using classical techniques. A force balance along the flight path was used to determine drag while a force balance perpendicular to that was used to determine lift D= cos(aest)FG-FRAM-FEDRAGGWAXW); L=sin(aest)FG+(GW"NXW). Three primary data reduction areas feed the performance mode: 1)Air Data, 2)IFT, and 3 )Accelerations. tal fuel accounting for crew weight. It also includes a calculation of an estimated alpha, aest, which is based on the trailing aircraft s pitch angle and the lead aircraft's flight path angle(aest= gtrail-ylead). This was required because the trailing aircraft's alpha probes are unusable during formation flight due to localized upwash influences of the lead aircraft. Because the lead aircraft flew at steady-state conditions(constant speed and altitude ), the flight path angle, lead, was always close model. The model calculated gross thrust (FG), ram drag(FRAM) and engine throttle dependent drag. (FEDRAG). Gross thrust is the primary force the engine produces out the tail pipe, FRAl The engine manufacturer's IFT model was used to calculate thrust on the F404-GE-400 engines installed in the trailing F/A-18 Aircraft The next chart describes the measurements used to run t presents the force loss due to the mom air, W1, entering the inlet, and FEDRAG accounts for the extemal drag forces associated with the engine nozzle and inlet spillage flow The INS was used to obtain vehicle acceleration data. This data was corrected for rotation effects due to not being mounted exactly on the center of gravity. It then was translated into the flight path (wind axis)coordinate system. Axial acceleration was used to compute vehicle excess thrust: FEX= GW'AXW The performance model used the information from the three paths described above to obtain lift, drag and respective coefficients. To obtain drag reduction values, data obtained during formation flight (vortex) was compared to baseline(non-vortex) points completed in a back-to-back fashion. Some formation flight test points did not include a slide-out maneuver to obtain baseline conditions For these few points, baseline data were estimated based on data trends in drag related to gross weight. A simple prediction model was used to calculate baseline lift and drag values to evaluate the reasonableness of the baseline data
Page 2 Autonomous Formation Flight Program NAS4-00041 TO-104 Lift and Drag Analysis Flight Test Database Air Data Engine Data INS Data In-Flight Thrust Model FG, FRAM, FEDRAG Wind Axis Accelerations AXW, AYW, AZW Air Data Computations Gross Weight, Vinf, Po Dest. = TTrail.- JLead Performance Model D = cos(Dest) FG – FRAM – FEDRAG - FEX CL, CD , CDi = CD– CD0 Vortex Effect = Vortex – Baseline %'CD, %'CDi, %'WFT Predicted Performance (outside vortex influence) CL, CD , CD0 FEX=GW*AXW Performance data was determined using classical techniques. A force balance along the flight path was used to determine drag while a force balance perpendicular to that was used to determine lift: D = cos(Dest) FG – FRAM – FEDRAG-(GW*AXW); L = sin(Dest) FG + (GW*NXW). Three primary data reduction areas feed the performance mode; 1) Air Data, 2) IFT, and 3) Accelerations. The Air Data model computes gross weight (GW) using empty weight and the remaining total fuel accounting for crew weight. It also includes a calculation of an estimated alpha, Dest, which is based on the trailing aircraft’s pitch angle and the lead aircraft’s flight path angle (Dest = qtrail-Jlead). This was required because the trailing aircraft’s alpha probes are unusable during formation flight due to localized upwash influences of the lead aircraft. Because the lead aircraft flew at steady-state conditions (constant speed and altitude), the flight path angle, Jlead, was always close to zero. The engine manufacturer’s IFT model was used to calculate thrust on the F404-GE-400 engines installed in the trailing F/A-18 Aircraft. The next chart describes the measurements used to run this model. The model calculated gross thrust (FG), ram drag (FRAM) and engine throttle dependent drag, (FEDRAG). Gross thrust is the primary force the engine produces out the tail pipe, FRAM represents the force loss due to the momentum of air, W1, entering the inlet, and FEDRAG accounts for the external drag forces associated with the engine nozzle and inlet spillage flow. The INS was used to obtain vehicle acceleration data. This data was corrected for rotation effects due to not being mounted exactly on the center of gravity. It then was translated into the flight path (wind axis) coordinate system. Axial acceleration was used to compute vehicle excess thrust: FEX = GW*AXW The performance model used the information from the three paths described above to obtain lift, drag and respective coefficients. To obtain drag reduction values, data obtained during formation flight (vortex) was compared to baseline (non-vortex) points completed in a back-to-back fashion. Some formation flight test points did not include a slide-out maneuver to obtain baseline conditions. For these few points, baseline data were estimated based on data trends in drag related to gross weight. A simple prediction model was used to calculate baseline lift and drag values to evaluate the reasonableness of the baseline data
Test point procedure · Pilot procedure Acquire and hold position within the influence of the vortex for 30 seconds of stable data Engage auto-throttle velocity-hold and maintain position for 20 seconds of stable data Laterally slide out of position (away from lead a/c), engage altitude-hold and stabilize outside of vortex for 20 seconds Technique provides direct comparison of performance data in and out of vortex Use of auto-pilot and auto-throttle significantly improved maneuver and data quality Each test point was conducted in the same way. Once both aircraft were on condition, the trail aircraft maintained its position behind the lead aircraft for 30 sec. During this ime, the pilot of the trail aircraft was controlling every aspect of his aircraft, including throttles. Because of the transient nature of the vortex effects, especially with significant wing overlap, the pilots throttle movements were, in some cases, coarse and over-corrective. This problem was exacerbated when combined with a significant longitudinal distance like 190 N2T, because maintaining longitudinal separation became especially difficult when the pilots did not have a good visual (close)reference After 30 sec of stable data, the pilot engaged the auto-throttle(AtC) velocity hold and held position for another 20 sec. More often than not, the atc would have to be set a few times before the N2T closure rate was small enough to call stable. After 20 sec of stable, ATC-engaged data, the control room gave the call for'slide out, at which time the pilot of the trail aircraft maneuvered laterally out of position to the right, engaged altitude- hold, and stabilized for another 20 sec outside of the vortex. The control room then gave a test point complete call at the appropriate time Following a video of an example test point, an explanation as to why the test point procedure was set up in this way will be given BELT Autonomous Formation Flight Page 3
Test Point Procedure • Pilot Procedure – Acquire and hold position within the influence of the vortex for 30 seconds of stable data – Engage auto-throttle velocity-hold and maintain position for 20 seconds of stable data – Laterally slide out of position (away from lead a/c), engage altitude-hold and stabilize outside of vortex for 20 seconds • Technique provides direct comparison of performance data in and out of vortex • Use of auto-pilot and auto-throttle significantly improved maneuver and data quality Each test point was conducted in the same way. Once both aircraft were on condition, the trail aircraft maintained its position behind the lead aircraft for 30 sec. During this time, the pilot of the trail aircraft was controlling every aspect of his aircraft, including throttles. Because of the transient nature of the vortex effects, especially with significant wing overlap, the pilot’s throttle movements were, in some cases, coarse and over-corrective. This problem was exacerbated when combined with a significant longitudinal distance like 190’ N2T, because maintaining longitudinal separation became especially difficult when the pilots did not have a good visual (close) reference. After 30 sec of stable data, the pilot engaged the auto-throttle (ATC) velocity hold and held position for another 20 sec. More often than not, the ATC would have to be set a few times before the N2T closure rate was small enough to call stable. After 20 sec of stable, ATC-engaged data, the control room gave the call for ‘slide out’, at which time the pilot of the trail aircraft maneuvered laterally out of position to the right, engaged altitude-hold, and stabilized for another 20 sec outside of the vortex. The control room then gave a ‘test point complete’ call at the appropriate time. Following a video of an example test point, an explanation as to why the test point procedure was set up in this way will be given. Page 3 Autonomous Formation Flight Program NAS4-00041 TO-104
Flight Test point matrix Condition 2 Condition 1 Mach 0.86 Mach 0.56 36,000ft25,000ft 50% 50% 25% Lead Aircraft -25% 5升N 75% 50%25% 0% 25% 50% Lateral Separation, Y, wingspan Real-time feedback in cockpit using ils needles (%Wingtip Separation N2T (X) position monitored through control room calls no To fully map the vortex, a grd of test points, or test matrix, was created direct feedback in cockpit) Several factors constrained this matrix, inclue imited test flights available he guidance(needle)display was limited to 60 target files. To maximize the resolution of the vortex mapping in the most efficient manner, the matrix was based on 1/8 of an FlA-18 wingspan, or just under 5 feet. ap the vortex, a grid of test points, or test matrix, was created. Because flight test time was limited, the number of matrix points had to be kept to a minimum without sacrificing the resolution of the vortex mapping. In addition, the guidance (needle)display used by the trailing pilot to fly each test point was limited to 60 target files. Designed within these boundaries, the test matrix was based on 1/8 of an FlA.18 wingspan( F/A-18=37.5 ft), or about 4.7 it a grid of equally-spaced points in the Y-and Z- axis was then set up using this parameter. An example of such a grid is shown above
Flight Test Point Matrix 4 ' 6 ' 8 ' 10 ' 12 ' -50% -25% 0% 25% 50% (% Wingtip Separation) Real-time feedback in cockpit using ILS needles N2T (X) position monitored through control room calls (no direct feedback in cockpit) 55 ft N2T Page 4 Autonomous Formation Flight Program NAS4-00041 TO-104 -75% -50% -25% 0% 25% 50% Lateral Separation, Y, % wingspan 50% 25% 0% -25% -50% Vertical Separation, Z, % wingspan Condition 2 Mach 0.86 36,000 ft Condition 1 Mach 0.56 25,000 ft To fully map the vortex, a grid of test points, or test matrix, was created. Several factors constrained this matrix, including: Limited test flights available The guidance (needle) display was limited to 60 target files. To maximize the resolution of the vortex mapping in the most efficient manner, the matrix was based on 1/8 of an F/A-18 wingspan, or just under 5 feet. To fully map the vortex, a grid of test points, or test matrix, was created. Because flight test time was limited, the number of matrix points had to be kept to a minimum without sacrificing the resolution of the vortex mapping. In addition, the guidance (needle) display used by the trailing pilot to fly each test point was limited to 60 target files. Designed within these boundaries, the test matrix was based on 1/8 of an F/A-18 wingspan (bF/A-18=37.5 ft), or about 4.7 ft. A grid of equally-spaced points in the Y- and Z- axis was then set up using this parameter. An example of such a grid is shown above
Nasa Autonomous Formation Flight Test Results Summary Of Phase 0-February 2001 Summary of Phase 1-august 2001 NASA Dryden Flight Research Center Photo Collection NASA http://www.dfrc.nasagovigallery/photo/indexhtml NASA Photo: EC01-00509 Date: February 21, 2001 Photo by: Jim Ross wo F/A- 18B aircraft involved in the AFF program return to base in close formation with the autonomou function disengaged
Autonomous Formation Flight Test Results Summary Of Phase 0 -February 2001 Summary Of Phase 1 -August 2001 Autonomous Formation Flight Test Results Summary Of Phase 0 -February 2001 Summary Of Phase 1 -August 2001
Phase 0 Control Experiment #1 20 Steady-State Tracking Extremely Accurate Position Outside The Vortex Demonstrated Winter 2001) Experiment: Dial In-75/-75 ft Translations, AFF Flight 715-February 21, 2001 2-Minute Tracking Task High Performance Gainst 20 Relative Lateral Position Error(fit) integrator was larger overshoots for this gainset than for the others. However, performance and stability were still well within the acceptable region for these gaffe in undesired side-effect of the e additional feedback of the integral of the position emor in the INTEGRAL gainst was very successful at elimina The HIGH PERFORMANcE gainst exhibited extremely good disturbance rejection capability. Position errors during steady-state tracking with these gains were approximately 1 foot both laterally
Phase 0 Control Experiment #1 Steady-State Tracking Phase 0 Control Experiment #1 Steady-State Tracking AFF Flight 715 - February 21, 2001 2-Minute Tracking Task High Performance Gainset 20 15 Relative Lateral Position Error (ft) Relative Vertical Position Error (ft) 010 5 -20 -15 -10 -5 -20 20 -15 -10 -5 0 5 10 15 Extremely Accurate Position Outside The Vortex Demonstrated (Winter 2001). (Experiment: Dial In –75/-75 ft Translations) Extremely Accurate Position Outside The Vortex Demonstrated (Winter 2001). (Experiment: Dial In –75/-75 ft Translations) Page 6 Autonomous Formation Flight Program NAS4-00041 TO-104 20 15 Relative Lateral Position Error (ft) Relative Vertical Position Error (ft) 010 5 -20 -15 -10 -5 -20 20 -15 -10 -5 0 5 10 15 AFF Flight 714 - February 21, 2001 2-Minute Tracking Task Integral Gainset The additional feedback of the integral of the position error in the INTEGRAL gainset was very successful at eliminating any steady-state offsets in position error. An undesired side-effect of the integrator was larger overshoots for this gainset than for the others. However, performance and stability were still well within the acceptable region for these gains. The HIGH PERFORMANCE gainset exhibited extremely good disturbance rejection capability. Position errors during steady-state tracking with these gains were approximately 1 foot both laterally and vertically
Drag Change Contour Plot Contour plots 0995590g5 · Provides a true perspective of the vortex s 3 20 …∴… influence on vehicle performance 5 Factors Number of test points 9m+9eA. △C1 Data smoothing percen bicubic spline · Extrapolation 15 missing data points30…………r Lateral Position, Percentage of Wing Span This particular contour plot( Mach 0.56, 25,000ft)contains 92 test points. BELT Autonomous Formation Flight Page 7
Drag Change Contour Plot Contour plots: • Provides a true perspective of the vortex's influence on vehicle performance Factors: • Number of test points • Data smoothing – bicubic spline • Extrapolation – missing data points ' C D, percent M=0.56, 25,000 ft altitude, 55ft nose-to-tail This particular contour plot (Mach 0.56, 25,000ft) contains 92 test points. Page 7 Autonomous Formation Flight Program NAS4-00041 TO-104
Actual Flight Test results validate Drag Bucket"Theorv Phase 0 Control Results 0o=+0 r,mel Phase 1 Control Requirement Lateral offset(△Yfet o感e BELT Autonomous Formation Flight Page 8
Actual Flight Test Results Validate “Drag Bucket” Theory P h as e 1 C o ntrol R e q u ir e m e n t A ct Phase 1 Control Phase 1 Control Requirement Requirement Phase 0 Control Phase 0 Control Results Results Drag Reduction % Lateral Offset ( Lateral Offset ( 'Y feet) Y feet) Vertical Offset ( Vertical Offset ('Z feet) Z feet) +'Y -'Z Xf Zf Yf -'X Page 8 Autonomous Formation Flight Program NAS4-00041 TO-104
Contour Plot of Multiple Data Points Vertical Z=0.37 △cD ●z=0.13 Wing-Span◆z=0.6 ◇Z=-0.06 z=0.13 Z=025 5% 20 25% 0.540.40.34020.100.10203040.5 Wing-Tip Separation( Y Relative Position), % Wing-Span 2010010 Lateral position, Percentage of Wing Span Percent change in drag versus position at M=0.56, 25,000ft, 55'N2T The change in lift and drag coefficient were evaluated for each maneuver by comparing drag while in the vortex to baseline values. In general very small changes(variations of less than 2%)in calculated lift coefficient were found for all conditions as predicted For the 55 feet N2T condition shown, up to 19% drag reduction was calculated with peak values between level and-13% vertical position and a lateral position of 10-20% wingtip overlap. Overall the data indicates a large region of significant gains. The data is not symmetric about the the peek position and shows increased sensitivity as the trailing aircraft moved inboard of the peak position as opposed to outboard of this position. In fact, drag increases were measured at some high wing overlap positions, verifying the importance of proper station-keeping to obtain the best results Data quality varied for each test point with the outboard data tending to have better quality than the inboard data, primarily due to the pilots ability to maintain stable conditions. Some inboard test points were very difficult to fly due to the lead aircraft's vortex impacting the tail or fuselage. Fortunately the region of best drag benefits was fairly stable and good data quality was obtained or most points. Atmospheric conditions also affected the data on some test points due to turbulence and vertical winds. The back-to-back comparison of vortex and baseline data helped to minimize these effects
Page 9 Autonomous Formation Flight Program NAS4-00041 TO-104 -25% -20% -15% -10% -5% 0% 5% 10% 15% -0.5 -0.4 -0.3 -0.2 -0.1 0 0.1 0.2 0.3 0.4 0.5 Wing-Tip Separation (Y Relative Position), % Wing-Span (CD-FF - CD-BL)/CD-BL Percent change in drag versus position at M=0.56, 25,000ft, 55’ N2T Vertical Separation, % Wing-Span Z=0.37 Z=0.25 Z=0.13 Z=0.06 Z=0.0 Z=-0.06 Z=-0.13 Z=-0.25 Z=-0.37 'CD +'Y -'Z Xf Zf Yf -'X Contour Plot of Multiple Data Points The change in lift and drag coefficient were evaluated for each maneuver by comparing drag while in the vortex to baseline values. In general very small changes (variations of less than 2%) in calculated lift coefficient were found for all conditions as predicted. For the 55 feet N2T condition shown,up to 19% drag reduction was calculated with peak values between level and -13% vertical position and a lateral position of 10-20% wingtip overlap. Overall the data indicates a large region of significant gains. The data is not symmetric about the the peek position and shows increased sensitivity as the trailing aircraft moved inboard of the peak position as opposed to outboard of this position. In fact, drag increases were measured at some high wing overlap positions, verifying the importance of proper station-keeping to obtain the best results. Data quality varied for each test point with the outboard data tending to have better quality than the inboard data, primarily due to the pilots ability to maintain stable conditions. Some inboard test points were very difficult to fly due to the lead aircraft’s vortex impacting the tail or fuselage. Fortunately the region of best drag benefits was fairly stable and good data quality was obtained on most points. Atmospheric conditions also affected the data on some test points due to turbulence and vertical winds. The back-to-back comparison of vortex and baseline data helped to minimize these effects
Total Drag Condition 2 △CD Mach 0.86 percent 36,000ft Condition 1 Mach 0.56 to percent 25,000ft Later l Pastina. Pereentage f Wing p罐 Lateral Putin, Pementagtd W吧pn 55 ft neT 110 ft n2T 190 ft n2T BUEWH onymous Formation Flight Page 10
Total Drag 'CD, percent 'CD, percent Condition 2 Mach 0.86 36,000 ft Condition 1 Mach 0.56 25,000 ft Page 10 Autonomous Formation Flight Program NAS4-00041 TO-104 55 ft N2T 110 ft N2T 190 ft N2T
