C-5A/orbiter wind tunnel testing and analysis: Piggyback ferry

CONTENTS Section Title Page LIST OF FIGURES iv SUMMARY vii 1.0 INTRODUCTION 1 2.0 ANALYSIS OF EXPERIMENTAL RESULTS 2 2.1 Stability and Control 2 2.1.1 Effects of Orbiter, Cruise Configuration 2 2.1.2 Effects of Orbiter Position, Cruise Configuration 3 2.1.3 Effect of Orbiter Incidence, Cruise Configuration 6 2.1.4 Effect of Afterbody Fairing, Cruise Configuration 6 2.1.5 Effect of Orbiter, Landing Configuration 7 2.1.6 Vertical Tail Development 8 2.1.7 Effect of Orbiter on Trim 9 2.2 Drag Characteristics 9 2.2.1 Effect of Orbiter 9 2.2.2 Effect of Orbiter Position 10 2.2.3 Effect of Afterbody Fairing 10 2.2.4 Effect of Orbiter Incidence 10 3.0 ASSESSMENT OF FULL SCALE FLIGHT FEASIBILITY 11 3.1 Stability and Control 11 3.1.1 Comparison of Wind Tunnel and Full Scale Directional Stability 11 3.1.2 Predicted Full Scale Directional Stability 11 3.2 Flying Qualities 12 efd Georgia Company II

CONTENTS (Continued) Page Section Title 3.3 Performance 14 3.3.1 Full Scale Drag Characteristics 14 3.3.2 Airfield Performance 3.3.3 Climb and Cruise Performance 15 3.3.4 Orbiter Ferry Capability 3.4 Flight Restrictions 4.0 CONCLUSIONS AND RECOMMENDATIONS 18 Appendix A WIND TUNNEL TEST DESCRIPTION AND PLOTTED DATA A-l Loc'tieea Geoigia Company HI LIST OF FIGURES Figure Title Page 1 Effect of Orbiter on Longitudinal Stability - Cruise Configuration 20 2 Effect of Orbiter on Longitudinal Stability - Cruise Configuration, C. -a, C. -C.i 21 3 Effect of Orbiter on Directional Stability - Cruise Configuration 22 4 Effect of Orbiter on Lateral Stability - Cruise Configuration 23 5 Effect of Orbiter on Side Force - Cruise Configuration 24 6 Effect of Orbiter Position on Longitudinal Stability -Cruise Configuration, C-.-O. 25 7 Effect of Orbiter Position on Longitudinal Stability - Cruise a Configuration, C. ~ , C -C., 26 8 Effect of Orbiter Position on Directional Stability - Cruise Configuration 27 9 Effect of Orbiter Position on Lateral Stability - Cruise Configuration 28 10 Effect of Orbiter Position on Sideforce - Cruise Configuration 29 11 Effect of Orbiter Incidence on Longitudinal Stability - Cruise Configuration, C.,-a 30 12 Effect of Orbiter Incidence on Longitudinal Stability - Cruise Configuration, C. -a, C. -C.. 31 13 Effect of Orbiter Incidence on Directional Stability - Cruise Configuration 32 14 Effect of Orbiter Incidence on Lateral Stability - Cruise Configuration 33 15 Effect of Orbiter Incidence on Sideforce - Cruise Configuration 34 16 Effect of Afterbody Fairing on Longitudinal Stability - Cruise Configuration 35 17 Effect of Afterbody Fairing on Directional Stability - Cruise Configuration 36 18 Effect of Orbiter on Longitudinal Stability - Landing Configuration, C.,-a 37 loclfieed Geoif.id Cuniyan-, IV LIST OF FIGURES (Continued) Figure Title Page 19 Effect of Orbiter on Longitudinal Stability - Landing Configuration, C. -a, C. -C.. 38 20 Effect of Orbiter on Directional Stability - Landing Configuration 39 21 Effect of Orbiter on Lateral Stability - Landing Configuration 40 22 Vertical Tail Development - Cruise Configuration, C^-P 41 23 Vertical Tail Development -Cruise Configuration, C.,-P 42 24 Vertical Tail Development - Landing Configuration, C^-P 43 25 Vertical Tail Development - Landing Configuration, Cy-P 44 26 Vertical Tail Development - Landing Configuration, Orbiter Off, C -P 45 27 Vertical Tail Development - Landing Configuration, Orbiter Off, C -P 46 28 Effects of Orbiter on Longitudinal Trim 47 29 Rudder Effectiveness, 6 = 10° 48 K 30 Effect of Orbiter on Drag - Cruise Configuration 49 31 Effect of Orbiter Position on Drag -Cruise Configuration 50 32 Effect of Afterbody Fairing Shape on Drag - Cruise Configuration 51 33 Effect of Orbiter Incidence on Drag - Cruise Configuration 52 34 Comparison of Wind Tunnel and Full Scale Directional Stability - Cruise Configuration, Orbiter Off 53 35 Comparison of Wind Tunnel and Full Scale Directional Stability - Landing Configuration, Orbiter Off 54 36 Predicted Full Scale Directional Stability - Cruise Configuration 55 37 Predicted Full Scale Directional Stability - Landing Configuration 56 38 C-5A/Orbiter Piggyback Flight Vehicle Lateral-Directional Data 57 39 Lateral-Directional Response Mode Data 58 40 C-5A Stability Augmentation and Autopilot Systems Approximations 59 41 Flight Vehicle Lateral Response Comparison for a 30 KTAS Side- Gust Disturbance (SAS on) 60 42 Flight Vehicle Response Comparison for a Control Wheel Input of 10.0 Degrees (SAS on) 61 tocKtieed GeorRta Company LIST OF FIGURES (Continued) Page Figure Title 43 Flight Vehicle Lateral Response Comparison for a 30 KTAS Side- Gust Disturbance (SAS and Autopilot on) 44 Comparison of Estimated and Wind Tunnel Test Drag Polars - Cruise Configuration 63 45 Comparison of L/D for the C-5 and C-5/Orbiter Piggyback 46 Comparison of Estimated and Wind Tunnel Test Drag Polars - Landing Configuration 65 47 Takeoff Distances 48 Landing Distances 49 One Engine Inoperative Climb Gradient 50 Cruise Ceilings 51 Altitude Speed Capability 70 52 Ferry Performance Summary 71 53 Flight Restrictions 72 54 Flight Restrictions Compared with Super Guppy 73 VI SUMMARY Wind tunnel testing and analytical studies of the feasibility of ferrying the NASA Shuttle Orbiter on the C-5A in a piggyback mode have been accomplished by the Lockheed-Georgia Company in response to NASA contract NAS9-13702. The study was managed by J. H. Paterson of the Flight Sciences Division. Testing was conducted in the Lockheed-California Company 8 x 1 2 foot low speed wind tunnel using an exist- ing Air Force 0.0399 scale C-5A model in conjunction with a NASA 0.0405 scale Orbiter model. Six component force and moment data were measured over a range of pitch and yaw angles to determine lift and drag characteristics, lateral/directional stability characteristics and longitudinal and directional control powers.

Appendix A contains a description of the wind tunnel test program with a run schedule and the complete plotted data for all the test runs. Initial emphasis was given to determining the effects of the Orbiter above the C-5A and the optimum location for minimum interference on C-5A characteristics. A comprehensive series of cruise configurations were tested including a range of Orbiter longitudinal and vertical locations, incidences, and afterbody fairings. Subsequently, a series of configura- tions were devised during the test program to determine means of recovering directional stability degradation due to Orbiter interference.

Extensions to the present C-5 vertical stabilizer were designed as were twin fins to be located at the tips of the horizontal stabilizer. Analytical studies subsequent to the test and based on test results indicate that these exterior changes should not be neces- sary as automatic flight controls provide satisfactory flying qualities.

Performance studies of the C-5A/Orbiter Piggyback show that the drag penalty of the Orbiter on the C-5A does not preclude non-stop, unrefueled ferry missions up to 2500 nautical miles. Some flight restrictions for the Piggyback are unavoidable; how- ever these are not considered unreasonable for the special nature of the mission. In short, ferrying the Shuttle Orbiter in a Piggyback mode on top of a C-5A appears feasible with minimum modifications to the basic C-5A.

need Geargu Compjn, V I I 1.0 INTRODUCTION Recent interest by NASA and Rockwell International in alternatives to the present Orbiter Airbreathing Propulsion System for ferry and flight test of the Space Shuttle Orbiter has led to a series of proposals, analytical studies and wind tunnel tests to determine the feasibility of alternate systems. The Lockheed-Georgia Company has actively participated in these studies because of the suitability of Lockheed's C-5A as a carrier system for the Orbiter and in an attempt to apply Lockheed's "big air- plane" talents and knowledge to this program.

In response to NASA RFP 9-BC451-M6-4-4P, regarding the feasibility of ferrying the Orbiter piggyback on top of a C-5A, Lockheed-Georgia submitted a proposal and subsequently was awarded NASA contract NAS9-13702 for a low speed wind tunnel test and analytical study of a C-5A/Orbiter Piggyback configuration . This report consti- tutes the final report for this contract work. Analysis of the wind tunnel test results and the feasibility of the C-5A Piggyback concept are contained in the main part of the report. Appendix A contains the final plotted results from the wind tunnel test.

locknceti Georgia Compart) 2.0 ANALYSIS OF EXPERIMENTAL RESULTS 2.1 STABILITY AND CONTROL 2.1.1 Effects of Orbiter, Cruise Configuration The small effect of the Orbiter on the C-5 longitudinal stability is demonstrated in Figure 1 . These data are for the forward, low position of the Orbiter where maximum interaction of the two wings should occur. A negative shift in CLA of 0.04 occurs at all angles of attack. Minor modifications of the medium angle of attack pitching moment, in the destabilizing sense, is apparent for the Orbiter configuration without a fairing due to wake impingement on the horizontal tail. At high angles of attack, beyond stall, the typical C-5 initial pitch up followed by a strong nose down pitch is modified by both Orbiter configurations in such a manner that the net result should be almost imperceptible to the pilot.

An increase in lift curve slope due to the presence of the Orbiter, as well as a small increase in C|_.. ... is demonstrated in Figure 2. A small further improvement in c ue ne a ^"l-MAy ' *° * ^ fairing is shown. The C^-C^ curves demonstrate the negative Cii shift and negligible change in neutral point due to the Orbiter. The C^pj shift is the equivalent of less than one degree of stabilizer angle.

The effect of the higher vertical center of gravity due to the Orbiter, approximately 60 inches, will result in a slight decrease in speed stability that will be most apparent in the landing approach mode. It is anticipated that this effect will require little more than pilot familiarization with tho new pitch response to engine power since the current aircraft already has a vertical e.g. range of 51 inches.

The major effect of the Orbiter on the C-5 aerodynamic data is the reduction of weathercock stability as reflected by Cf^j . Figure 3 demonstrates this effect for £ the most critical configuration forward and low with a negative shuttle incidence.

This loss of directional stability is primarily a result of the Orbiter's influence on the air flow at the C-5 vertical tail. There is also a secondary destabilizing effect with lockneea Georgia Company this Obiter location due to Orbiter side area that is ahead of the C-5 center of gravity.

The prime effect, however, occurs because of the flow bending, caused by the Orbiter body. As a result, the C-5 vertical tail does not experience the full yaw angle seen by the forward fuselage. This reduction in yaw angle, as seen by the fin is approxi- mately 30% of the nominal value. As shown in Figure 3, the afterbody fairing re- sulted in an improvement in stability at high sideslip angles but delayed the turnover point.

Lateral stability, represented by dihedral effect, is little affected by the Orbiter as shown in Figure 4. A small reduction in Cn occurs through 15 of sideslip accom- panied by a linearization of the higher sideslip angle data due to the Orbiter wing configuration effect on the C-5 wing. The aft fairing causes further increases in C* at high sideslip angles due to the fin effectiveness.

f /3 Figure 5 shows that a large increase in C ^ occurs due to the presence of the Orbiter as a result of the side area increase, as would be expected. The aft fairing causes a small increase in sideforce at sideslip angles greater than 15 and no effect at lesser angles. It is somewhat surprising that more sideforce does not result from the added side area of the fairing. Apparently this area is not effective in sideforce due to the very thick boundary layer or there is a compensating flow change at the fin, or both.

2. 1.2 Effect of Orbiter Position, Cruise Configuration The effect on longitudinal stability of Orbiter fore and aft and vertical position rela- tive to the C-5 is demonstrated in Figure 6. This comparison is made with the Gelac fairing No. 1 on the Orbiter and with the Orbiter at an incidence of 0.5 degrees.

The destabilizing effect of the Orbiter in the forward high position is due to the com- bined effect of the Orbiter lifting moment and the interference with the flow at the C-5 horizontal stabilizer. The aft low position represents a significant improvement; showing a small negative A C . , shift that remains constant until the stall is reached.

The pitch down tendency beyond stall of the basic C-5 has been reduced slightly. A small reduction in stability occurs in the aft high position with more pitch up at the loaneffl Georgia Coitpinj stall than that for the low position due to the increased stabilizer interference: how- ever, the stability change relative to the basic C-5, below stall is negligible.

Figure 7 shows that Orbiter position has little effect on the lift curve slope and only a small effect on C|_.. . : the highest C|_.. . y occurs with the Orbiter in the aft high y position. From a longitudinal stability point of view it is apparent that the aft low position would be the best with the aft high position a second choice.

The effect of Orbiter position on directional stability is very pronounced as shown in Figure 8. The major change that occurs with aft movement is due to the tail-off stability increase as the body side area is moved aft of the reference e.g. A small stabilizing change in fin effectiveness occurs with aft movement of the Orbiter.

Again these data demonstrate the ability of the Orbiter body to reduce the local flow angle at the fin relative to the free stream angle through - 15 degrees. The sensitivity of the C-5 weathercock stability to the presence of the Orbiter is largely due to the equal magnitudes of tail-off instability and tail-on stability. Thus, a 50 percent loss of fin effectiveness will cause a 100 percent loss of stability. The aft, high position of the Orbiter has the best directional stability characteristic but is still slightly un- stable through small sideslip angles.

A significant change in dihedral effect occurs as a function of Orbiter position as shown in Figure 9. The major effect is due to Orbiter height above the C-5, showing larger C- for increased height. Fore and aft position does not appear to have much influence, showing a small increase in Cn for aft movement of the Orbiter. It would 'P appear that the major effect on Cp« is probably due to the freeing effect of moving the wings apart thus allowing full development of the normal lift change due to side- slip on both wings.

Orbiter position has a negligible effect on the net sideforce due to sideslip as shown in Figure 10. A large increase in Cy o is, of course, present due to the side area of the Orbiter configuration.

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2.1.3 Effect of Orbiter Incidence, Cruise Configuration The effect of Orbiter incidence on the C-5 pitching moment is shown in Figure 1 1 for the aft, high position of the Orbiter with the aft fairing (Test Fairing No. 3). In- creasing the Orbiter incidence relative to the C-5 reduced the pitching moment shift through nominal angles of attack, and at high angles of attack, increased the pitch up with a slightly less stable pitch out.

Small shifts in a with no change in lift curve slope due to Orbiter incidence are shown in Figure 12. Increasing Orbiter incidence results in small increases in net e c a a r e e c ne ^LMAy ^ ^M ~ ^L ' * f' * * expected shift in C^_. with essentially no change in neutral point.

Insignificant changes in directional stability resulted from Orbiter incidence variation as shown in Figure 13. The basic configuration for the Orbiter was the aft, high position with the No. 3 fairing.

Only minor changes in Ci^. occur due to Orbiter incidence as shown in Figure 14.

A small increase in C^ at the higher sideslip angles, as Orbiter incidence increases, is apparent.

No change in C due to Orbiter incidence is apparent, as shown in Figure 15.

Vo 2.1.4 Effect of After-body Fairing, Cruise Configuration The effect of various after-body fairing changes on pitching moment is shown in Figure 16. These fairing modifications were aimed at improving directional stability characteristics and have little direct influence on longitudinal stability other than through the drag changes.

is unstable through small angles for C-5 Orbiter combinations with both the Gelac No. 1 fairing and the Rockwell fairing. Attempts to reshape the aft fairing to improve the flow field at the vertical tail are shown in Figure 17. Small improvements Locineec Gwigu Comwnj were obtained with fairing No. 2 and 3 but are not sufficient in themselves to cure the problem.

The effects of after-body shape on Q, and C are negligible, demonstrating the R lack of load producing area in the aft body region.

2.1.5 Effect of Orbiter, Landing Configuration The effect of the Orbiter on the longitudinal characteristics of the C-5A in the landing mode is similar to that of the clean configuration. A larger negative pitching moment shift due to the Orbiter is apparent - Figure 18. (The Orbiter is in high aft position.)

A slight increase in stability is also noted.

Little or no change in lift curve slope occurs, as shown in Figure 19. A small neutral point shift in the stable sense is predictable from the C^ - Ci curves of this figure.

These data were obtained without the uprigged spoilers normally used for the C-5 landing configuration, however, little or no influence is expected.

In the landing configuration, the airflow at the C-5 vertical tail is not as restricted as in the clean configuration due to the large downflow, away from the fin, caused by the flaps. As shown in Figure 20, the net result is a more stable CKJ level than for the clean airplane even though a small "flat spot" still occurs at small sideslip angles. The shape of the basic C-5 curve is predicated by fin stall at large sideslip angles. Since the air flow at the fin is restricted by the Orbiter in the Piggyback mode, the fin never experiences stall in the tested sideslip range, hence the more linear yawing moment at large angles.

The major effect of the Orbiter on C, is to delay the fin stall at high sideslip angles so that an increase in rolling moment occurs. This effect is shown in Figure 21.

Little or no effect on C^ in the small sideslip angle range is noted.

J Compjr,, The effect of the Orbiter on C parallels that obtained in the clean configuration Y [3 and the levels at each sideslip angle are almost identical.

2.1.6 Vertical Tail Development The effects of the addition of a central fin to the C-5 horizontal tail bullet and the addition of tip fins to the horizontal tail are demonstrated in Figure 22. The Orbiter position is aft and high with the ^3 aft fairing. This position results in a negligible change in tail-off CKI except at the higher sideslip angles, since the Orbiter fin- P body area is well aft of the e.g.

The addition of a center fin above the C-5 tail produces sufficient stability beyond 5 of sideslip but is influenced by the Orbiter body effect at smaller angles. The addition of twin fins at the horizontal stabilizer tips successfully achieves the same stability level as the basic C-5 throughout the small angle range, and a much increas- ed level at the higher sideslip angles.

The additional sideforce developed in sideslip by tip and center fins required for directional stability is shown in Figure 23. These large values are not desirable because of gust response and turn coordination, especially in light of the already large increase in sideforce due to the Orbiter.

As shown earlier, the directional stability in the landing configuration in the presence of the Orbiter, is better than that for the clean configuration. As a result the fins, as sized, represent an excess capability as shown in Figure 24.

The sideforce due to sideslip is approximately the same as for the clean configuration shown in Figure 25.

The effectiveness of the center and tip fins, without the Orbiter, are shown in Figure 26. The center fin retains its effectiveness at high sideslip angles to a higher degree than the tip fins. They are equally effective at small angles.

locfcnted Georgu Compan* Similar data for the sidefbrce characteristics are shown in Figure 27.

2.1.7 Effect of Orbiter on Longitudinal Trim The Orbiter, in the aft high position, has a negligible influence on the dynamic pressure of the airflow at the C-5 horizontal stabilizer and only a minor influence on the downwash. The net effect is shown in Figure 28 for the cruise configuration. It may be noted that the "Orbiter on" data of this figure also have the vertical center fin whereas the "Orbiter-off" data do not. Although not shown here, the data in the appendix demonstrates that there is no effect in pitch due to the center fin, thus the comparison is valid.

Data are not available for the landing configuration. The influence of the Orbiter on trim effectiveness is anticipated to be even less than that for the cruise configuration because of the downward depression of the wing wake, away from the tail, caused by the flaps.

A small loss in dynamic pressure at the fin occurs due to the presence of the Orbiter and a reduction of local yaw angle relative to the free stream yaw in steady sideslip, as previously demonstrated. The net effect on rudder power for trim is shown in Figure 29. The Orbiter on data also include the effects of a center fin extension as discussed in 2.1.6.

The incremental effectiveness of the rudder in yaw is not anticipated to be affected by flap deflection.

2.2 DRAG CHARACTERISTICS 2.2.1 Effect of Orbiter Figure 30 illustrates the magnitude of the effect of the Orbiter on C-5 drag. At a cruise C. of 0.5, the drag of the Piggyback configuration is 70% greater than the basic C-5 level. By enclosing the bluff aft end of the Orbiter, the drag level of the locineed Georgia C Piggyback is reduced to a level about 40% above that of the C-5 at C = 0.5. Undoubt- edly, the skin friction drag of this very long fairing offsets some of the potential reduc- tion in Orbiter base drag.

2.2.2 Effect of Orbiter Position The effect on Piggyback cruise configuration drag of Orbiter location is shown by Figure 31. In general, the drag is seen to be insensitive to position for the locations tested, except for the aft - low position, which carries slightly lower drag up to a C|_ of 0.6. The drag of the aft high position is about the same as that of the forward positions at all C^'s. These results indicate that interference drag is a very small con- tributor to total drag.

2.2.3 Effect of Afterbody Fairing Figure 32 compares drag for the various afterbodies tested. Not a great deal of signif- icance can be attached to these results. As expected, the increased afterbody fineness ratio of Gelac fairing *1 improved the flow relative to a blunter Rockwell fairing, how- ever, the increased skin friction drag due to additional wetted area almost negates this as the decrease in cruise drag is only about 2 percent.

2.2.4 Effect of Orbiter Incidence Figure compares drag results for the two Orbiter incidence angles and substantiates no change in Piggyback drag due to incidence over the range from -1.5 to 0.5 .

i Company *^== 3.0 ASSESSMENT OF FULL SCALE FLIGHT FEASIBILITY 3.1 STABILITY AND CONTROL 3.1.1 Comparison of Wind Tunnel and Full Scale Directional Stability The wind tunnel data, obtained from this test, are compared with the published, full- scale, levels for the C-5A to establish the base for the incremental data obtained in the presence of the Orbiter. The full-scale data are based upon the correlation of flight test data, obtained during the C-5A development program and the design wind tunnel data.

The cruise data for yaw due to sideslip are shown in Figure 34. The major change from the wind tunnel data is an extension of the fin sideforce capability to a higher yaw angle and a slightly more effective fin. There is also a more linear continuation of the tail-off yawing moment through high sideslip angles.

The landing flap data, shown in Figure 35, demonstrate further differences from the wind tunnel data. These differences are largely due to a change in the aft body interference with flaps down, that resulted in a less stable airplane than predicted by the wind tunnel. As may be noted, the net fin effectiveness, full scale, is con- siderably less than the wind tunnel level. These data are for landing flaps with the gear up. When the gear is down a higher CM is realized due to the effect of the P gear on the afterbody interference.

3.1.2 Predicted Full-Scale Directional Stability Using incremental tail-off and tail-on data for the effect of the Orbiter, the full- scale predicted levels of weathercock stability are shown in Figure 36 compared with the basic C-5 in the cruise configuration. As may be noted, the Orbiter/C-5 com- bination is neutrally stable through - 15 degrees of sideslip.

The full-scale prediction for the landing configuration, gear up, is shown in Figure 37.

if M Geoffia Com can-, These data reflect the low full-scale fin effectiveness discussed in the previous para- graph. The configuration is predicted to be neutrally stable through -2 of sideslip and lightly stable at higher angles. Although not shown here, the gear-down landing configuration will be more stable.

3.2 FLYING QUALITIES The wind tunnel test results have shown that the present C-5A longitudinal aerodynamic characteristics would not be critically affected by the piggyback shuttle installation.

Evidently such would not be the case for the lateral-directional characteristics, parti- cularly in the cruise configuration. The C-5A with the Orbiter in position exhibits an increase in sideforce due to sideslip, C =-1.39/radian compared to -0.80/ yp radian for the C-5A. The directional stability level is reduced to nil, C_ =0 n g composed with 0.0728/radian for the basic airplane. These predicted characteristics pertain to the M=0.52 at 20,000 feet flight condition. A cursory analysis was com- pleted to assess the impact of these aerodynamic changes on C-5A flying qualities.

Pertinent flight vehicle data are tabulated in Figure 38. The reference gross weight is 704,626 pounds, which represents a 550,000 -pound airplane (no pay load) with either a 154,626-pound cargo or the present design piggyback installation of the Orbiter vehicle. The Orbiter center of mass is considered to be 11.55 feet behind and 28.02 feet above the C-5A mass center.

Modal response data are presented in Figure 39. The aerodynamic changes due to the Orbiter installation result in a re-distribution of the total airplane damping due to C , Q, and C . The spiral mode is more stable, now characterized by a 15.2 y p n second time constant. The dutch roll mode is now unstable and the period of these oscillations is doubled, £<j = -.023 and Tj = 18.3 seconds. The C-5A airplane incorporates a full-time stability augmentation system (SAS) on roll and yaw axes, and thus this unstable condition would not be experienced in flight. The Orbiter ferry mission may be completed with the autopilot also operative in cruise.

Flight vehicle response data were obtained using a digital computer program to evaluate responses to a 30KTAS lateral gust disturbance and a 10.0-degree lateral control wheel input. The program considers the solution of the three lateral equations of motion with respect to the usual linear assumptions. SAS and pertinent autopilot functions were included on a simple gain basis. The various high-order filters and the 0.25-second servo time constants were neglected such that the problem reduced to the control loop closures indicated in Figure 40. It is noted that the autopilot control command loops are excluded. The lateral stability functions have been included to enable an evaluation of flight vehicle response to external gust disturbances. Bank angles are presumed to be less than 7.0 degrees and thus the heading stability elements of the autopilot may also be excluded.

Figure 41 presents sideslip and bank angle responses to a continuous step gust of 30 KTAS. The lack of directional stability with the piggyback Orbiter installation results in a reluctance of the flight vehicle to naturally crab into the wind. Figure 42 provides a comparison of flight vehicle responses to control wheel throw. The excellent turn coordination characteristics of the basic C-5A airplane are somewhat degraded by the Orbiter installation. It is evident from the foregoing material that the aerodynamic changes associated with Orbiter installation may require a re-tuning of the basic C-5A stability augmentation system gains for cruise flight. The autopilot will probably be activated for the cruise condition of the Orbiter ferry mission. These would be an associated tightening of the lateral stability loop for the autopilot opera- tive mode. The data presented in Figure 43 indicate that the flight vehicle responses to lateral gust disturbances would be stabilized, although still greater than for the basic C-5A airplane.

As stated earlier, the present analysis was of a cursory nature. The guarded conclusion is that the ferry cruise of the piggyback C-5A/Orbiter flight vehicle may not require significant C-5A flight control modifications. A continuation of studies to a greater depth than those described herein is recommended. The effects of flight vehicle vertical center of gravity location should receive attention. It is acknowledged that (*"e*G lie;:;.3 Com pan-.

an upward shift of the flight vehicle mass center will result in a reduction of the effective dihedral. The impact of c.g position on longitudinal characteristics should also be evaluated. Low speed, flaps-down, flight should also receive analytical attention.

3.3 PERFORMANCE 3.3.1 Full-Scale Drag Characteristics Figure 44 compares estimated and wind tunnel drag for the C-5/Orbiter Piggyback.

Test results on the isolated C-5 have been summed with wind tunnel data for an iso- lated Orbiter at the same test Reynolds number. The addition was accomplished at constant angle of attack. The excellent agreement between these two drag polars implies an absence of any net interference drag in the cruise configuration. Therefore, for purposes of this analysis, full-scale drag at flight Reynolds number for the Piggy- back configuration has been defined by summing the estimated full-scale drag of an isolated Orbiter with C-5 flight test correlated drag. Resulting lift-to-drag ratios for the Piggyback at a typical cruise Mach number of 0.6 are shown compared with the C-5 in Figure 45.

Figure 46 shows a drag comparison, similar to Figure 44, for the landing configuration.

The net interference drag in this case is seen to be equal to about 75% of the isolated Orbiter drag. Therefore, the low-speed, flaps-down drag data used for airport per- formance analyses reported herein have been increased to account for this effect.

3.3.2 Airfield Performance Figures 47 and 48 show the takeoff and landing distances for the C-5/Orbiter Piggyback at varying gross weights. These data represent standard C-5 takeoff and land distances increased slightly to account for drag due to the Orbiter. Runway conditions for an airfield pressure altitude of 2000 feet and standard-day temperatures have been used for these as well as all other airfield performance data presented.

Loc'fees GtOfgu Compjr.j Takeoff flap setting for the C-5A is 16 degrees with a takeoff speed of 1.2 V-.... .

For a long-range ferry mission takeoff gross weight of 700,000 pounds, takeoff ground roll is seen to be 7230 feet with a total distance of 8640 feet to clear a 50-foot obstacle. Engine-out climb capability of the C-5 Piggyback configuration may res- trict operations at these conditions such that increased takeoff speeds and distances may be required. However, operation fromairfields with runway lengths of 10,000 feet should not be prohibited.

Landing flaps for the C-5A are set at 40 degrees and approach speeds are normally 1.3 V,.,... . For an aborted mission after takeoff at 700,000 pounds, a landing ground roll of 3250 feet is indicated by Figure 48. Normal ferry mission landing weights would be approximately 550,000 pounds, for which a landing ground roll of 2200 feet and total landing distance from a 50-foot obstacle of 3580 feet would be expected.

3.3.3 Climb and Cruise Performance One-engine-inoperative climb gradients for the C-5 Piggyback at several takeoff speeds and with the landing gear retracted are shown in Figure 49 for standard-day, 2000- foot pressure altitude conditions. Since Piggyback climb gradients are reduced relative to those of the basic C-5A, consideration has been given to increasing the takeoff speeds to improve climbout performance. As can be seen, an increase from 1 .2 V_ . . to 1 .3 V^, .. .

T increases the gradient by about 0.35 percent, or for a constant climb gradient, the takeoff weight is increased by about 23,000 pounds. This amounts to approximately a 10 percent increase in fuel for long-range ferry missions.

Cruise ceilings for the C-5 Piggyback are shown in Figure 50 for several rates of climb. Long-range cruise performance calculated for the ferry mission is based on the altitudes for the 300-feet-per-minute ceiling shown for normal rated thrust (NRT).

The cruise ceilings with military rated thrust (MR T) are useful for determining maximum speed-altitude capability of the C-5 Piggyback.

tocinted Gto'gn Compjn> Figure 51 summarizes the speed-altitude capability at MRT of the Piggyback for weights corresponding to both an empty and fully loaded Orbiter. Also shown are data for the case of an Orbiter configuration without an afterbody fairing. At 25,000 feet, the maximum speed attainable is 259 KEAS with a faired afterbody, fully loaded Orbiter and 266.5 KEAS with an empty Orbiter.

3.3.4 Orbiter Ferry Capability Figure 52 summarizes the capability of the C-5A to ferry the Orbiter in the Piggyback mode as a function of military critical field length and takeoff ground roll. These data are shown for takeoff speeds of 1.2, 1.25 and 1.3 times the stall speed and for three values of one-engine inoperative climb gradient. A climb gradient of 2.3% is the current minimum allowable gradient for the C-5A. Reducing the climb gradient to 1.8% improves the range by 240 miles while increasing takeoff distance by less than 1000 feet. Similarly, increasing takeoff speed from 1.2 to 1.3 VC A i i increases T range by 160 miles but increases takeoff distance by 2500 feet.

For a special-purpose airplane it appears quite reasonable to accept lower climb gradients as a means of increasing range, provided there are no obstacles in the take- off path. Alternately, it is not necessary to resort to lower climb gradients, since the C-5A's inflight refueling capabilities make its range essentially unlimited.

3.4 FLIGHT RESTRICTIONS Flight restrictions for the Piggyback are summarized in Figure 53 for two configurations, the C-5 with and without tail modifications. As discussed previously in subsection 3.2, ferry flight without any extension modifications to the C-5 tail can be accom- plished with reliance on automatic flight controls, and flight restrictions listed here are given only as a matter of interest.

These restrictions have been established such that no structural modification to the C-5A is necessary other than that required to mount the Orbiter. The "fuselage fuel" Lockneed Georgu Compjnj included in the weights breakdown represents an amount of ballast required for the Orbiter mounted in the aft position. This position in 10 feet aft relative to the base- line location, and the ballast is required to bring the e.g. within the current aft limit of the C-5A. The operating weights shown include the weight of the fuselage fuel tank.

Flight restrictions for the C-5/Orbiter Piggyback are compared with those of the Super Guppy in Figure 54. As can be seen, they are quite comparable. The only condition in which the C-5A is restricted more than the Super Guppy is in touchdown rate of sink. This is insignificant, since the design weights can be lowered somewhat and still allow the ferry-range performance shown in subsection 3.3.4. Design speeds and gust weights are naturally considerably greater for the C-5A as represented by the 300 KCAS level-flight maximum speed for the C-5A/Orbiter Piggyback, compared with 219 KCAS for the Super Guppy, and a maximum gross weight of 865,000 pounds for the Piggyback compared with 162,000 pounds for the Super Guppy. Maneuver- load factors for cruise are about the same: 2.0 for the Piggyback and 2.2 for the Super Guppy.

lockrtced Gecfgu C 4.0 CONCLUSIONS AND RECOMMENDATIONS Wind tunnel testing of the C-5A/Orbiter Piggyback configuration has demonstrated that the major effect of the Orbiter on the aerodynamics of the C-5A is a loss of direc- tional stability due primarily to airflow losses at the vertical tail, and to an increase in overall side area and side forces. The effects of the Orbiter on longitudinal stability are almost negligible as evidenced by a C shift due to the Orbiter equivalent to m o less than one degree of horizontal stabilizer incidence. The effect on drag, as ex- pected, is significant, but the drag level of the Piggyback configuration can be reduced to a level about 40% above C-5A cruise configuration drag with an Orbiter afterbody fairing. Interference effects from a drag standpoint appear from the test results to be insignificant for the ferry cruise configuration.

Variations in Orbiter longitudinal and vertical locations showed that the aft high position was the best, primarily because the losses in directional stability were minimized by moving the side areas aft of the reference e.g. The effects of varying the Orbiter incidence relative to the C-5A were, from any viewpoint, inconsequential for the range tested (-1.5 and 0.5 ). A Lockheed-Georgia afterbody designed for the Orbiter to improve the flow at the empennage and the directional stability proved insufficient, although a slight drag reduction was noticed for the Lockheed-Georgia fairing.

During the wind tunnel test, several empennage modifications were designed and tested to remedy the directional stability problems. These modifications included a control fin addition above the present horizontal stabilizer, and twin fin additions to the horizontal stabilizer tip.

These were successful in restoring the stability level of the Piggyback to that of the basic C-5A so that, if desired, external modifications could be defined that would provide satisfactory flying qualities. Cursory analytical studies indicate that the C-5A automatic controls can be modified to fly the Piggyback configuration in a ferry operation without external modifications and with only minor modifications to the flight control systems.

Performance analyses revealed the feasibility of trans-continental unrefueled distances for the C-5A ferrying the Shuttle Orbiter. Airfield performance assures operation from fields of less than 10,000 feet where minimum takeoff climbout gradients can be toler- ated. In total, the feasibility of the C-5A/Orbiter Piggyback ferry concept appears excellent and the following recommendations are respectfully submitted: o Development of the C-5A ferry vehicle should be initiated as soon as possible, o A wind tunnel test program of the airlaunch configuration should be initiated.

o Studies of airlaunch concepts and separation trajectory analyses should be made in conjunction with the wind tunnel program.

o More detailed, flying-qualities studies of the C-5/Orbiter Piggyback configuration should be conducted to identify potential modifications of the C-5A automatic flight control systems.

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C-5A/ORBITER PIGGYBACK FLIGHT VEHICLE DATA V =319 KTAS M = 0.52@20,000ft /• Wing area, S =6200 ft Wing span, b = 219 ft #2 FLIGHT CASE #1 #3 Sross Weight ~lbs 550,000 704,626 704,626 2 -A 3.89 1 ~ slugs ft x }Q 3.43 3.45 XX 2 -6 1 ~ slugs ft x 10 6.07 6.11 5.48 ZZ 2 6 3.97 1 -slugs ft x 10 2.20 2.42 xz Stability and Control Derivatives -1.34 C ~ /rod. -.802 -.802 Sideslip Cfl ~ /rod. -.0771 -.0997 -.0997 C ~/rad. .0728 .0728 0 n C ~/rad. -.390 -.390 -.390 Roll P rate C ~/rad. -.081 -.081 -.081 n P C ~/rad . .510 .500 .500 X r Yaw C. ~/rad. .177 .199 .199 rate r C ~/rad . -.180 -.180 -.180 n r C, ~/rad. -.0319 -.0319 -.0319 Aileron 'd a angle C ~/rad.

0 0 0 n^ ' a C ~ /rod. .2006 .2006 .2006 y Rudder C ~ /rod. .0210 .0181 .0181 angle r C ~/rad . -.1031 -.1031 -.1031 \ r C ~/rad. -.0573 -.0573 -.0573 y .0268 .0268 .0268 C^ ~/rad.

Spoiler angle & s C ~/rad. .0057 .0057 .0057 FIGURE 38 LATERAL-DIRECTIONAL RESPONSE MODE DATA Characteristic Equation: 2 2 (s + 2 w s + w ) (s + I/ r ) (s + I/ ) = 0 fd d d R TS # Item or Parameter Case ^2 Case #3 Case 1 # G.W. (no payload) 550,000 550,000* 550,000* # - - cargo weight 154,626 - arbiter weight - 154,626* # gross weight 704,626 704,626* 550,000* Dutch Roll Mode frequency, w - rad/sec .344 .634 .620 damping ratio, f .118 .076 -.023 period, T , - sees. 9.98 10.2 18.3 time to 1/2-ampl, t , - sees, 14.6 -89 9.18 cycles 1/2-ampl, C. ,_ 1.44 -4.9 .92 Roll Convergence Mode time constant, - sees 1.42 1.45 1.35 T D K time to 1/2-ampl, t /- - sees .98 .93 ] 1.0 Spiral Mode time constant, r - sees 694 15.2 time to 1/2-ampl, t. ,„ - sees 479 149 10.5 NOTE: Negative values signify an unstable dutch roll mode.

FIGURE 39 C-5A STABILITY AUGMENTATION & AUTOPILOT SYSTEMS APPROXIMATIONS Stability Augmentation Elements Aileron: 8 =0.055(0) a Spoiler: 5 =0 Rudder: & = - . 4 8 2 ( p ) -. 101(0) + 1 . 0 ( r ) Incremental Elements for Autopilot Operative* Aileron: A8 = 2.25 (p) + 3.22 (0) Q Spoiler: A8 = -.786 (0) Rudder: A 8 =0 Control command and heading stability (0 < 7 ) elements are excluded.

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APPENDIX A WIND TUNNEL TEST DESCRIPTION AND PLOTTED DATA A-l

I L

CONTENTS

Section Title Page I MODEL DESCRIPTION A-2 II TEST FACILITY A-3 III TEST CONDITIONS A-4 IV DATA REDUCTION A-5 V REFERENCES A-6 VI MODEL CONFIGURATION SYMBOLS A-7 VII MODEL DIMENSIONAL DATA A-10 VIII RUN SCHEDULE A-31 loc'-need Geoigid Com pan,

A-la

LIST OF FIGURES Figure Title Poge A-l Stabilizer Effectiveness, Obiter Off A-35 A-2 Effect of Orbiter Position, Tail On and Tail Off A-39 A-3 Effect of Orbiter Afterbody Fairing Shape A-43 A-4 ' Effect of Orbiter - Afterbody Fairing On and Off A-47 A-5 Effect of Orbiter Incidence Angle and Position A-51 A-6 Effect of Orbiter Position and Afterbody Fairing Shape A-55 A-7 Spoiler Effectiveness, Orbiter On and Off A-59 A-8 Effect of Center Vertical Stabilizer Extensions A-63 A-9 Effect of Orbiter Incidence Angle - Orbiter in Aft High Position A-67 A-10 Stabilizer Effectiveness - Orbiter On A-71 A-ll Effect of Orbiter Incidence and Position A-75 A-l 2 Rudder Effectiveness - Orbiter Off A-79 A-l3 Effect of Orbiter Afterbody Fairing A-83 A-14 Effect of Orbiter - Tail On and Tail Off A-85 A-15 Vertical Tail Effectiveness with Center Vertical Extension A-90 A-l6 Effect of Vertical Tail Modifications - Orbiter Off - Flaps Extended A-94 A-l7 Vertical Tail Effectiveness with Center and Stabilizer Tip Vertical Extensions - Flaps Extended A-99 A-18 Effect of Orbiter Tail On and Off - Flaps Extended A-103 A-19 Rudder Effectiveness with Center Vertical Tail Extension A-108 A-20 Effect of Orbiter Position and Incidence A-l 12 A-21 Effect of Orbiter Incidence and Fairing Shape A-l 16 A-22 Vertical Tail Center Extension Effectiveness, Flaps Extended, Orbiter Off A-l 20 A-23 Vertical Tail Center Extension Effectiveness, Flaps Extended, Orbiter On A-l 24 A-24 Effect of Flaring Shape and Orbiter Location A-l 28 A-25 Effect of Orbiter, Tail On and Tail Off A-l32 A-lb

I t

LIST OF FIGURES (Continued) Figure Tit-le Page A-26 Effect of Orbiter Location A-136 A-27 Effect of Orbiter Afterbody Fairing On and Off A-140 A-28 Effect of Center Vertical Tail Extension, Flaps Extended, Orbiter Off A-142 A-29 Effect of Center Vertical Tail Extension, Flaps Retracted, Orbiter Off A-146 A-30 Effect of Center Vertical Tail Extension, Flaps Retracted, Orbiter Off A-151 A-31 Effect of Vertical Tail Extensions at Stabilizer Tips A-156 oifia Company A-lc

I I

I - MODEL DESCRIPTION The C-5A Piggyback model is a combination of.the Rockwell International 0.0405 Shuttle Orbiter model and the Lockheed-Georgia 0.0399-scale low speed C-5A model joined with suitable attach fittings.

The Orbiter model is fabricated from wood and metal and incorporates adjustable control surfaces. Provision was made for the installation of various afterbody fairings. Five afterbody fairing shapes were available for test. The basic Rockwell International fairing is denoted by a superscript 1 . The original Lockheed-Georgia fairing, denoted by a superscript 2, was designed to minimize the afterbody drag.

Fairings 3, 4, and 5 were fabricated by cutting away various portions of the fairing in an attempt to improve the flow at the C-5A tail.

The 0.0399-scale, low speed C-5A model is assembled from numerous components that allow the simulation of configurations encompassing the entire flight regime of the aircraft. The model is fabricated primarily from aluminum with some steel and plastic parts. All control surfaces are adjustable, and the landing gears and cargo doors may be positioned in increments from fully retracted to fully extended.

A symbol list of all the model components used in this test is presented in Section VI, A-2 II - TEST FACILITY The C-5A - Orbiter Piggyback combination tests were conducted in the Lockheed- California Company 8 X 12-Foot Low Speed Wind Tunnel. The tunnel is a con- ventional, low speed, single-return type with the test section vented to atmospheric pressure. Details of the facility are presented in Reference 1 .

The model is supported in the upright position by a three-support fork. The fork is connected to an external, six-component, pyramidal-type balance located below the floor of the test section. The balance transmits loads from the model and support to an electrical readout system. Raw data are converted to punched cards using an IBM 1442 card reader punch. The raw data cards are input to the IBM 1131 Processor computer, which converts these data into coefficient form for output as tabulated data and provides the input for the Calcomp 565 plotter which produced the finished data plots presented in this appendix.

a Gffltf'.ta Coalman; A-3 III - T E S T CONDITIONS All runs with flaps deflected were made at a dynamic pressure of 40 P.S.F. Flaps up runs were made at a dynamic pressure of 60 P.S.F. These dynamic pressures correspond to Mach Numbers of 0.165 and 0.201, and Reynolds Numbers of 1.436 X 10 and 1.758 X ICr, respectively. Reynolds Numbers are based on the C-5A model M.A.C.

Icc'-tieeo Groifia Compan, A-4

I I

IV - DATA REDUCTION Six-component force data were measured during all runs. The data were reduced to coefficient form and transferred to the stability axis system coincident at the ref- erence moment center (F.S. 53.762, W.L. 10.578, BL 0.000). Corrections applied to the six-component data include tunnel wall corrections, blockage, buoyancy drag, and support tare and interference corrections.

The support tare and interference were obtained in a previous test of a similar model (Reference 2). The correction values applied to the longitudinal components data were taken from faired plots of the tare and interference corrections, whereas the values applied to the lateral component data were taken directly from the computed results.

The six-component data reduction constants are listed below.

Wing Area, square feet 9.878 Wing Span, inches 104.997 Wing Mean Aerodynamic Chord, inches 14.817 Wing Mean Aerodynamic Chord Location F.S. 53.762 W.L. 12.577 B.L. 21.654 Front Trunnion Location F.S. 53.742 W.L 4.328 Moment Reference Center F.S. 53.762 W.L. 10.578 toc'»heea Georgia C A-5 V - REFERENCES 1. "Wind Tunnel Computing Handbook," Lockheed California Company Report LALI, 15 June 1955.

2. "C-141: Investigation of the Low Speed Characteristics of the Pro- duction Airplane Configuration Using a 0.044 Scale Model in the Lockheed-California Company 8 X 12-Foot Wind Tunnel," Tests L-45-1, II, and III; Report No. ER 5071, June 1963.

A-6

I t

VI - MODEL CONFIGURATION SYMBOLS

a Aileron, Simple hinge, sealed. Deflection 07-C5A-0 197-1 10 range - 25 , denoted by subscripts.

angles set with protractor.

A Orbiter, Shuttle - 0.0405 Scale Rockwell International Model.

with/without aft fairing; capability of being located at 4 position on C-5A model, 3 angles of attack (ref. Orbiter FRL; -1 1/2°, + 1/2°, +2 1/2°) Superscripts: Afterbody fairing shape and Orbiter location are denoted by number and letter superscripts, respectively. Lack of a number superscript indicates afterbody fairing removed.

1 Rockwell International fairing 2 Lockheed-Georgia fairing 3 Lockheed-Georgia fairing modified to lower surface upsweep 4 Lockheed-Georgia fairing modified to upper surface downsweep.

5 Lockheed-Georgia fairing modified to shorter fairing with horizontal knife edge Orbiter e.g. locations in terms of C-5 Fuselage Stations Longitudinal Position Vertical Position A (BASE) 54.424" Base B 56.819" Base C 59.235" Base D 54.424" Base +2. 395" E 56.819" Base +2. 395" F 59.235" Base +2. 395"

A -7

Subscripts: Orbiter incidence in degrees referenced to C-5A FRL is denoted by a subscript 1A (i.e. A, - - Orbiter with Rockwell International afterbody fairing located in the base position at 1 .5 incidence)

07-C5A-0181-200

B Fuselage

07-C5A-0182-403

Bullet

07-C5A-0181-402A

Dorsa I

07-C5A-0198-401

Elevator Inboard. Simple hinge, hinge line gap sealed. Deflection range -25 +15 ; denoted by subscripts. Set with protractor.

Elevator Outboard. Simple hinge, hinge line 07-C5A-0198-401

gap sealed. Deflections range - 25 15 , denoted by subscripts. Set with protractor.

,37

07-C5A-0198-105

Flaps, T.E. Fowler. Six sections/side, 0° and 40° (Idg.) to be tested.

07-C5A-0198-401-

H Horizontal Stabilizer. Incidence settings

-0195-400

capability. 0 , - 4 , - 6 , -8 , -12 ; set with pin in push rod in vertical.

.,24A 20A , ..

M

K N Pylon/Nacelles 07-C5A-0197-300

07-C5A-0197-109

Slats, Leading Edge. 14% C , 3 section/side.

Q Inboard 2 section sealed to wing and pylon, outboard section 1.25% G W T.E. gap and sealed to pylons.

Deflection 20 inboard sections, 20 outboard; denoted by subscript "20".

A-8

I t

7 8 r r Rudders, lower and upper, respectively. Simple 07-C5A-0192-402 hinge, hinge line gap sealed; deflections + 30°.

V Vertical Stabilizer 07-C5A-0182-402 1U W Wing 07-C5A-0197-100 f6 Z Flap Track Fairing 07-C5A-0197-106 g2?

Z Nose Landing Gear Fairing 07-C5A-0197-201 g23 Z Main Landing Gear Fairing 07-C5A-0151-204 w27 Z Wing - Fuselage Fillet- Alum, and Plastic; 07-C5A-0197-200 OA Composed of Z fwd. fillet and w 2 2 Z a f t . fillet.

1 22 11A w27 24A 20A f6 g27 923 S =B W A K N Z Z Z V Center Vertical Fin Extension - Alum, plate, cut to match L.E - T.E vertical stabilizer sweep and tip chord of vertical (V ), span 6", attached to top of horizontal bullet fairing.

h Horizontal Stab. Fins. - end plates on tips of horizontal stabilizer 1" inbd. from horizontal tips, 4" chord, 8" span (or height).

A-9

VII - MODEL DIMENSIONAL DATA

Aileron, (a ) 0.039916 Scale

Area per side, square 0.188

feet

Span, inches 10.651

Chord lengths, inches

Inboard 2.978

Outboard 2.337

Mean (RMS, streamwise)

Sweep of hinge line, degrees 20.4 17

Deflection limits, degrees j^25

22.

Fuselage, (B ;

Length, inches 110.487(9.207')

Maximum frontal area, 126.60

square inches

Equivalent maximum 12.69

diameter, inches

Fuselage reference line W.L. 7.983

Nose location F.S. 6.387

Wetted area, square 25.223

feet (imprints not removed)

Volume, Cu. Ft. 5.379

Bullet, (b )

Length, inches 21.44

Maximum frontal area, 3.22

square inches

Equivalent diameter, 2.03

inches

Wetted area, square 0.541

feet (Exposed Only) A-10

i i

MODEL DIMENSIONAL DATA (CONT.)

_ . 8 mod.

m x Dorsal, (D ) Wetted area, square 0.129 feet Imprint area, square 0.075 feet (On fuselage) Elevator, Inboard (e j Area per side, square 0.1434 feet Root chord, inches 3.227 Tip chord, inches 2.228 Mean chord length 2.773 (RMS), inches Span per side, inches 7.569 Hinge line, % horizontal 66.000 chord Deflections, degrees +30.000 Elevator, Outboard (e ) Area per side, square 0.0624 feet Root chord, inches 2.228 Tip chord, inches 1.609 Mean chord length 1 .943 (RMS), inches Span per side, inches 4.684 Hinge line, % horizontal 66.000 c hord Deflections, degrees +33.000 Trailing Edge Fowler Flaps, (f )* Panel 1 (Inboard) Area per side, square feet 0.214 *AII dimensions given in Wing Reference Plane.

A-ll

I £

MODEL DIMENSIONAL DATA (CONT.)

Trailing Edge Fowler Flaps, (Cont.)

Span, inches 7.085 Sweep of leading edge, 9.832 degrees Chord lengths, inches Root 4.410 Tip 4.410 Average 4.410 Mean (RMS) 4.410 Chord locations, inches Root W.S. 5.620

Tip W.S. 12.705

Average W.S. 9.163 Mean (RMS) W.S. 9,163 Maximum deflection, 40.000 degrees Panel 2 Area per side, square 0.173 feet Span, inches 5.718 Sweep of leading edge, 9.832 degrees Chord lengths, inches Root 4.410 Tip 4.410 Average 4.410 Mean (RMS) 4.410 Chord locations, inches Root W.S. 13.344 Tip W.S. 19.062 Average W . S . 16.203 A-12

MODEL DIMENSIONAL DATA (CQNT.)

Trailing Edge Fowler Flaps, (Cont.)

Mean (RMS W . S . 16.203 Maximum deflection, 40 .000 degrees Panel 3 Area per side, square 0.125 feet Span, inches 4.845 Sweep of leading edge, 12.364 degrees Chord lengths, inches Root 3.804 Tip 3.804 Average 3.804 Mean (RMS) 3.804 Chord locations, inches Root W.S. 19.701 Tip W.S. 24.546 Average W.S. 22.123 Mean (RMS) W.S. 22.123 Maximum deflection 40 .000 Panel 4 Area per side, square 0.094 feet Span, inches 4.497 Sweep of leading 17.033 edge, degrees Chord lengths, inches Root 3.133 Tip 3.133 Average 3.133 Mean (RMS) 3.133

A-13

MODEL DIMENSIONAL DATA (CONT.)

Trailing Edge Fowler Flaps (Cont.)

Chord locations, inches Root W.S. 25.184 Tip W . S . 29.681 Average W.S. 27.433 Mean (RMS) W.S. 27.433 Maximum deflection, 40 .000 degrees Panel 5 Area per side, square 0.079 feet Span, inches 3.802 Sweep of leading ._ 17.033 edge, degrees Chord lengths, inches Root 3.133 Tip 3.133 Average 3.133 Mean (RMS) 3.133 Chord locations, inches Root W.S. 30.321 Tip W.S. 34.135 Average W.S. 32.222 Mean (RMS) W.S. 32.222 Maximum deflection, 40.000 degrees Panel 6 Area per side, square 0.082 feet Span, inches 3.938 Sweep of leading edge, 17.033 degrees

A-14

I i

MODEL DIMENSIONAL DATA (CONT.)

Trailing Edge Fowler Flaps (Cont.)

Chord lengths, inches Root 3.133 Tip 3.133 Average 3.133 Mean (RMS) 3.133 Chord locations, inches

Root W.S. 34.773

Tip W . S . 38.712

Average W.S. 36.742

Mean (RMS) W.S. 36.742

Maximum deflection, 40.000 degrees Q Horizontal Stabilizer, (H ) Airfoil Section NACA 0010.5-0.833-0.40/1.432 (modified) Area - projected square 1.539 feet - wetted, square 2.910 feet (Exposed only) Span 32.397 Chord lengths -MAC, 7.322 inches Root, 9.985 inches Tip, inches 3.695 Aspect ratio 4.736 Taper ratio 0.370 Sweep of 25% chord 24.583 line, degrees 25% MAC Location F.S. 115.533 locitited Georgia Company

A-15

I i

MODEL DIMENSIONAL DATA (CONT.)

Horizontal Stabilizer, (Cont.)

B.L. 6.858 Volume coefficient 0.629 Tail length, inches 60.028 24A Pylon, (K ) Sweep of I.E., degrees 71 .504 Chord length, inches 13.073 Taper ratio 0.876 Airfoil section NACA

0008-1.100-0.335/1.575

(modified) Wing intersection, 1.4 % wing chord Toe -in, degrees 1 .0 Wing intersection, B.L. 19.122 inboard Wing intersection, B.L. 29.781 outboard 20A Nacelle, (N ) Length, inches 9.228 Maximum diameter, 4.091 inches Duct diameter, inches 3.409 Fineness ratio 2.256 Area, square feet Maximum frontal 0.091 area In let area 0.075 Side area 0.249 Toe-in angle, degrees 1 .0 Incidence, degrees 2.0 A-16

I t

MODEL DIMENSIONAL DATA (CONT.)

Nacelle (Cont.)

Inlet location Inboard nacelle F.S. 41.970 W.L. 8.864 B.L. 18.999 Outboard nacelle F.S. 47.490 W.L. 7.891

B.L. 29.658

Leading Edge Slat, (Q ) Section I (outboard) l-l/4%C gap w Area, square feet 0.191 Span, inches 19.556 Chord length - root, 1.698 inches tip, 1.135 inches average, 1 .417 inches Chord location - root B.L. 29.291

tip B.L. 48.788

Angle from stowed position, degrees 22.0 Section II (mid section), sea Ied Area, square feet 0.079 Span, inches 6.336 Chord length - root, 1 .881 inches tip, inches 1 .698 average, 1.790 inches cHtesd Gfflif 13 Company

A-17

I i

MODEL DIMENSIONAL DATA (CONT.)

Leading Edge Skit (Cont.)

22.974 B.L.

Chord location - root 29.291 tip B.L.

20.0

Angle from stowed position, degrees Section III (inboard), sealed

0.261

Area, square feet 16.377 Span

2.714

Chord length - root, inches

1.881

tip, inches 2.298 average, inches B.L.

Chord location - root 6.646 B.L. 22.974 tip

Angle from stowed 20.0

position, degrees 7 Q r Rudder, (r , r ) (Upper) (Lower) 0.161 0.203 Area, square feet Location

Lower end W.L. 22.924 15.793

29.119 22.924

Upper end W.L.

Hinge line, percent 71 71

vertical chord

7.133

Span, inches 6.195

Deflection limits, +30 + 30

degrees

Root chord, inches 4.278

3.097

Tip chord, inches 3.585 3.907

A-18

I i

MODEL DIMENSIONAL DATA (CONT.)

Rudder, (Cont.)

Mean chord length 3.750 4.097 (RMS), inches Mean chord location W.L. 25.949 19.272 Percent of vertical 10.5 13.1 tail Spoiler, ( )* a Panel I (Inboard Section) Area per side, square feet 0.0514 Span, inches 3.606 Sweep of hinge line, degrees 9.832 Chord lengths, inches Root 2.206 Tip 2.026 Average Mean (RMS) Chord locations, inches Root W . S . 5.550 Tip W . S . 9.157 Average W . S .

Mean (RMS) W . S .

Maximum deflection, 60.000 degrees Panel 2 Area per side, square feet 0.0508 Span, inches 3.610 Sweep of hinge line, degrees 9.832 Chord lengths, inches Root 2.026 *AII dimensions given in wing reference plane.

iic-a Oni.-ia C- A-19

I L

MODEL DIMENSIONAL DATA (CONT.)

Spoiler, (Cont.)

Tip 2.026 Average 2.026 Mean (RMS) 2.026 Chord locations, inches Root W.S. 9.169 Tip W.S. 12.779 Average W.S.

Mean (RMS) W.S.

Maximum deflection, 60.000 degrees Panel 3 Area per side, square feet 0.0412 Span, inches 2.927 Sweep of hinge line, degrees 9.832 Chord lengths, inches Root 2.026 Tip " • 2.026 Average 2.026 Mean (RMS) 2.026 Chord locations, inches Root W.S. 13.270 Tip W.S. 16.197 Average W.S.

Mean (RMS) W.S.

Maximum deflection, degrees 60.000 Panel 4 Area per side, square feet 0.04 12 Span, inches 2.927 Sweep of hinge line, degrees 9.832 A-20

k £

MODEL DIMENSIONAL DATA (CONT.)

Spoiler (Cont.)

Chord lengths, inches

Root 2.026

Tip 2.026

Average 2.026

Mean (RMS) 2.026

Chord locations, inches

Root W.S. 16.209

Tip W.S. 19.136

Average W.S.

Mean (RMS) W.S.

Maximum deflection, 60.000

degrees

Panel 5

Area per side, square feet 0.0272

Span, inches 2.490

Sweep of hinge line, degrees 12.347

Distance hinge line forward 0.336

of leading edge, inches

Chord lengths, inches

Root 1.910

Tip 1.910

Average 1.910

Reference** 2.247

**Reference chord is defined as twice the distance from the hinge line to

spoiler t.e. minus the average chord.

A-21

I I

MODEL DIMENSIONAL DATA (CONT.)

Spoiler (Cont.)

Chord locations, inches Root W.S. 19.627 Tip W.S. 22.117 Average W.S.

Reference W.S.

Maximum deflection, degrees 60.000 Panel 6 Area per side, square feet 0.0272 Span, inches 2.490 Sweep of hinge line, degrees 12.347 Distance hinge line forward of 0.336 leading edge, inches Chord lengths, inches Root 1.910 Tip 1.910 Average 1.910 References** 2.247 Chord locations, inches Root W.S. 22.129 Tip W.S. 24.620 Average W.S.

Reference W.S.

Maximum deflection, degrees 60.000 **Reference chord is defined as twice the distance from the hinge line to spoiler t.e. minus the average chord.

Geoifia Cui

A-22

k i

MODEL DIMENSIONAL DATA (CONT.)

Spoiler (Cant.)

Panel 7 Area per side, square feet 0.0399 Span, inches 4.645 Distance hinge line forward 0.260 of leading edge, inches Sweep of hinge line, degrees 17.033 Chord lengths, inches Root 1.236 Tip 1.236 Average 1.236 Reference** 1.757 Chord locations, inches Root W.S. 25.111 Tip W.S. 29.756 Average W.S.

Reference W.S.

Maximum deflection, degrees 60.000 Panel 8 Area per side, square feet 0.0340 Span, inches 3.962 Sweep of hinge line, degrees 17.033 Distance hinge line forward 0.260 of leading edge, inches Chord lengths, inches Root 1.236 Tip 1.236 Average 1.236 Reference** 1.757 **Reference chord is defined as twice the distance from the hinge line to spoiler t.e. minus the average chord.

a Cumuan, A-23

I i

MODEL DIMENSIONAL DATA (CONT.)

Spoiler (Cont.)

Chord locations, inches Root W.S. 30.247 Tip W.S. 34.209 Average W.S.

Ref erenc e W.S.

Maximum deflection, degrees 60.000 Panel 9 Area per side, square feet 0.0336 Span, inches 4.086 Sweep of hinge line, degrees 17.033 Distance hinge line forward 0.247 of leading edge, inches Chord lengths, inches Root 1.184 Tip 1.184 Average 1.184 Reference** 1.679 Chord locations, inches Root W.S. 34.700

Tip W.S. 38.785

Average W.S.

Reference W.S.

Maximum deflection, degrees 60.000 **Reference chord is defined as twice the distance from the hinge line to spoiler t.e. minus the average chord.

A-24

MODEL DIMENSIONAL DATA(CONT.)

Vortex Generator

, (u )

Height, inches Superscript A 0.08 Superscripts 0.10 Superscript C 0.12 Width, inches Superscript A 0.16 Superscripts 0.20 Superscript C 0.24 Angle to freestream, degrees 15.00 Chordwise location (centerline 15.00 of generator), % of t.e. flap Span wise location, inches from flap tip chord Subscript I 0.18 Subscript 2 0.23 Subscript 3 0.28 Subscript 4 0.33 Subscripts 0.38 Vertical Stabilizer, (V ) Airfoil section NACA 0013-1.1-0.40/1.575 (modified) Areas (theoretical), square feet Projected 1.531 Wetted (exposed only) 2.848 Span, inches 16.535 Chord lengths, MAC, inches 13.390 Root, inches 14.817 Tip, inches 11.853 Aspect ratio 1 .240 I Ml heed Georgia Company

A-25

I i

MODEL DIMENSIONAL DATA (CONT.)

Vertical Stabilizer (Cont.)

Taper ratio 0.800 Sweep of 25% chord line, degrees 34.931 25% MAC location F.S. 107.992 W.L. 23.388 Volume coefficient 0.079 Tail length, inches 53.246 11A 2 Wing, (W ) (6204.601 ft Full Scale) Data Area, square feet Reduction Planform, theoretical 9.8857 9.878 Planform, exposed 8.4930 (Outboard of B.L.)

Wetted, exposed 16.504 (Outboard ot B.L.)

Volume, Cu. Ft. 0.770 Span, inches (8.749') 104.997 104.997 MAC chord length, inches 14.826 14.817 Location of 0.25 chord MAC F.S. 53.765 W.L. 12.557 B.L. 21.658 Aspect ratio 7.744 Taper ratio, theoretical 0.371 Taper ratio, exposed 0.401 Dihedral (0.25 chord), degrees 3.500 Sweep angle, degrees Panel I (Inboard) Leading edge 28.449 0.25 chord 24.268 Trailing edge 10.046 lm • Hrr-'J Cfi.:r:j C A-26

k t

MODEL DIMENSIONAL DATA(CONT.)

Wing (Cont.)

Panel 2

28.449

Leading edge

24.803

0.25 chord

Trailing edge 12.581

Panel 3

Leading edge 27.382

0.25 chord 23.954

Trailing edge 12.581

Panel 4

Leading edge 27.382

0.25 chord

25.001

Trailing edge 17.298

Chord length, inches

Root 21.806

Break station, inboard 14.826

Break Station, Mid

13.606

Break Station, Outboard 13.018

Tip

7.332

Chord location, inches

Root B.L. 0

Break Station, Inboard

B.L. 19.144

Break Station, Mid

B.L. 22.973

Break Station, Outboard

B.L. 24.970

Tip

B.L. 52.498

Geometric twist, degrees Root Break Station, Inboard 1.132 Break Station, Mid 1.500 Break Station, Outboard 1.576 li« ''tieeo Georgia Companv A-27

I i

MODEL DIMENSIONAL DATA (CONT.)

Wing (Cont.)

Tip 3.500 Flop Trock Fairing, Z Centerline locations W.S.

W.S.

W.S.

W.S.

W . S .

W . S .

W.S.

C* 97 Nose Landing Gear Fairing (Z ) .....

Maximum length, inches 11.30 Maximum frontal area, square inches 6.06 Wetted area, square feet 0.6311 Imprint area on fuselage, square feet 0.5886 Main Landing Gear Fairing, (Z ) , Maximum length, inches 33.290 35.36 Maximum frontal area, square feet 0.125 0.135 Wetted area, square feet 4.364 4.366 Imprint area on fuselage, square feet 3.513 Maximum width, inches 14.078 14.18 Wing - Fuselage Fillet, (Z ) Maximum length 41.71 Wetted area 3.638 Imprint area of fuselage and fillet 3.582 on Wing Side area 1.10 A-28

I i

MODEL DIMENSIONAL DATA (CONT.)

Wing - Fuselage Fillet (Cont.)

Imprint area of wing and 3.013 fillet on fuselage Location:

Most forward point F.S. 32.13

Most aft point F.S. 73.84

o Main Landing Gear Outer Door (FWD), (d ) Length, inches 6.63 Reference area, square feet 0.1932 Span, inches 4.19 Deflections, % Open 0,10,25,50, 75,100 m4* Main Landing Gear Outer Door (AFT), (d ) Length, inches 6.63 Reference area, square feet 0.1932 Span, inches 4.19 Deflections, % Open 0,10,25,50, 75,100 Main Landing Gear Inner Door (FWD), (d ) Length, inches 4.23 Reference area, square feet 0.0311 Span, inches 1 .06 Deflections, degrees 0,3,6,29,75,95 Main Landing Gear Inner Door (AFT), (d ; Length, inches 4.23 Reference area, square feet 0.0311 Span, inches 1 .06 Deflections, degrees 0,3,6,29,75,95 A-29

k t

MODEL DIMENSIONAL DATA (CONT.)

Nose Landing Gear Inner Door, (d ) Length, inches 5.35 Reference area, square feet 0.0884 Span, inches 2.37 Deflections, % Open 0,10,25,50,75,100 Nose Landing Gear Outer Door, (d ) Length, inches 4.18 Reference area, square feet 0.0371 Span, inches 1 .28 Deflections, % Open 0,10,25,50,75,100 Ref. Moment Center: F.S. 53.762 W.L. 10.578

B.L. 0.000

A-30

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A-31

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