The F-15B Lifting Insulating Foam Trajectory (LIFT) Flight Test

NASA/TM-2006-213674

The F-15B Lifting Insulating Foam Trajectory

(LIFT) Flight Test

Stephen Corda, Donald Whiteman, Ting Tseng NASA Dryden Flight Research Center Edwards, California Ricardo Machin NASA Johnson Space Center Houston, Texas June 2006 NASA STI Program ... in Profile Since its founding, NASA has been dedicated CONFERENCE PUBLICATION. Collected • to the advancement of aeronautics and space papers from scientific and technical science. The NASA scientific and technical conferences, symposia, seminars, or other information (STI) program plays a key part in meetings sponsored or cosponsored by helping NASA maintain this important role.

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NASA/TM-2006-213674

The F-15B Lifting Insulating Foam Trajectory

(LIFT) Flight Test

Stephen Corda, Donald Whiteman, Ting Tseng NASA Dryden Flight Research Center Edwards, California Ricardo Machin NASA Johnson Space Center Houston, Texas National Aeronautics and Space Administration Dryden Flight Research Center Edwards, California 93523-0273

June 2006

NOTICE

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Available from the following: NASA Center for AeroSpace Information National Technical Information Service 7121 Standard Drive 5285 Port Royal Road Hanover, MD 21076-1320 Springfield, VA 22161-2171 (301) 621-0390 (703) 605-6000

Contents ABSTRACT . 1 NOMENCLATURE . 1 INTRODUCTION . 2 AIRCRAFT . AND . AERODYNAMIC . FLIGHT . TEST . FIXTURE . DESCRIPTIONS . 5 LIFTING . INSULATING . FOAM . TRAJECTORY . CONFIGURATION . DESCRIPTION . 7 High-Speed . Digital . Video . Camera . System . 16 Divot . Photogrammetry . and . Trajectories . 16 Camera . Pods . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17 Foam . Sheets . 20 Divot . Ejection . System . 23 Synchronization . System . 24 FLIGHT . TEST . CONDITIONS . AND . MANEUVERS . 26 HAZARDS . ASSESSMENT . 2 7 Divot . Recontact . 2 7 Divot . Ground . Impact . 29 RESULTS . AND . DISCUSSION . 31 Flight . Test . Conditions . 3 1 Divot . Ejection . 3 6 Divot . Structural . Integrity . 4 5 Divot . Shape . and . Size . 4 5 Divot . Aerodynamic . Stability . 4 7 Divot . Photogrammetry . and . Trajectories . 48 CONCLUDING . REMARKS . 49 REFERENCES . 50 i t ables 1. The F-15B aircraft Lifting Insulating Foam Trajectory divot ejection flight test conditions .

. and . results . summary . 10 2 . Divot . cases . 23 3 . Divot . ejection . pneumatic . system . components . 24 4 . Comparison . of . maximum . strain . energy . density . of . foam . and . aircraft . materials . 29 5 . Flight . test . divot . sizes . 38 Figures 1 . Cohesive-adhesive . strength . failure . of . shuttle . external . tank . insulating . foam . 3 2. The NASA F-15B research test bed aircraft in flight . 4 3 . The . NASA . F-15B . research . test . bed . aircraft, . shown . with . the . Aerodynamic . Flight . Test .

. Fixture . and . air . data . boom . 5 4 . Details . and . dimensions . of . the . Aerodynamic . Flight . Test . Fixture . 6 5 . High-speed . camera . system . components . and . interfaces . 16 6 . Camera . pod . views . relative . to . the . Aerodynamic . Flight . Text . Fixture . 17 7. The F-15B aircraft computational fluid dynamics analysis . 18 8 . Aerodynamic . Flight . Test . Fixture . foam . plate . dimensions . and . divot . sizes . 21 9 . Divot . ejection . pneumatic . system . 23 10 . Lifting . Insulating . Foam . Trajectory . aft . cockpit . control . panel . 25 11. The F-15B aircraft Lifting Insulating Foam Trajectory flight test envelope and test .

. points . 26 12. Comparison of dynamic pressures from the Lifting Insulating Foam Trajectory flight test .

. point . and . shuttle . ascent . 27 13. Divot recontact computational fluid dynamics predictions . 28 ii 14. Altitude as a function of Mach number for supersonic acceleration (flight 9) . 32 15. Specific excess power and altitude as a function of Mach number (flight 9) . 32 16. Angle of attack as a function of Mach number (flight 9) . 33 17. Angle of sideslip as a function of Mach number (flight 9) . 34 18 . Comparison . of . Mach . numbers . from . the . aircraft . and . Aerodynamic . Flight . Test .

Fixture (flight 9) . 35 19. Comparison of foam temperatures from the Lifting Insulating Foam Trajectory flight test .

. and . shuttle . ascent . 36 20. Composite digital video frame captures (at 2,000 pps) of divot ejections . 42 21. Composite digital video frame captures (at 2,000 pps) of divot ejections, forward camera .

. view . 43 22. Foam panels after the in-flight divot ejection . 44 23. Postflight close-up view of foam panel . 44 24 . Divot . ejection . pressure . as . a . function . of . divot . void . diameter . 45 25. Definition of divot geometry . 46 26 . Divot . aspect . ratio . as . a . function . of . divot . thickness . 47 27 . Composite . digital . video . frame . captures . of . divot . ejection, . forward . camera . view . looking . aft .

at 2,000 pps, time and number of frames from ejection (divot C). . 48 28. Comparison of F-15B flight data with one-degree-of-freedom prediction (flight 10, divot C, .

. Mach . 1 6, . altitude . of . 36,120 . ft, . and . dynamic . pressure . of . 850 . lbf/ft ) . 49 iii

abstra C t A series of flight tests has been performed to assess the structural survivability of space shuttle external tank debris, known as divots, in a real flight environment. The NASA F-15B research test bed aircraft carried the Aerodynamic Flight Test Fixture configured with a shuttle foam divot ejection system. The divots were released in flight at subsonic and supersonic test conditions . matching . points . on . the . shuttle . ascent . trajectory . Very . high-speed . digital . video . cameras .

recorded the divot trajectories. The objectives of the flight test were to determine the structural survivability of the divots in a real flight environment, assess the aerodynamic stability of the divots, . and . provide . divot . trajectory . data . for . comparison . with . debris . transport . models . A . total . of .

10 flights to Mach 2 were completed, resulting in 36 successful shuttle foam divot ejections. High- speed . video . was . obtained . at . 2,000 . pictures . per . second . for . all . of . the . divot . ejections . The . divots .

that . were . cleanly . ejected . remained . structurally . intact . The . conical . frustum-shaped . divots . tended . to .

aerodynamically trim in both the subsonic and supersonic free-stream flow.

no M en C lature 2 2 A . divot . cross-sectional . area, . ft . (m ) cross-section AFTF . . Aerodynamic . Flight . Test . Fixture C . . divot . flat . plate . drag . coefficient d C . . divot . moment . coefficient m CFD . . computational . fluid . dynamics deg . . degree DTA . . debris . transport . analysis E energy required to compress the foam, in-lbf (J) cr 2 2 g . . gravitational . acceleration, . ft/s . (m/s ) GPS . . global . positioning . system H pressure altitude, ft (m) HT . . divot . thickness, . in IRIG . . Inter-Range . Instrumentation . Group KCAS knots calibrated airspeed, nmi/h (m/s) KE divot kinetic energy, ft-lbf (J) L . . largest . divot . diameter, . in lbm . . pound-mass LIFT . . Lifting . Insulating . Foam . Trajectory M . . Mach . number 2 2 MEOP .. maximum . expected . operating . pressure, . lbf/in . (N/m ) NPT . . National . Pipe . Thread

pps . . pictures . per . second P specific excess power, ft/s (m/s) s psia . . pounds . per . square . inch, . absolute 2 2 q . . dynamic . pressure, . lbf/ft . (N/m ) STS . . Space . Transportation . System TC . . thermocouple TPS . . thermal . protection . system V divot terminal velocity, ft/s (m/s) term W divot weight, lb (kg) 1-DOF .. one-degree-of-freedom 3-D . . three-dimensional α . . angle . of . attack, . deg β . . angle . of . sideslip, . deg 3 3 ρ . . air . density, . slugs/ft . (kg/m ) air intro D u C tion The loss of the Space Shuttle Columbia (STS-107), caused by debris shed from the external tank . during . shuttle . ascent, . has . prompted . a . large . effort . by . NASA . to . understand . this . debris . transport .

phenomenon (ref. 1). The shuttle external tank thermal protection system (TPS) consists of a spray - on . insulating . foam . The . TPS . protects . the . tank . from . heating . during . shuttle . ascent . and . reduces .

the formation of ice after the tank is filled with cryogenic propellants.

One . of . the . TPS . failure . modes . involves . cohesive-adhesive . strength . failure . of . the . foam .

Because . of . imperfections . in . the . external . tank . foam . application, . air . is . trapped . in . voids . underneath .

the . foam . During . ascent, . decreasing . atmospheric . pressure . causes . an . increasing . pressure . differential .

in . the . trapped . volume . of . air . The . force . caused . by . the . pressure . differential . may . exceed . the .

cohesive - adhesive . strength . of . the . foam, . which . causes . the . foam . piece . over . the . air . void . to . shed .

This type of TPS foam shedding is referred to as “divoting” (fig. 1(a)), and the shed foam debris is called a “divot” (fig. 1(b)).

NASA . is . conducting . extensive . research . to . understand . the . TPS . debris . transport . The . debris .

transport analysis (DTA) includes definitions of the debris geometries, release initial conditions, and locations of debris release. Maximum expected flaw or defect characteristic lengths have been determined . from . dissections . of . sprayed . foam . layups . and . actual . shuttle . external . tanks . Based . on .

these . investigations, . a . typical . divot . has . a . conical . frustum . shape . with . a . base . diameter . of . many .

3 3 inches . Because . the . TPS . foam . density . is . roughly . 2 . lbm/ft . (32 04 . kg/m ), the divot mass typically is . very . low (a) Divoting near forward bipod area from STS-114.

(b) Typical shed foam debris known as a divot.

Figure 1. Cohesive-adhesive strength failure of shuttle external tank insulating foam.

 From . these . typical . divot . shapes . and . sizes, . aerodynamic . and . structural . divot . models . have .

been . developed . for . use . in . computational . tools . to . predict . the . divot . trajectories . To . understand .

divot “flight” after shedding, a critical unknown must be answered: does the divot trim (that is, assume a stable orientation with respect to the free-stream flow) or tumble? The aerodynamics of the trim configuration as opposed to the tumble configuration significantly influences the divot trajectory and cross range. Ultimately this configuration translates into where the shuttle orbiter can . be . struck . by . a . divot . from . the . external . tank . Another . critical . unknown . is . whether . the . divot .

remains . structurally . intact . or . fragments . after . release To . obtain . divot . data, . tests . have . been . conducted . by . several . means, . including . wind . tunnel, .

ballistic tunnel, and flight. This report presents the results of the flight tests performed with the NASA F - 15B . research . test . bed . aircraft . The . F-15B . aircraft . carried . a . centerline-mounted . Aerodynamic .

Flight Test Fixture (AFTF, previously called the Flight Test Fixture-II) configured with a divot ejection system (fig. 2). Sheets of external tank TPS foam were mounted on the side of the AFTF and back pressured to eject divots in flight at subsonic and supersonic speeds to near Mach 2.

Forward camera pod Aft camera pod AFTF with two foam sheets 050609 Figure 2. The NASA F-15B research test bed aircraft in flight (NASA Photo EC0030 - 12).

The primary objective of the flight tests was to determine the structural survivability of the divots in a real flight environment, matching the Mach number and dynamic pressure at discrete . points . along . the . shuttle . ascent . trajectory . Very . high-speed . digital . video . cameras . were .

used . to . visually . assess . the . structural . survivability . of . the . divot . The . stability . of . the . divot . trim .

configuration as opposed to the divot tumble configuration also was assessed. A secondary objective . was . to . quantify . the . divot . trajectories . through . the . use . of . photogrammetry . techniques

air C ra F t an D aero DY na M i C F lig H t test F i X ture D es C ri P tions The F-15B aircraft is a two-seat fighter-trainer version of the F-15A air-superiority fighter built by McDonnell Douglas Aircraft Company (now The Boeing Company, St. Louis, Missouri).

The F-15B airplane has a wingspan of 42.8 ft (13.05 m), height of 18.7 ft (5.7 m) and length of 63.7 ft (19.4 m), excluding the air data nose boom (fig. 3). The aircraft has a high-mounted swept main wing with a modified delta shape, twin vertical tails, all-moving horizontal stabilators, and twin turbofan jet engines. Primary flight control surfaces are controlled by a hydromechanical system and an electrical control augmentation system (CAS). The F-15B aircraft is capable of dash speeds in excess of Mach 2 and level flight at altitudes to 60,000 ft (18,288 m).

42.8 ft 18.7 ft 63.7 ft 050475 Figure . 3 . The . NASA . F-15B . research . test . bed . aircraft, . shown . with . the . Aerodynamic . Flight . Test .

Fixture . and . air . data . boom The . F-15B . aircraft . is . powered . by . two . F100-PW-100 . turbofan . engines . (Pratt . & . Whitney, . West .

Palm Beach, Florida). Each engine produces an uninstalled sea level static thrust of approximately 25,000 lbf (11,340 kgf) in full afterburner. The aircraft has a fully fueled takeoff weight of approximately 42,000 lb (19,051 kg) and a landing weight of approximately 32,000 lb (14,515 kg).

The . aircraft . has . aerial . refueling . capability . for . extended-duration . research . missions Modifications made to the NASA F-15B aircraft to convert it from an air-superiority fighter to . a . supersonic . research . test . bed . include . the . installation . of . research . systems . for . instrumentation, .

digital data recording, telemetry, in-flight video, and global positioning system (GPS) information.

A significant research feature of the aircraft is the ability to carry large experiment test fixtures on the . lower . fuselage . centerline . pylon The AFTF is the second-generation aerodynamic flight test fixture that was built to replace the first flight test fixture flown on a NASA F-104 aircraft (refs. 2 and 3). The AFTF is a low - aspect - ratio, rectangular fin shape that is mounted underneath the aircraft on the fuselage centerline pylon (fig. 3). It has an elliptical nose section and a blunt, squared-off base. Constructed of all composite materials, . the . AFTF . has . a . modular . structure . with . four . upper . and . four . lower . internal . bays . The .

bays are accessible through removable side panels. The AFTF is 107 in. (2.718 m) long, 32 in. .

(0.8128 m) high, and 8 in. (0.2032 m) wide. The maximum weight is approximately 500 lb .

(186.6 kg). Figure 4 shows the details and dimensions of the AFTF.

Top view Sway braces/ Nose section load cells (4) Suspension lugs (2) 8.0 in.

Splitter plate Left side view Aft view I H G 10.0 in.

32.0 in. Bay 3A 30.9 in.

13.0 in.

F E D C B A Bay 2B Bay 3B 18.8 in. 35.7 in. 52.5 in.

107.0 in.

050476 Figure . 4 . Details . and . dimensions . of . the . Aerodynamic . Flight . Test . Fixture The AFTF complements the current inventory of F-15B experimental flight test fixtures, which includes the Propulsion Flight Test Fixture (PFTF) and Centerline Instrumented Pylon (CLIP) (refs. 4 and 5). The PFTF is designed to conduct advanced propulsion experiments and the CLIP is a new fixture designed to accommodate larger span models underneath the aircraft.

Standard National Advisory Committee for Aeronautics (NACA) air data nose booms have been . mounted . on . the . F-15B . airplane . nose . and . AFTF . leading . edge . Each . air . data . nose . boom . measures .

the local total pressure, static pressure, angle of attack, and angle of sideslip. Conventional flow angle . vanes, . mounted . downstream . of . the . static . pressure . ports, . measure . angle . of . attack . and . angle . of .

sideslip . Total . temperature . is . measured . by . probes . mounted . on . both . the . airplane . and . AFTF . aft . right .

side . Linear . and . angular . accelerations . are . measured . near . the . centerline . of . the . aircraft . just . forward .

and . above . the . aircraft . center . of . gravity

All . data . were . digitally . recorded . on . board . the . aircraft . and . telemetered . in . real . time . to .

ground - based . recorders . and . control . room . displays . Two . very . high-speed . digital . video . cameras, .

mounted . to . the . forward . and . aft . fuselage . missile . rails, . were . aimed . at . the . AFTF . and . monitored . in .

the . aft . cockpit . and . control . room . Data . were . collected . continuously . from . takeoff . to . landing The . high-speed . digital . video . was . recorded . on . board . the . aircraft . Real-time . video . was .

transmitted . to . the . control . room . and . displayed . on . the . F-15B . aft . cockpit . video . monitor . at . a .

comparatively slower rate of 30 pictures per second (pps).

The . telemetry . system . includes . two . telemetry . transmitters . and . data . streams, . one . each . for . the .

F - 15B . aircraft . and . AFTF . The . F-15B . telemetry . system . monitors . aircraft . instrumentation . including .

onboard GPS, nose boom air data, and other aircraft performance and flying quality parameters.

The . AFTF . instrumentation . system . provided . data . directly . related . to . the . Lifting . Insulating . Foam .

Trajectory (LIFT) experiment, in addition to the AFTF air data parameters. The high-speed camera . system . provided . digital . storage . of . video . data . for . each . divot . ejection . event, . and . data . were .

downloaded for processing after each flight. All parameters, including the video, were correlated with onboard GPS–synchronized, Inter-Range Instrumentation Group–B (IRIG-B) timing.

The . control . panel . for . the . experiment . was . located . in . the . rear . cockpit . of . the . F-15B . aircraft .

Details . of . the . control . panel . switches . and . operation . are . provided . in . the . succeeding . section, .

“Synchronization . System ” li F ting insulating F oa M tra J e C tor Y C on F iguration D es C ri P tion The LIFT flight test required the development of two new systems: a very high-speed digital video . camera . system . and . a . divot . ejection . system . These . systems . were . developed . and . tested . during .

a 3-month ground test effort preceding the flight test program.

In selecting the best divot ejection system for flight, four different systems were designed, fabricated, . ground . tested, . and . evaluated . Three . of . these . systems, . the . burst . disk, . needle-guided .

pneumatic, . and . mechanical . piston, . ejected . a . preformed . divot, . whereas . one . system, . called . the .

pressure-failed . sheet . system, . produced . divots . from . a . solid . sheet . of . foam . The . three . systems . that .

used . a . preformed . divot . held . the . divot . in . a . cylindrical . chamber . and . ejected . it . either . mechanically .

or . pneumatically In . the . burst . disk . system, . a . metal . burst . disk . was . ruptured, . allowing . high-pressure . gas . to . eject .

the . divot . from . the . cylindrical . chamber . In . the . needle-guided . pneumatic . system, . the . center . of . the .

divot . was . pierced . with . a . thin, . needle-like . metal . spike . to . help . guide . the . divot . during . ejection . High- pressure . gas, . injected . at . the . back . of . the . divot . by . means . of . a . simple . solenoid . valve . system, . was .

used . to . eject . the . divot . In . the . mechanical . piston . system, . the . divot . was . ejected . by . means . of . a . piston .

that . pushed . the . divot . out . of . the . cylinder . A . mechanical . claw . was . used . to . pressurize . and . release . the .

back . face . of . the . piston .

. .

The . pressure-failed . sheet . system . used . a . sheet . of . foam . that . was . back . pressured . to . produce . irregularly .

shaped divots. This system ultimately was selected for flight because of its simplicity of operation, and . because . the . divots . were . more . representative . of . those . in . the . actual . event . on . the . shuttle . external .

tank The . pressure-failed . sheet . divot . ejection . system . was . comprised . of . sheets . of . shuttle . external .

tank . insulating . foam . attached . to . the . side . of . the . AFTF . and . a . nitrogen . gas . pneumatic . system . mounted .

inside . the . AFTF . Precut . voids . in . the . back . of . the . foam . sheets . were . pressurized . with . a . nitrogen . gas .

from . the . pneumatic . system . The . back . pressuring . of . the . foam . sheets . caused . the . foam . to . fail . and .

created . the . divots . that . were . ejected . from . the . AFTF The . back-pressuring . mechanism . used . to . create . divots . on . the . AFTF . is . similar . to . the . actual .

situation . on . the . shuttle . The . foam . surface . temperature . conditions . for . the . AFTF . are . not . the . same . as .

those . for . the . shuttle . For . the . shuttle, . the . foam . inner . surface . is . exposed . to . cryogenic . temperatures, .

and . the . foam . outer . surface . is . aerodynamically . heated . to . high . temperatures . during . ascent . For . the .

AFTF, . the . foam . inner . surface . was . near . ambient . temperature . during . the . experiment . Because . the .

AFTF . could . not . match . the . shuttle . ascent . trajectory, . the . aerodynamic . heating . experienced . by . the .

F-15B . foam . outer . wall . was . much . less . than . that . experienced . by . the . shuttle . Table . 1 . presents . the .

AFTF . foam . outer . wall . temperatures . at . the . test . conditions The high-speed camera system was used to obtain digital video of the divots ejected from the AFTF. The camera heads were housed inside two camera pods mounted on the left side fuselage missile rails of the F-15B aircraft. The next section, “High-Speed Digital Video Camera System,” presents an overview of the high-speed camera and synchronization systems. Reference 6 (to be published) provides further details about the high-speed camera system.

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Table 1. The F-15B aircraft Lifting Insulating Foam Trajectory divot ejection flight test conditions and . results . summary Foam Aircraft flight conditions AFTF flight conditions Temperature Flight .

KCAS, .

Dynamic . Static .

No Mach . Altitude, . KCAS, . Mach . Altitude, . TC . 1, . TC . 2, .

nmi/h pressure, . pressure, .

number ft nmi/h number ft °F °F lbf/ft . psia 1 0 71 20,012 330 0 0 70 20,007 337 5 329 7 6 8 55 2 53 1 0 71 20,021 329 8 0 70 20,032 337 5 329 8 6 7 52 0 51 3 0 72 20,020 332 0 0 71 20,050 341 2 332 2 6 7 51 1 49 5 0 72 10,240 398 0 0 71 10,432 509 7 401 4 9 9 86 8 89 4 0 71 10,245 396 3 0 71 10,422 505 2 399 1 9 9 85 9 84 0 0 72 10,243 398 7 0 71 10,436 510 3 401 7 9 9 87 7 89 4 2 0 71 10,224 396 8 0 71 10,395 506 5 399 7 10 0 88 1 86 5 0 72 10,215 399 9 0 72 10,405 512 8 402 9 9 9 88 1 85 6 0 71 10,223 397 2 0 71 10,407 507 1 400 0 9 9 87 2 83 8 0 61 7,749 350 9 0 61 7,940 406 3 354 9 10 9 79 6 78 2 0 61 7,741 352 7 0 61 7,950 410 4 357 0 10 9 79 6 78 7 0 60 7,698 347 2 0 60 7,818 396 8 349 7 11 0 79 2 76 9 3 1 22 28,483 501 7 1 23 28,926 700 6 506 0 4 6 63 9 58 8 1 22 28,471 502 8 1 23 28,832 701 9 506 4 4 6 62 6 62 0 1 21 28,469 500 6 1 23 28,932 698 6 505 4 4 6 63 9 62 5 1 21 28,469 496 6 1 22 28,986 687 2 501 7 4 6 68 4 66 7 1 21 28,467 499 2 1 22 28,898 692 6 503 4 4 6 69 3 67 6 1 21 28,470 496 4 1 22 28,963 684 6 500 8 4 6 69 3 67 6 4 1 57 38,293 536 6 1 63 40,206 722 9 552 9 2 7 108 1 110 6 1 57 38,286 536 5 1 63 40,201 722 8 552 8 2 7 108 9 114 2 1 57 38,287 535 9 1 63 40,258 721 4 552 5 2 7 109 8 114 2 1 57 38,290 536 3 1 63 40,228 722 0 552 6 2 7 112 9 114 0 1 57 38,295 535 7 1 63 40,291 720 1 552 1 2 7 112 0 111 8 1 57 38,300 536 3 1 63 40,226 721 7 552 5 2 7 111 1 108 7 Table . 1 . Continued Divot . ejection . conditions Results Tank . Tank . Line . Line .

Divot . Divot .

Divot pressure, . temp , pressure, . temp , Comments ejection stability psia . °F psia . °F A 296 1 57 0 69 0 62 9 OK NA B 287 0 60 2 62 7 54 8 OK NA C 280 9 49 2 62 7 52 9 OK NA G 273 7 54 7 64 4 53 4 OK NA No . divot .

H 268 4 44 2 65 0 52 0 NA ejection I 263 8 56 5 65 9 56 6 OK NA A 296 9 53 3 76 7 64 7 OK Trim B 285 2 66 0 68 2 69 2 OK Trim Very small foam piece flaked off at C 277 7 56 5 68 2 70 1 OK Trim ejection No . Divot may have tumbled after first full G 265 2 69 6 70 0 71 9 OK Trim oscillation No .

I 258 8 69 6 70 0 71 9 OK Divot tumbled after first full oscillation.

Trim No . divot .

H 248 1 70 1 70 0 71 9 NA ejection H 296 1 44 2 77 8 53 4 OK Trim G 288 4 56 5 67 1 54 8 OK Trim I 281 2 55 2 67 0 54 8 OK Trim A 272 6 55 2 67 6 54 8 OK Trim B 268 4 55 2 67 5 53 8 OK Trim Very small foam piece flaked off at C 260 6 55 2 67 1 54 8 OK Trim ejection H 299 0 47 0 75 1 49 7 OK Trim G 291 3 54 3 62 8 50 2 OK Trim I 285 2 54 3 62 7 50 2 OK Trim A 277 4 53 3 62 7 49 7 OK Trim B 269 4 54 3 62 8 50 2 OK Trim C 263 8 54 3 63 4 50 2 OK Trim Table . 1 . Continued Foam .

Aircraft flight conditions AFTF flight conditions temperature Flight .

Dynamic . Static .

No Mach . Altitude, . KCAS, . Mach . Altitude, . KCAS, . TC . 1, . TC . 2, .

pressure, . pressure, .

number ft nmi/h number ft nmi/h °F °F lbf/ft psia 5 1 74 45,866 508 0 1 79 47,722 605 5 533 4 1 9 138 4 144 0 1 75 45,803 509 8 1 79 47,599 610 7 535 7 1 9 138 4 145 3 1 75 45,725 511 7 1 79 47,332 616 8 537 7 1 9 138 8 145 7 1 76 45,650 513 8 1 79 47,213 622 4 540 1 1 9 142 8 143 0 1 76 45,581 515 7 1 79 47,030 628 2 542 3 1 9 143 6 142 1 1 77 45,516 518 2 1 80 47,018 634 9 545 9 1 9 144 5 143 0 6 1 20 27,108 506 4 1 20 27,358 710 7 508 6 4 9 90 7 76 5 1 21 27,118 510 2 1 22 27,699 729 5 514 9 4 8 92 1 79 1 1 20 27,120 509 5 1 22 27,743 728 1 514 5 4 8 91 6 78 3 7 1 19 27,051 505 0 1 19 27,278 707 2 507 3 4 9 64 8 60 2 1 20 27,063 506 8 1 20 27,339 713 0 509 4 4 9 63 9 63 8 1 20 27,065 508 8 1 21 27,428 721 4 512 2 4 9 65 2 63 8 8 1 20 27,165 508 1 1 22 27,786 723 8 513 1 4 8 78 2 74 7 1 20 27,168 508 4 1 22 27,768 724 5 513 4 4 8 78 2 76 5 1 20 27,173 508 6 1 22 27,761 724 8 513 5 4 8 78 2 77 4 9 1 97 47,191 554 9 1 99 48,243 730 4 611 9 1 8 152 4 164 6 1 97 47,257 555 9 1 99 48,242 734 1 613 9 1 8 152 4 164 6 1 98 47,299 556 7 2 00 48,354 735 5 615 8 1 8 153 3 165 9 Table . 1 . Continued Divot . ejection . conditions Results Tank . Tank . Line . Line .

Divot . Divot .

Divot pressure, . temp , pressure, . temp , Comments ejection stability psia . °F psia . °F H 295 0 50 6 77 6 46 5 OK Trim G 287 6 50 6 61 8 46 5 OK Trim I 281 2 51 1 61 8 47 5 OK Trim A 274 8 51 5 62 3 46 5 OK Trim B 267 6 51 5 61 8 47 5 OK Trim C 260 9 51 1 62 2 46 5 OK Trim No . divot .

D 292 6 64 2 33 0 64 7 NA ejection No . divot .

E 292 9 69 2 24 3 67 4 NA ejection No . divot .

F 292 9 67 8 19 9 66 5 NA ejection Divot . recontacted . with . parent . hole . and .

broke . upon . recontact . Large . portion . of .

Partial .

divot . appeared . to . remain . in . hole . Leading .

D 294 7 60 2 49 5 61 1 divot . NA and . trailing . edges . broke . off . and . traveled .

ejection downstream E 290 5 60 6 39 9 61 1 OK Trim Same . as . divot . D . More . of . the . divot .

appeared . to . break . free . and . travel .

Partial .

downstream . Some . of . the . debris .

F 287 6 60 6 40 5 61 1 divot . NA recontacted . the . partial . divot . D . and .

ejection shattered. Difficult to determine whether or . not . the . pieces . trimmed D 301 4 61 1 62 6 63 8 OK Trim E 295 0 64 2 54 3 62 0 OK Trim F 290 5 64 2 54 2 62 9 OK Trim D 299 3 49 7 59 1 49 7 OK Trim E 292 9 53 3 50 6 49 7 OK Trim F 287 0 51 1 50 3 49 7 OK Trim Table . 1 . Continued Foam .

Aircraft flight conditions AFTF flight conditions temperature Flight .

Dynamic . Static .

No Mach . Altitude, . KCAS, . Mach . Altitude, . KCAS, . TC . 1, . TC . 2, .

pressure, . pressure, .

number ft nmi/h number ft nmi/h °F °F lbf/ft psia 10 1 53 34,026 569 4 1 60 36,186 841 2 585 5 3 3 107 2 108 0 1 53 34,021 569 5 1 60 36,195 841 7 585 7 3 3 107 2 108 9 1 53 34,011 569 8 1 60 36,174 843 0 586 1 3 3 106 7 110 2 1 53 34,002 570 3 1 60 36,201 844 2 586 5 3 3 108 5 108 7 1 53 33,992 571 1 1 60 36,086 845 7 586 8 3 3 107 6 107 8 1 53 33,989 571 7 1 60 36,124 848 1 587 7 3 3 106 7 106 9 Table . 1 . Concluded Divot . ejection . conditions Results Tank . Tank . Line . Line .

Divot . Divot .

Divot pressure, . temp , pressure, . temp , Comments ejection stability psia . °F psia . °F H 317 9 63 8 73 5 61 1 OK Trim G 309 9 65 6 64 3 62 0 OK Trim I 303 5 64 2 64 4 61 1 OK Trim A 293 9 65 6 64 3 62 0 OK Trim B 289 7 65 6 65 0 62 0 OK Trim Passed through the first oscillation before C 282 8 65 6 64 4 62 0 OK Trim trimming . with . large . diameter . forward High- s peed Digital Video Camera s ystem Two flight-qualified, high-speed camera systems were required to achieve the experiment objectives . The . camera . systems . were . synchronized . with . the . divot . ejection . system, . which . was .

simultaneously . triggered . by . an . aft . cockpit . switch . to . capture . the . images . and . record . them . on . a .

solid - state . recorder . within . the . camera . systems . The . two . camera . systems . and . data . acquisition .

systems . were . correlated . with . the . onboard . GPS-synchronized, . IRIG-B . time . The . high-speed . digital .

camera system was comprised of the following components: camera controller and recorder units, camera . heads, . synchronization . and . divot . ejection . system . interface . relay . circuitry, . and . M-Hub .

junction box (fig. 5).

The . camera . controller . and . recorder . units . were . located . in . the . ammunition . bay . pallet . accessible .

from . underneath . the . aircraft, . forward . of . the . AFTF . leading . edge . The . synchronization . card . and .

relay cards were located in the upper aft instrumentation bay of the AFTF. To survive in the flight environment, . this . equipment . was . shock . mounted . on . vibration . isolators Divot ejection system Camera control Camera Synchronization Cockpit and recorder switch head system unit 1 M-Hub IRIG-B time code junction box 28 Vdc Camera control Camera and recorder head unit 2 050477 Figure . 5 . High-speed . camera . system . components . and . interfaces For all of the flight tests, the camera configuration was set at a frame rate of 2,000 pps, exposure rate of 50 microseconds, resolution of 1280 by 512 pixels ranged over the field of view, and . software . gain . of . +6 . dB . The . camera . system . had . the . capability . to . record . at . a . maximum . frame . rate .

of . 10,000 . pps; . however, . operating . at . this . rate . reduced . the . record . time . and . resolution . Each . camera .

had . a . cumulative . video . recording . capacity . of . 9 6 . seconds . Based . on . extensive . ground . testing, .

this camera configuration provided optimum conditions for capturing the divot ejection video.

The high-speed video captured images of the divot release and to approximately 5 ft (1.524 m) downstream . of . the . release . point . The . focal . lengths . of . the . forward . and . aft . camera . lenses . were . 25 .

and 11.6 mm, respectively. Only in-flight ambient lighting was used for all video recordings.

Divot Photogrammetry and t rajectories Standard . photogrammetry . analysis . techniques, . which . use . photographic . images . to . obtain .

measurements . of . position, . were . used . to . estimate . the . divot . trajectories . from . the . high-speed . digital .

video. The divot spatial position, rotation, and velocity were estimated for the in-flight ejections.

Multiple . reference . points . on . the . divot . were . used . to . determine . rotation . Velocity . was . calculated . by .

differentiating . the . spatial . coordinates . with . respect . to . time . Reference . 7 . provides . details . and . results .

of . the . photogrammetry . technique Camera Pods The . high-speed . camera . heads . were . housed . in . camera . pods . mounted . on . the . forward . and . aft .

fuselage left side missile rail stations (fig. 2). The camera pods are made of aluminum and have a wedge shape. The forward camera pod has a length of 46 in. (1.17 m) and a leading-edge wedge angle of 13.4°. The aft camera pod has a length of 50.8 in. (1.29 m) and a leading-edge wedge angle of 13.88°. Each camera pod has a removable 1/4-in-thick (6.35 mm), borosilicate crown optical glass window with an antireflective coating.

The forward camera had a field of view of approximately 34° and a view width of approximately 5 ft (1.524 m) at an object distance of roughly 10 ft (3.048 m). The aft camera had a field of view of approximately 67° and a view width of approximately 5 ft (1.524 m) at an object distance of roughly 4 ft (1.22 m). The camera head mount in the camera pod permitted fine adjustment of the viewing . area . of . approximately . –2° . to . +7° . up . and . down . and . ±10° . left . to . right . Figure . 6 . shows . the .

forward . and . aft . camera . viewing . angles . relative . to . the . AFTF AFTF Flow 34° field of view 67° field of view Aft camera pod Forward camera pod 050478 Figure . 6 . Camera . pod . views . relative . to . the . Aerodynamic . Flight . Text . Fixture Several . factors . were . considered . in . the . aerodynamic . design . of . the . camera . pods . The . right .

inboard side of each camera pod was parallel to the free-stream flow to reduce shock wave impingement . on . the . AFTF . The . camera . pod . leading-edge . turning . angle . was . on . the . left . outboard .

side . of . each . pod . The . forward . camera . pod . leading-edge . wedge . angle . was . selected . to . ensure . an .

attached . shock . wave . at . Mach . 1 6 . and . greater For . structural . design . considerations, . aerodynamic . analyses . were . conducted . on . the . camera .

pods . to . assess . leading-edge . shock . wave . locations, . base . region . reattachment . shock . locations, .

base region wake flow impingement, and pressure distributions. Analyses included application of simple shock-expansion wave theory and three-dimensional ( 3 -D), inviscid computational fluid dynamics (CFD).

The . 3-D . CFD . was . for . the . complete . F-15B . aircraft . with . the . AFTF . and . camera . pods . attached .

The CFD calculations were made for the Mach 1.2, 1.6, and 2.0 supersonic flight test conditions.

Figure . 7 . shows . the . results . for . the . Mach . 2 . case . For . all . of . the . supersonic . CFD . cases, . no . indication .

of . shock . impingement . on . the . AFTF . was . observed . from . the . camera . pods (a) Complete aircraft.

Figure 7. The F-15B aircraft computational fluid dynamics analysis.

(b) Bottom of aircraft.

(c) Left side of aircraft.

Figure . 7 . Concluded The . impact . of . the . camera . pods . on . aircraft . stability . and . control . was . predicted . based . on .

comparison . to . other . stores . The . main . area . of . concern . was . degradation . in . the . lateral-directional .

stability of the F-15B aircraft at high supersonic speeds. During a previous flight experiment, a . longer . camera . pod . with . a . shortened . span . had . been . successfully . carried . on . the . F-15B . missile .

rail stations. This pod was flown to Mach 2 with no stability and control issues. The size of the present . camera . pods . also . was . compared . to . that . of . the . AIM-7 . Sparrow . air-to-air . missile . (Raytheon . .

Company, . Waltham, . Massachusetts, . and . General . Dynamics . Propulsion, . now . Aerojet-General .

Corporation, Rancho Cordova, California). Two AIM-7 Sparrow missiles can be carried simultaneously . on . the . F-15B . forward . and . aft . missile . rails . and . carriage . of . the . missiles . is . cleared . for .

the full F-15B flight envelope. The side force area for the two camera pods, which impact lateral- directional . stability, . was . approximately . 60-percent . less . than . the . area . for . the . two . AIM-7 . Sparrow .

missiles. No stability and control issues were anticipated based on the significantly smaller side force area, and none were encountered in flight.

Static . and . dynamic . structural . analyses . were . performed . to . verify . the . structural . integrity . of . the .

camera . pods . The . camera . pods . had . positive . static . structural . margins . of . safety, . with . a . 2 25 . factor . of .

safety, for the worst-case pressure loads at 600 kn (308.7 m/s), sea level altitude.

An aluminum conduit, 0.75 in. (0.1905 m) in diameter and approximately 12 ft (3.66 m) long, . was . externally . mounted . from . the . aft . end . of . the . forward . missile . rail . to . the . forward . end . of .

the . aft . missile . rail . to . house . the . camera . system . video . cable . Simple . band . clamps . and . 12 . fasteners .

were . used . to . attach . the . external . conduit . to . the . aircraft . fuselage . A . MIL-S-8802 . fuel . tank . sealant- adhesive . was . used . to . fair . the . sides . of . the . conduit . into . the . fuselage .. Aerodynamic . pressure . and .

skin . friction . loads . were . calculated . for . the . external . conduit . to . ensure . that . the . fastener . arrangement .

was . adequate Foam s heets The . foam . sheets . were . constructed . out . of . aluminum . plates . that . had . been . sprayed . with . shuttle .

 external . tank . Stepanfoam BX-265 (Stepan Company, Northfield, Illinois) TPS insulating foam.

The foam thickness was 2 in. (0.0508 m). Three AFTF bays were configured to carry a foam sheet, bays 2B, 3A, and 3B (fig. 4). Six foam sheets were sprayed for each AFTF bay. One sheet from each bay was used for ground testing and the remaining five from each bay were used for .

flight testing.

 .

In . addition . to . the . Stepanfoam . BX-265 . foam, . an . inexpensive . closed-cell . blue . Styrofoam (Dow Chemical Company, Midland, Michigan) was used for the ground test and early flight test.

3 3 The . Styrofoam . has . a . density . of . 2 0 . lbm/ft . (32 0 . kg/m ), similar to the density of the Stepanfoam BX-265 . foam . The . Styrofoam . was . an . easily . available, . inexpensive . substitute . for . the . actual . shuttle .

foam . for . development . of . the . divot . ejection . system . and . functional . checkouts . The . Stepanfoam . BX- 265 . foam . was . sprayed . on, . whereas . the . blue . Styrofoam . was . simply . cut . to . the . desired . dimensions .

and . bonded . onto . the . aluminum . backing . plates . with . a . MIL-S-8802 . adhesive . Despite . the . difference .

in . the . way . in . which . the . foam . was . attached . to . the . backing . plates, . the . divots . produced . from . the . blue .

Styrofoam . were . very . similar . to . those . produced . from . the . Stepanfoam . BX-265 . foam Cylindrical . voids . were . precut . into . the . back . of . the . foam . sheets . adjacent . to . the . aluminum . plate, .

simulating . an . air . void . on . the . shuttle . external . tank . Nitrogen . gas . was . used . to . pressurize . the . void .

through . a . hole . in . the . aluminum . plate . The . foam . next . to . the . void . was . scored . to . assist . in . creating .

a . fracture . line . when . the . foam . was . back . pressured . Each . sheet . had . three . voids . and . therefore . was .

capable . of . producing . three . divots Figure . 8 . and . table . 2 . show . the . foam . plate . dimensions . and . divot . sizes, . respectively, . including .

the void diameters and depths. The divot sizes used in the flight test were selected based on predictions . for . the . largest . expected . voids . at . various . locations . on . the . shuttle . external . tank . The .

void diameter varied from 0.31 to 1.68 in. (0.007874 to 0.017272 m). The void depth varied from 0.89 to 1.35 in. (0.02261 to 0.03429 m). The predicted divot diameter produced from these void sizes varied from 2.5 to 5.5 in. (0.0635 to 0.1397 m). Foil thermocouples were used to measure foam . surface . temperatures . at . two . locations . on . each . foam . sheet . The . thermocouples . were . bonded .

to . the . sheet . near . the . corner . edges . of . the . panels . so . as . not . to . aerodynamically . interfere . with . the . .

divot . ejection 17.75 in.

17.38 in.

13.41 in.

10 x 0.199-in.

8.88 in.

diameter 0.75 in.

4.16 in.

5.50-in.

0.38 in. 5.00-in.

3.40-in.

diameter 0.38 in. diameter diameter 1.25 in.

6.25 in.

7.63 in.

8.00 in.

3 x 0.50-in. diameter 5.31 in. 5.31 in.

4 x 0.25-in. radius NPT far side 050482 (a) Bay 2B foam sheet.

Figure . 8 . Aerodynamic . Flight . Test . Fixture . foam . plate . dimensions . and . divot . sizes 18.00 in.

9.00 in.

6.75 in.

4 x 0.25-in.

0.375 in. radius 4.50 in.

24 x 0.199-in.

0.375 in. diameter 2.25 in.

0.75 in.

5.50-in.

3.00 in.

3 x 0.50-in. diameter diameter C NPT far side L 5.75 in.

11.50 in.

2.50-in.

5.00-in.

diameter diameter 050483 (b) Bay 3A foam sheet.

20.00 in.

10.00 in.

0.375 in.

4.20-in.

4.00 in. 0.75 in.

diameter 5.00-in.

5.50-in.

0.375 in.

diameter diameter 1.50 in.

5.00 in.

4.00 in.

C L 8.00 in.

3 x 0.50-in. diameter 5.87 in. 5.87 in.

4 x 0.25-in. radius NPT far side 050484 10 x 0.199-in. diameter (c) Bay 3B foam sheet.

Figure . 8 . Concluded Table . 2 . Divot . cases Case Void . diameter, . in Void . depth, . in 1 0 31 1 35 3 1 30 1 00 5 0 56 1 17 7 1 67 0 89 9 1 68 1 05 Divot e jection s ystem The . pressure . required . to . eject . divots . from . the . foam . sheets . was . supplied . by . a . pneumatic .

system . carried . in . the . aft . section . of . the . AFTF . Major . components . of . the . pneumatic . system . included .

a reservoir tank, manual fill valve, manual final stop valve, pressure regulator, electric solenoid valves, and associated fluid lines. Figure 9 illustrates the pneumatic system, and table 3 provides details . of . the . pneumatic . system . components . Pneumatic . system . measurements . included . tank .

pressure, . tank . wall . temperature, . regulated . pressure, . and . gas . temperature . downstream . of . the .

pressure . regulator Valves (AFTF Bay 2A) Fill Tank wall valve temperature measurement Manual Temperature stop Regulator measurement valve Reservoir tank 10 micron Pressure Pressure filter measurement measurement Valves (AFTF Bay 2B) Valves (AFTF Bay 3B) 050485 Figure . 9 . Divot . ejection . pneumatic . system Table . 3 . Divot . ejection . pneumatic . system . components Burst . Operating .

Model . / MEOP, Component Manufacturer pressure, temperature, part . No lbf/in lbf/in °F Lockheed . Martin .

Tank Type . 30-40, .

Corporation, . (Bethesda, . 3,000 6,650 –65 . to . +325 (200 . in ) MIL-R-8573A(ASG) Maryland) Fill . needle Alta-Robbins .

SSKG250-4T 6,000 24,000 –40 . to . +250 valve (Lindon, Utah) Manual . stop Swagelok . Company .

SS-43TS4 3,000 NA –65 . to . +150 ball . valve (Solon, Ohio) SS-HFML3B- Regulator Swagelok . Company 500 NA –10 . to . +150 VCR4-P-BK Solenoid Marotta . Controls, . Inc . MV510H/ 6,000 18,000 –65 . to . +165 valve (Montville, New Jersey) 805764-3312 Fluid . lines NA 6061T6 . aluminum 1,500 6,000 NA Wintec . Industries . .

Filter 12267-556 3,000 NA –423 . to . +800 (Fremont, California) The . pneumatic . system . complied . with . MIL-STD-1552A, . “General . Requirements . for . Safe .

Design and Operation of Pressurized Systems” (ref. 8). No relief valves were required in the system, because the maximum expected operating pressure (MEOP) could not be exceeded by design.

The . 200-cubic-in . (0 00328 . m ) reservoir tank was filled with nitrogen gas to a nominal operating . pressure . of . approximately . 300 . psia . (2,068,427 . N/m ). The reservoir tank was connected to . a . pressure . regulator . that . reduced . the . 300-psia . source . pressure . to . the . required . divot . ejection .

pressure, . nominally . 40 . to . 80 . psia . (275,790 . to . 551,581 . N/m ). The regulator pressure was set to the desired ejection pressure on the ground before flight. The regulator was plumbed to solenoid valves . that . were . attached . to . the . backs . of . the . foam . sheets . The . solenoid . valve . was . opened . for . .

300 . milliseconds, . which . back . pressured . the . foam . sheet, . resulting . in . a . divot . ejection s ynchronization s ystem A . system . was . developed . to . synchronize . the . high-speed . camera . and . divot . ejection . systems .

The . synchronization . system . digital . card . triggered . the . high-speed . camera . and . the . divot . ejection .

in . a . timed . and . sequential . order . A . relay . card . was . developed . to . provide . the . switching . interface .

between . the . digital . card . and . divot . ejection . system . The . synchronization . system . instrumentation .

included . monitoring . of . the . digital . card . trigger . impulses Two . types . of . divot . ejection . triggering . actions . were . possible, . a . single-mode . trigger . and . a .

multiple - mode . trigger . The . single . trigger . option . provided . a . single . divot . ejection . and . video . recording .

for each individual trigger action. This option provided the flexibility to eject single divots at various test conditions during the same flight or to simply allow time between divot ejections. The multiple . trigger . option, . which . was . used . for . the . supersonic . test . points, . ejected . multiple . divots, . with .

a preset 3-second delay between ejections. The 3-second delay between firings allowed the volume of . gas . between . the . regulator . and . solenoid . valves . to . recover . to . the . set . regulator . pressure The . digital . logic . card . used . a . programmable . logic . device . to . provide . the . desired . sequencing, .

delays, . and . holds . The . trigger . pulse . time . duration . for . activation . of . the . divot . ejection . system . solenoid .

valves . was . 300 . milliseconds . This . value . was . set . by . counters . that . were . compared . to . a . reference .

number . The . trigger . pulse . time . duration . then . could . be . easily . changed . by . simply . changing . the .

reference . number . in . the . synchronization . software . code . The . relay . card . operated . in . conjunction .

with . the . digital . card . to . drive . the . divot . ejection . system . solenoid . valves The . control . panel . that . included . the . trigger . function . was . located . in . the . F-15B . aft . cockpit . and .

contained the following switches: initiate, mode, arm, and trigger (fig. 10). The initiate switch was a . locking . toggle . switch . that . provided . power . to . the . digital . card . The . mode . switch . selected . either . the .

single . or . multiple . trigger . mode . The . arm . switch . was . a . locking . toggle . switch . that . provided . power .

to . the . relay . interface . card . A . control . panel . lamp . was . illuminated . when . the . arm . mode . was . activated .

The . trigger . switch . was . a . guarded . momentary . switch . that . triggered . the . digital . card . to . initiate . the .

high-speed . video . and . divot . ejection . sequence LIFT CONTROL PANEL SINGLE MULTI MODE ARMED ON ON ON ARM TRIGGER INITIATE 050486 Figure . 10 . Lifting . Insulating . Foam . Trajectory . aft . cockpit . control . panel Health . monitoring . outputs . from . the . synchronization . system . were . telemetered . in . real . time . to .

the . control . room . These . outputs . included . the . digital . card . “heartbeat” . to . verify . initialization . and .

correct . function, . camera . record . status, . and . arm . switch . status

F lig H t test C on D itions an D M aneu V ers

The flight envelope for the present flight test was bound by the operating limitations imposed by carriage of the AFTF (fig. 11). These limitations included an absolute speed limit of 600 kn (308.7 m/s), maximum Mach number of 2 because of heating of the AFTF composite structure, 2 2 maximum . dynamic . pressure . of . 1,100 . lbf/ft . (52,668 . N/m ), and a maximum value of the product 2 2 of . sideslip . angle . and . dynamic . pressure, . β q , . of . 5,500 . deg-lbf/ft . (263,341 . deg-N/m ). For aircrew safety, a maximum altitude of 50,000 ft (15,240 m) was imposed. All of the divot ejection flight test points were flown within the AFTF flight envelope.

70,000 STS-107 F-15B aircraft Shuttle ascent test point 60,000 F-15B aircraft flight envelope 50,000 AFTF flight AFTF flight M = 2.3 envelope envelope 40,000 Altitude, M = 2.0 ft 30,000 M = 1.8 20,000 M = 1.5 10,000 M = 0.5 M = 1.0 600 700 800 900 400 500 100 300 200 Airspeed, KCAS 050487 Figure 11. The F-15B aircraft Lifting Insulating Foam Trajectory flight test envelope and .

test . points

The flight test conditions included subsonic and supersonic test points to approximately .

Mach 2 and an altitude of 50,000 ft (15,240 m), which matched discrete points along the shuttle ascent trajectory (fig. 11). The test point at 330 knots calibrated airspeed (KCAS) (557 m/s), Mach 0.7, and an altitude of 20,000 ft is a “heart of the envelope” condition that was the first test point flown. Figure 12 compares the dynamic pressures from the LIFT test point and shuttle ascent.

Because the local flow conditions measured at the AFTF differ slightly from the aircraft free- stream conditions, the aircraft flight condition was adjusted as required to match the AFTF local conditions with the desired shuttle ascent condition. The divot ejection was initiated at a flight test condition . in . which . the . wings . were . level, . altitude . was . constant, . and . Mach . number . was . constant .

Table 1 presents the flight test points flown and the local flow conditions of both the aircraft and AFTF. For reference, the STS-107 mishap flight test condition was at Mach 2.5, an altitude of 2 2 69,000 ft (21,031 m), and a dynamic pressure of 425 lbf/ft . (20,349 . N/m ).

STS-114 F-15B aircraft Dynamic pressure, lbf/ft 0 0.5 1.0 1.5 2.0 2.5 Mach number 050488 Figure 12. Comparison of dynamic pressures from the Lifting Insulating Foam Trajectory flight test . point . and . shuttle . ascent The divot ejection flight test conditions can be separated into three general categories: .

(1) maximum Mach number, (2) maximum dynamic pressure, and (3) maximum aerodynamic torque on the divot, embodied by the product of the divot moment coefficient and dynamic pressure, . C q . The . maximum . Mach . number . test . condition . was . at . Mach . 2 0, . an . altitude . of . 48,354 . ft . .

m 2 2 (14,738 m), and a dynamic pressure of 736 lbf/ft . (35,240 . N/m ). The maximum dynamic pressure test condition was at Mach 1.60, an altitude of 36,124 ft (11,011 m), and a dynamic 2 2 pressure . of . 848 . lbf/ft . (40,602 . N/m ). The maximum aerodynamic torque test condition was nominally at Mach 1.2, an altitude of 27,400 ft (8,352 m), and a dynamic pressure of 710 lbf/ft . .

(33,995 . N/m ).

H a Z ar D s assess M ent Two major hazards associated with the present flight test were the recontact of the ejected divot . with . the . aircraft . and . divot . ground . impact . issues . Assessment . of . the . recontact . hazard . focused .

on . the . potential . damage . to . the . F-15B . aircraft, . whereas . assessment . of . the . ground . impact . hazard .

focused . on . the . potential . damage . to . property . or . injury . to . persons . on . the . ground Divot r econtact The divot recontact hazard was investigated from two approaches: predicting the trajectories of the ejected . divots . and . evaluating . the . possible . structural . impact . damage . Two . techniques . were . used . to .

predict the divot trajectories: a simple one-degree-of-freedom (1-DOF) trajectory calculation and a simplified trajectory calculation through a 3-D CFD flow field.

The 1-DOF calculation modeled the divot as a point mass. The divot drag was specified as a . function . of . Mach . number . only . The . divot . cross . range . was . generated . as . a . function . of . down- range . distance . by . means . of . six-degree-of-freedom . CFD . models . The . 1-DOF . analysis . predicted .

possible . divot . recontact . with . the . aircraft . with . a . lower . probability . of . contact . with . the . stabilizer . The .

prediction . indicated . no . contact . with . the . leading . edge . of . the . stabilizer The 3-D CFD analysis calculated the flow field around the complete F-15B aircraft with the AFTF . attached . The . divot . trajectory . then . was . computed . as . the . divot . moved . through . this . steady . .

3-D flow field (fig. 13). The lift vector, moving the divot through the flow field, was calculated under the assumption of a worst-case alignment of the divot to produce the maximum lift. The 3-D flow fields about the AFTF were calculated for Mach 1.2, 1.6, and 2.0. The 3-D CFD analysis predicted some . upwash . in . the . aft . region . of . the . AFTF, . downstream . of . divot . release . point . The . predictions .

indicated . that . the . foam . divots . could . recontact . the . aft . underside . of . the . aircraft . and . stabilizers Figure 13. Divot recontact computational fluid dynamics predictions.

Because . both . trajectory . analyses . predicted . that . the . divots . could . recontact . the . aircraft, . a .

structural damage assessment was performed, based on a previous analysis by Ko (ref. 9) for another F-15 aircraft flight test in which shuttle foam was used. The first step in the damage assessment . was . to . identify . the . structure . and . materials . in . the . potential . recontact . area . These . materials .

included various types of aluminum (2024-T81, 2024-T62, and 7075-T76) and titanium (6Al-4V). .

Two . types . of . foam . materials . were . assumed, . shuttle . external . tank . foam . and . blue . Styrofoam . The analysis . was . conducted . assuming . the . worst . case . of . a . complete . sheet . of . foam . departing . the . AFTF .

and . recontacting . the . aircraft The . maximum . strain . energy . density . of . the . foam . was . compared . to . that . of . the . aircraft . structural .

materials (table 4). The maximum strain energy density of the shuttle foam is 0.65 in-lbf/in . .

(4,482 . J/m ), more than two orders of magnitude lower than the values for the aircraft materials.

Based . on . this . comparison, . the . foam . was . predicted . to . fail . before . the . aircraft . structure . failed Table . 4 . Comparison . of . maximum . strain . energy . density . of . foam . and . aircraft . materials Strain . energy Material density, in-lbf/in Foam 0 65 Aluminum (2024-T81) 166 Aluminum (7075-T76) 166 Aluminum (2024-T62) 105 Titanium (6Al-4V) 496 A . 1-DOF . trajectory . analysis . was . used . to . predict . the . kinetic . energy . of . the . foam . at . a . point . 40 .

ft (12.2 m) downstream of the ejection point. This prediction provides a conservatively large value of . the . kinetic . energy, . because . the . distance . from . the . divot . ejection . point . to . the . trailing . edge . of . the .

aircraft stabilizer is approximately 26 ft (7.92 m). The assumed dimensions of the foam sheet are 18 by 6.5 by 2 in. (0.4572 by 0.1651 by 0.1656 m). The kinetic energy of the foam sheet, 40 ft (1.016 m) downstream of the ejection point, is approximately 311 ft-lbf (421.7 J).

The . energy . required . to . compress . the . foam, . E , . when . a . 90-percent . void . content . is . assumed, . is .

cr 5,265 in-lbf (594.9 J) for the blue foam and 8,887 in-lbf (1,004 J) for the shuttle foam. Because the E . is . much . larger . than . the . worst-case . foam . sheet . kinetic . energy, . the . foam . is . capable . of . absorbing .

cr the . impact . kinetic . energy . by . compacting . This . analysis . is . even . further . conservative, . because . it .

assumes . that . the . foam . hits . the . aircraft . skin . at . an . angle . of . 90 ° . as . opposed . to . an . oblique . impact The . possibility . of . denting . the . aircraft . skin . also . was . evaluated . The . aluminum . honeycomb .

2 7 backing . behind . the . aircraft . skin . has . a . yield . stress . between . 14,000 . and . 37,000 . lbf/in . (9 653x10 .

8 2 and . 2 551x10 . N/m ). The stress induced by the honeycomb on the skin, resulting from the pressure 2 2 force . that . the . foam . exerts . on . the . skin . from . a . foam . impact, . is . 1,190 . lbf/in . (56,978 . N/m ) for the 2 2 blue . foam . and . 2,009 . lbf/in . (96,191 . N/m ) for the shuttle foam. Therefore, the possibility of denting the aircraft skin was predicted to be very small. Postflight visual inspections were performed and no . damage . from . the . divot . impacts . was . found Divot g round i mpact When . the . divot . was . ejected . from . the . aircraft, . the . resulting . ground . impact . possibly . could . have .

posed . a . hazard . to . people . or . property, . based . on . the . kinetic . energy . of . the . divot . or . debris . at . ground .

impact . Based . on . the . local . test . range . predetermined . criteria, . ground . safety . was . not . an . issue . if . the .

kinetic energy of the debris at impact was less than 11 ft-lbf (14.9 J). Two approaches were used to analyze this hazard: a simple analytical calculation of the divot kinetic energy based on its terminal velocity, . and . a . more . complex . computation . based . on . a . 1-DOF . trajectory . simulation Equation . 1 . was . used . to . calculate . the . divot . impact . kinetic . energy, . KE W 1 KE V = (1) term g 2 The . kinetic . energy . was . calculated . under . the . assumption . of . the . largest . expected . divot . diameter, .

3 3 7 . in , . resulting . in . a . divot . volume . of . 0 0223 . ft . (0 000631 . m ) and a divot weight, W , . of . 0 0446 . lb .

(0.02023 kg). Equation 2 was used to calculate the divot terminal velocity, V term W 2 (2) V = term C A ρ d cross tion − air sec The divot flat plate drag coefficient, C , was assumed to be 1.28 (ref. 10). A worst - case d situation in which the divot falls from an altitude of 50,000 ft (15,240 m) was used with a 3 3 conservative . constant . value . of . the . air . density, . ρ , . of . 0 0003639 . slugs/ft . (0 1877 . kg/m ). Two air models . of . the . divot . cross-sectional . area . were . used, . the . area . assuming . the . divot . falls . “edge . on” . .

2 2 (7 . in , . 0 00452 . m ), and the area assuming the divot falls with its planform perpendicular to the 2 2 flow (38.48 in , . 0 02483 . m ).

When . these . assumptions . are . used, . the . terminal . velocity . and . impact . kinetic . energy . for . the .

divot that falls edge on are 62.77 ft/s (19.13 m/s) and 2.73 ft-lbf (3.70 J), respectively. For the divot that falls with its planform perpendicular to the flow, the terminal velocity and impact kinetic energy are 26.77 ft/s (8.159 m/s) and 0.496 ft-lbf (0.672 J), respectively. The impact kinetic energy for both cases is well below the limit of 11 ft-lbf (14.9 J).

The 1-DOF computation is a higher fidelity prediction than the simple analytical calculation.

It . includes . the . proper . variation . of . air . density . with . altitude . and . an . aerodynamic . model . that . accounts .

for . the . tumbling . of . the . foam . piece . Two . foam . densities . were . considered, . a . “light” . foam . density .

3 3 3 3 of . 1 8 . lbm/ft . (28 8 . kg/m ), and a “heavy” foam density of 2.6 lbm/ft . (41 6 . kg/m ). The F-15B computation . is . a . worst-case . situation . in . which . the . entire . foam . sheet . is . assumed . to . depart . the .

aircraft . The . highest . terminal . velocity . and . kinetic . energy . obtained . from . the . 1-DOF . computations .

are 23 ft/s (7.01 m/s) and 3 ft-lbf (4.07 J), respectively.

Both . types . of . calculations . predicted . ground . impact . kinetic . energies . well . below . the . limit . of .

11 ft-lbf (14.9 J). Therefore, the divot ground impact hazard was not deemed a safety issue.

results an D D is C ussion A total of 10 divot ejection flights were completed, resulting in 41 successful, 5 failed, and 2 . partial . divot . ejections . The . blue . Styrofoam . was . used . in . 5 . of . the . successful . divot . ejections . and .

1 . of . the . failed . ejections . The . shuttle . foam . was . used . in . 36 . successful, . 4 . failed, . and . 2 . partial . divot .

ejections. Table 1 provides a summary of the divot ejection flight conditions and results.

Flight t est Conditions The flying qualities of the aircraft when the AFTF is attached are well known from the initial AFTF flight test program (ref. 3) and many subsequent flight tests in which the AFTF was attached.

The new larger camera pods were not expected to cause any significant changes in the aircraft flying qualities. An envelope expansion process was followed to ensure that no adverse flying qualities . were . encountered, . especially . for . high . supersonic . Mach . numbers . in . which . the . F-15B .

lateral-directional . stability . decreases Table 1 presents the details of the aircraft and AFTF flight test conditions for all of the test points. Representative flight condition data are discussed for LIFT flight 9. The divot ejection flight test condition for this flight was nominally at Mach 2 and a pressure altitude of 48,250 ft 2 2 (14,707 m), with a dynamic pressure of approximately 730 lbf/ft . (34,953 . N/m ). Three shuttle foam . divots . were . successfully . ejected . at . the . Mach . 2 . test . condition Figure . 14 . shows . altitude . as . a . function . of . Mach . number . for . the . acceleration . from . approximately .

Mach 0.9 to 2.0. Because of significant added drag from the AFTF and two camera pods, an optimum trajectory had to be followed to reach Mach 2. Figure 15 shows the specific excess power, P , . overlaid . on . the . altitude–Mach . number . plot . The . P . has . been . normalized . by . the . maximum . P .

s s s value . during . the . acceleration . The . magnitude . of . the . P . is . an . indicator . of . the . energy . available . to .

s accelerate . and/or . climb 49,000 48,000 47,000 46,000 Altitude, 45,000 ft 44,000 43,000 42,000 41,000 0.9 1.0 1.1 1.2 1.3 1.4 1.5 1.6 1.7 1.8 1.9 2.0 Mach number 050490 Figure 14. Altitude as a function of Mach number for supersonic acceleration (flight 9).

1.0 49,000 - Altitude Specific 0.9 - 48,000 excess power P ( ) s 0.8 - 47,000 0.7 - 46,000 0.6 - Altitude, P P , s / s 45,000 0.5 - ft max 0.4 - 44,000 0.3 - 43,000 0.2 - 42,000 0.1 - 41,000 0 - 0.9 1.0 1.1 1.2 1.3 1.4 1.5 1.6 1.7 1.8 1.9 2.0 Mach number 050491 Figure 15. Specific excess power and altitude as a function of Mach number (flight 9).

In . a . region . of . lower . P , . the . aircraft . dove . through . to . supersonic . speeds . from . approximately .

s Mach 0.9 at an altitude of 46,000 ft (14,021 m) to roughly Mach 1.27 at an altitude of 42,000 ft (12,802 m). At an altitude of approximately 42,000 ft (12,802 m), with a high value of P , . the .

s aircraft accelerated in level flight to roughly Mach 1.41. As the P . decreased, . the . aircraft . climbed .

s to an altitude of approximately 45,000 ft (13,716 m).

Figures . 16 . and . 17 . show . the . aircraft . angle . of . attack . and . angle . of . sideslip, . respectively, . as .

a . function . of . Mach . number . The . angle . of . attack . varied . from . a . maximum . of . approximately . 4 7° .

at . roughly . Mach . 0 95 . to . a . minimum . of . approximately . 0 5° . at . Mach . 2 . The . angle . of . sideslip . was .

negative (nose right) below Mach 1.36 and slightly positive (nose left) above Mach 1.36. The angle . of . sideslip . was . approximately . 0 2° . at . the . Mach . 2 . test . point Angle of attack, deg –1 0.9 1.0 1.1 1.2 1.3 1.4 1.5 1.6 1.7 1.8 1.9 2.0 Mach number 050492 Figure 16. Angle of attack as a function of Mach number (flight 9).

1.0 0.5 Angle of sideslip, –0.5 deg –1.0 –1.5 –2.0 0.9 1.0 1.1 1.2 1.3 1.4 1.5 1.6 1.7 1.8 1.9 2.0 Mach number 050493 Figure 17. Angle of sideslip as a function of Mach number (flight 9).

Figure . 18 . compares . the . Mach . numbers . from . the . aircraft . and . the . AFTF . nose . booms . Note .

that . the . AFTF . local . Mach . number . was . different . from . the . aircraft . Mach . number . by . as . much . as . 0 1 .

Below . approximately . Mach . 1 5, . the . AFTF . measured . Mach . number . was . lower . than . the . aircraft . Mach .

number . At . an . aircraft . Mach . number . between . roughly . 1 54 . and . 1 62, . the . AFTF . Mach . number . was .

constant . at . approximately . Mach . 1 53, . then . increased . discontinuously . to . match . the . aircraft . value . at .

roughly . Mach . 1 62 . The . lower . AFTF . Mach . number . in . this . region . was . caused . by . the . passage . of . the .

aircraft . inlet . shock . wave . over . the . AFTF . nose . boom . static . pressure . ports . When . the . Mach . number .

of . the . aircraft . was . greater . than . approximately . 1 7, . the . Mach . number . of . the . AFTF . was . greater . than .

that . of . the . aircraft, . and . the . difference . decreased . to . zero . at . roughly . Mach . 1 9 2.0 Aircraft 1.9 AFTF 1.8 1.7 1.6 1.5 Mach 1.4 number 1.3 1.2 1.1 1.0 0.9 0.8 50 0 100 150 200 250 300 350 400 450 Time, s 050494 Figure . 18 . Comparison . of . Mach . numbers . from . the . aircraft . and . Aerodynamic . Flight . Test . Fixture .

(flight 9).

Figure . 19 . compares . the . foam . surface . temperatures . of . the . LIFT . test . points . and . shuttle . trajectory .

as . a . function . of . Mach . number . Recall . that . the . F-15B . aircraft . test . points . were . on . the . shuttle . ascent .

trajectory, matching the Mach number and altitude, but the aircraft was not able to fly along the shuttle . ascent . trajectory . to . the . test . points . For . this . reason, . the . LIFT . foam . temperatures . at . the . test .

points . were . lower . than . the . shuttle . foam . temperatures Thermocouple 1 Thermocouple 2 Thrust panel acreage bipod 7406 Thrust panel flange bipod 7436 Stringer flange bipod 7430 Foam temperature, °F 0.5 0 1.0 1.5 2.0 2.5 Mach number 050495 Figure 19. Comparison of foam temperatures from the Lifting Insulating Foam Trajectory flight test . and . shuttle . ascent Divot e jection Of the 42 shuttle foam divot ejections that were attempted in flight, 36 resulted in successfully . ejected . divots . In . addition, . two . partial . divot . ejections . occurred . in . which . fractured .

pieces of foam (rather than a whole divot) separated from the sheet. Table 5 summarizes the divot ejection . results

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Table . 5 . Flight . test . divot . sizes Divot . Divot . Divot .

Flight . Mach . Divot .

Divot radius . 1 . radius . 2 . R1/R2 average .

No number width, . in (R1), in. (R2), in. radius, . in 270 1 23 H 0 30 1 0520 0 7295 1 44 0 891 271 1 63 H 0 30 0 9560 0 6195 1 54 0 788 272 1 79 H 0 30 0 9475 0 7450 1 27 0 846 277 1 60 H 1 30 0 7585 1 0170 0 75 0 888 274 1 20 E 0 56 1 1380 1 5115 0 75 1 325 275 1 22 E 0 56 0 9220 1 4055 0 66 1 164 276 1 99 E 0 56 1 1280 1 4875 0 76 1 308 269 0 72 B 1 30 2 0460 1 7970 1 14 1 922 270 1 22 B 1 30 2 0080 2 1390 0 94 2 074 271 1 63 B 1 30 2 1235 2 2035 0 96 2 164 272 1 79 B 1 30 2 1370 2 0770 1 03 2 107 277 1 60 B 1 30 2 0640 1 9525 1 06 2 008 269 0 71 C 1 67 2 3680 2 1690 1 09 2 269 270 1 22 C 1 67 2 4435 2 4195 1 01 2 432 271 1 63 C 1 67 2 4975 2 6120 0 96 2 555 272 1 80 C 1 67 2 5565 2 7075 0 94 2 632 277 1 60 C 1 67 2 6570 2 6165 1 02 2 637 274 1 19 D 1 67 2 7810 2 7105 1 03 2 746 275 1 22 D 1 67 2 4625 2 4750 0 99 2 469 276 1 99 D 1 67 2 6170 2 4375 1 07 2 527 269 0 61 I 1 67 2 2545 1 7920 1 26 2 023 Table . 5 . Continued Largest .

Divot . Divot .

divot . Divot . Divot . Divot .

thickness . L/HT volume, .

diameter . angle, . deg weight, . lb area, . ft ( HT ), in. ft ( L ), in.

1 78 0 65 2 74 41 3 0 00037 0 0008 0 04 1 58 0 65 2 42 45 5 0 00030 0 0006 0 03 1 69 0 65 2 60 43 0 0 00034 0 0007 0 04 1 78 1 00 1 78 76 6 0 00108 0 0023 0 06 2 65 0 83 3 19 38 5 0 00111 0 0023 0 09 2 33 0 83 2 80 43 2 0 00088 0 0019 0 07 2 62 0 83 3 15 38 9 0 00108 0 0023 0 08 3 84 1 00 3 84 38 2 0 00325 0 0069 0 18 4 15 1 00 4 15 35 1 0 00368 0 0078 0 21 4 33 1 00 4 33 33 5 0 00394 0 0083 0 22 4 21 1 00 4 21 34 5 0 00378 0 0080 0 21 4 02 1 00 4 02 36 4 0 00349 0 0074 0 20 4 54 1 11 4 09 37 8 0 00520 0 0110 0 25 4 86 1 11 4 38 34 8 0 00581 0 0123 0 28 5 11 1 11 4 60 32 8 0 00629 0 0133 0 31 5 26 1 11 4 74 31 7 0 00661 0 0139 0 33 5 27 1 11 4 75 31 6 0 00663 0 0140 0 33 5 49 1 11 4 95 30 2 0 00708 0 0150 0 35 4 94 1 11 4 45 34 2 0 00596 0 0126 0 29 5 05 1 11 4 55 33 3 0 00618 0 0131 0 30 4 05 1 11 3 65 43 0 0 00436 0 0092 0 21 Table . 5 . Continued Divot . Divot . Divot .

Flight . Mach . Divot .

Divot radius . 1 . radius . 2 . R1/R2 average .

No number width, . in (R1), in. (R2), in. radius, . in 270 1 22 A 1 68 2 1955 2 3200 0 95 2 258 271 1 63 A 1 68 2 2180 2 3170 0 96 2 268 272 1 79 A 1 68 2 2350 2 3890 0 94 2 312 277 1 60 A 1 68 2 3125 2 2055 1 05 2 259 274 1 21 F 1 68 2 2390 2 1680 1 03 2 204 275 1 22 F 1 68 2 2745 2 2460 1 01 2 260 276 2 00 F 1 68 2 2835 2 1905 1 04 2 237 269 0 61 G 1 68 1 9425 2 0020 0 97 1 972 270 1 23 G 1 68 2 2460 2 2310 1 01 2 239 271 1 63 G 1 68 2 2460 2 2810 0 98 2 264 272 1 79 G 1 68 2 0845 2 0815 1 00 2 083 277 1 60 G 1 68 2 3595 2 2975 1 03 2 329 Table . 5 . Concluded Largest .

Divot . Divot .

divot . . Divot . Divot . Divot .

thickness . L/HT volume, .

diameter . angle, . deg weight, . lb area, . ft (HT), in. ft (L), in.

4 52 0 95 4 75 33 8 0 00443 0 0094 0 24 4 54 0 95 4 77 33 6 0 00446 0 0094 0 24 4 62 0 95 4 87 32 8 0 00460 0 0097 0 25 4 52 0 95 4 76 33 8 0 00444 0 0094 0 24 4 41 0 95 4 64 34 9 0 00427 0 0090 0 23 4 52 0 95 4 76 33 8 0 00444 0 0094 0 24 4 47 0 95 4 71 34 2 0 00437 0 0092 0 24 3 94 0 95 4 15 40 0 0 00360 0 0076 0 19 4 48 0 95 4 71 34 2 0 00437 0 0092 0 24 4 53 0 95 4 77 33 7 0 00445 0 0094 0 24 4 17 0 95 4 39 37 4 0 00391 0 0083 0 21 4 66 0 95 4 90 32 5 0 00465 0 0098 0 26 All . of . the . divot . ejections . were . captured . with . high-speed . digital . video . at . 2,000 . pps . Figures .

20 and 21 show composite digital video frame captures of divot ejections from flight 9 at .

Mach 1.99, an altitude of 48,240 ft (14,704 m), and a dynamic pressure of 730 lbf/ft . (34,953 .

N/m ). Figure 20(a) shows the forward camera view, and figure 20(b) shows the aft camera view.

The . conical . frustum-shaped . divot . was . cleanly . ejected . from . the . AFTF . and . trimmed . with . its . small .

diameter . facing . upstream (a) Forward camera view (looking downstream).

(b) Aft camera view (looking upstream).

Figure 20. Composite digital video frame captures (at 2,000 pps) of divot ejections.

Figure 21. Composite digital video frame captures (at 2,000 pps) of divot ejections, forward camera view . Figures 22 and 23 show examples of the shuttle foam panels after the in-flight divot ejection. Figure 22 shows the foam panels in the AFTF bays 3A and 3B after LIFT flight 4 at Mach 2 2 1.63, an altitude of 40,200 ft (12,253 m), and a dynamic pressure of 720 lbf/ft . (34,474 . N/m ). All six divots were successfully ejected. Figure 23 shows a postflight close-up picture of the shuttle foam panel 3A from LIFT flight 3 at Mach 1.2, an altitude of 29,000 ft (8,839 m), and a dynamic 2 2 pressure . of . 700 . lbf/ft . (33,516 . N/m ).

Figure 22. Foam panels after the in-flight divot ejection.

.

Figure 23. Postflight close-up view of foam panel.

.

Figure . 24 . shows . the . pressure . required . to . cleanly . eject . divots . from . the . foam . sheets . as . a . function .

of divot void diameter. The divot ejection pressure is shown as a gauge pressure defined as the difference . between . the . absolute . pressure . behind . the . divot . and . the . free-stream . static . pressure . on .

the external surface of the divot. The figure shows both ground and flight test data. The ejection pressure . did . not . appear . to . affect . the . divot . geometry Ground (ejected) Ground (not ejected) Flight (ejected) Flight (not ejected) 50 Ejection pressure minus ambient pressure, psia 0.2 0.3 0.4 0.5 0.6 0.7 0.8 0.9 1.0 1.1 1.2 1.3 1.4 1.5 1.6 1.7 1.8 Divot void diameter, in.

050496 Figure . 24 . Divot . ejection . pressure . as . a . function . of . divot . void . diameter Divot s tructural i ntegrity Thirty-six . divots . survived . the . aerodynamic . deceleration . associated . with . ejection . into . the .

flow field. Of the three divots generated from the lowest successful ejection pressure, two rotated back . into . the . foam . sheet . As . a . result . of . this . recontact . with . the . sheet, . these . divots . fractured . into .

several . pieces Divot s hape and s ize The . ejected . divots . had . a . conical . frustum . shape, . in . which . the . larger . and . smaller . diameters .

are defined as the divot diameter and void diameter, respectively. Figure 25 defines the divot geometry . Most . of . the . ejected . divots . had . a . near . circular . shape, . except . for . the . divots . obtained .

from the smallest void (see table 2, divot case 1). The smallest voids, nominally less than 1/2 in. .

(12.7 mm) in diameter, tended to produce elliptical outer surface divots.

Divot diameter Divot Divot thickness Divot angle Foam thickness Void diameter Void depth 050497 Figure 25. Definition of divot geometry.

Figure 26 shows the divot aspect ratio, defined as the largest divot diameter, L , . divided . by .

divot . thickness, . HT , as a function of divot thickness. Divot data are presented for the F-15B flight test and for various ground tests. The aspect ratios of the divots obtained from the flight tests all were . smaller . than . those . obtained . from . ground . tests . In . the . ground . tests, . the . outer . and . inner . foam .

surface . temperatures . were . matched . to . the . shuttle . conditions, . hot . wall . outer . surface . and . cold . wall .

inner surface. The outer temperature of the flight test foam was colder and the inner temperature was . warmer . than . the . temperatures . of . the . ground . test . foam . This . mismatch . in . foam . temperature .

resulted in different divot wall angles between ground and flight. Table 5 presents details of the flight test divot sizes.

F-15B aircraft flight data 6.0 7.4 5.1 Aspect ratio, L/HT 4.8-in. diameter divot 4 2.5-in. diameter 5.0-in.

divot diameter 5.5-in. diameter divot divot 0.6 0.4 0.8 1.0 1.2 1.4 Divot thickness, HT, in.

050498 Figure . 26 . Divot . aspect . ratio . as . a . function . of . divot . thickness Divot a erodynamic s tability The DTA assumes three potential aerodynamic models for the divot shape: the static or high drag . model . in . which . the . divot . trims . with . its . small . diameter . into . the . wind, . the . nominal . drag . model .

in . which . the . divot . oscillates . back . and . forth . about . its . static . trim . point, . and . the . tumble . or . low . drag .

model . in . which . the . divot . tumbles . The . static . and . tumbling . models . are . considered . to . bound . the .

potential deceleration of the divot when it is released into the flow field.

As shown in figure 20, the divots tended to trim with their small diameters forward (facing into the wind). All 31 of the supersonic divots trimmed. Of the five subsonic divots, two tumbled after . one . oscillation . Figure . 27 . shows . the . situation . in . which . the . divot . trimmed . with . its . large .

diameter facing upstream during flight 10 at Mach 1.6, an altitude of 36,124 ft (11,011 m), and 2 2 a . dynamic . pressure . of . 848 . lbf/ft . (40,602 . N/m ). The divot did not cleanly eject from the AFTF; instead, . the . downstream . edge . of . the . divot . fractured . into . several . smaller . pieces . The . resulting .

asymmetrical . divot . was . ejected . and . passed . through . a . yawing . oscillation . before . trimming . with . its .

large . diameter . facing . upstream . The . estimated . divot . dimensions . are . 5 3 . by . 1 7 . by . 1 1 . in . (0 135 . by .

0.0432 by 0.0279 m), with a mass of 0.014 lbm (0.00635 kg).

Figure . 27 . Composite . digital . video . frame . captures . of . divot . ejection, . forward . camera . view . looking .

aft at 2,000 pps, time and number of frames from ejection (divot C).

Divot Photogrammetry and t rajectories Standard . photogrammetry . analysis . techniques . were . used . to . estimate . the . divot . trajectories . from .

the . high-speed . digital . video . Photogrammetry . uses . photographic . images . to . obtain . measurements .

of . position . and . rotation . Velocity . can . be . calculated . by . differentiating . the . spatial . coordinates . with .

respect to time. The divot spatial position, rotation, and velocity were estimated for the in-flight ejections . Reference . 7 . presents . details . and . results . of . the . photogrammetry . technique Figure 28 shows trajectories calculated from the F-15B flight data by means of photogrammetry and . from . a . 1-DOF . DTA . prediction . Downrange . distance . is . plotted . as . a . function . of . time . from .

ejection. Three 1-DOF trajectories are shown based on the assumed divot stability-drag model: the nominal . drag . model . in . which . the . divot . oscillates . about . its . static . trim . point, . the . static . or . high . drag .

model . in . which . the . divot . trims . with . its . small . diameter . facing . upstream, . and . the . tumble . or . low . drag .

model . in . which . the . divot . tumbles . The . trajectory . based . on . the . nominal . drag . model . appears . to . most .

closely match the flight data. The trajectories based on the static (high drag) and tumble (low drag) models appear to properly bound the nominal trajectory and flight data.

3.0 Nominal 1-DOF Static 1-DOF Tumble 1-DOF 2.5 F-15B aircraft 2.0 Downrange 1.5 distance, ft 1.0 0.5 0 0.001 0.002 0.003 0.004 0.005 0.006 Time, s 050499 Figure 28. Comparison of F-15B flight data with one-degree-of-freedom prediction (flight 10, divot . C, . Mach . 1 6, . altitude . of . 36,120 . ft, . and . dynamic . pressure . of . 850 . lbf/ft ).

C on C lu D ing re M ar K s Flight . tests . were . conducted . on . the . NASA . F-15B . aircraft . to . assess . the . structural . survivability .

of Space Shuttle external tank foam debris or “divots” in a real flight environment. Divots were ejected . from . foam . sheets . mounted . on . the . side . of . the . Aerodynamic . Flight . Test . Fixture, . which . was .

carried . underneath . the . F-15B . aircraft . Divots . were . ejected . at . subsonic, . transonic, . and . supersonic .

speeds to Mach 2. A total of 10 divot ejection test flights were flown with 36 successful in-flight divot ejections. The divot ejection flight conditions matched the altitude and Mach number at discrete . points . along . the . shuttle . ascent . trajectory . High-speed . digital . video . was . used . to . capture . the .

divot . ejections . and . trajectories The . divots . remained . structurally . intact . if . they . were . cleanly . ejected, . that . is, . if . the . divots . did .

not . recontact . with . the . foam . sheets . or . Aerodynamic . Flight . Test . Fixture . In . general, . the . cleanly .

ejected divots had a nearly circular planform shape. The aspect ratios of the divots obtained in flight were . smaller . than . those . obtained . from . ground . tests . The . difference . in . aspect . ratio . is . the . result . of .

a mismatch in wall temperature conditions between the flight and ground tests. After ejection, the conical . frustum-shaped . divots . tended . to . trim . with . their . large . diameters . facing . upstream . All . 31 . of .

the supersonically ejected divots trimmed. Of the five divots ejected at subsonic speeds, two tumbled after one oscillation. The flight data were compared with debris transport analysis predictions of the . divot . trajectories . for . various . divot . drag . models . Photogrammetric . techniques . were . used . to .

estimate divot position, orientation, and speed from the flight data. The divot trajectories based on the nominal drag model of the debris transport analysis most closely matched the flight data.

re F eren C es Columbia Accident Investigation Board: Report, Volume 1, Aug. 2003. 1.

Richwine, David M., F-15B/Flight Test Fixture II: A Test Bed for Flight Research , 2.

NASA TM - 4782, 1996.

Meyer, Robert R., Jr., A Unique Flight Test Facility: Description and Results , 3.

NASA TM - 84900, 1982.

Corda, Stephen, M. Jake Vachon, Nathan Palumbo, Corey Diebler, Ting Tseng, Anthony Ginn, 4.

and David Richwine, The F-15B Propulsion Flight Test Fixture: A New Flight Facility for Propulsion Research , NASA TM-2001-210395, 2001.

Palumbo, Nathan, Timothy R. Moes, and M. Jake Vachon, Initial Flight Tests of the NASA 5.

F - 15B Propulsion Flight Test Fixture , NASA TM-2002-210736, July 2002.

Tseng, Ting, Matthew Reaves, and Kendall Mauldin, High Speed Video for Airborne 6.

Instrumentation Application (to be published).

Smith, Mark S., Photogrammetric Trajectory Estimation of Foam Debris Ejected from an F-15 7.

Aircraft , NASA TM-2006-213675, 2006.

Standard General Requirements for Safe Design and Operation of Pressurized Missile and 8.

Space Systems , MIL-STD-1522A, 1984.

Ko, William L., Impacts of Space Shuttle Thermal Protection System Tile on an F-15 Aircraft 9.

Vertical Tail , NASA TM-85904, 1985.

Hoerner, Sighard F., Fluid-Dynamic Drag: Practical Information on Aerodynamic Drag and 10.

Hydrodynamic Resistance , Dr.-Ing. S. F. Hoerner, Midland Park, New Jersey, 1965.

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REPORT DOCUMENTATION PAGE

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2. REPORT TYPE 1. REPORT DATE (DD-MM-YYYY) 3. DATES COVERED (From - To)

28-06-2006 Technical Memorandum

4. TITLE AND SUBTITLE 5a. CONTRACT NUMBER

The F-15B Lifting Insulating Foam Trajectory (LIFT) Flight Test

5b. GRANT NUMBER 5c. PROGRAM ELEMENT NUMBER 6. AUTHOR(S) 5d. PROJECT NUMBER Corda, Stephen, Whiteman, Donald, Tseng, Ting,

and Machin, Ricardo 5e. TASK NUMBER

5f. WORK UNIT NUMBER 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) 8. PERFORMING ORGANIZATION REPORT NUMBER

NASA Dryden Flight Research Center

P.O. Box 273

Edwards, California 93523-0273

H-2627

9. SPONSORING/MONITORING AGENCY NAME(S) AND ADDRESS(ES) 10. SPONSORING/MONITOR'S ACRONYM(S)

National Aeronautics and Space Administration

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NASA/TM-2006-213674

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Subject Category 015 Availability: NASA CASI (301) 621-0390 ! Distribution: Standard

13. SUPPLEMENTARY NOTES

Corda, Whiteman, Tseng, Dryden Flight Research Center; Machin, Johnson Space Center

An electronic version can be found at the NASA Dryden Flight Research Center Web Site, under Technical Reports.

14. ABSTRACT A series of flight tests has been performed to assess the structural survivability of space shuttle external tank debris, known as divots, in a real flight environment. The NASA F-15B research test bed aircraft carried the Aerodynamic Flight Test Fixture configured with a shuttle foam divot ejection system. The divots were released in flight at subsonic and supersonic test conditions matching points on the shuttle ascent trajectory. Very high-speed digital video cameras recorded the divot trajectories. The objectives of the flight test were to determine the structural survivability of the divots in a real flight environment, assess the aerodynamic stability of the divots, and provide divot trajectory data for comparison with debris transport models. A total of 10 flights to Mach 2 were completed, resulting in 36 successful shuttle foam divot ejections. High-speed video was obtained at 2,000 pictures per second for all of the divot ejections. The divots that were cleanly ejected remained structurally intact. The conical frustum-shaped divots

tended to aerodynamically trim in both the subsonic and supersonic free-stream flow. !!

15. SUBJECT TERMS Divot Aerodynamics, F-15 Flight Test Fixture, Foam Divot, Photogrammetry, Space Shuttle External Tank 19a. NAME OF RESPONSIBLE PERSON 18. NUMBER 17. LIMITATION OF 16. SECURITY CLASSIFICATION OF: OF ABSTRACT

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