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AIAA-2002-0417
Reynolds Number Effects on the
Stability & Control Characteristics of a
Supersonic Transport (Invited)
L. R. Owens, and R. A. Wahls
NASA Langley Research Center
Hampton, Virginia
M. B. Elzey
Boeing Commercial Airplane Group
Seattle, Washington
M. P. Hamner
LeaTech, LLC
Baltimore, Maryland
40th AIAA Aerospace Sciences
Meeting & Exhibit
14-17 January 2002 / Reno, NV
For permission to copy or republish, contact the copyright owner named on the first page. For AIAA-held copy-right, write to AIAA, Permissions Department, 1801 Alexander Bell Drive, Suite 500, Reston, VA 20191-4344 AIAA-2002-0417
REYNOLDS NUMBER EFFECTS ON THE
STABILITY & CONTROL CHARACTERISTICS OF A
SUPERSONIC TRANSPORT
L. R. Owens " and R. A. Wahls t Aerodynamics, Aerothermodynamics, and Acoustics Competency NASA Langley Research Center Hampton, Virginia M. B. Elzey:t: M. P. Hamner § Boeing Commercial Airplane Group LeaTech, LLC Seattle, Washington Baltimore, Maryland ABSTRACT with higher Reynolds numbers for both the landing A High Speed Civil Transport (HSCT) and transonic configurations. The stabilizer configuration was tested in the National Transonic effectiveness increased with Reynolds number for both configurations. The directional stability also Facility at the NASA Langley Research Center as part of NASA's High Speed Research Program. A increased with Reynolds number for both series of tests included longitudinal and configurations. The landing configuration without lateral/directional studies at transonic and low- forebody chines exhibited a large yawing-moment speed, high-lift conditions across a range of departure at high angles of attack and zero Reynolds numbers from that available in sideslip that varied with. increasing Reynolds numbers. This departure characteristic nearly conventional wind tunnels to near flight conditions.
disappeared when forebody chines were added.
Results presented focus on Reynolds The landing configuration's rudder effectiveness number sensitivities of the stability and control also exhibited sensitivities to changes in Reynolds characteristics at Mach 0.30 and 0.95 for a complete HSCT aircraft configuration including number.
empennage. The angle of attack where the pitching-moment departure occurred increased .
INTRODUCTION Aerospace Engineer, Flow Physics and Control Branch, Ground-to-flight scaling remains one of Senior Member, AIAA many challenges facing today's designers of t Assistant Head, Configuration Aerodynamics Branch, aerospace vehicles. The goal of ground-to-flight Associate Fellow, AIAA scaling is the preflight prediction of multiple key :t:Senior Principal Engineer, Boeing Commercial Airplane Group aerodynamic characteristics with sufficient §Principal, LeaTech, LLC, Senior Member, AIAA accuracy to meet both performance guarantees Copyright © 2002 by the American Institute of Aeronautics and Astronautics, Inc. No copyright is asserted in the United States and certification requirements. The designer must under Title 17, U. S. Code. The U. S. Government has a strive to know the performance of a vehicle with royalty-free license to exercise all rights under the copyright high confidence prior to flight, thus enabling claimed herein for Governmental Purposes. All other rights are reserved by the copyright owner.
American Institute of Aeronautics and Astronautics AIAA-2002-0417 optimal design trades prior to flight and elimination Technology tasks was the development of of costly fixes to the aircraft after initial flight tests. practical concepts and design and analysis Specific challenges, experiences, and methods to allow the HSCT to operate safely and suggested approaches to ground-to-flight scaling efficiently. Towards this goal, a scaling effort was have been documented extensively over the years defined to reduce the risk in the design process by for a variety of vehicle classes (refs. 1, 2, among identifying those physical features of an actual many others). Reynolds number effects are flight vehicle that would contribute to stability and foremost among many factors affecting successful control differences between it and wind-tunnel ground-to-flight scaling (refs. 3 - 5). The Reynolds models of various scale. Figure 1 shows the number is the ratio of inertial to viscous forces, nominal mission profile for the baseline reference and is the primary aerodynamic scaling parameter configuration used in the HSR program, and a used to relate sub-scale wind tunnel models to full- comparison to the capability of several wind scale aircraft in flight. The challenge of Reynolds tunnels. The baseline reference configuration, number scaling increases with the size of a full- known as Reference H, was provided by Boeing scale aircraft as the Reynolds number increment and represented a Mach 2.4, 300 passenger between that obtainable in conventional wind aircraft with a 5000 nautical mile range.
tunnels and flight conditions expands. A series of wind tunnel tests was Additionally, the challenge for both wind tunnel conducted in the National Transonic Facility (NTF) and computational approaches increases as flow at the NASA Langley Research Center (LaRC) features become dominated by viscous-sensitive across a wide range of Reynolds numbers. The phenomena such as boundary-layer transition, Reynolds numbers ranged from that available in shock/boundary-layer interaction, and separation conventional wind tunnels to near flight condition onset and progression. at subsonic and transonic Mach numbers. The The present investigation was conducted tests included longitudinal and lateral/directional in support of NASA's High Speed Research (HSR) studies with and without an empennage at Program, Phase II, which was conducted from transonic and low-speed, high-lift conditions. This 1993-1999 (ref. 6). The objective of this program, paper presents results focused on the Reynolds which was NASA sponsored and jointly executed number sensitivities of the stability and control with US industry, was to develop critical high-risk characteristics at Mach numbers of 0.30 and 0.95 airframe and propulsion technologies to enable for a complete HSCT aircraft configuration industry development of an economically viable including empennage.
and environmentally acceptable second generation, high speed civil transport (HSCT).
Aerodynamic performance, one of several broad TERMS. ABBREVIATIONS. & ACRONYMS airframe technology areas, included tasks to ARC NASA Ames Reseach Center address Configuration Aerodynamics for high- BL butt-line, model coordinates, inches speed conditions and High-Lift Technology for Cl 95% confidence interval g5 take-off and landing. These elements c local chord length, inches encompassed not only the challenge of efficient Co drag coefficient supersonic cruise flight, but also the off-design C lift coefficient L rolling-moment coefficient referenced to challenges (ref. 7) of efficient transonic cruise and C 0.50 mac acceleration and quiet high-performance take-off and landing. The objective of both the C pitching-moment coefficient referenced to M Configuration Aerodynamics and the High-Lift 0.50 mac American Institute of Aeronautics and Astronautics AIAA-2002-0417 C pitching-moment coefficient at C =0 to full-scale flight values, depending on the aircraft M L o type and size. The facility (fig. 3) is a fan-driven, C longitudinal stability derivative M C L closed circuit, continuous-flow, pressurized wind stabilizer effectiveness, per deg M6 C ,stab tunnel capable of operating in either dry air at directional stability derivative, per deg CnB warm temperatures or nitrogen from warm to C rudder effectiveness, per deg cryogenic temperatures. The test section is 8.2 ft nb,rud by 8.2 ft in cross section and 25 ft in length. The C yawing-moment coefficient referenced to n test section floor and ceiling are slotted (6 percent 0.50 mac open), and the sidewalls are solid. Freestream C side-force coefficient y turbulence is damped by four screens and a ETW European Transonic Wind tunnel 14.95:1 contraction ratio from the settling chamber FS fuselage station, model coordinates, to the test section. Fan-noise effects are inches minimized by an acoustic treatment both upstream neutral pOint, fraction of mac hnp and downstream of the fan. A detailed HSCT High Speed Civil Transport assessment of the dynamic flow quality in the NTF HSR High Speed Research is reported in reference 9, and reconfirmed with LaRC NASA Langley Research Center recent measurements shown in reference 10. The LE leading edge NTF is capable of an absolute pressure range UD lift-to-drag ratio from 15 psia to 125 pSia, a temperature range M Mach number from -320°F to 150°F, a Mach number range from mac mean aerodynamic chord, inches 0.2 to 1.2, and a maximum Reynolds number of NTF NASA's National Transonic Facility 146x10 per ft at Mach 1. Typical tests use a P total pressure, pSia T q dynamic pressure, psf temperature range from -250°F to 120°F. Further Rn Reynolds number based on mac facility details can be found in reference 11.
local leading-edge radius, inches local maximum airfoil thickness, Model Description inches The wind-tunnel model is a 2.2% scale TE trailing edge representation of the HSR baseline configuration total temperature, of known as Reference H. Although the model
TT
without the empennage was tested in the NTF WL waterline, model coordinates, inches during the HSR program, the present paper a angle of attack, deg focuses on results obtained for the full angle of sideslip, deg configuration with empennage. Figure 4 shows a downwash angle, deg planform and sideviell)' sketch of the model with non-dimensional semi-span station several reference locations noted.
The model has a cranked-delta wing planform with an aspect ratio of 2.367, a span of EXPERIMENTAL APPROACH 34.23 inches, and a mac of 22.71 inches. The inboard wing (11 s 0.522) has a blunt (ric,..., 0.0025 Facility Description to 0.0030), subsonic LE with a sweep change from The NTF (ref. 8) is a unique national 76 to 68.5 deg at 11 = 0.226, a twist varying from facility (fig. 2) that enables tests of aircraft
approximately 1 deg near 11 = 0.10 to -2 deg near
configurations at conditions ranging from subsonic to low supersonic speeds at Reynolds numbers up 11 = 0.50, and variable thickness ratio (tma/c) from American Institute of Aeronautics and Astronautics AIAA-2002-0417 0.043 to 0.024. The outboard, supersonic LE is configuration only. The chine LE was located at sharp, swept 48 deg, has a constant twist of -1.6 FS 9.900, and the chine semi-span was 0.265 deg for 'Y] ~ 0.65, and a constant thickness ratio of inches from the side of the body. The chine root 0.024. The reference area for the model is 3.436 chord was 1.98 inches and it had a tip chord of ft2. Table 1 provides several important ratios 1.367 inches.
The model fuselage had an upswept, relating the model size to the NTF test section.
closed aft body of the Reference H configuration.
The overall body length was 83.060 inches.
The model's horizontal tail had an ref. area / NTF cross sectional area 0.0515 exposed area of 0.338 fe and an aspect ratio of model span / NTF width 0.3478 1.845 (based on exposed area and span). The solid blockage ratio, a = 0 deg 0.0022 horizontal tail span was 9.476 inches. The LE Table 1. Model size relative to the NTF test section.
sweep was 53.5 deg and the TE sweep was -27.4 deg. The airfoil section was a wedge-slab-wedge The model has multiple inboard LE and type.
TE parts and multiple outboard wing panels each The model's vertical tail had an exposed with different LE and TE deflections. It also area of 0.199 fe and an aspect ratio of 0.869 included four detachable, 8.43 inch long, constant (based on exposed area and span). The vertical internal diameter (1.236 inches), circular flow- tail span was 4.990 inches. The LE sweep was 51 through nacelles with boundary-layer diverters deg and the TE sweep was -11.5 deg. The airfoil located between the wing and nacelle. The section was also a wedge-slab-wedge type. Two inboard nacelles are rigged with toe-in and pitch vertical tails were built, one with an undeflected (nose down) angles of 1 and 4.17 deg, rudder and the other with a +30-deg rudder respectively; the outboard nacelles are rigged with deflection (TE toward the left wing). The rudder toe-in and pitch angles of 2.4 and 2.84 deg, had an exposed area of 0.054 ft2. The rudder respectively. The multiple LE and TE parts in hingeline was a vertical line located at FS 73.618.
combination with the multiple outboard panels The tip chord of the rudder is 47.5% of the vertical enabled testing of a variety of configurations tail tip chord, and the rudder root chord is 22.08% including the supersonic cruise, take-off, landing, of the vertical tail root chord. The rudder stall recovery, and transonic cruise configurations.
deflection was only tested at the low speed Results for the transonic cruise and landing wing conditions.
configurations with the nacelle/diverters are The model was instrumented with 17 aft included herein. Table 2 includes wing flap body pressures distributed circumferentially at FS deflections for these two configurations.
65.306 and 6 pressures distributed in a row on the port side (45-deg up from bottom-dead-center) just Designation LE Deflection, deg TE Deflection, deg below the horizontal tail location. These pressures Inboard/Outboard Inboard/Outboard were used for a limited computational study (Euler Landing 30/30 20/20 calculations) that investigated the blade-sting interference effects for symmetric flow conditions Transonic Cruise 0/10 0/3 only. This computational study showed a small, Table 2. Wing flap configurations.
lower-surface compression increase in the wing TE region near the blade sting entry, which A forebody chine for the Reference H provided a small interference effect at transonic geometry was tested with the high-lift, landing conditions. The results of this study are American Institute of Aeronautics and Astronautics AIAA-2002-0417 documented in reference 12. Cavity pressures to have some unknown level of bias in the were also measured just inside the model near the absolute data levels that are consistent with blade-sting seal. These pressures were used to standard sideslip testing techniques. However, monitor the integrity of the seal during testing.
steps were taken to ensure that the Rn effects None of the pressure data obtained during these would be primarily indicative of changes in the tests will be presented in this report. model flowfield only. Boundary-layer transition trip The model was designed and constructed strips were placed near the leading edge of the specifically for testing in the cryogenic, blade sting to minimize the change in this pressurized conditions of the NTF. The model jig interference effect with Rn. By tripping the shape was that of the Mach 2.4 cruise design boundary layer on the blade sting, the Rn effects point. The model was built of maraging steel with were assumed to be produced mainly by the a surface finish of 8-16",,-inches (root mean model configuration. In addition, directional square) and a contour tolerance of ±0.005 inches. derivatives were calculated over a limited j3 range The model is shown in figure 5 mounted in the (-4 deg < ~ < 4 deg) in an attempt to minimize the NTF test section on a lower-swept blade sting, blade-sting, interference effect.
which has a NACA 0012 airfoil section normal to The entry point of the blade sting into the the blade sting's swept LE. The sting mounts to a fuselage is shown generally in figure 4. The non- 6-deg offset stub sting, which in turn mounts to the metric blade sting needed a clearance gap to facility arcsector resulting in a model a range from prevent fouling as the model/balance deflected -4 to 24 deg. The j3 range varied depending on under load. An unsealed gap would have allowed flow to enter the fuselage cavity and thus would the angle-of-attack setting. At lower a's, j3 varied have affected the measured forces and moments.
between -12 and 12 deg. At higher a's, j3 was During testing of this configuration, different seals limited to a range of -8 to 8 deg.
(manufacturing techniques, size, stiffness, and material thickness) were developed in an attempt Model Support System to find the best method for consistently developing Testing on a blade sting support in a good seal (minimal flow into the fuselage cavity) sideslip is not an ideal way to obtain that also produced minimal fouling loads. This lateral/directional data. However, the research turned out to be a somewhat difficult task, goal of trying to model the aftbody closure with especially considering the range of test conditions minimal geometry modification for accommodating that needed to be covered. Monitoring of local the support sting led to the use of a blade sting cavity pressures during testing provided an support. The blade-portion of this sting in sideslip indication of seal integrity. The results of a seal produced a pressure field on the aft body and loading study indicated a small amount of fouling vertical tail, which generated a positive, was present, but this fouling was considered directional-stability interference effect. A negligible.
comparison of the directional stability was made at However, a problem was encountered in low Reynolds number for a similar configuration maintaining the integrity of the strut seal during the with a single post mounting system tested in testing. The local aerodynamic loads on the seal, NASA LaRC's 14-by-22 Foot Subsonic Tunnel.
especially for the transonic conditions, and the This comparison showed that the NTF measured exposure of the seals to the cryogenic directional stability was somewhat higher, but the environment made it necessary to manufacture post mount also causes some interference effect.
new seals frequently. The process for controlling From this comparison, the NTF data was assumed American Institute of Aeronautics and Astronautics AIAA-2002-0417 the manufacturing of new seals was carefully complete description of these measurements and considered and executed to try to maintain good subsequent calculations is given in reference 14.
quality seals. An improved sealing technique is needed for any future testing on a blade-sting Data Reduction and Corrections model support of this size. Information on the various instrumentation devices, the data acquisition and control Instrumentation computers, and the data reduction algorithms for Aerodynamic force and moment data were the different measurement systems is provided in obtained with an internal, unheated, six- reference 14. Standard balance, a, and tunnel component, strain gauge balance. The balance parameter corrections have been applied. Note used was one of the NTF-113-class balances that the use of unheated balances in the cryogenic having the load capacity and accuracy shown in environment requires additional attention towards table 3. An internal, heated accelerometer temperature compensation. The temperature package was used to measure the onboard angle compensation methods are designed to correct balance output due to thermal loads (refs. 14, 15).
Component Full-Scale Nominal Accuracy Body cavity pressures were used to calculate Load 95% confidence corrections to normal and axial forces and pitching moment to adjust the internal cavity pressure Normal,lbs ±D.09% full-scale ±6500 condition to freestream static. Nacelle internal Axial,lbs ±400 ±D.33% full-scale drag and base pressure corrections were only Side,lbs ±4000 ±D.19% full-scale
applied to the j3 = 0 deg, a sweep data based on
Pitch, in-Ibs ±D.11 % full-scale ±13000 the measurements described previously in the wing/body testing (refs. 16, 17). The angle of Yaw, in-Ibs ±6500 ±D.23% full-scale attack was corrected for flow angularity (upflow Roll, in-Ibs ±9000 ±D.35% full-scale only) by measurement of both upright and inverted Table 3. NTF-113 balance capacity and accuracy.
model normal force data for a given configuration and flow condition. No consistent technique or
of attack for j3 = 0 deg a sweeps; quoted accuracy
data was available to characterize the tunnel side of the package under smooth operating wind flow and no attempt was made to correct the flow tunnel conditions is ±0.01 deg (ref. 13). For angularity for this component. Wall and model sideslip conditions, arcsector measured pitch and support interference effects have not been roll angles plus calibrated sting bending (including accounted for in the data. The wall effects were non-metric bending from blade loading) were used minimized through model sizing (table 1).
to determine a and j3. The onboard accelerometer used could not measure angles out of the tunnel Test Conditions vertical plane of symmetry. Angles measured The NTF allows testing across a wide using the arcsector angles plus sting bending range of Rn's from that available in conventional technique are not as accurate as those measured wind tunnels to near flight conditions at subsonic by an onboard accelerometer, but are generally and transonic M's. Tests of the 2.2% Reference H considered of the same order of accuracy.
model spanned M from 0.30 to 1.10, and Rn's The primary measured flow variables from 4.5 to 120 million based on the mac. The include both the total and static pressures and the present paper focuses on both the low-speed and total temperature. Mach number, Rn, and q are transonic regimes representative of landing and calculated from these measured parameters. A American Institute of Aeronautics and Astronautics AIAA-2002-0417 transonic cruise.
tunnel; the NTF model is usually tested at a The landing configuration data was matching low Rn condition with the boundary-layer obtained at M=0.30 for a Rn range from 4.5 to 90 tripping (forced transition) strategy used in that million. The transonic cruise configuration data facility. The data for the 2.2% Reference H full was obtained at M=0.95 for a Rn range from 10.2 configuration model was not acquired with fixed to 80 million. Figure 1 indicates the relationship of transition on the wing or the empennage. This the NTF test conditions to flight, and figure 6 was primarily due to the. potential at the time for a provides the NTF operational envelopes for one-third-scale flight test (which never occurred) M=0.30 and M=0.95 with specific test pOints anticipated to fly at conditions susceptible to identified. Full-scale flight Rn's were not transitional flow. No data with fixed transition on obtainable due to the large size .of the full-scale the wing or tail surfaces is available for aircraft, model size and other limitations. For the configurations presented herein.
M=0.30 test condition, the Rn was limited by the Transition was consistently fixed on the maximum P sustainable cryogenic forebody with a ring of carborundum grit located T for reliable, operations (i.e., 100 psia). The other limit was in 1.5 inches from the nose, and on the nacelle part driven by the requirement of testing the same internal surface to facilitate the internal nacelle model at transonic conditions. Testing of the full drag correction. As previously discussed, configuration on a blade support sting imposed transition was also fixed on the blade sting to additional load limits at M=0.95 (q=1800 psf minimize the dependence of the blade sting boundary in fig. 1). interference effect on Rn variation. All trips were The goals of assessing Rn scale effects sized and located based on traditional criteria (ref.
and extrapolation to flight conditions required a 18).
series of intermediate conditions to better identify trends. As seen in figure 6, the desired Rn range could not be covered at a constant, PT level (q RESULTS & DISCUSSION level). However, the independent control of PT, TT, The purpose of this paper is to document and fan speed in the NTF allow the isolation of the Rn sensitivities of stability and control pure Rn effects, pure static aeroelastic (q) effects, characteristics for a relevant, supersonic transport and pure compressibility (M) effects. Several configuration at conditions representative of conditions at each M are used to isolate static landing and transonic cruise, M = 0.30 and 0.95.
aeroelastic effects from the Rn effects as shown in Note that in the discussion of these data, the figure 6. During Rn sweeps, the ratio of dynamic landing configuration has wing landing flap pressure (q) to the model material modulus of deflections (see table 2), forebody chines, a elasticity (E) is held constant. This is done to vertical tail with no rudder deflection, and a maintain a constant, static aeroelastic state (q/E) horizontal stabilizer (stab) setting of 0 deg. Any due to the variability of the modulus of elasticity changes to this baseline configuration are referred over the temperature range of the NTF. to as the "landing (change)". For example, if data was obtained for a landing configuration with no Boundary-Layer Transition horizontal stabilizer, then this data will be identified as "landing (no stab)". Similarly, the transonic A basic strategy used in the NTF includes testing at high Rn conditions with free transition. cruise configuration has wing transonic cruise flap The high Rn test condition typically corresponds to deflections, a vertical tail with no rudder deflection, a design flight condition. To anchor the NTF data and a horizontal stab setting of 0 deg. Any to low Rn data obtained in a conventional wind changes to this baseline configuration will be American Institute of Aeronautics and Astronautics AIAA-2002-0417 referred to as "transonic (change)". Note that the interval for each configuration. The 95% transonic configurations were always tested confidence interval is interpreted as the bounds without forebody chines during this investigation. about an estimated mean (average of multiple, Figure 7a presents representative repeat polars) that encompasses the true mean longitudinal data for the landing configurations at a value with a chance of 95%. A number of repeat Rn of 90 million. The figure illustrates the basic,
runs were obtained for longitudinal runs with 13 = 0
longitudinal aerodynamic characteristics with deg to provide the average values of the 95% different horizontal stabilizer configurations. confidence interval for each force and moment Figure 7b presents similar data for the transonic coefficient. Since only a few repeat runs were cruise configurations at a Rn of 80 million. These made for the lateral/directional data runs, the data are shown to give the reader a general idea averages listed do not include any of these data.
of the overall character of the forces and moments Table 4 below lists these values for the from which the longitudinal stability and control longitudinal repeat runs.
parameters were calculated. Note that the effects of adding and deflecting the stab are clearly seen Landinq Transonic Cruise in the forces and moments.
±0.0014 ±0.0020 C L Figure 8a presents representative ± 0.0003 ± 0.0004 CD lateral/directional data for the landing configuration ±0.0003 ±0.0007 C M at various a's and a Rn of 90 million. Figure 8b C ± 0.0005 ± 0.0006 y presents similar data for the transonic C ± 0.0002 ± 0.0003 n configuration at a Rn of 80 million. These data are C ± 0.0002 ± 0.0001 shown to give the reader a general idea of the Table 4. Average CI for each configuration for overall character of the forces and moments from longitudinal repeat runs.
which the directional stability and control parameters were calculated.
Static Aeroelastic Effects The data as acquired, and presented in Achieving high Rn's approaching those figures 7 and 8, include the combined effects of characteristics of flight requires the manipulation static aeroelastic deformation and Rn effects. In of both the T T and P , as seen in figure 6. As a T general, addressing static aeroelastic effects is result, the static aeroelastic deformation of the necessary as a means to isolate and more model, in particular the wing, under load must be properly address Rn effects. However, the static considered when attempting to isolate Rn effects.
aeroelastic corrections are not included for the Previous reports for high aspect ratio subsonic data with 13 because only a very limited set of transport configurations have shown the static static aeroelastic data was acquired for these aeroelastic effects to be on the order of Rn effects.
runs.
Often these aeroelastic effects are opposite in sense to that of Rn trends, thus masking the Rn Repeatability effects (refs. 20, 21). Like the subsonic transport Data presented herein were acquired configurations, the current low aspect ratio HSCT across two wind-tunnel tests of the model within model is flexible under load, most notably on the several months of each other. This section thin outboard wing panel and empennage (refs.
provides a list of short-term repeatability estimates 16,17).
(within test / Mach series), as defined in reference The effects of static aeroelastic wing and 19, quantified in terms of a 95% confidence empennage bending were obtained with constant American Institute of Aeronautics and Astronautics AIAA-2002-0417 Rn at high and low q test conditions, as shown in that departure to progress more rapidly at even figure 6. Adjustments for these effects were made higher a's. Overall, the pitch stability appears to
to the a sweep (13 = 0 deg) data only because
improve with increasing Rn. Later, the discussion limited resources and test plan priorities did not will look at the longitudinal stability as a function of permit the acquisition of aeroelastic effects for 13 Rn in greater detail.
sweep runs. For the longitudinal data presented, The transonic configuration demonstrated the sensitivity to aeroelastic effects for lift and a more positive C as compared to the landing M o pitching-moment coefficients were obtained and configuration, which is produced by the smaller
used to shift these data to a wind-off condition (q =
amount of wing camber associated with the
o psf). This adjustment was used to obtain results
outboard wing flap deflections. This C is also M o for the rigid, non-deformed model shape most somewhat insensitive to any variation in Rn as frequently used in computational simulations. The shown in figure 9b. The transonic configuration correction procedure is similar to that discussed in exhibits stable pitch stability up to a C of about L references 16 and 17. However, the correction 0.5, above which the nonlinear progression begins procedure used in the current paper adjusted the due to the same factors discussed above for the coefficient data to the rigid model shape instead of landing configuration. At lower a's, the the lowest dynamic pressure level as described in longitudinal stability is less sensitive to Rn change these references.
transonically. As observed for the landing configuration, the model nose up onset with C is L Reynolds Number Effects delayed as the Rn increases.
The following discussion will examine the Some of the basic longitudinal stability Rn trends for pertinent longitudinal stability and and control parameters calculated from data (with control characteristics (with static aeroelastic static aeroelastic corrections) for the landing corrections) and directional stability and control configurations are shown in figure 10. In the characteristics (without static aeroelastic upper left portion of this figure, the wing induced corrections) .
downwash angle affecting the horizontal Longitudinal Characteristics. Figure 9 stabilizer's performance is shown as a function of presents the Rn effects on pitChing-moment Rn for specific a'S. As a increases, the downwash characteristics for both the landing and the angle increases because the inboard wing transonic configurations. These data include generates more lift, which results in a larger corrections for static aeroelastic effects, thus turning angle in the oncoming flow. The providing better isolation of Rn effects. In figure downwash angle increases slightly as Rn 9a, the landing configuration has the expected increases, which presumably is the result of more negative C produced by the increased wing M o efficient turning of the flow by the wing and TE camber from the inboard/outboard, wing flap flaps. This efficiency increase is probably due to a deflections. The C is somewhat constant as the M combination of decreases in the wing boundary- o Rn increases. The landing configuration exhibits layer thickness (local camber increase and pitch stability up to a C of about 0.45. Above this healthier boundary-layer approaching flap) as well L C level, the stability degrades as the configuration as improvements in the TE flap performance L experiences the typical high attitude phenomenon caused by local separation delays.
associated with increasing outboard wing panel In the lower left plot of figure 10, the separations. The increase in Rn delays the onset stabilizer effectiveness for the landing of the pitching-moment departure, but also causes configuration is shown as a function of Rn. At American Institute of Aeronautics and Astronautics AIAA-2002-0417 each Rn, the stabilizer effectiveness decreases as configurations are shown in figure 11. The layout
a increases as expected. The data at a = 20 deg
of this figure is the same as that discussed in figure 10 for the landing configuration.
are an exception to this trend, which is attributed The downwash angle, the pitch stability, to the error associated with trying to use steady state data to characterize a highly unsteady flow and the neutral point show no significant state. A consistent pattern of stabilizer dependence on Rn. However, the stabilizer effectiveness emerges with Rn, if the a = 20 deg effectiveness for this configuration shows a strong dependence on Rn. This trend with Rn is also data are ignored. At a's of 8 and 12 deg, the consistent with the stabilizer effectiveness stabilizer effectiveness increases on the order of increase seen in subsonic transport data 5% as the Rn's increase toward that of flight.
presented in reference 22. Both the values shown These a's would be typical for the landing here and those presented in this reference configuration. These results are consistent with demonstrate 10% increases in stabilizer the stabilizer effectiveness results for subsonic effectiveness as the Rn increases from that of a transports shown in reference 22. The results in wind tunnel model scale to that of flight scale. A this reference compare the stabilizer effectiveness possible cause for this increase in stabilizer calculated from both wind tunnel and flight data.
effectiveness at higher Rn may be the result of The plots on the right side of figure 10 thinning fuselage and stabilizer boundary layers.
present the local pitch stability and neutral point The thinner boundary layers may expose more of trends with Rn for constant values of C • For C L L the actual horizontal tail geometry to the flow field values larger than 0.45, the landing configuration potentially making it more effective (ref 22).
exhibits an unstable longitudinal condition as Directional Characteristics. Next, the discussed previously for figure 9a. Just before the discussion focuses on the directional stability and onset of the model nose up condition, the local control characteristics. The reader is reminded pitch stability is insensitive to changes with that none of the data that follows has any static
Reynolds number as is shown for C = 0.4 in figure
L aeroelastic corrections because the resource
10. Right after the nonlinear onset, C = 0.45, the
L limitations did not allow a complete set of these increase in Rn produces increased pitch stability data to be collected.
as the onset of the nose up pitching moment is Figure 12 shows the Rn effects on the delayed. Moving deeper into the pitch non- yawing moment for the landing configuration with linearity, the local C values are changing rapidly M and without chines. These a sweeps were and the local pitching-moment slopes should be
obtained with !3 = 0 deg. The landing configuration
viewed more qualitatively. However, from this tested without the forebody chines demonstrated a qualitative viewpoint, the pitch stability also strong yawing moment departure at high a's that appears to be increasing with increases in Rn.
was dependent on Rn. The forebody flow The neutral point behavior with changing Rn is shown in the last plot of figure 10. Note that asymmetry dependence on Rn is typical for these values were calculated from the local smooth-sided forebodies (ref. 23). The forebody flow field is symmetric at a Rn = 4.5 million and no pitching-moment slopes and the same qualitative view should be considered for the higher C yawing-moment departure is observed for the L values. given a range. Increasing the Rn to 10 million The basic longitudinal stability and control causes a strong yawing-moment asymmetry to parameters calculated from data (with static develop. Further increases in Rn moves the onset aeroelastic corrections) for the transonic of this departure characteristic to lower a's. The American Institute of Aeronautics and Astronautics AIAA-2002-0417 lack of a fixed separation line associated with the shown in figure 15. These slopes were calculated chines for the shed forebody vortices produces from data similar to that presented in figures 13 this dependence on Rn. The addition of forebody and 14. As mentioned previously, these chines to the landing configuration greatly reduced derivatives were calculated based on data from a the magnitude of the yawing moment departure limited f3 range (-4 deg < 13 < 4 deg). One factor in and reduced the Rn dependence by providing a limiting this range was the consideration of the fixed line of separation for the fore body vortices at positive interference effect caused by the high a.
presence of the blade support sting for the model.
The Rn effects on the directional It is assumed that the blade sting with forced characteristics for several landing configurations boundary-layer transition will produce an are shown in figure 13 at a= 12 deg, before the interference effect that will have minimal Rn dependence.
onset of the previously discussed yawing-moment departure. It is obvious that at this a the vertical Figure 15a presents the directional stability derivatives for the landing configuration as tail provides a strong input to the directional a function of a and Rn. In general, the increase in stability. Also note that with the vertical tail on, the Rn tended to provide on the order of a 10% forebody chines provide an additional increase in increase in the directional stability for this the directional stability. However, the directional stability drops significantly when the vertical tail is configuration. Note that due to the nonlinear and highly unsteady nature of the a = 20 deg flow field, removed. At this a, these configurations show small Rn effects that tend to be greater at the the data at this a should only be considered larger j3's. qualitatively. Figure 15b presents similar data for the transonic configuration. This data also shows The Rn effects on the directional increases in directional stability on the order of characteristics for the same set of landing 10% with increases in Rn approaching flight configurations are shown in figure 14 at a= 20 conditions. The increase in the directional stability deg, after the onset of the previously discussed at higher Rn's may be the result of thinning yawing-moment departure. The configurations fuselage and vertical tail boundary layers exposing without the forebody chines exhibit a severe more of the actual vertical tail geometry to the flow directional instability at 13 = 0 deg. The addition of field potentially making it more effective.
the chines eliminates the strong instability at j3 = 0 The Rn effects on the directional deg, but the vertical tail is still necessary to give characteristics of the landing configuration with the configuration any directional stability at all.
and without rudder deflection are presented in However, at this a the vertical tail does not appear
figure 16. The en data for the configuration with
to be as effective in providing directional stability no rudder deflection exhibits a slight non-linearity as it was at lower a's presumably due to the near j3 = 0 deg that tends to go away as the Rn blanketing effect attributed to the wakes of both increases. For the configuration with a +30-deg the fuselage and wing. The Rn effects on the rudder deflection, a stability reversal occurs at f3 = directional characteristics are seen throughout the 1 deg, which disappears as the Rn approaches 13 range at this a. However, these effects should flight conditions. The source of this non-linearity in be carefully considered because of the highly the rudder-deflected data is believed to be unsteady nature of the flow field at this high a associated with a hingeline separation on the condition.
rudder at low Rn that goes away at Rn's The directional stability derivatives for both approaching flight. The significance of this data is the landing and the transonic configurations are American Institute of Aeronautics and Astronautics AIAA-2002-0417 that it is the first high Rn testing of a rudder 2. The stabilizer effectiveness increased with configuration, not just for the HSR program, but as higher Rn's for both the landing and the far as is known, for any Boeing Commercial wind transonic configurations. This increase was tunnel model. larger for the transonic configuration.
Finally, the effects of Rn on the rudder 3. The forebody chines supplied a strong favorable increment to C at higher a's in the effectiveness are presented in figure 17 for the n ~ landing configuration. In general, the rudder landing configuration. The a where the onset effectiveness decreases slightly as the Rn of a strong yawing-moment departure increases. This decrease may possibly be the occurred decreased with higher Rn's for the result of static aeroelastic deformation of the landing configuration without chines at 13=0 vertical tail at the higher Rn conditions. The deg.
variation of the rudder effectiveness with 13 tends 4. Directional stability increased with higher Rn's to decrease at the higher Rn test conditions.
for both the landing and the transonic However, this variation shows a significant configurations.
increase at a Rn = 10 million, especially at the
5. Directional stability in the landing configuration larger 13 values. By looking back at the C data for n was somewhat non-linear in 13 with the rudder the rudder configuration shown in figure 16, the Rn deflected +30 deg and reverses between 13 of
= 10 million data appears to have a second non-
1 and 2 deg at lower Rn. This non-linearity is linear break at a 13 > 6 deg. This second break in eliminated at the highest Rn tested.
the C data appears to be source of the increased n 6. The Rn effects on the stability and control rudder effectiveness at Rn = 10 million. Since this characteristics for these configurations were second break occurs at higher 13 values, the consistent and considered reasonable.
potential for some strong interaction with the blade However, the development of better test sting, interference flow field must be considered.
techniques (i.e., model support sting system) However, there may also be some transitional to obtain high Rn, high load data is needed for boundary-layer flow effects that are contributing to future testing efforts.
this variation in rudder effectiveness.
ACKNOWLEDGEMENTS CONCLUDING REMARKS The authors would like to thank our many Wind tunnel tests with a 2.2% scale HSCT partners from industry and the staff of the NTF for model were conducted in the NTF at NASA LaRC making these tests successful. In particular, we across a wide range of Rn's. These Rn's ranged would like to acknowledge Chet Nelson (Boeing) from that available in conventional wind tunnels to and Susan Williams (NASA-retired) who invested near flight condition at subsonic and transonic considerable effort over many years towards the Mach numbers. Results were presented that focus development and testing of this model. Also, on the Rn sensitivities of the stability and control discussions with Dave Bogue (Boeing) were very
characteristics at M = 0.30 and 0.95 for the full
helpful in the analysis of the data in this paper.
configuration with the empennage. General Finally, we would like to thank Elwood Putnam conclusions are summarized as follows: (NASA-retired) for his leadership and encouragement (especially to publish) without 1. The a where the pitching-moment departure which this work would have been greatly limited.
occurred increased with higher Rn's.
American Institute of Aeronautics and Astronautics
AIAA-2002-0417
REFERENCES 13; Finley, t.D. and Tcheng, P.: "Model Attitude 1. McKinney, L.W. and Baals, D.O. (editors): Measurements at NASA Langley Research "Wind-Tunnel/Flight Correlation - 1981," Center," AIAA Paper 92-0763, 1992.
NASA CP 2225, November 1981. 14. Foster, J.M. and Adcock, J.B.: "User's Guide 2. Haines, AB.: "Scale Effects on Aircraft and for the National Transonic Facility Research Weapon Aerodynamics," AGARD AG-323, Data System," NASA TM-110242, April 1996.
1994. 15. Williams, M.S.: "Experience with Strain Gage 3. Goldhammer, M.E. and Steinle, F.W. Jr.: Balances for Cryogenic Wind Tunnels," "Design and Validation of Advanced Transonic AGARD-R-774, 1989, pp. 18.1-18.14.
Wings Using CFD and Very High Reynolds 16. Owens, L. R., and Wahls, A.A: "Reynolds Number Wind Tunnel Testing," 17th ICAS Number Effects on a Supersonic Transport at Congress, September 1990. Subsonic Conditions," AIAA Paper 2001-0911, 4. Lynch, F.T.: "Experimental Necessities for January 2001.
Subsonic Transport Configuration Develop- 17. Wahls, R.A., Owens, L.A., and Rivers, S.M.B.: ment," AIAA Paper 92-0158, January 1992. "Reynolds Number Effects on a Supersonic 5. Bushnell, D.M., Yip, L.P., Yao, C.S., Lin, J.C., Transport at Transonic Conditions," AIAA Lawing, P.L., Batina, J.T., Hardin, J.C., Paper 2001-0912, January 2001.
Horvath, T.J., Fenbert, J.W., and Domack, 18. Braslow, A.L., and Knox, E.C.: "Simplified C.S.: "Reynolds Number Influences in Method for Determination of Critical Height of Aeronautics," NASA TM 107730, May 1993. Distributed Roughness Particles for Boundary- 6. Wilhite, A W., and Shaw, A. J.: "An Overview Layer Transition at Mach Numbers from 0 to of NASA's High-Speed Research Program," 5," NACA TN-4363, 1958.
th 20 ICAS Congress, Paper 112, August 2000. 19. Wahls, A.A, Adcock, J.B., Witkowski, D.P., 7. Nelson, C.P.: "Effects of Wing Planform on and Wright, F.L.: "A Longitudinal HSCT Off-Design Aerodynamics," AIAA Paper Aerodynamic Data Repeatability Study for a 92-2629, June 1992. Commercial Transport Model in the National 8. Gloss, B. B.: "Current Status and Some Future Transonic Facility," NASA TP-3522, August Test Directions for the US National Transonic 1995.
Facility," Wind Tunnels and Wind Tunnel Test 20. Wahls, A.A, Gloss, B.B., Flechner, S.G., Techniques, R. Aeronaut. Soc., 1992, pp. 3.1- Johnson, W.G.,Jr., Wright, F.L., Nelson, C.P., 3.7. Nelson, A.S., Elzey, M.B., and Hergert, D.W.: 9. Igoe, W.B.: "Analysis of Fluctuating Static "A High Reynolds Number Investigation of a Pressure Measurements in the National Commercial Transport Model in the National Transonic Facility," NASA TP-3475, March Transonic Facility," NASA TM-4418, April 1996. 1993.
10. Bobbitt, C.W., Hemsch, M.J., and Everhart, 21. AI-Saadi, J.A.: "Effect of Reynolds Number, J.L.: "NTF Characterization Status," AIAA Boundary-Layer Transition, and Aeroelasticity Paper 2001-755, January 2001. on Longitudinal Aerodynamic Characteristics 11. Fuller, D.E.: "Guide for Users of the National of a Subsonic Transport Wing," NASA TP- Transonic Facility," NASA TM-83124, 1981. 3655, September 1997.
12. Londenberg, W.K.: "Computational Assess- 22. Reichenbach, S.H., and McMasters, J.H.: "A ment of Aft-Body Closure for the Reference H Semiempirical Interpolation Technique for Configuration," NASA CR-1999-209521, Predicting Full-Scale Flight Characteristics," November 1999.
AIAA Paper 87-0427, January 1987.
American Institute of Aeronautics and Astronautics AIAA-2002-0417 23 . Owens , L.R ., Jr ., Hemsch , M.J ., and Low-speed diffu ser Popernack , T . G. , Jr. : "Reynolds Number Turn 3 Tu rn 2 Effects on Advanced Slender Forebodies for Angles of Attack Up to 27° at Mach 0.2," NASA TP-3493, August 1994. 48.6
~~~~
25 d i a ~ ~ ,- .;:.. " -/') J1 ~;;i; ii ;:J; I ~~
_ Hi gh-speed diffuser Tu rn 4 / / I - Screens 27:"dia plen um 2.6 half-a ng le " Cooling coil L Sl otted test sec ti on L Wide-angle diffu ser 8.2 by 8.2 350 ~---~-~-----~ Figure 3. NTF circuit diagram (lin ear dimensions in ft).
B. L. 0.000
0- ~ - '''~ 1\
I \ I LE :::~~ ' - s ) -- I \
I 50 _ -;. '-- ____ ~ ETW , -250 ° F: 1. 94 % scale 2. 81 % scale II I I, _---""' La ""' RC 16- ft: 3. 80 % scale
o
1 (,: .;~ ~ I~ ~:~:;'
0 .8 1.2 1. 6 2.0 2.4 0.0 0.4 Mach ,I / \ "':~~.
4 ~ (500 ,. mac) Figure 1. Nominal HSCT mission profile and wind tunnel capabilities ( mode l scale adjusted to test section size,
~-=---1_ Cl J [l 1'--- IF.S . 55 .825
F.S. 58.396 · ··- -l.lLr - - L.r-~) Blade
2.2% scale in the NTF is the baseline size ).
,., H .'" ro, J, 1 :~~~69135
~ ~ Tail LEs s harp
-'-.. IJ / --- F.S . 83 .060 - ··- V Figure 4. Model sketch with reference loca ti ons (linear dimensions in inches) .
Figure 2. External view of the NTF.
American Institute of Aeronautics and Astronautics AIAA-2002-0417 11 0 :? 70 : VI - Co - --;:. 60 0..
a) Front 3/4 view • Test Conditions 100 ~ ' ~~~ 2~ 0 ~~~~ 40 ~~~~ 60 ~~~~ 80 ~~~~ 100 Rn (millions) a) Mach = 0.30 • Test Conditions b) Side view 40 60 80 100 Rn (millions) c) Rear 3/4 view b) Mach = 0.95 Figure 5. 2.2% Reference H model in the NTF.
Figure 6. NTF operational envelopes with TT lines and test conditions.
American Institute of Aeronautics and Astronautics AIAA-2002-0417 LAND ING (NO STAB) l. 2 .8 TRANSON IC (NO ST A B) 0 LA DING (STAB=O DEG) TRANSON IC (STAB=O DEG) LAND ING (STAB=- IO DEG ) l.0 .6 TRAN SON IC (S TA B= -5 DEG)
&~
.8 .4 -l -l - - --
-
.6 .2 U U .4 0 -..@ V .2 - .2 0 -.4 .4 . 10 O ~ ~ --- .08 .3
i -
.06 Cl Cl .2 U U -{) .04 OD .1 .02
~dJ
~
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O
.04
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:/ ' 10 ~ ~
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-. 04 ~ ~ §
D ~ -.04
-.08 -.12 -.08 ]2 e -.....
.....l .....l 0 0 -5
~
{} 0 -10 ]6 -4 0 4 8 12 20 24 -4 0 4 8 12 a , d eg a , d eg
a) M = 0.30, Rn = 90 million b) M = 0.95, Rn = 80 million
Figure 7. Basic Longitudinal Force and Moment Data.
American Institute of Aeronautics and Astronautics AIAA-2002-0417 .12 .04 .03 a. d eg .08 a. deg ~ .02 o -OJ o -0.2 - o 8. 1 .04 o 3.8 .O ] ~ o 12 .0 o 7.8 >- t:, 1 9.8 >- - U 0 I--------~--- u
2~~
-.01
-. tr
-.04 -.02
-- ~~
-.08 -.03 ~ -.12 -.04 .04 .02 D .03 .02 . 01 .0] c U -.0] -.02 .04 ,------, .02 .03 - -r .0] - .02 (La-
.0 ] '<.)"""""""7""--; 6> 0>
6 o n ~ r& () () 0 r--~~~ ~~~~~~~~-- <Y 0OO 0 0 0 o D -.0] DO 0 --- - -.0] -.02 ~ -.03 -.04 L..L..L....l..---1--L-L-..L....l...-l..-l......l-L..L-....l..-L...l._L..L.L- -.02
- 12 -8 -4 0 4 8 12 -6 -4 -2 o 2 6
~ , d eg ~ , de g a) Landing (Vert on , Chine on), M = 0.30, Rn = 90 million b) Tr ansonic (Vert on) , M = 0.95, Rn = 80 million Figure 8. Basic LateralfDirectional Force and Moment Data.
American Institute of Aeronautics and Astronautics AIAA-2002-0417 0 .02 ---- Rn (million s) q. p f Rn (m illi on s) q. psf ...00- -.0 1 .01 4. -1 153.7 0 10 .5 1020 .
~ 0 10.0 260 .6 0 30.1 105 9.
-.02 29.9 269.9 0 30.0 175 8.
29.9 8 16.9 6 79.8 1796 .
r::,.
89.9 8 -1 2.3 -. 03 -.01 ~-- 2 2 U -.04 U -.02 ~
-.05 - -.03
Nose Up Delay -.06 1-.04 -.07 -.05 -.08 -.06 0 .2 .4 .6 .8 1.0 1. 2 -.4 -.2 0 .2 .4 .6 .8 C C L L a) M =0 .30, Landing (Stab=O deg) b) M=0.95, Transonic (Stab=O deg ) Figure 9. Rn eff e cts on pitch i ng mom ent ( corrected for st at ic aeroelastics).
20 - ~~- .50 -- 6.
6.
6. .45 1\ .40
Stab = 0 deg I ~
16 .35
I
I .30 .25
00 0
U
0 0 <1.l
::;: 12 ::- .20 - --- 6 - ~ u . 15 10 ~ "6.
LJ =-- ,-
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r
0- 0 - 8- 8
~ -- g -.----O -.05 4 -. 10 -.
o. de g ,....
-.0030 - .8 -- 4.00 0 0.40 i
-. 0035 8.00 -
0 0.55 I 0
.7 -
12.00 0.70
<>
-.0040
?? 20.00
!:,.
0.90 .6 -.0045 <1.l n Increasing Effectiveness .5
'0 .I
I ~ -.0050
°- 8- 8
., c.
{) S- --,- .4 -.0055 .r:."" 8 ~ - 6.
i
- ~ - § - e --
~ ~ -. 0060 ::;: ~ .3 I ., , u , - -.0065 .2 I -.0070 Stab = 0 deg ~ .1 -.00 75
I
~
-.0080 I 5 I 5 1 1 10 ° 10 10 Rn (million ) Rn (millions) F igure 10. Rn eff e ct s on longitudinal stability and contr ol parameters for landing co nfiguration ( corrected for static aeroelastic eff ec ts) at M=O.30 .
American In st itute of Aeronautics and Astronautics AIAA-2002-0417 14 . 50 AS Stab = 0 deg --- 10 .35 L .30 -
- 0
<> - 0
0.00 .25 0 "..) - On 0.25 :2 0 6 .20
'" ~
-- U
0.50 w . 15 -- 0 . 10 ,..,..-
2 - ---
.05 --+ <) 0 0
0 0
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-.05
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ld 0
a. d eg -.0030 .8 0.00 .... -+--r--+ -.0035 .\.00 .7 .00 -.0040 -~-
~
....
.6
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-.0045 On
~ 0
Increasing E Hec tiveness .5
<> '0
'"
-.0050 t g.
S oC ... ...
... ~ ...
-.0055 A <> ..c: ~ .0 -.0060 - :2 .. r - ....... "'"'+- .3 u -.0065 -- 0 Stab = 0 deg 0 ~ ...............
.2 -.0070 -- 0 -T .1 ;---,-~ -.0075 -.0080 0 I 5 5 I 5 5 I 2 2 10 ° 10' 10 10° 10' 10 Rn (mi ll ion) Rn (millions) Figure 11. Rn effects on longitudinal stability and control parameters for tr ansonic configuration (corrected for static aeroelastic effects) at M=O .95.
Chines OH Ch ines On .040 .040 Rn ( milli on s) .032 .032 4 .5 10.0 ...
t .024 .024 23.0 /:;.
29.9 t:" d: .016 90.7 }--- - ---------- ...............
U .0 16
.008 1- ---+-- .008 °bxOX~ ~~~ ~Dxe- -.00 -.008 LULL......LJ.-L........L..L ...l.-...L..L .L.......L.l..... '------I...LJ. .........J.. ...LL..
-4 o 4 8 12 16 20 24 -4 0 4 8 12 16 20 24
a, deg a , deg Figure 12. Rn and fo rebody chine effects on yawing moment departure for landing configurations , M=O.30, f.\ = 0 deg .
American Institute of Aeronautics and Astronautics AIAA-2002-0417 LANDING (VERT 0 , CHINES ON ) LANDING (NO VERT , CHINES ON ) .032 Rn ( milli on s) -- .024 4.4 10.0 D .016 23 .1
<>
[::, 30.0 .008 ~ 90.9 c:: u oj- - -.008 - -.016
- ~
-.024
r
-.032 --- LANDI G ( VERT 0 , NO CHINES ) LANDING (NO VERT , NO CHINES) .032 .024 .016 .008
U Ot- ----~~~N=-- -----
-.008
~~
-.016 -.024 -.032 -4
-12 -8 -4 0 4 8 12 - 12 -8 o 4 8 12
~,deg ~ , d eg Figure 13. Rn effects on yawing momen t coefficien t for various high-lift configurations , a = 12 deg , M=O .30.
American In stitute of Aeronautics and Astronautics AIAA-2002-0417 LANDING (VERT 0 ,CHINES ON) LANDING ( 0 VERT, CHINES 0 ) .032 Rn (mi llion ) .024 4.4 10.0 --+-- - -- - - - - .016 23.1
<>
b.
30.0 ~ 91.0 .008 c
u
-.008 -- ~
- ~
-.0]6 -.024 -.032 LANDING (VERT ON, NO CHINES) LANDING (NO VERT, NO CHINES) .032 - .024 .016 .008 c u -.008 -.016 -.024 -.032 -4 12
-12 -8 -4 o 4 8 12 -12 -8 o 4 8
~ , deg ~ , deg Figure 14. Rn effects on yawing moment coefficient for various high-lift configurations , a = 20 deg , M=0.30.
American Institute of Aeronautics and Astronautics --- -- -- -- ------ AIAA-2002-0417 Q. d eg u. deg 0.02 -0.13 8.00 0 .004 .004 3.92 11 .94 7.92
D. /9.80 0
.003 .003
ca
~ -8
-m
09 0
00 00 0) 0)
~ -+--
.002 .002 -0 -0 ....
0) ~ a.
-:; -......- D,D.
""- _ ...............
= c::
-§ .00 1 -
- IT- .00 1 d= U ----+ + + 0 0
6+-
-.00 1 -.00 1 I 5 5 5 I 5 I 1 2 1 2 10° ]0 10 10 10 ° 10 Rn ( milli on s) Rn ( million s) a) Landing (Vert On, Chines On), M=O . 30 b) Transonic (Vert On , No Chines) , M=O .95 Figure 15. Rn effect on directional stability (no static aeroelastic corrections) .
Rn ( milli on s) 4A LANDING (RUD=O DEG ) 10.0 L AN DING (RUD=O DEG ) 90 .7 LA DI G (RU D= O DEG) ~.4 LA DI NG (RUD=+ 30 DEG) I 1.9 deg a = 8.0 de g a= ~ 10.0 LA DI G (RUD=+ 30 DEG) D 9 1. 0 LAND ING (R D=+30 D EG ) .04 .-------- ~D .03 . 03
- @~
.02 1 --+---- .02
- ~~ -
.0 1 .O J
~
~
o r-------~~~---~ ~ ----
b..
b..
-.OJ ~ -. 01 :: uC:: u -.02 -.02 ~ -
-.03 ~ fJ -.03
-.04 1--- -.04 ~ ------ -.05 -.05 + -.06 -.06 -. 07 -.07 L...L..-'--'- _L...L...L.--'---J'--'--'--'--'---''---'---'----'-----'--.I.......L.- - ]2 4 -12 4 -8 -4 o 8 J2 -8 -4 0 8 J2 ~ , de g ~ , de g Fi gure 16. Rn effects on yawing moment coefficient for landing configurations (Chines On) withlwithout rudder deflection , M=O.30.
American In stitute of Aeronautics and Astronautics .- .. _-- ---------, AIAA-2002-0417 a= 8.0 deg a= 11. 9 deg -.0006 -.0006 ~ ~ ~ Increasing Effectiveness -.0007 -.0007 -.0008 -.0008
@
- ~
-.0009 -.0009
~ - e 8
~ -
-~ @ bo bo
~
0) 0) ~~ --+- -0 -0 -.00 10 -.0010 ...
0) i3 ~ ~ +-------T 5 ... 5 -.001 1 -.00 11 ~ deg "C >2 c: ..0 ..0 -.0012 -.0012
" C
0-
d= 0.0 U ,
-.0013 -.00 13 -
4.0 D ~ --~ -.0014 -.00 14 8.0
<>
-+ -.0015 -.00 15 ~- ~ 10.0 -.0016 -.00 16 I 5 1 5 I 5 I 5 1 2 1 2 10 ° 10 10 10 ° 10 10 Rn ( million s) Rn ( millions ) Figure 17. Rn effects on rudder effectiveness for landing (Chines On) configurations , M=O.30.
American Institute of Aeronautics and Astronautics