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Sample records for supersonic aerodynamic characteristics

  1. Aerodynamic characteristics of reentry vehicles at supersonic velocities

    NASA Astrophysics Data System (ADS)

    Adamov, N. P.; Kharitonov, A. M.; Chasovnikov, E. A.; Dyad'kin, A. A.; Kazakov, M. I.; Krylov, A. N.; Skorovarov, A. Yu.

    2015-09-01

    Models of promising reentry vehicles, experimental equipment, and test program are described. The method used to determine the total aerodynamic characteristics of these models on the AB-313 mechanical balance in the T-313 supersonic wind tunnel and the method used for simulations are presented. The aerodynamic coefficients of the examined objects in wide ranges of Mach numbers and angles of attack are obtained. The experimental data are compared with the results of simulations.

  2. Supersonic Aerodynamic Characteristics of Proposed Mars '07 Smart Lander Configurations

    NASA Technical Reports Server (NTRS)

    Murphy, Kelly J.; Horvath, Thomas J.; Erickson, Gary E.; Green, Joseph M.

    2002-01-01

    Supersonic aerodynamic data were obtained for proposed Mars '07 Smart Lander configurations in NASA Langley Research Center's Unitary Plan Wind Tunnel. The primary objective of this test program was to assess the supersonic aerodynamic characteristics of the baseline Smart Lander configuration with and without fixed shelf/tab control surfaces. Data were obtained over a Mach number range of 2.3 to 4.5, at a free stream Reynolds Number of 1 x 10(exp 6) based on body diameter. All configurations were run at angles of attack from -5 to 20 degrees and angles of sideslip of -5 to 5 degrees. These results were complemented with computational fluid dynamic (CFD) predictions to enhance the understanding of experimentally observed aerodynamic trends. Inviscid and viscous full model CFD solutions compared well with experimental results for the baseline and 3 shelf/tab configurations. Over the range tested, Mach number effects were shown to be small on vehicle aerodynamic characteristics. Based on the results from 3 different shelf/tab configurations, a fixed control surface appears to be a feasible concept for meeting aerodynamic performance metrics necessary to satisfy mission requirements.

  3. Subsonic/supersonic aerodynamic characteristics for a tactical supercruiser

    NASA Technical Reports Server (NTRS)

    Capone, F. J.; Bare, E. A.; Hollenback, D.; Hutchison, R.

    1984-01-01

    A series of cooperative NASA-Langley/Boeing experimental investigations have been conducted to determine the aeropropulsive characteristics of an advanced tactical fighter designed for supersonic cruise. These investigations were conducted in the Langley 16-Foot Transonic and Lewis 10 x 10-Foot Supersonic Wind Tunnels at Mach numbers from 0.60 to 2.47. This fighter is a Mach 2.0, 49,000 pound class vehicle that features a close-coupled canard and underwing propulsion units that utilize multifunction two-dimensional exhaust nozzles. Tests were conducted to determine the basic aerodynamic characteristics of the configuration with flow-through nacelles in which the spillage effects of representative inlets were measured. The effects of thrust-induced forces on overall aerodynamic performance were evaluated with a series of multifunction nozzles installed on air-powered nacelles. An axisymmetric nozzle configuration was also tested to obtain comparative aeropropulsive performance. Trim aerodynamic characteristics for the flow-through and powered configurations and the effect of thrust vectoring at subsonic speeds are presented.

  4. Supersonic Aerodynamic Characteristics of Blunt Body Trim Tab Configurations

    NASA Technical Reports Server (NTRS)

    Korzun, Ashley M.; Murphy, Kelly J.; Edquist, Karl T.

    2013-01-01

    Trim tabs are aerodynamic control surfaces that can allow an entry vehicle to meet aerodynamic performance requirements while reducing or eliminating the use of ballast mass and providing a capability to modulate the lift-to-drag ratio during entry. Force and moment data were obtained on 38 unique, blunt body trim tab configurations in the NASA Langley Research Center Unitary Plan Wind Tunnel. The data were used to parametrically assess the supersonic aerodynamic performance of trim tabs and to understand the influence of tab area, cant angle, and aspect ratio. Across the range of conditions tested (Mach numbers of 2.5, 3.5, and 4.5; angles of attack from -4deg to +20deg; angles of sideslip from 0deg to +8deg), the effects of varying tab area and tab cant angle were found to be much more significant than effects from varying tab aspect ratio. Aerodynamic characteristics exhibited variation with Mach number and forebody geometry over the range of conditions tested. Overall, the results demonstrate that trim tabs are a viable approach to satisfy aerodynamic performance requirements of blunt body entry vehicles with minimal ballast mass. For a 70deg sphere-cone, a tab with 3% area of the forebody and canted approximately 35deg with no ballast mass was found to give the same trim aerodynamics as a baseline model with ballast mass that was 5% of the total entry mass.

  5. Aerodynamic characteristics of supersonic fighter airplane configurations based on Soviet design concepts

    NASA Technical Reports Server (NTRS)

    Spearman, M. L.; Fournier, R. H.; Lamb, M.

    1977-01-01

    The aerodynamic, stability, and control characteristics of several supersonic fighter airplane concepts are examined. The configurations, which are based on Soviet design concepts, include fixed-wing aircraft having delta wings, swept wings, and trapezoidal wings, and a variable wing-sweep aircraft. Each concept employs aft tail controls. The concepts vary from lightweight, single-engine, air superiority, point interceptor, or ground attack types to larger twin-engine interceptor and reconnaissance designs. Analytical and experimental results indicate that careful application of the transonic or supersonic area rule can provide nearly optimum shaping for minimum drag for a specified Mach number requirement. In addition, through the proper location of components and the exploitation of interference flow fields, the concepts provide linear pitching moment characteristics, high control effectiveness, and reasonably small variations in aerodynamic center location with a resulting high potential for maneuvering capability.

  6. Supersonic aerodynamic characteristics of an advanced F-16 derivative aircraft configuration

    NASA Technical Reports Server (NTRS)

    Fox, Mike C.; Forrest, Dana K.

    1993-01-01

    A supersonic wind tunnel investigation was conducted in the NASA Langley Unitary Plan Wind Tunnel on an advanced derivative configuration of the United States Air Force F-16 fighter. Longitudinal and lateral directional force and moment data were obtained at Mach numbers of 1.60 to 2.16 to evaluate basic performance parameters and control effectiveness. The aerodynamic characteristics for the F-16 derivative model were compared with the data obtained for the F-16C model and also with a previously tested generic wing model that features an identical plan form shape and similar twist distribution.

  7. Nonaxisymmetric Body Supersonic, Aerodynamic Prediction

    DTIC Science & Technology

    1987-08-01

    wing - tail configuration are compared in Figure 27. CN comparisons are good. C. is a sensitive computation for xcp close to x’. 7.2...Analytical and Experimental Supersonic Aerodynamic Characteristics of a Forward Control Missile , AIAA Paper No. 81-0398, AIAA 19th Aerospace Sciences...body diameter. The next computational example is for a body- wing - tail configuration from Reference 32 A body-alone comparison has been made earlier in

  8. Aerodynamic characteristics of the orbital reentry vehicle experimental probe fins in a supersonic flow

    NASA Astrophysics Data System (ADS)

    Watanabe, Mitsunori; Sekine, Hideo; Tate, Atsushi; Noda, Junichi

    1994-04-01

    The aerodynamic characteristics of probe fins with a sweep angle of 60 deg, which are equipped on the Orbital Reentry Experiment (OREX) vehicle to measure the surrounding ionized gas temperature and electron number density distributions in the high temperature communication black out regions, have been measured in the supersonic wind tunnel of the National Aerospace Laboratory and compared with those of the fins of 0 deg sweep angles. Since the probes are to be embedded in the boundary layer where the local Mach number is less than 2.5 over the OREX surface at a hypersonic flight speed, the aerodynamic characteristics in supersonic regions are needed to estimate the rolling moments of fins caused by the error of the installation angles. The lift coefficient slope of the probe fins decreases as the Mach number increases, being less than the values for the 0 deg sweep fins. The drag coefficient depends highly on the sweep angle of the fins in Mach number regions less than 2.5.

  9. Supersonic Parachute Aerodynamic Testing and Fluid Structure Interaction Simulation

    NASA Astrophysics Data System (ADS)

    Lingard, J. S.; Underwood, J. C.; Darley, M. G.; Marraffa, L.; Ferracina, L.

    2014-06-01

    The ESA Supersonic Parachute program expands the knowledge of parachute inflation and flying characteristics in supersonic flows using wind tunnel testing and fluid structure interaction to develop new inflation algorithms and aerodynamic databases.

  10. A Fundamental Study for Aerodynamic Characteristics of Supersonic Biplane Wing and Wing-Body Configurations

    NASA Astrophysics Data System (ADS)

    Odaka, Yusuke; Kusunose, Kazuhiro

    In order to develop a quiet supersonic transport, it is necessary to reduce shock waves around the transport. Shock waves, in general, are the cause of the airplane's sonic boom. Authors have been studying an aerodynamic feasibility of supersonic biplanes based on the concept of the Busemann biplane. In this paper, the three dimensional effect of wing geometries on their wave drags, including wing tip effects and the interference effects between the wing and a body (Wing-Body configurations) are investigated, using CFD code in Euler (inviscid) mode. As a result, we can conclude that the supersonic biplane wings at their design Mach number (M∞=1.7) are still capable of reducing wave drag significantly similar to that of the 2-D supersonic biplane.

  11. Supersonic aerodynamics of delta wings

    NASA Technical Reports Server (NTRS)

    Wood, Richard M.

    1988-01-01

    Through the empirical correlation of experimental data and theoretical analysis, a set of graphs has been developed which summarize the inviscid aerodynamics of delta wings at supersonic speeds. The various graphs which detail the aerodynamic performance of delta wings at both zero-lift and lifting conditions were then employed to define a preliminary wing design approach in which both the low-lift and high-lift design criteria were combined to define a feasible design space.

  12. Aerodynamics of sounding rockets at supersonic speeds

    NASA Astrophysics Data System (ADS)

    Vira, N. R.

    This dissertation presents a practical and low cost method of computing the aerodynamic characteristics of vehicles such as sounding rockets, high speed bombs, projectiles and guided missiles in supersonic flight. The vehicle configuration consists of a slender axisymmetric body with a conical or ogive noise, cylinders, shoulders and boattails, if any, and have sets of two, three or four fins. Geometry of the fin cross section can be single wedge, double wedge, modified single wedge or modified double wedge. First the aerodynamics of the fins and the body are analyzed separately; then fin body and fore and aft fin interferences are accounted for when they are combined to form the total vehicle. Results and formulas documented in this work are the basis of the supersonic portion of the Theoretical Aerodynamic Derivatives (TAD) computer program operating at the NASA Goddard Space Flight Center.

  13. Aerodynamics of Supersonic Lifting Bodies

    DTIC Science & Technology

    1981-02-01

    verso of front cover. 19 Y WOROS (Continue on rt.’,;erso side i recessary and identily by block number) Theoretical Aerodynamics Lifting Bodies Wind ...waverider solution, developed from the supersonic wedge flow solution, is then i Fused to fashion vertLcal stabilizer-likh control surfaces. Wind ...served as Project Engineers ror thE wind tunnel work. Important contributions were also made bv: Mr. iis±ung Miin; Lee, -M. Beom-Soo Kim, Mtr. Martin Weeks

  14. Supersonic aerodynamic characteristics of a tail-control cruciform maneuverable missile with and without wings

    NASA Technical Reports Server (NTRS)

    Spearman, M. L.; Fournier, R. H.

    1978-01-01

    The aerodynamic characteristics for a winged and a wingless cruciform missile are examined. The body was an ogive-cylinder with a 3.5 caliber forebody; an overall length-to-diameter ratio of 11.667; and has cruciform tails that were trapexoidal in planform. Tests were made both with and without 72.9 deg cruciform delta wings. The investigation was made for Mach numbers from 1.50 to 4.63, roll attitudes of 0 and 45 deg, angles of attack from -40 to 22 deg, and tail control deflections from 10 to -40 deg. The purpose is to determine the influence of the aerodynamic behavior on the design choice for maneuverable missiles intended primarily for air-to-air or surface-to-surface missions. The results indicate that the winged missile with its more linear aerodynamic characteristics and higher lift-curve slope, should provide the highest maneuverability over a large operational range.

  15. Transonic aerodynamic characteristics of a supersonic cruise aircraft research model with the engines suspended above the wing

    NASA Technical Reports Server (NTRS)

    Mercer, C. E.; Carson, G. T., Jr.

    1979-01-01

    The influence of upper-surface nacelle exhaust flow on the aerodynamic characteristics of a supersonic cruise aircraft research configuration was investigated in a 16 foot transonic tunnel over a range of Mach numbers from 0.60 to 1.20. The arrow-wing transport configuration with engines suspended over the wing was tested at angles of attack from -4 deg to 6 deg and jet total pressure ratios from 1 to approximately 13. Wing-tip leading edge flap deflections of -10 deg to 10 deg were tested with the wing-body configuration. Various nacelle locations (chordwise, spanwise, and vertical) were tested over the ranges of Mach numbers, angles of attack, and jet total-pressure ratios. The results show that reflecting the wing-tip leading edge flap from 0 deg to -10 deg increased the maximum lift-drag ratio by 1.0 at subsonic speeds. Jet exhaust interference effects were negligible.

  16. Supersonic aerodynamic characteristics of some reentry concepts for angles of attack to 90 deg

    NASA Astrophysics Data System (ADS)

    Spearman, M. L.

    1985-11-01

    Past studies of reentry vehicles tested to high angles of attack (up to 90 deg) in the Mach number range from 2 to 4.8 are reviewed. Two basic planforms are considered: highly-swept deltas and circular. The delta concepts include variations in cross section (and thus volume) and in camber distribution. The effectiveness of various types of aerodynamic control devices is also included. The purpose of the paper is to examine the characteristics of the vehicles with a view toward the potential usefulness of such concepts in a flight regime that would include reentry from space into the atmosphere followed by a transition to sustained atmospheric flight.

  17. Effect of milling machine roughness and wing dihedral on the supersonic aerodynamic characteristics of a highly swept wing

    NASA Technical Reports Server (NTRS)

    Darden, Christine M.

    1989-01-01

    An experimental investigation was conducted to assess the effect of surface finish on the longitudinal and lateral aerodynamic characteristics of a highly-swept wing at supersonic speeds. A study of the effects of wing dihedral was also made. Included in the tests were four wing models: three models having 22.5 degrees of outboard dihedral, identical except for surface finish, and a zero-dihedral, smooth model of the same planform for reference. Of the three dihedral models, two were taken directly from the milling machine without smoothing: one having a maximum scallop height of 0.002 inches and the other a maximum scallop height of 0.005 inches. The third dihedral model was handfinished to a smooth surface. Tests were conducted in Test Section 1 of the Unitary Plan Wind Tunnel at NASA-Langley over a range of Mach numbers from 1.8 to 2.8, a range of angle of attack from -5 to 8 degrees, and at a Reynolds numbers per foot of 2 x 10(6). Selected data were also taken at a Reynolds number per foot of 6 x 10(6). Drag coefficient increases, with corresponding lift-drag ratio decreases were the primary aerodynamic effects attributed to increased surface roughness due to milling machine grooves. These drag and lift-drag ratio increments due to roughness increased as Reynolds number increased.

  18. Numerical Analysis on Aerodynamic Characteristics of Delta Wing with Variable Geometry Device in Supersonic Flow

    NASA Astrophysics Data System (ADS)

    Kanamori, Masashi; Imamura, Osamu; Suzuki, Kojiro

    The application of the variable geometry (VG) wing to a lifting re-entry body is expected to enhance the control capability of its aerodynamic characteristics and, as a result, to widen the corridor for the flight trajectory. In the present study, the flow field around a plain delta wing having three chord-wise hinges, one is on the wing root and the others on both sides of the mid-span of the wing, at Mach number 3 is numerically investigated by solving the Euler equations. The effects of the angle of attack and the “tip-down” bending angles around these hinges are clarified. The results show that the lift-to-drag ratio is hardly affected by the tip-down angle and that the overall lift and drag forces vary almost proportional to the change in the projected wing area by taking the tip-down configuration. The center of pressure moves backward by the tip-down effect.

  19. Effects of upper-surface blowing and thrust vectoring on low-speed aerodynamic characteristics of a large-scale supersonic transport model

    NASA Technical Reports Server (NTRS)

    Coe, P. L., Jr.; Mclemore, H. C.; Shivers, J. P.

    1975-01-01

    Tests were conducted in the Langley full-scale tunnel to determine the low-speed aerodynamic characteristics of a large-scale arrow-wing supersonic transport configured with engines mounted above the wing for upper surface blowing, and conventional lower surface engines with provisions for thrust vectoring. A limited number of tests were conducted for the upper surface engine configuration in the high lift condition for beta = 10 in order to evaluate lateral directional characteristics, and with the right engine inoperative to evaluate the engine out condition.

  20. Transonic and supersonic ground effect aerodynamics

    NASA Astrophysics Data System (ADS)

    Doig, G.

    2014-08-01

    A review of recent and historical work in the field of transonic and supersonic ground effect aerodynamics has been conducted, focussing on applied research on wings and aircraft, present and future ground transportation, projectiles, rocket sleds and other related bodies which travel in close ground proximity in the compressible regime. Methods for ground testing are described and evaluated, noting that wind tunnel testing is best performed with a symmetry model in the absence of a moving ground; sled or rail testing is ultimately preferable, though considerably more expensive. Findings are reported on shock-related ground influence on aerodynamic forces and moments in and accelerating through the transonic regime - where force reversals and the early onset of local supersonic flow is prevalent - as well as more predictable behaviours in fully supersonic to hypersonic ground effect flows.

  1. Supersonic aerodynamic characteristics of a variable-geometry spacecraft designed for high hypersonic performance

    NASA Technical Reports Server (NTRS)

    Spencer, B., Jr.; Fournier, R. H.

    1973-01-01

    An investigation was made in the high Mach number test section of the Langley Unitary Plan wind tunnel on a variable-geometry high hypersonic performance spacecraft concept at Mach numbers from 2.30 to 4.63. The basic lifting body is designed for hypersonic lift-drag ratio near 3.0. The variable-geometry feature is a single-pivot two-position high wing which is deployed at subsonic speeds to improve vehicle landing characteristics. For the present investigation the wing was maintained in a stowed position, and the effects of horizontal stabilizer dihedral, elevon control effectiveness, and the addition of either a conventional single vertical tail or dorsal-fin-type vertical stabilizers on the longitudinal and lateral-directional stability and control characteristics were studied.

  2. Aerodynamic Design Opportunities for Future Supersonic Aircraft

    NASA Technical Reports Server (NTRS)

    Wood, Richard M.; Bauer, Steven X. S.; Flamm, Jeffrey D.

    2002-01-01

    A discussion of a diverse set of aerodynamic opportunities to improve the aerodynamic performance of future supersonic aircraft has been presented and discussed. These ideas are offered to the community in a hope that future supersonic vehicle development activities will not be hindered by past efforts. A number of nonlinear flow based drag reduction technologies are presented and discussed. The subject technologies are related to the areas of interference flows, vehicle concepts, vortex flows, wing design, advanced control effectors, and planform design. The authors also discussed the importance of improving the aerodynamic design environment to allow creativity and knowledge greater influence. A review of all of the data presented show that pressure drag reductions on the order of 50 to 60 counts are achievable, compared to a conventional supersonic cruise vehicle, with the application of several of the discussed technologies. These drag reductions would correlate to a 30 to 40% increase in cruise L/D (lift-to-drag ratio) for a commercial supersonic transport.

  3. Supersonic aerodynamic characteristics of a series of wrap-around-fin missile configurations

    NASA Technical Reports Server (NTRS)

    Fournier, R. H.

    1977-01-01

    A parametric study of wrap-around-fin missile configurations was conducted at Mach numbers from 1.60 to 2.86 in the Langley Unitary Plan wind tunnel. The fin configurations investigated included variations in chord length, leading edge sweep, thickness ratio, and leading edge shape. The investigation also included a smooth and a stepped-down afterbody required for flush retraction of the wrap-around-fin configuration. The investigation indicated no unusual longitudinal characteristics; however, all the wrap-around-fin configurations tested indicated erratic lateral behavior, particularly in the form of induced roll at zero angle of attack and irregular variations of roll with angle of attack and Mach number. The magnitude of rolling moment at an angle of attack of 0 deg is estimated to represent approximately 0.25 deg or less roll control deflection. The stepped-down afterbody has a marked effect on reducing the induced roll.

  4. Low-speed aerodynamic characteristics from wind-tunnel tests of a large-scale advanced arrow-wing supersonic-cruise transport concept

    NASA Technical Reports Server (NTRS)

    Smith, P. M.

    1978-01-01

    Tests have been conducted to extend the existing low speed aerodynamic data base of advanced supersonic-cruise arrow wing configurations. Principle configuration variables included wing leading-edge flap deflection, wing trailing-edge flap deflection, horizontal tail effectiveness, and fuselage forebody strakes. A limited investigation was also conducted to determine the low speed aerodynamic effects due to slotted training-edge flaps. Results of this investigation demonstrate that deflecting the wing leading-edge flaps downward to suppress the wing apex vortices provides improved static longitudinal stability; however, it also results in significantly reduced static directional stability. The use of a selected fuselage forebody strakes is found to be effective in increasing the level of positive static directional stability. Drooping the fuselage nose, which is required for low-speed pilot vision, significantly improves the later-directional trim characteristics.

  5. High supersonic aerodynamic characteristics of five irregular planform wings with systematically varying wing fillet geometry tested in the NASA/LaRC 4-foot UPWT (LEG 2) (LA45A/B)

    NASA Technical Reports Server (NTRS)

    1976-01-01

    An experimental and analytical aerodynamic program to develop predesign guides for irregular planform wings is reported. The benefits are linearization of subsonic lift curve slope to high angles of attack and avoidance of subsonic pitch instabilities at high lift by proper tailoring of the planform fillet wing combination while providing the desired hypersonic trim angle and stability. The two prime areas of concern are to optimize shuttle orbiter landing and entry characteristics. Basic longitudinal aerodynamic characteristics at high supersonic speeds are developed.

  6. Post-Flight Aerodynamic and Aerothermal Model Validation of a Supersonic Inflatable Aerodynamic Decelerator

    NASA Technical Reports Server (NTRS)

    Tang, Chun; Muppidi, Suman; Bose, Deepak; Van Norman, John W.; Tanimoto, Rebekah; Clark, Ian

    2015-01-01

    NASA's Low Density Supersonic Decelerator Program is developing new technologies that will enable the landing of heavier payloads in low density environments, such as Mars. A recent flight experiment conducted high above the Hawaiian Islands has demonstrated the performance of several decelerator technologies. In particular, the deployment of the Robotic class Supersonic Inflatable Aerodynamic Decelerator (SIAD-R) was highly successful, and valuable data were collected during the test flight. This paper outlines the Computational Fluid Dynamics (CFD) analysis used to estimate the aerodynamic and aerothermal characteristics of the SIAD-R. Pre-flight and post-flight predictions are compared with the flight data, and a very good agreement in aerodynamic force and moment coefficients is observed between the CFD solutions and the reconstructed flight data.

  7. Hydrodynamic and Aerodynamic Characteristics of a Model of a Supersonic Multijet Water-Based Aircraft Equipped with Supercavitating Hydrofoils

    NASA Technical Reports Server (NTRS)

    McKann, Robert E.; Blanchard, Ulysse J.; Pearson, Albin O.

    1960-01-01

    The hydrodynamic and aerodynamic characteristics of a model of a multijet water-based Mach 2.0 aircraft equipped with hydrofoils have been determined. Takeoff stability and spray characteristics were very good, and sufficient excess thrust was available for takeoff in approximately 32 seconds and 4,700 feet at a gross weight of 225,000 pounds. Longitudinal and lateral stability during smooth-water landings were good. Lateral stability was good during rough-water landings, but forward location of the hydrofoils or added pitch damping was required to prevent diving. Hydrofoils were found to increase the aerodynamic lift-curve slope and to increase the aerodynamic drag coefficient in the transonic speed range, and the maximum lift-drag ratio decreased from 7.6 to 7.2 at the cruise Mach number of 0.9. The hydrofoils provided an increment of positive pitching moment over the Mach number range of the tests (0.6 to 1.42) and reduced the effective dihedral and directional stability.

  8. Aerodynamic characteristics at Mach numbers from 0.6 to 2.16 of a supersonic cruise fighter configuration with a design Mach number of 1.8

    NASA Technical Reports Server (NTRS)

    Shrout, B. L.

    1977-01-01

    An investigation was made in the Langley 8-foot transonic tunnel and the Langley Unitary Plan wind tunnel, over a Mach number range of 0.6 to 2.16, to determine the static longitudinal and lateral aerodynamic characteristics of a model of a supersonic-cruise fighter. The configuration, which is designed for efficient cruise at Mach number 1.8, is a twin-engine tailless arrow-wing concept with a single rectangular inlet beneath the fuselage and outboard vertical tails and ventral fins. It had untrimmed values of lift-drage ratio ranging from 10 at subsonic speeds to 6.4 at the design Mach number. The configuration was statically stable both longitudinally and laterally.

  9. Aerodynamic characteristics of a supersonic cruise airplane configuration at Mach numbers of 2.30, 2.96, and 3.30. [Langley Unitary Plan wind tunnel test

    NASA Technical Reports Server (NTRS)

    Shrout, B. L.; Fournier, R. H.

    1979-01-01

    An investigation was made in the Langley Unitary Plan wind tunnel at Mach numbers of 2.30, 2.96, and 3.30 to determine the static longitudinal and lateral aerodynamic characteristics of a model of a supersonic cruise airplane. The configuration, with a design Mach number of 3.0, has a highly swept arrow wing with tip panels of lesser sweep, a fuselage chine, outboard vertical tails, and outboard engines mounted in nacelles beneath the wings. For wind tunnel test conditions, a trimmed value above 6.0 of the maximum lift-drag ratio was obtained at the design Mach number. The configuration was statically stable, both longitudinally and laterally. Data are presented for variations of vertical-tail roll-out and toe-in and for various combinations of components. Some roll control data are shown as are data for the various sand grit sizes used in fixing the boundary layer transition location.

  10. Experimental study of aerodynamic characteristics of a reentry vehicle on a setup with free oscillations at supersonic velocities

    NASA Astrophysics Data System (ADS)

    Adamov, N. P.; Kharitonov, A. M.; Chasovnikov, E. A.; Dyad'kin, A. A.; Krylov, A. N.; Aleksandrov, E. N.

    2016-11-01

    A setup with free oscillations containing a transverse sting for holding the test model and possible test regimes are described. The method of testing and data processing is presented. Aerodynamic characteristics of the pitching moment of the model in a wide range of Mach numbers are obtained. Comparisons of quasi-steady data with numerical predictions and of damping derivatives with those obtained previously in tests of the model mounted on the base sting and with calculated results are performed. The model is found to be statically and dynamically stable except for regimes with M = 1.75 and 2.25, where nondecaying oscillations are excited.

  11. Low-Speed Aerodynamic and Hydrodynamic Characteristics of a Proposed Supersonic Multijet Water-Based Hydro-Ski Aircraft with Upward-Rotating Engines

    NASA Technical Reports Server (NTRS)

    Petynia, William W.; Croom, Delwin R.; Davenport, Edwin E.

    1958-01-01

    The low-speed aerodynamic and hydrodynamic characteristics of a proposed multijet water-based aircraft configuration for supersonic operation have been investigated. The design features include upward-rotating engines, body indentation, a single hydro-ski, and a wing with an aspect ratio of 3.0, a taper ratio of 0.143, 36.90 sweepback of the quarter-chord line, and NACA 65AO04 airfoil sections. For the aerodynamic investigation, with the flaps retracted, the model was longitudinally and directionally stable up to the stall. The all-movable horizontal tail was capable of trimming the model up to a lift coefficient of approximately 0.87. All flap configurations investigated had a tendency to become longitudinally unstable at stall. The effectiveness of the all-movable horizontal tail increased with increasing lift coefficient for all flap configurations investigated; however, with the large static margin of the configuration with the center of gravity at 0.25 mean aerodynamic chord, the all-movable horizontal tail was not powerful enough to trim all the various flapped configurations investigated throughout the angle-of-attack range. For the hydrodynamic investigation, longitudinal stability during take-offs and landings was satisfactory. Decreasing the area of the hydro-ski 60 percent increased the maximum resistance and emergence speed 40 and 70 percent, respectively. Without the jet exhaust, the resistance was reduced by simulating the vertical-lift component of the forward engines rotated upward. However, the jet exhaust of the forward engines increased the maximum resistance approximately 60 percent. The engine inlets and horizontal tail were free from spray for all loads investigated and for both hydro-ski sizes.

  12. Aerodynamic detuning analysis of an unstalled supersonic turbofan cascade

    NASA Technical Reports Server (NTRS)

    Hoyniak, D.; Fleeter, S.

    1985-01-01

    An approach to passive flutter control is aerodynamic detuning, defined as designed passage-to-passage differences in the unsteady aerodynamic flow field of a rotor blade row. Thus, aerodynamic detuning directly affects the fundamental driving mechanism for flutter. A model to demonstrate the enhanced supersonic aeroelastic stability associated with aerodynamic detuning is developed. The stability of an aerodynamically detuned cascade operating in a supersonic inlet flow field with a subsonic leading edge locus is analyzed, with the aerodynamic detuning accomplished by means of nonuniform circumferential spacing of adjacent rotor blades. The unsteady aerodynamic forces and moments on the blading are defined in terms of influence coefficients in a manner that permits the stability of both a conventional uniformally spaced rotor configuration as well as the detuned nonuniform circumferentially spaced rotor to be determined. With Verdon's uniformly spaced Cascade B as a baseline, this analysis is then utilized to demonstrate the potential enhanced aeroelastic stability associated with this particular type of aerodynamic detuning.

  13. Supersonic Flight Dynamics Test 2: Trajectory, Atmosphere, and Aerodynamics Reconstruction

    NASA Technical Reports Server (NTRS)

    Karlgaard, Christopher D.; O'Farrell, Clara; Ginn, Jason M.; Van Norman, John W.

    2016-01-01

    The Supersonic Flight Dynamics Test is a full-scale flight test of aerodynamic decelerator technologies developed by the Low Density Supersonic Decelerator technology demonstration project. The purpose of the project is to develop and mature aerodynamic decelerator technologies for landing large-mass payloads on the surface of Mars. The technologies include a Supersonic Inflatable Aerodynamic Decelerator and supersonic parachutes. The first Supersonic Flight Dynamics Test occurred on June 28th, 2014 at the Pacific Missile Range Facility. The purpose of this test was to validate the test architecture for future tests. The flight was a success and, in addition, was able to acquire data on the aerodynamic performance of the supersonic inflatable decelerator. The Supersonic Disksail parachute developed a tear during deployment. The second flight test occurred on June 8th, 2015, and incorporated a Supersonic Ringsail parachute which was redesigned based on data from the first flight. Again, the inflatable decelerator functioned as predicted but the parachute was damaged during deployment. This paper describes the instrumentation, analysis techniques, and acquired flight test data utilized to reconstruct the vehicle trajectory, main motor thrust, atmosphere, and aerodynamics.

  14. Supersonic Flight Dynamics Test: Trajectory, Atmosphere, and Aerodynamics Reconstruction

    NASA Technical Reports Server (NTRS)

    Kutty, Prasad; Karlgaard, Christopher D.; Blood, Eric M.; O'Farrell, Clara; Ginn, Jason M.; Shoenenberger, Mark; Dutta, Soumyo

    2015-01-01

    The Supersonic Flight Dynamics Test is a full-scale flight test of a Supersonic Inflatable Aerodynamic Decelerator, which is part of the Low Density Supersonic Decelerator technology development project. The purpose of the project is to develop and mature aerodynamic decelerator technologies for landing large mass payloads on the surface of Mars. The technologies include a Supersonic Inflatable Aerodynamic Decelerator and Supersonic Parachutes. The first Supersonic Flight Dynamics Test occurred on June 28th, 2014 at the Pacific Missile Range Facility. This test was used to validate the test architecture for future missions. The flight was a success and, in addition, was able to acquire data on the aerodynamic performance of the supersonic inflatable decelerator. This paper describes the instrumentation, analysis techniques, and acquired flight test data utilized to reconstruct the vehicle trajectory, atmosphere, and aerodynamics. The results of the reconstruction show significantly higher lofting of the trajectory, which can partially be explained by off-nominal booster motor performance. The reconstructed vehicle force and moment coefficients fall well within pre-flight predictions. A parameter identification analysis indicates that the vehicle displayed greater aerodynamic static stability than seen in pre-flight computational predictions and ballistic range tests.

  15. Nonlinear potential analysis techniques for supersonic-hypersonic aerodynamic design

    NASA Technical Reports Server (NTRS)

    Shankar, V.; Clever, W. C.

    1984-01-01

    Approximate nonlinear inviscid theoretical techniques for predicting aerodynamic characteristics and surface pressures for relatively slender vehicles at supersonic and moderate hypersonic speeds were developed. Emphasis was placed on approaches that would be responsive to conceptual configuration design level of effort. Second order small disturbance and full potential theory was utilized to meet this objective. Numerical codes were developed for relatively general three dimensional geometries to evaluate the capability of the approximate equations of motion considered. Results from the computations indicate good agreement with experimental results for a variety of wing, body, and wing-body shapes.

  16. Supersonic/hypersonic aerodynamic methods for aircraft design and analysis

    NASA Technical Reports Server (NTRS)

    Torres, Abel O.

    1992-01-01

    A methodology employed in engineering codes to predict aerodynamic characteristics over arbitrary supersonic/hypersonic configurations is considered. Engineering codes use a combination of simplified methods, based on geometrical impact angle and freestream conditions, to compute pressure distribution over the vehicle's surface in an efficient and timely manner. These approximate methods are valid for both hypersonic (Mach greater than 4) and lower speeds (Mach down to 2). It is concluded that the proposed methodology enables the user to obtain reasonable estimates of vehicle performance and engineering methods are valuable in the design process of these type of vehicles.

  17. Experimental Study on Control of Supersonic Aerodynamic Characteristics of a Vertical Landing Rocket by Using Opposing Jet

    NASA Astrophysics Data System (ADS)

    Akita, Daisuke; Yamada, Kazuhiko; Suzuki, Kojiro

    In order to control a reentry trajectory of a vertical landing rocket, an opposing jet system is experimentally tested at supersonic speeds. Experiments are conducted at Mach number 4.0 in the supersonic wind tunnel of ISAS. Supersonic nozzles of the exit Mach number 2.4 are installed on a blunt nose of a rocket model. The most significant drag reduction due to the jet-spike effect is obtained when the jet is injected from the stagnation point of the body in the opposite direction to the freestream. Various types of the nozzle arrangement are investigated and the method to increase the L/D is discussed. It is found that the off-axis jet is effective both to increase the L/D and to enable a vehicle to be trimmed at an intended attack angle.

  18. Wind-Tunnel Investigation of Subsonic Longitudinal Aerodynamic Characteristics of a Tiltable-Wing Vertical-Take-Off-and-Landing Supersonic Bomber Configuration Including Turbojet Power Effects

    NASA Technical Reports Server (NTRS)

    Thompson, Robert F.; Vogler, Raymond D.; Moseley, William C., Jr.

    1959-01-01

    Jet-powered model tests were made to determine the low-speed longitudinal aerodynamic characteristics of a vertical-take-off and-landing supersonic bomber configuration. The configuration has an unique engine-wing arrangement wherein six large turbojet engines (three on each side of the fuselage) are buried in a low-aspect-ratio wing which is tilted into the vertical plane for take-off. An essentially two-dimensional variable inlet, spanning the leading edge of each wing semispan, provides air for the engines. Jet flow conditions were simulated for a range of military (nonafterburner) and afterburner turbojet-powered flight at subsonic speeds. Three horizontal tails were tested at a station down-stream of the jet exit and at three heights above the jet axes. A semi-span model was used and test parameters covered wing-fuselage incidence angles from 0 deg to 15 deg, wing angles of attack from -4 deg to 36 deg, a variable range of horizontal-tail incidence angles, and some variations in power simulation conditions. Results show that, with all horizontal tails tested, there were large variations in static stability throughout the lift range. When the wing and fuselage were alined, the model was statically stable throughout the test range only with the largest tail tested (tail span of 1.25 wing span) and only when the tail was located in the low test position which placed the tail nearest to the undeflected jet. For transition flight conditions, none of the tail configurations provided satisfactory longitudinal stability or trim throughout the lift range. Jet flow was destabilizing for most of the test conditions, and varying the jet-exit flow conditions at a constant thrust coefficient had little effect on the stability of this model. Wing leading-edge simulation had some important effects on the longitudinal aerodynamic characteristics.

  19. Study of changes in the aerodynamic characteristics of the axisymmetric supersonic vehicle in case of gas blowing from the lateral surface

    NASA Astrophysics Data System (ADS)

    Kislovskiy, V. A.; Zvegintsev, V. I.

    2016-10-01

    Consider the flow around the axisymmetric supersonic vehicle with the use of gas jet blowing from the lateral surface. The blowing is made in series of points at different distances from the nose fairing. The aim of the work was to determine the changes in aerodynamic forces and the formation of the moment, when jet of gas blowing in different parts of the supersonic vehicle. The study was conducted by numerical modeling of different cases of injection. As a result, data were obtained which showed the degree of influence not only jet thrust from the jet flow, but same the impact of the redistribution of the flow by body surface on the formation of aerodynamic forces and moments.

  20. Supersonic Aerodynamic Characteristics of a Low-Drag Aircraft Configuration having an Arrow Wing of Aspect Ratio 1.86 and a Body of Fineness Ratio 20

    NASA Technical Reports Server (NTRS)

    Gillespie, Warren, Jr.

    1960-01-01

    A free-flight rocket-propelled-model investigation was conducted at Mach numbers of 1.2 to 1.9 to determine the longitudinal and lateral aero-dynamic characteristics of a low-drag aircraft configuration. The model consisted of an aspect-ratio -1.86 arrow wing with 67.5 deg. leading-edge sweep and NACA 65A004 airfoil section and a triangular vertical tail with 60 deg. sweep and NACA 65A003 section in combination with a body of fineness ratio 20. Aerodynamic data in pitch, yaw, and roll were obtained from transient motions induced by small pulse rockets firing at intervals in the pitch and yaw directions. From the results of this brief aerodynamic investigation, it is observed that very slender body shapes can provide increased volumetric capacity with little or no increase in zero-lift drag and that body fineness ratios of the order of 20 should be considered in the design of long-range supersonic aircraft. The zero-lift drag and the drag-due-to-lift parameter of the test configuration varied linearly with Mach number. The maximum lift-drag ratio was 7.0 at a Mach number of 1.25 and decreased slightly to a value of 6.6 at a Mach number of 1.81. The optimum lift coefficient, normal-force-curve slope, lateral-force-curve slope, static stability in pitch and yaw, time to damp to one-half amplitude in pitch and yaw, the sum of the rotary damping derivatives in pitch and also in yaw, and the static rolling derivatives all decreased with an increase in Mach number. Values of certain rolling derivatives were obtained by application of the least-squares method to the differential equation of rolling motion. A comparison of the experimental and calculated total rolling-moment-coefficient variation during transient oscillations of the model indicated good agreement when the damping-in-roll contribution was included with the static rolling-moment terms.

  1. Some effects of wing and body geometry on the aerodynamic characteristics of configurations designed for high supersonic Mach numbers

    NASA Technical Reports Server (NTRS)

    Spearman, M. L.; Tice, David C.; Braswell, Dorothy O.

    1992-01-01

    Experimental and theoretical results are presented for a family of aerodynamic configurations for flight Mach numbers as high as Mach 8. All of these generic configurations involved 70-deg sweep delta planform wings of three different areas and three fuselage shapes with circular-to-elliptical cross sections. It is noted that fuselage ellipticity enhances lift-curve slope and maximum L/D, while decreasing static longitudinal stability (especially with smaller wing areas).

  2. Aerodynamic Models for the Low Density Supersonic Declerator (LDSD) Supersonic Flight Dynamics Test (SFDT)

    NASA Technical Reports Server (NTRS)

    Van Norman, John W.; Dyakonov, Artem; Schoenenberger, Mark; Davis, Jody; Muppidi, Suman; Tang, Chun; Bose, Deepak; Mobley, Brandon; Clark, Ian

    2015-01-01

    An overview of pre-flight aerodynamic models for the Low Density Supersonic Decelerator (LDSD) Supersonic Flight Dynamics Test (SFDT) campaign is presented, with comparisons to reconstructed flight data and discussion of model updates. The SFDT campaign objective is to test Supersonic Inflatable Aerodynamic Decelerator (SIAD) and large supersonic parachute technologies at high altitude Earth conditions relevant to entry, descent, and landing (EDL) at Mars. Nominal SIAD test conditions are attained by lifting a test vehicle (TV) to 36 km altitude with a large helium balloon, then accelerating the TV to Mach 4 and and 53 km altitude with a solid rocket motor. The first flight test (SFDT-1) delivered a 6 meter diameter robotic mission class decelerator (SIAD-R) to several seconds of flight on June 28, 2014, and was successful in demonstrating the SFDT flight system concept and SIAD-R. The trajectory was off-nominal, however, lofting to over 8 km higher than predicted in flight simulations. Comparisons between reconstructed flight data and aerodynamic models show that SIAD-R aerodynamic performance was in good agreement with pre-flight predictions. Similar comparisons of powered ascent phase aerodynamics show that the pre-flight model overpredicted TV pitch stability, leading to underprediction of trajectory peak altitude. Comparisons between pre-flight aerodynamic models and reconstructed flight data are shown, and changes to aerodynamic models using improved fidelity and knowledge gained from SFDT-1 are discussed.

  3. Extreme shockwave systems in problems of external supersonic aerodynamics

    NASA Astrophysics Data System (ADS)

    Uskov, V. N.; Chernyshov, M. V.

    2014-01-01

    The stationary shockwave systems (the sequences of shocks, isentropic expansion and compression waves), which arise at a planar supersonic flow of perfect inviscid gas around the bodies are investigated theoretically. The domains of the existence of shockwave systems under consideration are found analytically and numerically for the model problems of supersonic aerodynamics (the flow around a single plate, the plate with the frontal shield, polygonal profiles), the parameters of systems are determined, which provide the extrema of the force and thermal loadings as well as of the aerodynamic coefficients of streamlined bodies.

  4. Aerodynamic Characteristics of the Close-Coupled Canard as Applied to Low-to-Moderate Swept Wings. Volume 3. Transonic-Supersonic Speed Regime

    DTIC Science & Technology

    1979-12-01

    Volume 3: Transonic-Supersonic Speed Regime; and Volume 4: F-4 Phantom II Aircraft. N ., ,..• .. c-", i. iii Im TABLE OF CONTENTS 4. Page LIST OF...numbers. Stability characteristics in the form of center of pressure, CM/CN, neutral point DCM !/CN and pif:ching moment slope CM are presented in

  5. A Wind-Tunnel Investigation of the Aerodynamic Characteristics of a Full-Scale Supersonic-Type Three-blade Propeller at Mach Numbers to 0.96

    NASA Technical Reports Server (NTRS)

    Evans, Albert J; Liner, George

    1958-01-01

    An investigation of the characteristics of a full-scale supersonic-type propeller has been made in the Langley 16-foot transonic tunnel with the 6000-horsepower propeller dynamometer. The tests covered a range of blade angles from 20.2 degrees to 60.2 degrees at forward Mach numbers up to 0.96. The results showed that envelope efficiency at an advance ratio of 2.8 decreased from 86 percent to 72 percent when the forward Mach number was increased from 0.70 to 0.96. A comparison of the experimental results with calculated results showed that maximum propeller efficiency can be calculated with good accuracy by using ordinary subsonic strip theory when the blade-section speeds are supersonic. The investigation also showed favorable power-absorption properties of the supersonic-type propeller at high speeds.

  6. Experimental aerodynamic characteristics at Mach numbers from 0.60 to 2.70 of two supersonic cruise fighter configurations

    NASA Technical Reports Server (NTRS)

    Dollyhigh, S. M.

    1979-01-01

    Two 0.085-scale full span wind-tunnel models of a Mach 1.60 design supercruiser configuration were tested at Mach numbers from 0.60 to 2.70. One model incorporated a varying dihedral (swept-up) wing to obtain the desired lateral-directional characteristics; the other incorporated more conventional twin vertical tails. The data from the wind-tunnel tests are presented without analysis.

  7. The aerodynamic characteristics of a supersonic aircraft configuration with a 40 degree sweptback wing through a Mach number range from 0 to 2.4 obtained from various sources

    NASA Technical Reports Server (NTRS)

    Spearman, M Leroy; Robinson, Ross B

    1952-01-01

    A summary and analysis have been made of the results of various investigations to determine the aerodynamic characteristics of a supersonic aircraft configuration. The configuration has a wing with 40 degree sweepback at the quarter-chord line, aspect ratio 4, taper ratio 0.5, and 10-percent-thick circular-arc sections normal to the quarter-chord line. Experimental data were available for a Mach number range from 0.16 to 2.32. Results obtained from wing-flow, rocket-model, transonic-bump, and tunnel tests are presented and, where possible, are supplemented by empirical and theoretical calculations.

  8. Three-dimensional aerodynamic shape optimization of supersonic delta wings

    NASA Technical Reports Server (NTRS)

    Burgreen, Greg W.; Baysal, Oktay

    1994-01-01

    A recently developed three-dimensional aerodynamic shape optimization procedure AeSOP(sub 3D) is described. This procedure incorporates some of the most promising concepts from the area of computational aerodynamic analysis and design, specifically, discrete sensitivity analysis, a fully implicit 3D Computational Fluid Dynamics (CFD) methodology, and 3D Bezier-Bernstein surface parameterizations. The new procedure is demonstrated in the preliminary design of supersonic delta wings. Starting from a symmetric clipped delta wing geometry, a Mach 1.62 asymmetric delta wing and two Mach 1. 5 cranked delta wings were designed subject to various aerodynamic and geometric constraints.

  9. Transonic Aerodynamic Characteristics of a Model of a Proposed Six-Engine Hull-Type Seaplane Designed for Supersonic Flight

    NASA Technical Reports Server (NTRS)

    Wornom, Dewey E.

    1960-01-01

    Force tests of a model of a proposed six-engine hull-type seaplane were performed in the Langley 8-foot transonic pressure tunnel. The results of these tests have indicated that the model had a subsonic zero-lift drag coefficient of 0.0240 with the highest zero-lift drag coefficient slightly greater than twice the subsonic drag level. Pitchup tendencies were noted for subsonic Mach numbers at relatively high lift coefficients. Wing leading-edge droop increased the maximum lift-drag ratio approximately 8 percent at a Mach number of 0.80 but this effect was negligible at a Mach number of 0.90 and above. The configuration exhibited stable lateral characteristics over the test Mach number range.

  10. Supersonic aerodynamic characteristics of a Sparrow 3 type missile model with wing controls and comparison with existing tail-control results

    NASA Technical Reports Server (NTRS)

    Monta, W. J.

    1977-01-01

    An experimental investigation was conducted on a model of a wing control version of the Sparrow III type missile to determine the static aerodynamic characteristics over an angle of attack range from 0 deg to 40 deg for Mach numbers from 1.50 to 4.60.

  11. Supersonic aerodynamic characteristics of a lifting-body orbiter model with a blunted delta planform at Mach 2.30 to 4.60

    NASA Technical Reports Server (NTRS)

    Blair, A. B., Jr.

    1972-01-01

    An investigation has been made in the Langley Unitary Plan wind tunnel to determine the aerodynamic characteristics of a lifting-body orbiter model with a blunted delta planform. The model was tested at Mach numbers from 2.30 to 4.60, at nominal angles of attack from -4 deg to 60 deg and angles of sideslip from -4 deg to 10 deg, and at a Reynolds number of 2.5 million per foot.

  12. Supersonic aerodynamic characteristics of a low-aspect-ratio missile model with wing and tail controls and with tails in line and interdigitated

    NASA Technical Reports Server (NTRS)

    Graves, E. B.

    1972-01-01

    A study has been made to determine the aerodynamic characteristics of a low-aspect ratio cruciform missile model with all-movable wings and tails. The configuration was tested at Mach numbers from 1.50 to 4.63 with the wings in the vertical and horizontal planes and with the wings in a 45 deg roll plane with tails in line and interdigitated.

  13. Subsonic and supersonic indicial aerodynamics and aerodynamic transfer function for complex configurations. [aerodynamic configurations for subsonic and supersonic speeds using the finite element method

    NASA Technical Reports Server (NTRS)

    Morino, L.

    1974-01-01

    A general theory for indicial-potential-compressible aerodynamics around complex configurations is presented. The motion is assumed to consist of constant subsonic or supersonic speed (steady state) and small perturbations around the steady state. Using the finite-element method to discretize the space problem, a set of differential-difference equations in time relating the potential to its normal derivative on the surface of the body was obtained. The aerodynamics transfer function was derived by using standard method of operational calculus.

  14. Forced response analysis of an aerodynamically detuned supersonic turbomachine rotor

    NASA Technical Reports Server (NTRS)

    Hoyniak, D.; Fleeter, S.

    1985-01-01

    High performance aircraft-engine fan and compressor blades are vulnerable to aerodynamically forced vibrations generated by inlet flow distortions due to wakes from upstream blade and vane rows, atmospheric gusts, and maldistributions in inlet ducts. In this report, an analysis is developed to predict the flow-induced forced response of an aerodynamically detuned rotor operating in a supersonic flow with a subsonic axial component. The aerodynamic detuning is achieved by alternating the circumferential spacing of adjacent rotor blades. The total unsteady aerodynamic loading acting on the blading, as a result of the convection of the transverse gust past the airfoil cascade and the resulting motion of the cascade, is developed in terms of influence coefficients. This analysis is used to investigate the effect of aerodynamic detuning on the forced response of a 12-blade rotor, with Verdon's Cascade B flow geometry as a uniformly spaced baseline configuration. The results of this study indicate that, for forward traveling wave gust excitations, aerodynamic detuning is very beneficial, resulting in significantly decreased maximum-amplitude blade responses for many interblade phase angles.

  15. Wind-tunnel/flight correlation study of aerodynamic characteristics of a large flexible supersonic cruise airplane (XB-70-1). 3: A comparison between characteristics predicted from wind-tunnel measurements and those measured in flight

    NASA Technical Reports Server (NTRS)

    Arnaiz, H. H.; Peterson, J. B., Jr.; Daugherty, J. C.

    1980-01-01

    A program was undertaken by NASA to evaluate the accuracy of a method for predicting the aerodynamic characteristics of large supersonic cruise airplanes. This program compared predicted and flight-measured lift, drag, angle of attack, and control surface deflection for the XB-70-1 airplane for 14 flight conditions with a Mach number range from 0.76 to 2.56. The predictions were derived from the wind-tunnel test data of a 0.03-scale model of the XB-70-1 airplane fabricated to represent the aeroelastically deformed shape at a 2.5 Mach number cruise condition. Corrections for shape variations at the other Mach numbers were included in the prediction. For most cases, differences between predicted and measured values were within the accuracy of the comparison. However, there were significant differences at transonic Mach numbers. At a Mach number of 1.06 differences were as large as 27 percent in the drag coefficients and 20 deg in the elevator deflections. A brief analysis indicated that a significant part of the difference between drag coefficients was due to the incorrect prediction of the control surface deflection required to trim the airplane.

  16. Overview of Supersonic Aerodynamics Measurement Techniques in the NASA Langley Unitary Plan Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.

    2007-01-01

    An overview is given of selected measurement techniques used in the NASA Langley Research Center (NASA LaRC) Unitary Plan Wind Tunnel (UPWT) to determine the aerodynamic characteristics of aerospace vehicles operating at supersonic speeds. A broad definition of a measurement technique is adopted in this paper and is any qualitative or quantitative experimental approach that provides information leading to the improved understanding of the supersonic aerodynamic characteristics. On-surface and off-surface measurement techniques used to obtain discrete (point) and global (field) measurements and planar and global flow visualizations are described, and examples of all methods are included. The discussion is limited to recent experiences in the UPWT and is, therefore, not an exhaustive review of existing experimental techniques. The diversity and high quality of the measurement techniques and the resultant data illustrate the capabilities of a ground-based experimental facility and the key role that it plays in the advancement of our understanding, prediction, and control of supersonic aerodynamics.

  17. Extended mapping and characteristics techniques for inverse aerodynamic design

    NASA Technical Reports Server (NTRS)

    Sobieczky, H.; Qian, Y. J.

    1991-01-01

    Some ideas for using hodograph theory, mapping techniques and methods of characteristics to formulate typical aerodynamic design boundary value problems are developed. The inverse method of characteristics is shown to be a fast tool for design of transonic flow elements as well as supersonic flows with given shock waves.

  18. Some unique characteristics of supersonic cruise vehicles and their effect on airport community noise

    NASA Technical Reports Server (NTRS)

    Driver, C.; Maglieri, D. J.

    1980-01-01

    The paper examines the differences between the supersonic and subsonic commercial aircraft in terms of their configuration, aerodynamic characteristics, propulsion systems, and the manner of operation. The unique characteristics of supersonic cruise vehicles should provide improved airport-community noise exposures if the vehicle is permitted to operate at its most efficient and effective flight modes. It is concluded that noise exposure levels for supersonic cruise vehicles can be comparable to those of its equivalent subsonic counterpart of that time period.

  19. Aerodynamic and performance characterization of supersonic retropropulsion for application to planetary entry and descent

    NASA Astrophysics Data System (ADS)

    Korzun, Ashley M.

    shock layer of a blunt body in supersonic flow. Although numerous wind tunnel tests of relevance to SRP have been conducted, the scope of the work is limited in the freestream conditions and composition, retropropulsion conditions and composition, and configurations and geometries explored. The SRP aerodynamic - propulsive interaction alters the aerodynamic characteristics of the vehicle, and models must be developed that accurately represent the impact of SRP on system mass and performance. Work within this thesis has defined and advanced the state of the art for supersonic retropropulsion. This has been achieved through the application of systems analysis, computational analysis, and analytical methods. The contributions of this thesis include a detailed performance analysis and exploration of the design space specific to supersonic retropropulsion, establishment of the relationship between vehicle performance and the aerodynamic - propulsive interaction, and an assessment of the required fidelity and computational cost in simulating supersonic retropropulsion flowfields, with emphasis on the effort required to develop aerodynamic databases for conceptual design.

  20. The aerodynamic design of the oblique flying wing supersonic transport

    NASA Technical Reports Server (NTRS)

    Vandervelden, Alexander J. M.; Kroo, Ilan

    1990-01-01

    The aerodynamic design of a supersonic oblique flying wing is strongly influenced by the requirement that passengers must be accommodated inside the wing. It was revealed that thick oblique wings of very high sweep angle can be efficient at supersonic speeds when transonic normal Mach numbers are allowed on the upper surface of the wing. The goals were motivated by the ability to design a maximum thickness, minimum size oblique flying wing. A 2-D Navier-Stokes solver was used to design airfoils up to 16 percent thickness with specified lift, drag and pitching moment. A new method was developed to calculate the required pressure distribution on the wing based on the airfoil loading, normal Mach number distribution and theoretical knowledge of the minimum drag of oblique configurations at supersonic speeds. The wing mean surface for this pressure distribution was calculated using an inverse potential flow solver. The lift to drag ratio of this wing was significantly higher than that of a comparable delta wing for cruise speeds up to Mach 2.

  1. Calculation of subsonic and supersonic steady and unsteady aerodynamic forces using velocity potential aerodynamic elements

    NASA Technical Reports Server (NTRS)

    Haviland, J. K.; Yoo, Y. S.

    1976-01-01

    Expressions for calculation of subsonic and supersonic, steady and unsteady aerodynamic forces are derived, using the concept of aerodynamic elements applied to the downwash velocity potential method. Aerodynamic elements can be of arbitrary out of plane polygon shape, although numerical calculations are restricted to rectangular elements, and to the steady state case in the supersonic examples. It is suggested that the use of conforming, in place of rectangular elements, would give better results. Agreement with results for subsonic oscillating T tails is fair, but results do not converge as the number of collocation points is increased. This appears to be due to the form of expression used in the calculations. The methods derived are expected to facilitate automated flutter analysis on the computer. In particular, the aerodynamic element concept is consistent with finite element methods already used for structural analysis. The method is universal for the complete Mach number range, and, finally, the calculations can be arranged so that they do not have to be repeated completely for every reduced frequency.

  2. Effect of conventional and square stores on the longitudinal aerodynamic characteristics of a fighter aircraft model at supersonic speeds. [in the langley unitary plan wind tunnel

    NASA Technical Reports Server (NTRS)

    Monta, W. J.

    1980-01-01

    The effects of conventional and square stores on the longitudinal aerodynamic characteristics of a fighter aircraft configuration at Mach numbers of 1.6, 1.8, and 2.0 was investigated. Five conventional store configurations and six arrangements of a square store configuration were studied. All configurations of the stores produced small, positive increments in the pitching moment throughout the angle-of-attack range, but the configuration with area ruled wing tanks also had a slight decrease on stability at the higher angles of attack. There were some small changes in lift coefficient because of the addition of the stores, causing the drag increment to vary with the lift coefficient. As a result, there were corresponding changes in the increments of the maximum lift drag ratios. The store drag coefficient based on the cross sectional area of the stores ranged from a maximum of 1.1 for the configuration with three Maverick missiles to a minimum of about .040 for the two MK-84 bombs and the arrangements with four square stores touching or two square stores in tandem. Square stores located side by side yielded about 0.50 in the aft position compared to 0.74 in the forward position.

  3. Vacuum chamber with a supersonic-flow aerodynamic window

    DOEpatents

    Hanson, C.L.

    1980-10-14

    A supersonic flow aerodynamic window is disclosed whereby a steam ejector situated in a primary chamber at vacuum exhausts superheated steam toward an orifice to a region of higher pressure, creating a barrier to the gas in the region of higher pressure which attempts to enter through the orifice. In a mixing chamber outside and in fluid communication with the primary chamber, superheated steam and gas are combined into a mixture which then enters the primary chamber through the orifice. At the point of impact of the ejector/superheated steam and the incoming gas/superheated steam mixture, a barrier is created to the gas attempting to enter the ejector chamber. This barrier, coupled with suitable vacuum pumping means and cooling means, serves to keep the steam ejector and primary chamber at a negative pressure, even though the primary chamber has an orifice to a region of higher pressure.

  4. Vacuum chamber with a supersonic flow aerodynamic window

    DOEpatents

    Hanson, Clark L.

    1982-01-01

    A supersonic flow aerodynamic window, whereby a steam ejector situated in a primary chamber at vacuum exhausts superheated steam toward an orifice to a region of higher pressure, creating a barrier to the gas in the region of higher pressure which attempts to enter through the orifice. In a mixing chamber outside and in fluid communication with the primary chamber, superheated steam and gas are combined into a mixture which then enters the primary chamber through the orifice. At the point of impact of the ejector/superheated steam and the incoming gas/superheated steam mixture, a barrier is created to the gas attempting to enter the ejector chamber. This barrier, coupled with suitable vacuum pumping means and cooling means, serves to keep the steam ejector and primary chamber at a negative pressure, even though the primary chamber has an orifice to a region of higher pressure.

  5. Rocket Sled Propelled Testing of a Supersonic Inflatable Aerodynamic Decelerator

    NASA Technical Reports Server (NTRS)

    Meacham, Michael B.; Kennett, Andrew; Townsend, Derik J.; Marti, Benjamin

    2013-01-01

    Decelerators (IADs) have traditionally been tested in wind tunnels. As the limitations of these test facilities are reached, other avenues must be pursued. The IAD being tested is a Supersonic IAD (SIAD), which attaches just aft of the heatshield around the perimeter of an entry body. This 'attached torus' SIAD is meant to improve the accuracy of landing for robotic class missions to Mars and allow for potentially increased payloads. The SIAD Design Verification (SDV) test aims to qualify the SIAD by applying a targeted aerodynamic load to the vehicle. While many test architectures were researched, a rocket sled track was ultimately chosen to be the most cost effective way to achieve the desired dynamic pressures. The Supersonic Naval Ordnance Research Track (SNORT) at the Naval Air Warfare Center Weapons Division (NAWCWD) China Lake is a four mile test track, traditionally used for warhead and ejection seat testing. Prior to SDV, inflatable drag bodies have been tested on this particular track. Teams at Jet Propulsion Laboratory (JPL) and NAWCWD collaborate together to design and fabricate one of the largest sleds ever built. The SDV sled is comprised of three individual sleds: a Pusher Sled which holds the solid booster rockets, an Item Sled which supports the test vehicle, and a Camera Sled that is pushed in front for in-situ footage and measurements. The JPL-designed Test Vehicle has a full-scale heatshield shape and contains all instrumentation and inflation systems necessary to inflate and test a SIAD. The first campaign that is run at SNORT tested all hardware and instrumentation before the SIAD was ready to be tested. For each of the three tests in this campaign, the number of rockets and top speed was increased and the data analyzed to ensure the hardware is safe at the necessary accelerations and aerodynamic loads.

  6. Handbook of Supersonic Aerodynamics. Section 8. Bodies of Revolution

    DTIC Science & Technology

    1961-10-01

    degree of accuracy the characteristics method requires much more time. Heaviside Operational Method (Ref. 78).--The Heaviside opera- tor is applied to...Supersonic Blunt-Body Flows. JPL/CIT Progress Report No. 20-372, February 1959. 88. Oliver , R. E. An Experimental Investigation of Flow over Simple Blunt...272 Newtonian-Prandtl-Meyer theory, 160 skin fric t ion, effect on, 233 normal force coefficient, 163 Heaviside operator, 62 angle-of-attack effect on

  7. Nonlinear potential analysis techniques for supersonic aerodynamic design

    NASA Technical Reports Server (NTRS)

    Shankar, V.; Szema, K. Y.

    1985-01-01

    A numerical method based on the conservation form of the full potential equation has been applied to the problem of three-dimensional supersonic flows with embedded subsonic regions. The governing equation is cast in a nonorthogonal coordinate system, and the theory of characteristics is used to accurately monitor the type-dependent flow field. A conservative switching scheme is employed to transition from the supersonic marching procedure to a subsonic relaxation algorithm and vice versa. The newly developed computer program can handle arbitrary geometries with fuselage, canard, wing, flow through nacelle, vertical tail and wake components at combined angles of attack and sideslip. Results are obtained for a variety of configurations that include a Langley advanced fighter concept with fuselage centerline nacelle, Rockwell's Advanced Tactical Fighter (ATF) with wing mounted nacelles, and the Shuttle Orbiter configuration. Comparisons with available experiments were good.

  8. A system for aerodynamic design and analysis of supersonic aircraft. Part 4: Test cases

    NASA Technical Reports Server (NTRS)

    Middleton, W. D.; Lundry, J. L.

    1980-01-01

    An integrated system of computer programs was developed for the design and analysis of supersonic configurations. The system uses linearized theory methods for the calculation of surface pressures and supersonic area rule concepts in combination with linearized theory for calculation of aerodynamic force coefficients. Interactive graphics are optional at the user's request. Representative test cases and associated program output are presented.

  9. Flight effects on the aerodynamic and acoustic characteristics of inverted profile coannular nozzles, volume 1. [supersonic cruise aircraft research wind tunnel tests

    NASA Technical Reports Server (NTRS)

    Kozlowski, H.; Packman, A. B.

    1978-01-01

    Jet noise spectra obtained at static conditions from an acoustic wind tunnel and an outdoor facility are compared. Data curves are presented for (1) the effect of relative velocity on OASPL directivity (all configurations); (2) the effect of relative velocity on noise spectra (all configurations); (3) the effect of velocity on PNL directivity (coannular nozzle configurations); (4) nozzle exhaust plume velocity profiles; and (5) the effect of relative velocity on aerodynamic performance.

  10. Estimation of Supersonic Stage Separation Aerodynamics of Winged-Body Launch Vehicles Using Response Surface Methods

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.; Deloach, Richard

    2008-01-01

    A collection of statistical and mathematical techniques referred to as response surface methodology was used to estimate the longitudinal stage separation aerodynamic characteristics of a generic, bimese, winged multi-stage launch vehicle configuration using data obtained on small-scale models at supersonic speeds in the NASA Langley Research Center Unitary Plan Wind Tunnel. The simulated Mach 3 staging was dominated by multiple shock wave interactions between the orbiter and booster vehicles throughout the relative spatial locations of interest. This motivated a partitioning of the overall inference space into several contiguous regions within which the separation aerodynamics were presumed to be well-behaved and estimable using cuboidal and spherical central composite designs capable of fitting full second-order response functions. The primary goal was to approximate the underlying overall aerodynamic response surfaces of the booster vehicle in belly-to-belly proximity to the orbiter vehicle using relatively simple, lower-order polynomial functions that were piecewise-continuous across the full independent variable ranges of interest. The quality of fit and prediction capabilities of the empirical models were assessed in detail, and the issue of subspace boundary discontinuities was addressed. The potential benefits of augmenting the central composite designs to full third order using computer-generated D-optimality criteria were also evaluated. The usefulness of central composite designs, the subspace sizing, and the practicality of fitting low-order response functions over a partitioned inference space dominated by highly nonlinear and possibly discontinuous shock-induced aerodynamics are discussed.

  11. Estimation of Supersonic Stage Separation Aerodynamics of Winged-Body Launch Vehicles Using Response Surface Methods

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.

    2010-01-01

    Response surface methodology was used to estimate the longitudinal stage separation aerodynamic characteristics of a generic, bimese, winged multi-stage launch vehicle configuration at supersonic speeds in the NASA LaRC Unitary Plan Wind Tunnel. The Mach 3 staging was dominated by shock wave interactions between the orbiter and booster vehicles throughout the relative spatial locations of interest. The inference space was partitioned into several contiguous regions within which the separation aerodynamics were presumed to be well-behaved and estimable using central composite designs capable of fitting full second-order response functions. The underlying aerodynamic response surfaces of the booster vehicle in belly-to-belly proximity to the orbiter vehicle were estimated using piecewise-continuous lower-order polynomial functions. The quality of fit and prediction capabilities of the empirical models were assessed in detail, and the issue of subspace boundary discontinuities was addressed. Augmenting the central composite designs to full third-order using computer-generated D-optimality criteria was evaluated. The usefulness of central composite designs, the subspace sizing, and the practicality of fitting lower-order response functions over a partitioned inference space dominated by highly nonlinear and possibly discontinuous shock-induced aerodynamics are discussed.

  12. Developing Supersonic Impactor and Aerodynamic Lens for Separation and Handling of Nano-Sized Particles

    SciTech Connect

    Goodarz Ahmadi

    2008-06-30

    A computational model for supersonic flows of compressible gases in an aerodynamic lens with several lenses and in a supersonic/hypersonic impactor was developed. Airflow conditions in the aerodynamic lens were analyzed and contour plots for variation of Mach number, velocity magnitude and pressure field in the lens were evaluated. The nano and micro-particle trajectories in the lens and their focusing and transmission efficiencies were evaluated. The computational model was then applied to design of a aerodynamic lens that could generate focus particle beams while operating under atmospheric conditions. The computational model was also applied to airflow condition in the supersonic/hypersonic impactor. Variations of airflow condition and particle trajectories in the impactor were evaluated. The simulation results could provide understanding of the performance of the supersonic and hypersonic impactors that would be helpful for the design of such systems.

  13. Low-speed aerodynamic test of an axisymmetric supersonic inlet with variable cowl slot

    NASA Technical Reports Server (NTRS)

    Powell, A. G.; Welge, H. R.; Trefny, C. J.

    1985-01-01

    The experimental low-speed aerodynamic characteristics of an axisymmetric mixed-compression supersonic inlet with variable cowl slot are described. The model consisted of the NASA P-inlet centerbody and redesigned cowl with variable cowl slot powered by the JT8D single-stage fan simulator and driven by an air turbine. The model was tested in the NASA Lewis Research Center 9- by 15-foot low-speed tunnel at Mach numbers of 0, 0.1, and 0.2 over a range of flows, cowl slot openings, centerbody positions, and angles of attack. The variable cowl slot was effective in minimizing lip separation at high velocity ratios, showed good steady-state and dynamic distortion characteristics, and had good angle-of-attack tolerance.

  14. Aerodynamic design and analysis system for supersonic aircraft. Part 1: General description and theoretical development

    NASA Technical Reports Server (NTRS)

    Middleton, W. D.; Lundry, J. L.

    1975-01-01

    An integrated system of computer programs has been developed for the design and analysis of supersonic configurations. The system uses linearized theory methods for the calculation of surface pressures and supersonic area rule concepts in combination with linearized theory for calculation of aerodynamic force coefficients. Interactive graphics are optional at the user's request. This part presents a general description of the system and describes the theoretical methods used.

  15. A computational system for aerodynamic design and analysis of supersonic aircraft. Part 2: User's manual

    NASA Technical Reports Server (NTRS)

    Middleton, W. D.; Lundry, J. L.; Coleman, R. G.

    1976-01-01

    An integrated system of computer programs was developed for the design and analysis of supersonic configurations. The system uses linearized theory methods for the calculation of surface pressures and supersonic area rule concepts in combination with linearized theory for calculation of aerodynamic force coefficients. Interactive graphics are optional at the user's request. This user's manual contains a description of the system, an explanation of its usage, the input definition, and example output.

  16. Pressure distribution and aerodynamic coefficients associated with heat addition to supersonic air stream adjacent to two-dimensional supersonic wing

    NASA Technical Reports Server (NTRS)

    Pinkel, I Irving; Serafini, John S; Gregg, John L

    1952-01-01

    The modifications in the pressure distributions and the aerodynamic coefficients associated with additions of heat to the two-dimensional supersonic in viscid flow field adjacetnt to the lower surface of of a 5-percent-thickness symmetrical circular-arc wing are presented in this report. The pressure distributions are obtained by the use of graphical method which gives the two-dimensional supersonic inviscid flow field obtained with moderate heat addition. The variation is given of the lift-drag ratio and of the aerodynamic coefficients of lift, drag, and moment with free stream Mach number, angle of attack, and parameters defining extent and amount of heat addition. The six graphical solutions used in this study included Mach numbers of 3.0 and 5.0 and angles of attack of 0 degrees and 2 degrees.

  17. Aerodynamic Models for the Low Density Supersonic Decelerator (LDSD) Test Vehicles

    NASA Technical Reports Server (NTRS)

    Van Norman, John W.; Dyakonov, Artem; Schoenenberger, Mark; Davis, Jody; Muppidi, Suman; Tang, Chun; Bose, Deepak; Mobley, Brandon; Clark, Ian

    2016-01-01

    An overview of aerodynamic models for the Low Density Supersonic Decelerator (LDSD) Supersonic Flight Dynamics Test (SFDT) campaign test vehicle is presented, with comparisons to reconstructed flight data and discussion of model updates. The SFDT campaign objective is to test Supersonic Inflatable Aerodynamic Decelerator (SIAD) and large supersonic parachute technologies at high altitude Earth conditions relevant to entry, descent, and landing (EDL) at Mars. Nominal SIAD test conditions are attained by lifting a test vehicle (TV) to 36 km altitude with a helium balloon, then accelerating the TV to Mach 4 and 53 km altitude with a solid rocket motor. Test flights conducted in June of 2014 (SFDT-1) and 2015 (SFDT-2) each successfully delivered a 6 meter diameter decelerator (SIAD-R) to test conditions and several seconds of flight, and were successful in demonstrating the SFDT flight system concept and SIAD-R technology. Aerodynamic models and uncertainties developed for the SFDT campaign are presented, including the methods used to generate them and their implementation within an aerodynamic database (ADB) routine for flight simulations. Pre- and post-flight aerodynamic models are compared against reconstructed flight data and model changes based upon knowledge gained from the flights are discussed. The pre-flight powered phase model is shown to have a significant contribution to off-nominal SFDT trajectory lofting, while coast and SIAD phase models behaved much as predicted.

  18. Integration of a supersonic unsteady aerodynamic code into the NASA FASTEX system

    NASA Technical Reports Server (NTRS)

    Appa, Kari; Smith, Michael J. C.

    1987-01-01

    A supersonic unsteady aerodynamic loads prediction method based on the constant pressure method was integrated into the NASA FASTEX system. The updated FASTEX code can be employed for aeroelastic analyses in subsonic and supersonic flow regimes. A brief description of the supersonic constant pressure panel method, as applied to lifting surfaces and body configurations, is followed by a documentation of updates required to incorporate this method in the FASTEX code. Test cases showing correlations of predicted pressure distributions, flutter solutions, and stability derivatives with available data are reported.

  19. X-34 Vehicle Aerodynamic Characteristics

    NASA Technical Reports Server (NTRS)

    Brauckmann, Gregory J.

    1998-01-01

    The X-34, being designed and built by the Orbital Sciences Corporation, is an unmanned sub-orbital vehicle designed to be used as a flying test bed to demonstrate key vehicle and operational technologies applicable to future reusable launch vehicles. The X-34 will be air-launched from an L-1011 carrier aircraft at approximately Mach 0.7 and 38,000 feet altitude, where an onboard engine will accelerate the vehicle to speeds above Mach 7 and altitudes to 250,000 feet. An unpowered entry will follow, including an autonomous landing. The X-34 will demonstrate the ability to fly through inclement weather, land horizontally at a designated site, and have a rapid turn-around capability. A series of wind tunnel tests on scaled models was conducted in four facilities at the NASA Langley Research Center to determine the aerodynamic characteristics of the X-34. Analysis of these test results revealed that longitudinal trim could be achieved throughout the design trajectory. The maximum elevon deflection required to trim was only half of that available, leaving a margin for gust alleviation and aerodynamic coefficient uncertainty. Directional control can be achieved aerodynamically except at combined high Mach numbers and high angles of attack, where reaction control jets must be used. The X-34 landing speed, between 184 and 206 knots, is within the capabilities of the gear and tires, and the vehicle has sufficient rudder authority to control the required 30-knot crosswind.

  20. Supersonic aerodynamic characteristics of hypersonic low-wave-drag elliptical body-tail combinations as affected by changes in stabilizer configuration

    NASA Technical Reports Server (NTRS)

    Spencer, B., Jr.; Fournier, R. H.

    1973-01-01

    An investigation has been made at Mach numbers from 1.50 to 4.63 to determine systematically the effects of the addition and position of outboard stabilizers and vertical- and vee-tail configurations on the performance and stability characteristics of a low-wave-drag elliptical body. The basic body shape was a zero-lift hypersonic minimum-wave-drag body as determined for the geometric constraints of length and volume. The elliptical cross section had an axis ratio of 2 (major axis horizontal) and an equivalent fineness ratio of 6.14. Base-mounted outboard stabilizers were at various dihedral angles from 90 deg to minus 90 deg with and without a single center-line vertical tail or a vee-tail. The angle of attack was varied from about minus 6 to 27 deg at sideslip angles of 0 and 5 deg and a constant Reynolds number of 4.58 x one million (based on body length).

  1. Effect of planform and body on supersonic aerodynamics of multibody configurations

    NASA Technical Reports Server (NTRS)

    Mcmillin, S. Naomi; Bauer, Steven X. S.; Howell, Dorothy T.

    1992-01-01

    An experimental and theoretical investigation of the effect of the wing planform and bodies on the supersonic aerodynamics of a low-fineness-ratio, multibody configuration has been conducted in the Langley Unitary Plan Wind Tunnel at Mach numbers of 1.60, 1.80, 2.00, and 2.16. Force and moment data, flow-visualization data, and surface-pressure data were obtained on eight low-fineness-ratio, twin-body configurations. These configurations varied in inboard wing planform shape, outboard wing planform shape, outboard wing planform size, and presence of the bodies. The force and moment data showed that increasing the ratio of outboard wing area to total wing area or increasing the leading-edge sweep of the inboard wing influenced the aerodynamic characteristics. The flow-visualization data showed a complex flow-field system of shocks, shock-induced separation, and body vortex systems occurring between the side bodies. This flow field was substantially affected by the inboard wing planform shape but minimally affected by the outboard wing planform shape. The flow-visualization and surface-pressure data showed that flow over the outboard wing developed as expected with changes in angle of attack and Mach number and was affected by the leading-edge sweep of the inboard wing and the presence of the bodies. Evaluation of the linear-theory prediction methods revealed their general inability to consistently predict the characteristics of these multibody configurations.

  2. Prediction of nacelle aerodynamic interference effects at low supersonic Mach numbers

    NASA Technical Reports Server (NTRS)

    Kulfan, R. M.

    1980-01-01

    The accuracy of analytical predictions of nacelle aerodynamic interference effects at low supersonic speeds are studied by means of test versus theory comparisons. Comparisons shown include: (1) isolated wing body lift, drag, and pitching moments; (2) isolated nacelle drag and pressure distributions; (3) nacelle interference shock wave patterns and pressure distributions on the wing lower surface; (4) nacelle interference effects on wing body lift, drag, and pitching moments; and (5) total installed nacelle interference effects on lift, drag, and pitching moment. The comparisons also illustrate effects of nacelle location, nacelle spillage, angle of attack, and Mach numbers on the aerodynamic interference. The initial results seem to indicate that the methods can satisfactorily predict lift, drag, pitching moment, and pressure distributions of installed engine nacelles at low supersonic Mach numbers with mass flow ratios from 0.7 to 1.0 for configurations typical of efficient supersonic cruise airplanes.

  3. X-33 Hypersonic Aerodynamic Characteristics

    NASA Technical Reports Server (NTRS)

    Murphy, Kelly J.; Nowak, Robert J.; Thompson, Richard A.; Hollis, Brian R.; Prabhu, Ramadas K.

    1999-01-01

    Lockheed Martin Skunk Works, under a cooperative agreement with NASA, will design, build, and fly the X-33, a half-scale prototype of a rocket-based, single-stage-to-orbit (SSTO), reusable launch vehicle (RLV). A 0.007-scale model of the X-33 604BOO02G configuration was tested in four hypersonic facilities at the NASA Langley Research Center to examine vehicle stability and control characteristics and to populate the aerodynamic flight database for the hypersonic regime. The vehicle was found to be longitudinally controllable with less than half of the total body flap deflection capability across the angle of attack range at both Mach 6 and Mach 10. Al these Mach numbers, the vehicle also was shown to be longitudinally stable or neutrally stable for typical (greater than 20 degrees) hypersonic flight attitudes. This configuration was directionally unstable and the use of reaction control jets (RCS) will be necessary to control the vehicle at high angles of attack in the hypersonic flight regime. Mach number and real gas effects on longitudinal aerodynamics were shown to be small relative to X-33 control authority.

  4. X-33 Hypersonic Aerodynamic Characteristics

    NASA Technical Reports Server (NTRS)

    Murphy, Kelly J.; Nowak, Robert J.; Thompson, Richard A.; Hollis, Brian R.; Prabhu, Ramadas K.

    1999-01-01

    Lockheed Martin Skunk Works, under a cooperative agreement with NASA, will build and fly the X-33, a half-scale prototype of a rocket-based, single-stage-to-orbit (SSTO), reusable launch vehicle (RLV). A 0.007-scale model of the X-33 604B0002G configuration was tested in four hypersonic facilities at the NASA Langley Research Center to examine vehicle stability and control characteristics and to populate an aerodynamic flight database i n the hypersonic regime. The vehicle was found to be longitudinally controllable with less than half of the total body flap deflection capability across the angle of attack range at both Mach 6 and Mach 10. At these Mach numbers, the vehicle also was shown to be longitudinally stable or neutrally stable for typical (greater than 20 degrees) hypersonic flight attitudes. This configuration was directionally unstable and the use of reaction control jets (RCS) will be necessary to control the vehicle at high angles of attack in the hypersonic flight regime. Mach number and real gas effects on longitudinal aerodynamics were shown to be small relative to X-33 control authority.

  5. X-33 Hypersonic Aerodynamic Characteristics

    NASA Technical Reports Server (NTRS)

    Murphy, Kelly J.; Nowak, Robert J.; Thompson, Richard A.; Hollis, Brian R.; Prabhu, Ramadas K.

    1999-01-01

    Lockheed Martin Skunk Works, under a cooperative agreement with NASA, will build and fly the X-33, a half-scale prototype of a rocket-based, single-stage-to-orbit (SSTO), reusable launch vehicle (RLV). A 0.007-scale model of the X-33 604B0002G configuration was tested in four hypersonic facilities at the NASA Langley Research Center to examine vehicle stability and control characteristics and to populate an aerodynamic flight database in the hypersonic regime. The vehicle was found to be longitudinally controllable with less than half of the total body flap deflection capability across the angle of attack range at both Mach 6 and Mach 10. At these Mach numbers, the vehicle also was shown to be longitudinally stable or neutrally stable for typical (greater than 20 degrees) hypersonic flight attitudes. This configuration was directionally unstable and the use of reaction control jets (RCS) will be necessary to control the vehicle at high angles of attack in the hypersonic flight regime. Mach number and real gas effects on longitudinal aerodynamics were shown to be small relative to X-33 control authority.

  6. X-33 Hypersonic Aerodynamic Characteristics

    NASA Technical Reports Server (NTRS)

    Murphy, Kelly J.; Nowak, Robert J.; Thompson, Richard A.; Hollis, Brian R.; Prabhu, Ramadas K.

    1999-01-01

    Lockheed Martin Skunk Works, under a cooperative agreement with NASA, will build and fly the X-33, a half-scale prototype of a rocket-based, single-stage-to-orbit (SSTO), reusable launch vehicle (RLV). A 0.007-scale model of the X-33 604B0002G configuration was tested in four hypersonic facilities at the NASA Langley Research Center to examine vehicle stability and control characteristics and to populate an aerodynamic flight database in the hypersonic regime, The vehicle was found to be longitudinally controllable with less than half of the total body flap deflection capability across the angle of attack range at both Mach 6 and Mach 10. At these Mach numbers, the vehicle also was shown to be longitudinally stable or neutrally stable for typical (greater than 20 degrees) hypersonic flight attitudes. This configuration was directionally unstable and the use of reaction control jets (RCS) will be necessary to control the vehicle at high angles of attack in the hypersonic flight regime. Mach number and real gas effects on longitudinal aerodynamics were shown to be small relative to X-33 control authority.

  7. CFD Simulations of the Supersonic Inflatable Aerodynamic Decelerator (SIAD) Ballistic Range Tests

    NASA Technical Reports Server (NTRS)

    Brock, Joseph; Stern, Eric; Wilder, Michael

    2017-01-01

    A series of ballistic range tests were performed on a scaled model of the Supersonic Flight Demonstration Test (SFDT) intended to test the Supersonic Inflatable Aerodynamic Decelerator (SIAD) geometry. The purpose of these experiments were to provide aerodynamic coefficients of the vehicle to aid in mission and vehicle design. The experimental data spans the moderate Mach number range, $3.8-2.0$, with a total angle of attack ($alpha_T$) range, $10o-20o$. These conditions are intended to span the Mach-$alpha$ space for the majority of the SFDT experiment. In an effort to validate the predictive capabilities of Computational Fluid Dynamics (CFD) for free-flight aerodynamic behavior, numerical simulations of the ballistic range experiment are performed using the unstructured finite volume Navier-Stokes solver, US3D. Comparisons to raw vehicle attitude, and post-processed aerodynamic coefficients are made between simulated results and experimental data. The resulting comparisons for both raw model attitude and derived aerodynamic coefficients show good agreement with experimental results. Additionally, near body pressure field values for each trajectory simulated are investigated. Extracted surface and wake pressure data gives further insights into dynamic flow coupling leading to a potential mechanism for dynamic instability.

  8. The predicted effect of aerodynamic detuning on coupled bending-torsion unstalled supersonic flutter

    NASA Technical Reports Server (NTRS)

    Hoyniak, D.; Fleeter, S.

    1986-01-01

    A mathematical model is developed to predict the enhanced coupled bending-torsion unstalled supersonic flutter stability due to alternate circumferential spacing aerodynamic detuning of a turbomachine rotor. The translational and torsional unsteady aerodynamic coefficients are developed in terms of influence coefficients, with the coupled bending-torsion stability analysis developed by considering the coupled equations of motion together with the unsteady aerodynamic loading. The effect of this aerodynamic detuning on coupled bending-torsion unstalled supersonic flutter as well as the verification of the modeling are then demonstrated by considering an unstable 12 bladed rotor, with Verdon's uniformly spaced Cascade B flow geometry as a baseline. However, with the elastic axis and center of gravity at 60 percent of the chord, this type of aerodynamic detuning has a minimal effect on stability. For both uniform and nonuniform circumferentially space rotors, a single degree of freedom torsion mode analysis was shown to be appropriate for values of the bending-torsion natural frequency ratio lower than 0.6 and higher 1.2. When the elastic axis and center of gravity are not coincident, the effect of detuning on cascade stability was found to be very sensitive to the location of the center of gravity with respect to the elastic axis. In addition, it was determined that when the center of gravity was forward of an elastic axis located at midchord, a single degree of freedom torsion model did not accurately predict cascade stability.

  9. The incorporation of plotting capability into the Unified Subsonic Supersonic Aerodynamic Analysis program, version B

    NASA Technical Reports Server (NTRS)

    Winter, O. A.

    1980-01-01

    The B01 version of the United Subsonic Supersonic Aerodynamic Analysis program is the result of numerous modifications and additions made to the B00 version. These modifications and additions affect the program input, its computational options, the code readability, and the overlay structure. The following are described: (1) the revised input; (2) the plotting overlay programs which were also modified, and their associated subroutines, (3) the auxillary files used by the program, the revised output data; and (4) the program overlay structure.

  10. Dynamic characteristics of pulsed supersonic fuel sprays

    NASA Astrophysics Data System (ADS)

    Pianthong, K.; Matthujak, A.; Takayama, K.; Milton, B. E.; Behnia, M.

    2008-06-01

    This paper describes the dynamic characteristics of pulsed, supersonic liquid fuel sprays or jets injected into ambient air. Simple, single hole nozzles were employed with the nozzle sac geometries being varied. Different fuel types, diesel fuel, bio-diesel, kerosene, and gasoline were used to determine the effects of fuel properties on the spray characteristics. A vertical two-stage light gas gun was employed as a projectile launcher to provide a high velocity impact to produce the liquid jet. The injection pressure was around 0.88-1.24 GPa in all cases. The pulsed, supersonic fuel sprays were visualized by using a high-speed video camera and shadowgraph method. The spray tip penetration and velocity attenuation and other characteristics were examined and are described here. An instantaneous spray tip velocity of 1,542 m/s (Mach number 4.52) was obtained. However, this spray tip velocity can be sustained for only a very short period (a few microseconds). It then attenuates very quickly. The phenomenon of multiple high frequency spray pulses generated by a single shot impact and the changed in the angle of the shock structure during the spray flight, which had already been observed in previous studies, is again noted. Multiple shock waves from the conical nozzle spray were also clearly captured.

  11. Genetic Evolution of Shape-Altering Programs for Supersonic Aerodynamics

    NASA Technical Reports Server (NTRS)

    Kennelly, Robert A., Jr.; Bencze, Daniel P. (Technical Monitor)

    2002-01-01

    Two constrained shape optimization problems relevant to aerodynamics are solved by genetic programming, in which a population of computer programs evolves automatically under pressure of fitness-driven reproduction and genetic crossover. Known optimal solutions are recovered using a small, naive set of elementary operations. Effectiveness is improved through use of automatically defined functions, especially when one of them is capable of a variable number of iterations, even though the test problems lack obvious exploitable regularities. An attempt at evolving new elementary operations was only partially successful.

  12. Aeroelastic stability consideration of supersonic flight vehicle using nonlinear aerodynamic response surfaces

    NASA Astrophysics Data System (ADS)

    Fathi Jegarkandi, M.; Nobari, A. S.; Sabzehparvar, M.; Haddadpour, H.

    2009-08-01

    Aeroelastic stability of a flexible supersonic flight vehicle is considered using nonlinear dynamics, nonlinear aerodynamics, and a linear structural model. Response surfaces including global multivariate orthogonal modeling functions are invoked to derive applied nonlinear aerodynamic coefficients. A modified Gram-Schmidt method is utilized to orthogonalize the produced polynomial multivariate functions, selected and ranked by predicted squared error metric. Local variation of angle-of-attack and side-slip angle is applied to the analytical model. Identification of nonlinear aerodynamic coefficients of the flight vehicle is conducted employing a CFD code and the required analytical model for simulation purposes is constructed. The method is used to determine the aeroelastic instability and response of a selected flight vehicle.

  13. Abort System Using Supersonic Aerodynamic Interaction for Capsule-Type Space Transportation System

    NASA Astrophysics Data System (ADS)

    小澤, 啓伺; 北村, 圭一; 花井, 勝祥; 三好, 理也; 森, 浩一; 中村, 佳朗

    The space transportation system using capsule/rocket configurations such as Apollo and Soyuz are simple compared with Space Shuttle, and have several merits from the viewpoint of reliability. The capsule/rocket system will take over the Space Shuttle, after it retires in 2010. As the Space Shuttle accidents had been caused by several factors, e.g., aerodynamic interaction of shock waves ahead of its wing, advanced abort systems such as LAS (Launch Abort System) are required for the capsule/rocket system. In the present study, as a baseline configuration, a combination of a cone and a cylinder is employed as a CEV (Crew Exploration Vehicle), which consists of a capsule (LAV: Launch Abort Vehicle) and a rocket (SM: Service Module). By changing the relative position of the two components as well as the profile area of the rocket, their effects on the capsule/rocket aerodynamic interaction and characteristics (drag and pitching moment) are experimentally and numerically investigated at a supersonic speed (M∞ = 3.0). It is found from the results that the clearance have little effects on the flow field for the case of the baseline configuration. The capsule always showed a positive drag (CD = 0.34), which means that thrust is required to overcome the drag. Otherwise the capsule will recontact the rocket. However in the case where the rocket contact area is 2.2 times as large as the capsule profile, more favorable effects were obtained. Especially in the case of a certain clearance (h/D = 0.40), the drag coefficient of the capsule is CD = -0.35, which means that the capsule suffers a thrust force from the aerodynamic interaction. Under this condition, if capsule has a pitch angle with 5 degrees instantaneously, then pitching moment coefficient becomes CMp = -0.41 therefore capsule stabilize. However, in the case of a very small clearance (h/D ∝ 0.00), the flow becomes unsteady involving pulsating shock wave, leading to a potentially risky separation of the capsule.

  14. Aerodynamic characteristics of a fixed arrow-wing supersonic cruise aircraft at Mach numbers of 2.30, 2.70, and 2.95. [Langley Unitary Plan wind tunnel tests

    NASA Technical Reports Server (NTRS)

    Morris, O. A.; Fuller, D. E.; Watson, C. B.

    1978-01-01

    Tests were conducted in the Langley Unitary Plan wind tunnel at Mach numbers of 2.30. 2.70, and 2.95 to determine the performance, static stability, and control characteristics of a model of a fixed-wing supersonic cruise aircraft with a design Mach Number of 2.70 (SCAT 15-F-9898). The configuration had a 74 deg swept warped wing with a reflexed trailing edge and four engine nacelles mounted below the reflexed portion of the wing. A number of variations in the basic configuration were investigated; they included the effect of wing leading edge radius, the effect of various model components, and the effect of model control deflections.

  15. Assessment of CFD-based Response Surface Model for Ares I Supersonic Ascent Aerodynamics

    NASA Technical Reports Server (NTRS)

    Hanke, Jeremy L.

    2011-01-01

    The Ascent Force and Moment Aerodynamic (AFMA) Databases (DBs) for the Ares I Crew Launch Vehicle (CLV) were typically based on wind tunnel (WT) data, with increments provided by computational fluid dynamics (CFD) simulations for aspects of the vehicle that could not be tested in the WT tests. During the Design Analysis Cycle 3 analysis for the outer mold line (OML) geometry designated A106, a major tunnel mishap delayed the WT test for supersonic Mach numbers (M) greater than 1.6 in the Unitary Plan Wind Tunnel at NASA Langley Research Center, and the test delay pushed the final delivery of the A106 AFMA DB back by several months. The aero team developed an interim database based entirely on the already completed CFD simulations to mitigate the impact of the delay. This CFD-based database used a response surface methodology based on radial basis functions to predict the aerodynamic coefficients for M > 1.6 based on only the CFD data from both WT and flight Reynolds number conditions. The aero team used extensive knowledge of the previous AFMA DB for the A103 OML to guide the development of the CFD-based A106 AFMA DB. This report details the development of the CFD-based A106 Supersonic AFMA DB, constructs a prediction of the database uncertainty using data available at the time of development, and assesses the overall quality of the CFD-based DB both qualitatively and quantitatively. This assessment confirms that a reasonable aerodynamic database can be constructed for launch vehicles at supersonic conditions using only CFD data if sufficient knowledge of the physics and expected behavior is available. This report also demonstrates the applicability of non-parametric response surface modeling using radial basis functions for development of aerodynamic databases that exhibit both linear and non-linear behavior throughout a large data space.

  16. Aerodynamic and Aeroelastic Characteristics of a Tension Cone Inflatable Aerodynamic Decelerator

    NASA Technical Reports Server (NTRS)

    Clark, Ian G.; Cruz, Juan R.; Hughes, Monica F.; Ware, Joanne S.; Madlangbayan, Albert; Braun, Robert D.

    2009-01-01

    The supersonic aerodynamic and aeroelastic characteristics of a tension cone inflatable aerodynamic decelerator were investigated by wind tunnel testing. Two sets of tests were conducted: one using rigid models and another using textile models. Tests using rigid models were conducted over a Mach number range from 1.65 to 4.5 at angles of attack from -12 to 20 degrees. The axial, normal, and pitching moment coefficients were found to be insensitive to Mach number over the tested range. The axial force coefficient was nearly constant (C(sub A) = 1.45 +/- 0.05) with respect to angle of attack. Both the normal and pitching moment coefficients were nearly linear with respect to angle of attack. The pitching moment coefficient showed the model to be statically stable about the reference point. Schlieren images and video showed a detached bow shock with no evidence of large regions of separated flow and/or embedded shocks at all Mach numbers investigated. Qualitatively similar static aerodynamic coefficient and flow visualization results were obtained using textile models at a Mach number of 2.5. Using inflatable textile models the torus pressure required to maintain the model in the fully-inflated configuration was determined. This pressure was found to be sensitive to details in the structural configuration of the inflatable models. Additional tests included surface pressure measurements on rigid models and deployment and inflation tests with inflatable models.

  17. A review of ONERA aerodynamic research in support of a future supersonic transport aircraft

    NASA Astrophysics Data System (ADS)

    Thibert, J. J.; Arnal, D.

    2000-11-01

    The ONERA activities concerning the aerodynamics of the future supersonic transport aircraft are reviewed. Section 1 is devoted to the performance prediction and detailed comparisons between CFD and wind-tunnel data are presented and discussed. Section 2 addresses the problem of the drag prediction in cruise flight conditions from wind-tunnel data. Skin friction coefficients values measured in flight are compared to the results of boundary layer computations. Section 3 is devoted to wing designs with numerical optimisation techniques. Several examples are presented and discussed. Results concerning riblets and laminar flow control are given in Section 4 part which also will present experiments carried out for attachment line contamination investigation. Results from basic research on supersonic laminar flows are also be presented. Section 5 deals with activities on air intake aerodynamics. After a brief recall of supersonic air intakes operational modes and a description of the Concorde air intake, comparisons between CFD and wind tunnel data on a generic 2D intake are presented. Basic experiments on intake internal flow are described and the problem of the internal shock control is addressed.

  18. Tribological study of an aerodynamic thrust bearing in the supersonic regime

    NASA Astrophysics Data System (ADS)

    Dupuy, F.; Bou-Saïd, B.; Garcia, M.

    2017-02-01

    Nowadays, aerodynamic air thrust bearing are mainly used over a large panel of turbo-machineries. These systems become increasingly faster and up to operate in supersonic regime. They have not reached a sufficient level of research in terms of high speed. The possibility of an aerodynamic thrust bearing operating in a supersonic regime is studied. First, the air film dynamic study for high Reynolds number is based on the “modified Reynolds” equations, which take into account the inertia terms, the viscosity’s variation in the film thickness, and the turbulence. It’s an extension of the traditional model used in lubrication called the generalized Reynolds equation. The results show that a depression occur at the location of the change of slope of the tapper flat geometry. The hypothesis of presence of shock or rarefaction waves shows that the pressure gradient in the film thickness may be no longer negligible. The modified Reynolds equation may be not enough to describe the problem. A new system is built to study these phenomena: the Navier-Stokes equation are adapted to the film’s geometry. The dynamic air film’s behavior study in supersonic regime requires a shock capturing scheme called WENO scheme (“Weighted Essentially Non Oscillatory”), mainly used in shock and turbulence studies. The numerical results give the film behavior modelling of a fixed air thrust bearing pad. The evolution of the quantities shows that shock wave can occur in a thin film.

  19. Application of a one-strip integral method to the unsteady supersonic aerodynamics of an inclined flat surface

    NASA Technical Reports Server (NTRS)

    Bennett, R. M.

    1972-01-01

    The method of integral relations is applied in a one-strip approximation to the perturbation equations governing small motions of an inclined, sharp-edged, flat surface about the mean supersonic steady flow. Algebraic expressions for low reduced-frequency aerodynamics are obtained and a set of ordinary differential equations are obtained for general oscillatory motion. Results are presented for low reduced-frequency aerodynamics and for the variation of the unsteady forces with frequency. The method gives accurate results for the aerodynamic forces at low reduced frequency which are in good agreement with available experimental data. However, for cases in which the aerodynamic forces vary rapidly with frequency, the results are qualitatively correct, but of limited accuracy. Calculations indicate that for a range of inclination angles near shock detachment such that the flow in the shock layer is low supersonic, the aerodynamic forces vary rapidly both with inclination angle and with reduced frequency.

  20. Aerodynamic analysis of the aerospaceplane HyPlane in supersonic rarefied flow

    NASA Astrophysics Data System (ADS)

    Zuppardi, Gennaro; Savino, Raffaele; Russo, Gennaro; Spano'Cuomo, Luca; Petrosino, Eliano

    2016-06-01

    HyPlane is the Italian aerospaceplane proposal targeting, at the same time, both the space tourism and point-to-point intercontinental hypersonic flights. Unlike other aerospaceplane projects, relying on boosters or mother airplanes that bring the vehicle to high altitude, HyPlane will take off and land horizontally from common runways. According to the current project, HyPlane will fly sub-orbital trajectories under high-supersonic/low-hypersonic continuum flow regimes. It can go beyond the von Karman line at 100 km altitude for a short time, then starting the descending leg of the trajectory. Its aerodynamic behavior up to 70 km have already been studied and the results published in previous works. In the present paper some aspects of the aerodynamic behavior of HyPlane have been analyzed at 80, 90 and 100 km. Computer tests, calculating the aerodynamic parameters, have been carried out by a Direct Simulation Monte Carlo code. The effects of the Knudsen, Mach and Reynolds numbers have been evaluated in clean configuration. The effects of the aerodynamic surfaces on the rolling, pitching and yawing moments, and therefore on the capability to control attitude, have been analyzed at 100 km altitude. The aerodynamic behavior has been compared also with that of another aerospaceplane at 100 km both in clean and flapped configuration.

  1. Laser velocimetry applied to transonic and supersonic aerodynamics

    NASA Technical Reports Server (NTRS)

    Johnson, D. A.; Bachalo, W. D.; Moddaress, D.

    1976-01-01

    As a further demonstration of the capabilities of laser velocity in compressible aerodynamics, measurements obtained in a Mach 2.9 separated turbulent boundary layer and in the transonic flow past a two-dimensional airfoil section are presented and compared to data realized by conventional techniques. In the separated-flow study, the comparisons were made against pitot-static pressure data. Agreement in mean velocities was realized where the pressure measurements could be considered reliable; however, in regions of instantaneous reverse velocities, the laser results were found to be consistent with the physics of the flow whereas the pressure data were not. The laser data obtained in regions of extremely high turbulence suggest that velocity biasing does not occur if the particle occurrence rate is low relative to the turbulent fluctuation rate. Streamwise turbulence intensities are also presented. In the transonic airfoil study, velocity measurements obtained immediately outside the upper surface boundary layer of a 6-inch chord MACA 64A010 airfoil are compared to edge velocities inferred from surface pressure measurements. For free-stream Mach numbers of 0.6 and 0.8, the agreement in results was very good. Dual scatter optical arrangements in conjunction with a single particle, counter-type signal processor were employed in these investigations. Half-micron-diameter polystyrene spheres and naturally occurring condensed oil vapor acted as light scatterers in the two respective flows. Bragg-cell frequency shifting was utilized in the separated flow study.

  2. Aerodynamic optimization of supersonic compressor cascade using differential evolution on GPU

    NASA Astrophysics Data System (ADS)

    Aissa, Mohamed Hasanine; Verstraete, Tom; Vuik, Cornelis

    2016-06-01

    Differential Evolution (DE) is a powerful stochastic optimization method. Compared to gradient-based algorithms, DE is able to avoid local minima but requires at the same time more function evaluations. In turbomachinery applications, function evaluations are performed with time-consuming CFD simulation, which results in a long, non affordable, design cycle. Modern High Performance Computing systems, especially Graphic Processing Units (GPUs), are able to alleviate this inconvenience by accelerating the design evaluation itself. In this work we present a validated CFD Solver running on GPUs, able to accelerate the design evaluation and thus the entire design process. An achieved speedup of 20x to 30x enabled the DE algorithm to run on a high-end computer instead of a costly large cluster. The GPU-enhanced DE was used to optimize the aerodynamics of a supersonic compressor cascade, achieving an aerodynamic loss minimization of 20%.

  3. Aerodynamic sensitivities from subsonic, sonic and supersonic unsteady, nonplanar lifting-surface theory

    NASA Technical Reports Server (NTRS)

    Yates, E. Carson, Jr.

    1987-01-01

    The technique of implicit differentiation has been used in combination with linearized lifting-surface theory to derive analytical expressions for aerodynamic sensitivities (i.e., rates of change of lifting pressures with respect to general changes in aircraft geometry, including planform variations) for steady or oscillating planar or nonplanar lifting surfaces in subsonic, sonic, or supersonic flow. The geometric perturbation is defined in terms of a single variable, and the user need only provide simple expressions or similar means for defining the continuous or discontinuous global or local perturbation of interest. Example expressions are given for perturbations of the sweep, taper, and aspect ratio of a wing with trapezoidal semispan planform. In addition to direct computational use, the analytical method presented here should provide benchmark criteria for assessing the accuracy of aerodynamic sensitivities obtained by approximate methods such as finite geometry perturbation and differencing. The present process appears to be readily adaptable to more general surface-panel methods.

  4. A computational system for aerodynamic design and analysis of supersonic aircraft. Part 1: General description and theoretical development

    NASA Technical Reports Server (NTRS)

    Middleton, W. D.; Lundry, J. L.

    1976-01-01

    An integrated system of computer programs was developed for the design and analysis of supersonic configurations. The system uses linearized theory methods for the calculation of surface pressures and supersonic area rule concepts in combination with linearized theory for calculation of aerodynamic force coefficients. Interactive graphics are optional at the user's request. Schematics of the program structure and the individual overlays and subroutines are described.

  5. A Study of the Motion and Aerodynamic Heating of Missiles Entering the Earth's Atmosphere at High Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Allen, H. Julian; Eggers, A. J., Jr.

    1953-01-01

    A simplified analysis is made of the velocity and deceleration history of missiles entering the earth's atmosphere at high supersonic speeds. It is found that, in general, the gravity force is negligible compared to the aerodynamic drag force and, hence, that the trajectory is essentially a straight line. A constant drag coefficient and an exponential variation of density with altitude are assumed and generalized curves for the variation of missile speed and deceleration with altitude are obtained. A curious finding is that the maximum deceleration is independent of physical characteristics of a missile (e.g., mass, size, and drag coefficient) and is determined only by entry speed and flight-path angle, provided this deceleration occurs before impact. This provision is satisfied by missiles presently of more usual interest.

  6. Effects of plasma aerodynamic actuation on oblique shock wave in a cold supersonic flow

    NASA Astrophysics Data System (ADS)

    Wang, Jian; Li, Yinghong; Cheng, Bangqin; Su, Changbing; Song, Huimin; Wu, Yun

    2009-08-01

    Wedge oblique shock wave control using an arc discharge plasma aerodynamic actuator was investigated both experimentally and theoretically. Schlieren photography measurements in a small-scale short-duration supersonic wind tunnel indicated that the shock wave angle decreased and its start point shifted upstream with the plasma aerodynamic actuation. Also the shock wave intensity weakened, as shown by the decrease in the gas static pressure ratio of flow downstream and upstream of the shock wave. Moreover, the shock wave control effect was intensified when a static magnetic field was applied. Under test conditions of Mach 2.2, magnetic control and input voltage 3 kV, the start point of the shock wave shifted 4 mm upstream, while its angle and intensity decreased 8.6% and 8.8%, respectively. A thermal choking model was proposed to deduce the change laws of oblique shock wave control by surface arc discharge. The theoretical result was consistent with the experimental result, which demonstrated that the thermal choking model can effectively forecast the effect of plasma actuation on an oblique shock wave in a cold supersonic flow.

  7. Aerodynamic Design and Numerical Analysis of Supersonic Turbine for Turbo Pump

    NASA Astrophysics Data System (ADS)

    Fu, Chao; Zou, Zhengping; Kong, Qingguo; Cheng, Honggui; Zhang, Weihao

    2016-09-01

    Supersonic turbine is widely used in the turbo pump of modern rocket. A preliminary design method for supersonic turbine has been developed considering the coupling effects of turbine and nozzle. Numerical simulation has been proceeded to validate the feasibility of the design method. As the strong shockwave reflected on the mixing plane, additional numerical simulated error would be produced by the mixing plane model in the steady CFD. So unsteady CFD is employed to investigate the aerodynamic performance of the turbine and flow field in passage. Results showed that the preliminary design method developed in this paper is suitable for designing supersonic turbine. This periodical variation of complex shockwave system influences the development of secondary flow, wake and shock-boundary layer interaction, which obviously affect the secondary loss in vane passage. The periodical variation also influences the strength of reflecting shockwave, which affects the profile loss in vane passage. Besides, high circumferential velocity at vane outlet and short blade lead to high radial pressure gradient, which makes the low kinetic energy fluid moves towards hub region and produces additional loss.

  8. Aerodynamic shape optimization directed toward a supersonic transport using sensitivity analysis

    NASA Technical Reports Server (NTRS)

    Baysal, Oktay

    1995-01-01

    This investigation was conducted from March 1994 to August 1995, primarily, to extend and implement the previously developed aerodynamic design optimization methodologies for the problems related to a supersonic transport design. These methods had demonstrated promise to improve the designs (more specifically, the shape) of aerodynamic surfaces, by coupling optimization algorithms (OA) with Computational Fluid Dynamics (CFD) algorithms via sensitivity analyses (SA) with surface definition methods from Computer Aided Design (CAD). The present extensions of this method and their supersonic implementations have produced wing section designs, delta wing designs, cranked-delta wing designs, and nacelle designs, all of which have been reported in the open literature. Despite the fact that these configurations were highly simplified to be of any practical or commercial use, they served the algorithmic and proof-of-concept objectives of the study very well. The primary cause for the configurational simplifications, other than the usual simplify-to-study the fundamentals reason, were the premature closing of the project. Only after the first of the originally intended three-year term, both the funds and the computer resources supporting the project were abruptly cut due to their severe shortages at the funding agency. Nonetheless, it was shown that the extended methodologies could be viable options in optimizing the design of not only an isolated single-component configuration, but also a multiple-component configuration in supersonic and viscous flow. This allowed designing with the mutual interference of the components being one of the constraints all along the evolution of the shapes.

  9. Aerodynamic characteristics of French consonants

    NASA Astrophysics Data System (ADS)

    Demolin, Didier; Hassid, Sergio; Soquet, Alain

    2004-05-01

    This paper reports some aerodynamic measurements made on French consonants with a group of ten speakers. Speakers were recorded while saying nonsense words in phrases such as papa, dis papa encore. The nonsense words in the study combined each of the French consonants with three vowels /i, a, u/ to from two syllables words with the first syllable being the same as the second. In addition to the audio signal, recordings were made of the oral airflow, the pressure of the air in the pharynx above the vocal folds and the pressure of the air in the trachea just below the vocal folds. The pharyngeal pressure was recorded via a catheter (i.d. 5 mm) passed through the nose so that its open end could be seen in the pharynx below the uvula. The subglottal pressure was recorded via a tracheal puncture between the first and the second rings of the trachea or between the cricoid cartilage and the first tracheal ring. Results compare subglottal presssure, pharyngeal pressure, and airflow values. Comparisons are made between values obtained with male and female subjects and various types of consonants (voiced versus voiceless at the same place of articulation, stops, fricatives, and nasals).

  10. Parametric study of supersonic STOVL flight characteristics

    NASA Technical Reports Server (NTRS)

    Rapp, David C.

    1985-01-01

    A number of different control devices and techniques are evaluated to determine their suitability for increasing the short takeoff performance of a supersonic short-takeoff/vertical landing (STOVL) aircraft. Analysis was based on a rigid-body mathematical model of the General Dynamics E-7, a single engine configuration that utilizes ejectors and thrust deflection for propulsive lift. Alternatives investigated include increased static pitch, the addition of a close-coupled canard, use of boundary layer control to increase the takeoff lift coefficient, and the addition of a vectorable aft fan air nozzle. Other performance studies included the impact of individual E-7 features, the sensitivity to ejector performance, the effect of removing the afterburners, and a determination of optional takeoff and landing transition methods. The results pertain to both the E-7 and other configurations. Several alternatives were not as well suited to the E-7 characteristics as they would be to an alternative configuration, and vice versa. A large amount of supporting data for each analysis is included.

  11. Wind-tunnel/flight correlation study of aerodynamic characteristics of a large flexible supersonic cruise airplane (XB-701) 2: Extrapolation of wind-tunnel data to full-scale conditions

    NASA Technical Reports Server (NTRS)

    Peterson, J. B., Jr.; Mann, M. J.; Sorrells, R. B., III; Sawyer, W. C.; Fuller, D. E.

    1980-01-01

    The results of calculations necessary to extrapolate performance data on an XB-70-1 wind tunnel model to full scale at Mach numbers from 0.76 to 2.53 are presented. The extrapolation was part of a joint program to evaluate performance prediction techniques for large flexible supersonic airplanes similar to a supersonic transport. The extrapolation procedure included: interpolation of the wind tunnel data at the specific conditions of the flight test points; determination of the drag increments to be applied to the wind tunnel data, such as spillage drag, boundary layer trip drag, and skin friction increments; and estimates of the drag items not represented on the wind tunnel model, such as bypass doors, roughness, protuberances, and leakage drag. In addition, estimates of the effects of flexibility of the airplane were determined.

  12. Spin-Entry Characteristics of a Large Supersonic Bomber as Determined by Dynamic Model Tests

    NASA Technical Reports Server (NTRS)

    Bowman, James S.

    1965-01-01

    An investigation has been conducted in the Langley spin tunnel and at a catapult launch facility of a 1/60-scale dynamic model to determine the spin-entry characteristics of a large supersonic bomber. Catapult tests indicated that spin-entry motions were obtainable for a center-of-gravity location of 0.21 mean aerodynamic chord but were not obtainable at a center-of-gravity location of 0.25 mean aerodynamic chord. Deflected ailerons were effective in promoting or preventing the spin- entry motion and this effect was qualitatively the same as it was for the fully developed spin. Varying the configuration had little significant effect on the spin-entry characteristics. Brief tests conducted with the model in the Langley spin tunnel indicated that fully developed spins were obtainable at the forward center-of-gravity location and that spins were highly unlikely at the rearward center-of-location.

  13. General purpose computer program for interacting supersonic configurations. User's manual. [determining unsteady aerodynamic foreces

    NASA Technical Reports Server (NTRS)

    Crill, W.; Dale, B.

    1977-01-01

    The input data required to execute the computer program ISCON are described. The program generates a numerical procedure for the determination of unsteady aerodynamic forces on arbitrarily interacting wings and tails in supersonic flow. A velocity potential gradient method is used. Constant Mach number is assumed throughout the flow field. Lifting surfaces are represented by trapezoidal elements which can be generated automatically by the program. The wake field is represented by rectangular strip elements. The formulation is reviewed as well as input overview and input format. Instruction on how to use ISCON, a sample problem, and the restart feature are discussed. Program size limitations, computer program flow, and error messages are also included along with a description of the SS31 program used to compute the coefficients of surface spline.

  14. Aerodynamic Shape Optimization of Supersonic Aircraft Configurations via an Adjoint Formulation on Parallel Computers

    NASA Technical Reports Server (NTRS)

    Reuther, James; Alonso, Juan Jose; Rimlinger, Mark J.; Jameson, Antony

    1996-01-01

    This work describes the application of a control theory-based aerodynamic shape optimization method to the problem of supersonic aircraft design. The design process is greatly accelerated through the use of both control theory and a parallel implementation on distributed memory computers. Control theory is employed to derive the adjoint differential equations whose solution allows for the evaluation of design gradient information at a fraction of the computational cost required by previous design methods. The resulting problem is then implemented on parallel distributed memory architectures using a domain decomposition approach, an optimized communication schedule, and the MPI (Message Passing Interface) Standard for portability and efficiency. The final result achieves very rapid aerodynamic design based on higher order computational fluid dynamics methods (CFD). In our earlier studies, the serial implementation of this design method was shown to be effective for the optimization of airfoils, wings, wing-bodies, and complex aircraft configurations using both the potential equation and the Euler equations. In our most recent paper, the Euler method was extended to treat complete aircraft configurations via a new multiblock implementation. Furthermore, during the same conference, we also presented preliminary results demonstrating that this basic methodology could be ported to distributed memory parallel computing architectures. In this paper, our concern will be to demonstrate that the combined power of these new technologies can be used routinely in an industrial design environment by applying it to the case study of the design of typical supersonic transport configurations. A particular difficulty of this test case is posed by the propulsion/airframe integration.

  15. Aerodynamic Shape Optimization of Supersonic Aircraft Configurations via an Adjoint Formulation on Parallel Computers

    NASA Technical Reports Server (NTRS)

    Reuther, James; Alonso, Juan Jose; Rimlinger, Mark J.; Jameson, Antony

    1996-01-01

    This work describes the application of a control theory-based aerodynamic shape optimization method to the problem of supersonic aircraft design. The design process is greatly accelerated through the use of both control theory and a parallel implementation on distributed memory computers. Control theory is employed to derive the adjoint differential equations whose solution allows for the evaluation of design gradient information at a fraction of the computational cost required by previous design methods (13, 12, 44, 38). The resulting problem is then implemented on parallel distributed memory architectures using a domain decomposition approach, an optimized communication schedule, and the MPI (Message Passing Interface) Standard for portability and efficiency. The final result achieves very rapid aerodynamic design based on higher order computational fluid dynamics methods (CFD). In our earlier studies, the serial implementation of this design method (19, 20, 21, 23, 39, 25, 40, 41, 42, 43, 9) was shown to be effective for the optimization of airfoils, wings, wing-bodies, and complex aircraft configurations using both the potential equation and the Euler equations (39, 25). In our most recent paper, the Euler method was extended to treat complete aircraft configurations via a new multiblock implementation. Furthermore, during the same conference, we also presented preliminary results demonstrating that the basic methodology could be ported to distributed memory parallel computing architectures [241. In this paper, our concem will be to demonstrate that the combined power of these new technologies can be used routinely in an industrial design environment by applying it to the case study of the design of typical supersonic transport configurations. A particular difficulty of this test case is posed by the propulsion/airframe integration.

  16. Interaction of aerodynamic noise with laminar boundary layers in supersonic wind tunnels

    NASA Technical Reports Server (NTRS)

    Schopper, M. R.

    1984-01-01

    The interaction between incoming aerodynamic noise and the supersonic laminar boundary layer is studied. The noise field is modeled as a Mach wave radiation field consisting of discrete waves emanating from coherent turbulent entities moving downstream within the supersonic turbulent boundary layer. The individual disturbances are likened to miniature sonic booms and the laminar boundary layer is staffed by the waves as the sources move downstream. The mean, autocorrelation, and power spectral density of the field are expressed in terms of the wave shapes and their average arrival rates. Some consideration is given to the possible appreciable thickness of the weak shock fronts. The emphasis in the interaction analysis is on the behavior of the shocklets in the noise field. The shocklets are shown to be focused by the laminar boundary layer in its outer region. Borrowing wave propagation terminology, this region is termed the caustic region. Using scaling laws from sonic boom work, focus factors at the caustic are estimated to vary from 2 to 6 for incoming shocklet strengths of 1 to .01 percent of the free stream pressure level. The situation regarding experimental evidence of the caustic region is reviewed.

  17. Subsonic-to-Hypersonic Aerodynamic Characteristics for a Winged, Circular-Body, Single-Stage-to-Orbit Spacecraft Configuration

    NASA Technical Reports Server (NTRS)

    Phillips, W. P.; Engelund, W. C.

    1995-01-01

    Experimental aerodynamic characteristics were obtained for a generic, winged, circular-body, single-stage-to-orbit spacecraft configuration. The baseline configuration was longitudinally stable and trimmable at almost all Mach numbers from 0.15 to 10.0--with the exception occurring at low supersonic speeds. Landing speed and subsonic-to-hypersonic longitudinal stability and control appear to be within design guidelines. Lateral-directional instabilities found over the entire speed range, however, create a problem area for this configuration. Longitudinal aerodynamic predictions made utilizing the Aerodynamic Preliminary Analysis System (APAS) were in qualitative, often quantitative agreement with experimental values.

  18. Aerodynamic characteristics of the HL-20

    NASA Astrophysics Data System (ADS)

    Ware, George M.; Cruz, Christopher I.

    1993-09-01

    Wind tunnel tests were made from subsonic to hypersonic speeds to define the aerodynamic characteristics of the HL-20 lifting-body configuration. The data have been assembled into an aerodynamic database for flight analysis of this proposed vehicle. The wind tunnel data indicates that the model is longitudinally and laterally stable (about a center-of-gravity location of 0.54 body length) over the test range from Mach 20 to 0.3. At hypersonic speeds, the HL-20 model trimmed at a lift/drag (L/D) ratio of 1.4. This value gives the vehicle a crossrange capability similar to that of the space shuttle. At subsonic speeds, the HL-20 has a trimmed L/D ratio of about 3.6. Replacing the flat-plate outboard fins with fins having an airfoil shape increased the maximum subsonic trimmed L/D to 4.2.

  19. Improvement in Capsule Abort Performance Using Supersonic Aerodynamic Interaction by Fences

    NASA Astrophysics Data System (ADS)

    Koyama, Hiroto; Wang, Yunpeng; Ozawa, Hiroshi; Doi, Katsunori; Nakamura, Yoshiaki

    The space transportation system will need advanced abort systems to secure crew against serious accidents. Here this study deals with the capsule-type space transportation systems with a Launch Abort System (LAS). This system is composed of a conic capsule as a Launch Abort Vehicle (LAV) and a cylindrical rocket as a Service Module (SM), and the capsule is moved away from the rocket by supersonic aerodynamic interactions in an emergency. We propose a method to improve the performance of the LAV by installing fences at the edges of surfaces on the rocket and capsule sides. Their effects were investigated by experimental measurements and numerical simulations. Experimental results show that the fences on the rocket and capsule surfaces increase the aerodynamic thrust force on the capsule by 70% in a certain clearance between the capsule and rocket. Computational results show the detailed flow fields where the centripetal flow near the surface on the rocket side is induced by the fence on the rocket side and the centrifugal flow near the surface on the capsule side is blocked by the fence on the capsule side. These results can confirm favorable effects of the fences on the performance of the LAS.

  20. Comparison of Various Supersonic Turbine Tip Designs to Minimize Aerodynamic Loss and Tip Heating

    NASA Technical Reports Server (NTRS)

    Shyam, Vikram; Ameri, Ali

    2012-01-01

    The rotor tips of axial turbines experience high heat flux and are the cause of aerodynamic losses due to tip clearance flows, and in the case of supersonic tips, shocks. As stage loadings increase, the flow in the tip gap approaches and exceeds sonic conditions. This introduces effects such as shock-boundary layer interactions and choked flow that are not observed for subsonic tip flows that have been studied extensively in literature. This work simulates the tip clearance flow for a flat tip, a diverging tip gap and several contoured tips to assess the possibility of minimizing tip heat flux while maintaining a constant massflow from the pressure side to the suction side of the rotor, through the tip clearance. The Computational Fluid Dynamics (CFD) code GlennHT was used for the simulations. Due to the strong favorable pressure gradients the simulations assumed laminar conditions in the tip gap. The nominal tip gap width to height ratio for this study is 6.0. The Reynolds number of the flow is 2.4 x 10(exp 5) based on nominal tip width and exit velocity. A wavy wall design was found to reduce heat flux by 5 percent but suffered from an additional 6 percent in aerodynamic loss coefficient. Conventional tip recesses are found to perform far worse than a flat tip due to severe shock heating. Overall, the baseline flat tip was the second best performer. A diverging converging tip gap with a hole was found to be the best choice. Average tip heat flux was reduced by 37 percent and aerodynamic losses were cut by over 6 percent.

  1. Unsteady aerodynamics of missiles. Part 3: Determination of the longitudinal stability of wings at high angles of attack in supersonic flight

    NASA Astrophysics Data System (ADS)

    Schneider, C. P.

    1980-05-01

    A theoretical method for the determination of unsteady aerodynamic coefficients associated with the longitudinal stability of slender wings in supersonic flight is presented. It is based on the indicial functional theory of Tobak. Extension to higher incidences is effected by combining the indicial functions with steady nonlinear coefficients derived from a semiempiricial procedure. The unsteady nonlinear aerodynamic coefficients are determined for delta wings with subsonic and supersonic leading edges, respectively.

  2. Experimental investigation of the aerodynamic characteristics for a winged-cone concept

    NASA Technical Reports Server (NTRS)

    Phillips, W. Pelham; Brauckmann, Gregory J.; Micol, John R.; Woods, William C.

    1987-01-01

    Experimental longitudinal and lateral-directional aerodynamics were obtained for a generic aerodynamics were obtaiend for a generic winged-cone configuration having possible application as a transatmospheric vehicle concept. Data were obtained at Mach numbers from 0.6 to 20.0; Reynolds numbers, based on model length, between 2.5 and 5.3 million; and angles of attack from -4 to 20 deg. Results indicate a longitudinal center-of-pressure travel of about 23 percent of the fuselage length for the test Mach number range, with longitudinal instabilities noted at high-supersonic to hypersonic Mach numbers. These instabilities are coupled with directional instability at similar Mach numbers. Predictions with analytic codes, namely, the USAF DATCOM and the tangent-cone option of the Hypersonic Arbitrary Body Program, provided fair agreement with the experimental aerodynamic characteristics at low angles-of-attack.

  3. Aerodynamic characteristics of the Scout 133R vehicle determined from wind tunnel tests

    NASA Technical Reports Server (NTRS)

    Abramson, F. B.; Muir, T. G., Jr.; Simmons, H. L.

    1972-01-01

    Bending moments and other associated parameters were measured on a Scout vehicle during a launch through high velocity horizontal winds. Comparison of the measured data with predictions revealed some unexplained discrepancies. Possible sources of error in the experimental data and predictions were considered; one of which is the predicted aerodynamic characteristics. A wind tunnel investigation was initiated, including supersonic force and pressure tests, to better define the aerodynamics. In addition to basic aerodynamic coefficients from the force test, detailed pressure and load distributions along the body were established from the pressure test. Pressure coefficients were integrated to determine normal load distributions, total normal force, and total pitching moment of the body. Comparison of the normal forces from pressure and force tests resulted in agreement within 15%. Comparison of pitching moment data from the two tests resulted in larger differences.

  4. Aerodynamic Characteristics and Glide-Back Performance of Langley Glide-Back Booster

    NASA Technical Reports Server (NTRS)

    Pamadi, Bandu N.; Covell, Peter F.; Tartabini, Paul V.; Murphy, Kelly J.

    2004-01-01

    NASA-Langley Research Center is conducting system level studies on an-house concept of a small launch vehicle to address NASA's needs for rapid deployment of small payloads to Low Earth Orbit. The vehicle concept is a three-stage system with a reusable first stage and expendable upper stages. The reusable first stage booster, which glides back to launch site after staging around Mach 3 is named the Langley Glide-Back Booster (LGBB). This paper discusses the aerodynamic characteristics of the LGBB from subsonic to supersonic speeds, development of the aerodynamic database and application of this database to evaluate the glide back performance of the LGBB. The aerodynamic database was assembled using a combination of wind tunnel test data and engineering level analysis. The glide back performance of the LGBB was evaluated using a trajectory optimization code and subject to constraints on angle of attack, dynamic pressure and normal acceleration.

  5. Aerodynamic yawing moment characteristics of bird wings.

    PubMed

    Sachs, Gottfried

    2005-06-21

    The aerodynamic yawing moments due to sideslip are considered for wings of birds. Reference is made to the experience with aircraft wings in order to identify features which are significant for the yawing moment characteristics. Thus, it can be shown that wing sweep, aspect ratio and lift coefficient have a great impact. Focus of the paper is on wing sweep which can considerably increase the yawing moment due to sideslip when compared with unswept wings. There are many birds the wings of which employ sweep. To show the effect of sweep for birds, the aerodynamic characteristics of a gull wing which is considered as a representative example are treated in detail. For this purpose, a sophisticated aerodynamic method is used to compute results of high precision. The yawing moments of the gull wing with respect to the sideslip angle and the lift coefficient are determined. They show a significant level of yaw stability which strongly increases with the lift coefficient. It is particularly high in the lift coefficient region of best gliding flight conditions. In order to make the effect of sweep more perspicuous, a modification of the gull wing employing no sweep is considered for comparison. It turns out that the unswept wing yields yawing moments which are substantially smaller than those of the original gull wing with sweep. Another feature significant for the yawing moment characteristics concerns the fact that sweep is at the outer part of bird wings. By considering the underlying physical mechanism, it is shown that this feature is most important for the efficiency of wing sweep. To sum up, wing sweep provides a primary contribution to the yawing moments. It may be concluded that this is an essential reason why there is sweep in bird wings.

  6. A study of the motion and aerodynamic heating of ballistic missiles entering the earth's atmosphere at high supersonic speeds

    NASA Technical Reports Server (NTRS)

    Allen, H Julian; Eggers, A J , Jr

    1958-01-01

    A simplified analysis of the velocity and deceleration history of ballistic missiles entering the earth's atmosphere at high supersonic speeds is presented. The results of this motion analysis are employed to indicate means available to the designer for minimizing aerodynamic heating. The heating problem considered involves not only the total heat transferred to a missile by convection, but also the maximum average and local time rates of convective heat transfer.

  7. Unstructured Grid Euler Method Assessment for Longitudinal and Lateral/Directional Aerodynamic Performance Analysis of the HSR Technology Concept Airplane at Supersonic Cruise Speed

    NASA Technical Reports Server (NTRS)

    Ghaffari, Farhad

    1999-01-01

    Unstructured grid Euler computations, performed at supersonic cruise speed, are presented for a High Speed Civil Transport (HSCT) configuration, designated as the Technology Concept Airplane (TCA) within the High Speed Research (HSR) Program. The numerical results are obtained for the complete TCA cruise configuration which includes the wing, fuselage, empennage, diverters, and flow through nacelles at M (sub infinity) = 2.4 for a range of angles-of-attack and sideslip. Although all the present computations are performed for the complete TCA configuration, appropriate assumptions derived from the fundamental supersonic aerodynamic principles have been made to extract aerodynamic predictions to complement the experimental data obtained from a 1.675%-scaled truncated (aft fuselage/empennage components removed) TCA model. The validity of the computational results, derived from the latter assumptions, are thoroughly addressed and discussed in detail. The computed surface and off-surface flow characteristics are analyzed and the pressure coefficient contours on the wing lower surface are shown to correlate reasonably well with the available pressure sensitive paint results, particularly, for the complex flow structures around the nacelles. The predicted longitudinal and lateral/directional performance characteristics for the truncated TCA configuration are shown to correlate very well with the corresponding wind-tunnel data across the examined range of angles-of-attack and sideslip. The complementary computational results for the longitudinal and lateral/directional performance characteristics for the complete TCA configuration are also presented along with the aerodynamic effects due to empennage components. Results are also presented to assess the computational method performance, solution sensitivity to grid refinement, and solution convergence characteristics.

  8. Aerodynamic Shape Optimization of a Dual-Stream Supersonic Plug Nozzle

    NASA Technical Reports Server (NTRS)

    Heath, Christopher M.; Gray, Justin S.; Park, Michael A.; Nielsen, Eric J.; Carlson, Jan-Renee

    2015-01-01

    Aerodynamic shape optimization was performed on an isolated axisymmetric plug nozzle sized for a supersonic business jet. The dual-stream concept was tailored to attenuate nearfield pressure disturbances without compromising nozzle performance. Adjoint-based anisotropic mesh refinement was applied to resolve nearfield compression and expansion features in the baseline viscous grid. Deformed versions of the adapted grid were used for subsequent adjoint-driven shape optimization. For design, a nonlinear gradient-based optimizer was coupled to the discrete adjoint formulation of the Reynolds-averaged Navier- Stokes equations. All nozzle surfaces were parameterized using 3rd order B-spline interpolants and perturbed axisymmetrically via free-form deformation. Geometry deformations were performed using 20 design variables shared between the outer cowl, shroud and centerbody nozzle surfaces. Interior volume grid deformation during design was accomplished using linear elastic mesh morphing. The nozzle optimization was performed at a design cruise speed of Mach 1.6, assuming core and bypass pressure ratios of 6.19 and 3.24, respectively. Ambient flight conditions at design were commensurate with 45,000-ft standard day atmosphere.

  9. Study of aerodynamic noise in low supersonic operation of an axial flow compressor

    NASA Technical Reports Server (NTRS)

    Arnoldi, R. A.

    1972-01-01

    A study of compressor noise is presented, based upon supersonic, part-speed operation of a high hub/tip ratio compressor designed for spanwise uniformity of aerodynamic conditions, having straight cylindrical inlet and exit passages for acoustic simplicity. Acoustic spectra taken in the acoustically-treated inlet plenum, are presented for five operating points at each of two speeds, corresponding to relative rotor tip Mach numbers of about 1.01 and 1.12 (60 and 67 percent design speed). These spectra are analyzed for low and high frequency broadband noise, blade passage frequency noise, combination tone noise and "haystack' noise (a very broad peak somewhat below blade passage frequency, which is occasionally observed in engines and fan test rigs). These types of noise are related to diffusion factor, total pressure ratio, and relative rotor tip Mach number. Auxiliary measurements of fluctuating wall static pressures and schlieren photographs of upstream shocks in the inlet are also presented and related to the acoustic and performance data.

  10. Aerodynamic characteristics of airplanes at high angles of attack

    NASA Technical Reports Server (NTRS)

    Chambers, J. R.; Grafton, S. B.

    1977-01-01

    An introduction to, and a broad overiew of, the aerodynamic characteristics of airplanes at high angles of attack are provided. Items include: (1) some important fundamental phenomena which determine the aerodynamic characteristics of airplanes at high angles of attack; (2) static and dynamic aerodynamic characteristics near the stall; (3) aerodynamics of the spin; (4) test techniques used in stall/spin studies; (5) applications of aerodynamic data to problems in flight dynamics in the stall/spin area; and (6) the outlook for future research in the area. Although stalling and spinning are flight dynamic problems of importance to all aircraft, including general aviation aircraft, commercial transports, and military airplanes, emphasis is placed on military configurations and the principle aerodynamic factors which influence the stability and control of such vehicles at high angles of attack.

  11. The Practical Calculation of the Aerodynamic Characteristics of Slender Finned Vehicles

    NASA Technical Reports Server (NTRS)

    Barrowman, James S.

    1967-01-01

    The basic objective of this thesis is to provide a practical method of computing the aerodynamic characteristics of slender finned vehicles such as sounding rockets, high speed bombs, and guided missiles. The aerodynamic characteristics considered are the normal force coefficient derivative, c(sub N(sub alpha)); center of pressure, bar-X; roll forcing moment coefficient derivative, c(sub l(sub delta)); roll damping moment coefficient derivative, c(sub l(sub p)); pitch damping moment coefficient derivative, c(sub mq); and the drag coefficient, c (sub D). Equations are determined for both subsonic and supersonic flow. No attempts is made to analyze the transonic region. The general configuration to which the relations are applicable is a slender axisymmetric body having three or four fins.

  12. Numerical Investigation of Aerodynamics of Canard-Controlled Missile Using Planar and Grid Tail Fins. Part 1. Supersonic Flow

    NASA Astrophysics Data System (ADS)

    DeSpirito, James; Vaughn, Milton E., Jr.; Washington, W. D.

    2002-09-01

    Viscous computational fluid dynamic simulations were used to predict the aerodynamic coefficients and flowfield around a generic canard-controlled missile configuration in supersonic flow. Computations were performed for Mach 1.5 and 3.0, at six angles of attack between 0 and 10, with 0 and 10 canard deflection, and with planar and grid tail fins, for a total of 48 cases. Validation of the computed results was demonstrated by the very good agreement between the computed aerodynamic coefficients and those obtained from wind tunnel measurements. Visualizations of the flowfield showed that the canard trailing vortices and downwash produced a low-pressure region on the starboard side of the missile that in turn produced an adverse side force. The pressure differential on the leeward fin produced by the interaction with the canard trailing vortices is primarily responsible for the adverse roll effect observed when planar fins are used. Grid tail fins improved the roll effectiveness of the canards at low supersonic speed. No adverse rolling moment was observed with no canard deflection, or at the higher supersonic speed for either tail fin type due to the lower intensity of the canard trailing vortices in these cases. Flow visualizations from the simulations performed in this study help in the understanding of the flow physics and can lead to improved canard and tail fin designs for missiles and rockets.

  13. Investigation of flow characteristics over missile bodies at supersonic speeds

    NASA Technical Reports Server (NTRS)

    Barger, R. L.; Sawyer, W. C.

    1979-01-01

    Three missile body shapes tested at Mach numbers of 1.50, 2.16, and 2.86 with angles of attack up to 30 degrees are described. The flow characteristics for each body shape are examined. The measured aerodynamic forces and moments are presented. The use of flow visualization techniques are described and the results such as vortex effects are discussed.

  14. Unsteady Aerodynamic Analysis of Supersonic Through-Flow Fan with Vibrating Blades Under Non-Zero Mean Loading

    NASA Astrophysics Data System (ADS)

    Hanada, T.; Namba, M.

    1996-08-01

    The double linearization concept is applied to a rotating annular cascade model operating at supersonic axial velocity. It is assumed that each blade vibrates with infinitesimal displacement amplitude under small but non-zero mean loading. Vibration modes normal and parallel to the blade chord are considered. Numerical results indicate that the mean loading effects play a crucial role on the aerodynamic instability of the blade motion. The bending motion can be unstable due to the presence of mean loading. Both the steady performance and the flutter boundary are highly sensitive to the blade camber. The bending motion instability is substantially influenced also by the chordwise component of the blade motion. Some numerical results compared with strip theory prediction demonstrate significant three-dimensional effects on the unsteady aerodynamic force under non-zero mean loading.

  15. Propulsion and airframe aerodynamic interactions of supersonic V/STOL configurations. Volume 1: Wind tunnel test pressure data report

    NASA Technical Reports Server (NTRS)

    Zilz, D. E.; Devereaux, P. A.

    1985-01-01

    A wind tunnel model of a supersonic V/STOL fighter configuration has been tested to measure the aerodynamic interaction effects which can result from geometrically close-coupled propulsion system/airframe components. The approach was to configure the model to represent two different test techniques. One was a conventional test technique composed of two test modes. In the Flow-Through mode, absolute configuration aerodynamics are measured, including inlet/airframe interactions. In the Jet-Effects mode, incremental nozzle/airframe interactions are measured. The other test technique is a propulsion simulator approach, where a sub-scale, externally powered engine is mounted in the model. This allows proper measurement of inlet/airframe and nozzle/airframe interactions simultaneously. This is Volume 1 of 2: Wind Tunnel Test Pressure Data Report.

  16. Study on the characteristics of supersonic Coanda jet

    NASA Astrophysics Data System (ADS)

    Matsuo, Shigeru; Setoguchi, Toshiaki; Kudo, Takemasa; Yu, Shen

    1998-09-01

    Techniques using Coanda effect have been applied to the fluid control devices. In this field, experimental studies were so far performed for the spiral jet obtained by the Coanda jet issuing from a conical cylinder with an annular slit, thrust vectoring of supersonic Coanda jets and so on. It is important from the viewpoints of effective applications to investigate the characteristics of the supersonic Coanda jet in detail. In the present study, the effects of pressure ratios and nozzle configurations on the characteristics of the supersonic Coanda jet have been investigated experimentally by a schlieren optical method and pressure measurements. Furthermore, Navier-Stokes equations were solved numerically using a 2nd-order TVD finite-volume scheme with a 3rd-order three stage Runge-Kutta method for time integration. k - ɛ model was used in the computations. The effects of initial conditions on Coanda flow were investigated numerically. As a result, the simulated flow fields were compared with experimental data in good agreement qualitatively.

  17. Aeroelastic characteristics of a cascade of mistuned blades in subsonic and supersonic flows. [turbofan engines

    NASA Technical Reports Server (NTRS)

    Kielb, R. E.; Kaza, K. R. V.

    1981-01-01

    The effects of mistuning on flutter and forced response of a cascade in subsonic in subsonic and supersonic flow were investigated. The aerodynamic and structural coupling between the bending and torsional motions and the aerodynamic coupling between the blades were studied. It is shown that frequency mistuning always has a beneficial effect on flutter. For the cascade considered, the potential for raising flutter speed is greater in subsonic than in supersonic flow. Preliminary results for structural damping mistuning show that there are no additional benefits over adding damping mistuning may have either a beneficial or an adverse effect on forced response, depending on the engine order of the excitation and Mach number.

  18. Predicted aerodynamic characteristics for HL-20 lifting-body using the aerodynamic preliminary analysis system (APAS)

    NASA Technical Reports Server (NTRS)

    Cruz, Christopher I.; Ware, George M.

    1992-01-01

    The aerodynamic characteristics of the HL-20 lifting body configuraiton obtained through the APAS and from wind-tunnel tests have been compared. The APAS is considered to be an easy-to-use, relatively simple tool for quick preliminary estimation of vehicle aerodynamics. The APAS estimates are found to be in good agreement with experimental results to be used for preliminary evaluation of the HL-20. The APAS accuracy in predicting aerodynamics of the HL-20 varied over the Mach range. The speed ranges of best agreement were subsonic and hypersonic, while least agreement was in the Mach range from 1.2 to about 2,5.

  19. Computer programs for calculating the static longitudinal aerodynamic characteristics of wing-body-tail configurations

    NASA Technical Reports Server (NTRS)

    Mendenhall, M. R.; Goodwin, F. K.; Dillenius, M. F. E.; Kline, D. M.

    1975-01-01

    Four computer programs developed to calculate the longitudinal aerodynamic characteristics of wing-body and wing-body-tail combinations are presented. The R1307 program is based on a linear method and is limited to the small range of angles of attack for which the lift and moment characteristics of wings and bodies are linear with angle of attack. The CRSFLW program is based on a crossflow method of predicting the forces and moments on bodies alone or wing-body combinations over a large angle of attack range. The SUBSON program predicts the longitudinal aerodynamic characteristics of wing-body-tail combinations at subsonic speeds and at angles of attack for which symmetrical pairs of vortices are shed from the body nose and the leading and side edges of the lifting surfaces. Program SUPSON predicts the longitudinal aerodynamic characteristics of wing-body-tail combinations at supersonic speeds in the same angle-of-attack range. A description of the use of each program, instructions for preparation of input, a description of the output, program listings, and sample cases for each program are included.

  20. Advanced supersonic cruise aircraft technology

    NASA Technical Reports Server (NTRS)

    Baber, H. T., Jr.; Driver, C.

    1977-01-01

    A multidiscipline approach is taken to the application of the latest technology to supersonic cruise aircraft concept definition, and current problem areas are identified. Particular attention is given to the performance of the AST-100 advanced supersonic cruise vehicle with emphasis on aerodynamic characteristics, noise and chemical emission, and mission analysis. A recently developed aircraft sizing and performance computer program was used to determine allowable wing loading and takeoff gross weight sensitivity to structural weight reduction.

  1. Burning of the Supersonic Propane-Air Mixture in the Aerodynamic Channel With the Stagnant Zone

    DTIC Science & Technology

    2007-11-02

    V.Chernikov, V.Shibkov, O.Surkont. Mechanisms of transversal electric discharge sustention in supersonic air and propane-air flows. -American Institute of Aeronautics and Astronautics, AIAA Paper, 2003, No.03-0872, p. 1 -6 .

  2. Aerodynamic design and analysis system for supersonic aircraft. Part 2: User's manual

    NASA Technical Reports Server (NTRS)

    Middleton, W. D.; Lundry, J. L.; Coleman, R. G.

    1975-01-01

    An integrated system of computer programs for supersonic configurations is described. An explanation of system usage, the input definitions, and example output are included. For Part 1, see N75-18185; for Part 3, see N75-18186.

  3. High Speed Aerodynamic Characteristics of the GAF0PH Aerofoil

    DTIC Science & Technology

    1980-09-01

    upper surface of the aerofoil for angles of incidence greater than 210. POSTAL ADDRESS: Chief Superintendent, Aeronautical Research Laboratories, Box...kCLAERO-.NOTE3 98 -AR-002-223 -LEVEL m DEPARTMENT OF DEFENCE 00 DEFENCE SCIENCE AND TECHNOLOGY ORGANISATION AERONAUTICAL RESEARCH LABORATORIES...MELBOURNE, VICTORIA AERODYNAMICS NOTE 398 ’,\\ HIGH SPEED AERODYNAMIC CHARACTERISTICS OF THE GAFPH AEROFOIL by ~B D :, . , .IR-© Approved for Public Release

  4. Subsonic, transonic, and supersonic stability and control characteristics of the -147B space shuttle orbiter

    NASA Technical Reports Server (NTRS)

    Mennell, R. C.

    1973-01-01

    Experimental aerodynamic investigations were conducted on 0.015 scale representations of two Space Shuttle Orbiter configurations in a trisonic wind tunnel from June 20, 1973 to June 30, 1973. The primary test objective was to define subsonic, transonic, and supersonic stability and control characteristics of the -147B Orbiter. Six-component aerodynamic force and moment data for the -147B Orbiter were recorded over an angle of attack range of -2 deg to 30 deg at Mach numbers of 0.6, 0.9, 1.2, 2.0, and 3.0. Reynolds numbers of 5.0, 7.0, 8.0, and 9.0 x 100000 6/ft were tested at Mach numbers less than 2.0 while testing at Mach 2.0 and 3.0 was conducted at a Reynolds number of 11.0 x 100000/ft. Eleven deflections of 0 deg, +15 deg, -20, deg and -40 deg; body flap deflections of 0 deg, +13.75 deg and -14.25 deg; and rudder flare angles of 24.92 deg and 54.92 deg were tested on the -147B Orbiter over the entire Mach number range. Testing of the -139B Orbiter was for data verification and configuration comparison purposes only.

  5. Influence of airfoil geometry on delta wing leading-edge vortices and vortex-induced aerodynamics at supersonic speeds

    NASA Technical Reports Server (NTRS)

    Wood, Richard M.; Byrd, James E.; Wesselmann, Gary F.

    1992-01-01

    An assessment of the influence of airfoil geometry on delta wing leading edge vortex flow and vortex induced aerodynamics at supersonic speeds is discussed. A series of delta wing wind tunnel models were tested over a Mach number range from 1.7 to 2.0. The model geometric variables included leading edge sweep and airfoil shape. Surface pressure data, vapor screen, and oil flow photograph data were taken to evaluate the complex structure of the vortices and shocks on the family of wings tested. The data show that airfoil shape has a significant impact on the wing upper surface flow structure and pressure distribution, but has a minimal impact on the integrated upper surface pressure increments.

  6. Comparison of Theoretical and Experimental Unsteady Aerodynamics of Linear Oscillating Cascade With Supersonic Leading-Edge Locus

    NASA Technical Reports Server (NTRS)

    Ramsey, John K.; Erwin, Dan

    2004-01-01

    An experimental influence coefficient technique was used to obtain unsteady aerodynamic influence coefficients and, consequently, unsteady pressures for a cascade of symmetric airfoils oscillating in pitch about mid-chord. Stagger angles of 0 deg and 10 deg were investigated for a cascade with a gap-to-chord ratio of 0.417 operating at an axial Mach number of 1.9, resulting in a supersonic leading-edge locus. Reduced frequencies ranged from 0.056 to 0.2. The influence coefficients obtained determine the unsteady pressures for any interblade phase angle. The unsteady pressures were compared with those predicted by several algorithms for interblade phase angles of 0 deg and 180 deg.

  7. Effects of Nozzle Geometry and Intermittent Injection of Aerodynamic Tab on Supersonic Jet Noise

    NASA Astrophysics Data System (ADS)

    Araki, Mikiya; Sano, Takayuki; Fukuda, Masayuki; Kojima, Takayuki; Taguchi, Hideyuki; Shiga, Seiichi; Obokata, Tomio

    Effects of the nozzle geometry and intermittent injection of aerodynamic tabs on exhaust noise from a rectangular plug nozzle were investigated experimentally. In JAXA (Japan Aerospace Exploration Agency), a pre-cooled turbojet engine for an HST (Hypersonic transport) is planned. A 1/100-scaled model of the rectangular plug nozzle is manufactured, and the noise reduction performance of aerodynamic tabs, which is small air jet injection from the nozzle wall, was investigated. Compressed air is injected through the rectangular plug nozzle into the atmosphere at the nozzle pressure ratio of 2.7, which corresponds to the take-off condition of the vehicle. Aerodynamic tabs were installed at the sidewall ends, and 4 kinds of round nozzles and 2 kinds of wedge nozzles were applied. Using a high-frequency solenoid valve, intermittent gas injection is also applied. It is shown that, by use of wedge nozzles, the aerodynamic tab mass flow rate, necessary to gain 2.3dB reduction in OASPL (Overall sound pressure level), decreases by 29% when compared with round nozzles. It is also shown that, by use of intermittent injection, the aerodynamic tab mass flow rate, necessary to gain 2.3dB reduction in OASPL, decreases by about 40% when compared with steady injection. By combination of wedge nozzles and intermittent injection, the aerodynamic tab mass flow rate significantly decreases by 57% when compared with the conventional strategy.

  8. Experimental transonic flutter characteristics of supersonic cruise configurations

    NASA Technical Reports Server (NTRS)

    Durham, Michael H.; Cole, Stanley R.; Cazier, F. W., Jr.; Keller, Donald F.; Parker, Ellen C.; Wilkie, W. Keats

    1990-01-01

    The flutter characteristics of a generic arrow-wing supersonic transport configuration are studied. The wing configuration has a 3 percent biconvex airfoil and a leading-edge sweep of 73 deg out to a cranked tip with a 60 deg leading-edge sweep. The ground vibration tests and flutter test procedure are described. The effects of flutter on engine nacelles, fuel loading, wing-mounted vertical fin, wing angle-of-attack, and wing tip mass and stiffness distributions are analyzed. The data reveal that engine nacelles reduce the transonic flutter dynamic pressure by 25-30 percent; fuel loadings decrease dynamic pressures by 25 percent; 4-6 deg wing angles-of-attack cause steep transonic boundaries; and 5-10 percent changes in flutter dynamic pressures are the result of the wing-mounted vertical fin and wing-tip mass and stiffness distributions.

  9. Upper surface blowing aerodynamic and acoustic characteristics

    NASA Technical Reports Server (NTRS)

    Ryle, D. M., Jr.; Braden, J. A.; Gibson, J. S.

    1977-01-01

    Aerodynamic performance at cruise, and noise effects due to variations in nacelle and wing geometry and mode of operation are studied using small aircraft models that simulate upper surface blowing (USB). At cruise speeds ranging from Mach .50 to Mach .82, the key determinants of drag/thrust penalties are found to be nozzle aspect ratio, boattailing angle, and chordwise position; number of nacelles; and streamlined versus symmetric configuration. Recommendations are made for obtaining favorable cruise configurations. The acoustic studies, which concentrate on the noise created by the jet exhaust flow and its interaction with wing and flap surfaces, isolate several important sources of USB noise, including nozzle shape, exit velocity, and impingement angle; flow pathlength; and flap angle and radius of curvature. Suggestions for lessening noise due to trailing edge flow velocity, flow pathlength, and flow spreading are given, though compromises between some design options may be necessary.

  10. Aerodynamic Characteristic of the Active Compliant Trailing Edge Concept

    NASA Astrophysics Data System (ADS)

    Nie, Rui; Qiu, Jinhao; Ji, Hongli; Li, Dawei

    2016-06-01

    This paper introduces a novel Morphing Wing structure known as the Active Compliant Trailing Edge (ACTE). ACTE structures are designed using the concept of “distributed compliance” and wing skins of ACTE are fabricated from high-strength fiberglass composites laminates. Through the relative sliding between upper and lower wing skins which are connected by a linear guide pairs, the wing is able to achieve a large continuous deformation. In order to present an investigation about aerodynamics and noise characteristics of ACTE, a series of 2D airfoil analyses are established. The aerodynamic characteristics between ACTE and conventional deflection airfoil are analyzed and compared, and the impacts of different ACTE structure design parameters on aerodynamic characteristics are discussed. The airfoils mentioned above include two types (NACA0012 and NACA64A005.92). The computing results demonstrate that: compared with the conventional plane flap airfoil, the morphing wing using ACTE structures has the capability to improve aerodynamic characteristic and flow separation characteristic. In order to study the noise level of ACTE, flow field analysis using LES model is done to provide noise source data, and then the FW-H method is used to get the far field noise levels. The simulation results show that: compared with the conventional flap/aileron airfoil, the ACTE configuration is better to suppress the flow separation and lower the overall sound pressure level.

  11. Development of the Orion Crew Module Static Aerodynamic Database. Par 2; Supersonic/Subsonic

    NASA Technical Reports Server (NTRS)

    Bibb, Karen L.; Walker, Eric L.; Brauckmann, Gregory J.; Robinson, Phil

    2011-01-01

    This work describes the process of developing the nominal static aerodynamic coefficients and associated uncertainties for the Orion Crew Module for Mach 8 and below. The database was developed from wind tunnel test data and computational simulations of the smooth Crew Module geometry, with no asymmetries or protuberances. The database covers the full range of Reynolds numbers seen in both entry and ascent abort scenarios. The basic uncertainties were developed as functions of Mach number and total angle of attack from variations in the primary data as well as computations at lower Reynolds numbers, on the baseline geometry, and using different flow solvers. The resulting aerodynamic database represents the Crew Exploration Vehicle Aerosciences Project's best estimate of the nominal aerodynamics for the current Crew Module vehicle.

  12. Aeroacoustic Characteristics of a Rectangular Multi-Element Supersonic Jet Mixer-Ejector Nozzle

    NASA Technical Reports Server (NTRS)

    Raman, Ganesh; Taghavi, Ray

    1996-01-01

    This paper provides a unique, detailed evaluation of the acoustics and aerodynamics of a rectangular multi-element supersonic jet mixer-ejector noise suppressor. The performance of such mixer-ejectors is important in aircraft engine application for noise suppression and thrust augmentation. In contrast to most prior experimental studies on ejectors that reported either aerodynamic or acoustic data, our work documents both types of data. We present information on the mixing, pumping, ejector wall pressure distribution, thrust augmentation and noise suppression characteristics of four simple, multi-element, jet mixer-ejector configurations. The four configurations included the effect of ejector area ratio (AR = ejector area/primary jet area) and the effect of non-parallel ejector walls. We also studied in detail the configuration that produced the best noise suppression characteristics. Our results show that ejector configurations that produced the maximum maximum pumping (entrained flow per secondary inlet area) also exhibited the lowest wall pressures in the inlet region, and the maximum thrust augmentation. When cases having the same total mass flow were compared, we found that noise suppression trends corresponded with those for pumping. Surprisingly, the mixing (quantified by the peak Mach number, and flow uniformity) at the ejector exit exhibited no relationship to the noise suppression at moderate primary jet fully expanded Mach numbers (Mj is less than 1.4). However, the noise suppression dependence on the mixing was apparent at higher Mj. The above observations are justified by noting that the mixing at the ejector exit is ot a strong factor in determining the radiated noise when noise produced internal to the ejector dominates the noise field outside the ejector.

  13. Integral-equation methods in steady and unsteady subsonic, transonic and supersonic aerodynamics for interdisciplinary design

    NASA Technical Reports Server (NTRS)

    Yates, E. Carson, Jr.

    1990-01-01

    Progress in the development of computational methods for steady and unsteady aerodynamics has perennially paced advancements in aeroelastic analysis and design capabilities. Since these capabilities are of growing importance in the analysis and design of high-performance aircraft, considerable effort has been directed toward the development of appropriate aerodynamic methodology. The contributions to those efforts from the integral-equations research program at the NASA Langley Research Center is reviewed. Specifically, the current scope, progress, and plans for research and development for inviscid and viscous flows are discussed, and example applications are shown in order to highlight the generality, versatility, and attractive features of this methodology.

  14. Program LRCDM2: Improved aerodynamic prediction program for supersonic canard-tail missiles with axisymmetric bodies

    NASA Technical Reports Server (NTRS)

    Dillenius, Marnix F. E.

    1985-01-01

    Program LRCDM2 was developed for supersonic missiles with axisymmetric bodies and up to two finned sections. Predicted are pressure distributions and loads acting on a complete configuration including effects of body separated flow vorticity and fin-edge vortices. The computer program is based on supersonic panelling and line singularity methods coupled with vortex tracking theory. Effects of afterbody shed vorticity on the afterbody and tail-fin pressure distributions can be optionally treated by companion program BDYSHD. Preliminary versions of combined shock expansion/linear theory and Newtonian/linear theory have been implemented as optional pressure calculation methods to extend the Mach number and angle-of-attack ranges of applicability into the nonlinear supersonic flow regime. Comparisons between program results and experimental data are given for a triform tail-finned configuration and for a canard controlled configuration with a long afterbody for Mach numbers up to 2.5. Initial tests of the nonlinear/linear theory approaches show good agreement for pressures acting on a rectangular wing and a delta wing with attached shocks for Mach numbers up to 4.6 and angles of attack up to 20 degrees.

  15. Aerodynamic Characteristics of a Model of an Inflatable-Sphere Launching Vehicle under Simulated Conditions of Mach Number and Altitude

    NASA Technical Reports Server (NTRS)

    Robinson, Ross B.; Morris, Odell A.

    1960-01-01

    An investigation has been conducted in the Langley 4- by 4-foot supersonic pressure tunnel to determine the aerodynamic characteristics in pitch of a two-stage-rocket model configuration which simulated the last two stages of the launching vehicle for an inflatable sphere. Tests were made through an angle-of-attack range from -6 deg to 18 deg at dynamic pressures of 102 and 255 pounds per square foot with corresponding Mach numbers of 1.89 and 1.98 for the model both with and without a bumper arrangement designed to protect the rocket casing from the outer shell of the vehicle.

  16. Aerodynamic Characteristics of Water Rocket and Stabilization of Flight Trajectory

    NASA Astrophysics Data System (ADS)

    Watanabe, Rikio; Tomita, Nobuyuki; Takemae, Toshiaki

    The aerodynamic characteristics of water rockets are analyzed experimentally by wind tunnel testing. Aerodynamic devices such as vortex generators and dimples are tested and their effectiveness to the flight performance of water rocket is discussed. Attaching vortex generators suppresses the unsteady body fluttering. Dimpling the nose reduces the drag coefficient in high angles of attack. Robust design approach is applied to water rocket design for flight stability and optimum water rocket configuration is determined. Semi-sphere nose is found to be effective for flight stability and it is desirable for the safety of landing point. Stiffed fin attachment is required for fins to work properly as aerodynamic device and it enhances the flight stability of water rockets.

  17. Aerodynamic Characteristics of Telescopic Aerospikes with Multiple-Row-Disk

    NASA Astrophysics Data System (ADS)

    Kobayashi, Hiroaki; Maru, Yusuke; Sato, Tetsuya

    This paper reports experimental studies on telescopic aerospikes with multiple disks. The telescopic aerospike is useful as an aerodynamic control device; however, changing its length causes a buzz phenomenon, which many researchers have reported. The occurrence of buzzing might be critical to the vehicle because it brings about severe pressure oscillations on the surface. Disks on the shaft produce stable recirculation regions by dividing the single separation flow into several conical cavity flows. The telescopic aerospikes with stabilizer disks are useful without any length constraints. Aerodynamic characteristics of the telescopic aerospikes were investigated through a series of wind tunnel tests. Transition of recirculation/reattachment flow modes of a plain spike causes a large change in the drag coefficient. Because of this hysteresis phenomenon and the buzzing, the plain spike is unsuitable for fine aerodynamic control devices. Adding stabilizer disks is effective for the improved control of aerospikes.

  18. Techniques for estimating Space Station aerodynamic characteristics

    NASA Technical Reports Server (NTRS)

    Thomas, Richard E.

    1993-01-01

    A method was devised and calculations were performed to determine the effects of reflected molecules on the aerodynamic force and moment coefficients for a body in free molecule flow. A procedure was developed for determining the velocity and temperature distributions of molecules reflected from a surface of arbitrary momentum and energy accommodation. A system of equations, based on momentum and energy balances for the surface, incident, and reflected molecules, was solved by a numerical optimization technique. The minimization of a 'cost' function, developed from the set of equations, resulted in the determination of the defining properties of the flow reflected from the arbitrary surface. The properties used to define both the incident and reflected flows were: average temperature of the molecules in the flow, angle of the flow with respect to a vector normal to the surface, and the molecular speed ratio. The properties of the reflected flow were used to calculate the contribution of multiply reflected molecules to the force and moments on a test body in the flow. The test configuration consisted of two flat plates joined along one edge at a right angle to each other. When force and moment coefficients of this 90 deg concave wedge were compared to results that did not include multiple reflections, it was found that multiple reflections could nearly double lift and drag coefficients, with nearly a 50 percent increase in pitching moment for cases with specular or nearly specular accommodation. The cases of diffuse or nearly diffuse accommodation often had minor reductions in axial and normal forces when multiple reflections were included. There were several cases of intermediate accommodation where the addition of multiple reflection effects more than tripled the lift coefficient over the convex technique.

  19. Supersonic compressor

    DOEpatents

    Roberts, II, William Byron; Lawlor, Shawn P.; Breidenthal, Robert E.

    2016-04-12

    A supersonic compressor including a rotor to deliver a gas at supersonic conditions to a diffuser. The diffuser includes a plurality of aerodynamic ducts that have converging and diverging portions, for deceleration of gas to subsonic conditions and then for expansion of subsonic gas, to change kinetic energy of the gas to static pressure. The aerodynamic ducts include vortex generating structures for controlling boundary layer, and structures for changing the effective contraction ratio to enable starting even when the aerodynamic ducts are designed for high pressure ratios, and structures for boundary layer control. In an embodiment, aerodynamic ducts are provided having an aspect ratio of in excess of two to one, when viewed in cross-section orthogonal to flow direction at an entrance to the aerodynamic duct.

  20. Bumblebee program, aerodynamic data. Part 2: Flow fields at Mach number 2.0. [supersonic missiles

    NASA Technical Reports Server (NTRS)

    Barnes, G. A.; Cronvich, L. L.

    1979-01-01

    Available flow field data which can be used in validating theoretical procedures for computing flow fields around supersonic missiles are presented. Tabulated test data are given which define the flow field around a conical-nosed cylindrical body in a crossflow plane corresponding to a likely tail location. The data were obtained at a Mach number of 2.0 for an angle of attack of 0 to 23 degrees. The data define the flow field for cases both with and without a forward wing present.

  1. The Experimental Measurement of Aerodynamic Heating About Complex Shapes at Supersonic Mach Numbers

    NASA Technical Reports Server (NTRS)

    Neumann, Richard D.; Freeman, Delma C.

    2011-01-01

    In 2008 a wind tunnel test program was implemented to update the experimental data available for predicting protuberance heating at supersonic Mach numbers. For this test the Langley Unitary Wind Tunnel was also used. The significant differences for this current test were the advances in the state-of-the-art in model design, fabrication techniques, instrumentation and data acquisition capabilities. This current paper provides a focused discussion of the results of an in depth analysis of unique measurements of recovery temperature obtained during the test.

  2. Complex conservative difference schemes for computing supersonic flows past simple aerodynamic forms

    NASA Astrophysics Data System (ADS)

    Azarova, O. A.

    2015-12-01

    Complex conservative modifications of two-dimensional difference schemes on a minimum stencil are presented for the Euler equations. The schemes are conservative with respect to the basic divergent variables and the divergent variables for spatial derivatives. Approximations of boundary conditions for computing flows around variously shaped bodies (plates, cylinders, wedges, cones, bodies with cavities, and compound bodies) are constructed without violating the conservation properties in the computational domain. Test problems for computing flows with shock waves and contact discontinuities and supersonic flows with external energy sources are described.

  3. Method of characteristics for three-dimensional axially symmetrical supersonic flows.

    NASA Technical Reports Server (NTRS)

    Sauer, R

    1947-01-01

    An approximation method for three-dimensional axially symmetrical supersonic flows is developed; it is based on the characteristics theory (represented partly graphically, partly analytically). Thereafter this method is applied to the construction of rotationally symmetrical nozzles. (author)

  4. Aerodynamic Characteristics of Tracheostomy Speaking Valves.

    ERIC Educational Resources Information Center

    Fornataro-Clerici, Lisa; Zajac, David J.

    1993-01-01

    Pressure-flow characteristics were determined for four different one-way valves (Kisner, Montgomery, Olympic, and Passy-Muir) used for speech production in tracheotomy patients. Results indicated significant differences in resistance among the valves, with the resistance of one valve substantially greater than that of the normal upper airways.…

  5. Investigation of aerodynamic characteristics of subsonic wings

    NASA Technical Reports Server (NTRS)

    Dejarnette, F. R.; Frink, N. T.

    1979-01-01

    An analytical strake design procedure is investigated. A numerical solution to the governing strake design equation is used to generate a series of strakes which are tested in a water tunnel to study their vortex breakdown characteristics. The strakes are scaled for use on a half-scale model of the NASA-LaRC general research fuselage with a 44 degrees trapezoidal wing. An analytical solution to the governing design equation is obtained. The strake design procedure relates the potential-flow leading-edge suction and pressure distributions to vortex stability. Several suction distributions are studied and those which are more triangular and peak near the tip generate strakes that reach higher angles of attack before vortex breakdown occurs at the wing trailing edge. For the same suction distribution, a conical rather than three dimensional pressure specification results in a better strake shape as judged from its vortex breakdown characteristics.

  6. Hypersonic and Supersonic Static Aerodynamics of Mars Science Laboratory Entry Vehicle

    NASA Technical Reports Server (NTRS)

    Dyakonov, Artem A.; Schoenenberger, Mark; Vannorman, John W.

    2012-01-01

    This paper describes the analysis of continuum static aerodynamics of Mars Science Laboratory (MSL) entry vehicle (EV). The method is derived from earlier work for Mars Exploration Rover (MER) and Mars Path Finder (MPF) and the appropriate additions are made in the areas where physics are different from what the prior entry systems would encounter. These additions include the considerations for the high angle of attack of MSL EV, ablation of the heatshield during entry, turbulent boundary layer, and other aspects relevant to the flight performance of MSL. Details of the work, the supporting data and conclusions of the investigation are presented.

  7. Numerical methods and a computer program for subsonic and supersonic aerodynamic design and analysis of wings with attainable thrust considerations

    NASA Technical Reports Server (NTRS)

    Carlson, H. W.; Walkley, K. B.

    1984-01-01

    This paper describes methodology and an associated computer program for the design of wing lifting surfaces with attainable thrust taken into consideration. The approach is based on the determination of an optimum combination of a series of candidate surfaces rather than the more commonly used candidate loadings. Special leading-edge surfaces are selected to provide distributed leading-edge thrust forces which compensate for any failure to achieve the full theoretical leading-edge thrust, and a second series of general candidate surfaces is selected to minimize drag subject to constraints on the lift coefficient and, if desired, on the pitching moment coefficient. A primary purpose of the design approach is the introduction of attainable leading-edge thrust considerations so that relatively mild camber surfaces may be employed in the achievement of aerodynamic efficiencies comparable to those attainable if full theoretical leading-edge thrust could be achieved. The program provides an analysis as well as a design capability and is applicable to both subsonic and supersonic flow.

  8. Wake shape and its effects on aerodynamic characteristics

    NASA Technical Reports Server (NTRS)

    Emdad, H.; Lan, C. E.

    1986-01-01

    The wake shape under symmetrical flight conditions and its effects on aerodynamic characteristics are examined. In addition, the effect of wake shape in sideslip and discrete vortices such as strake or forebody vortex on lateral characteristics is presented. The present numerical method for airplane configurations, which is based on discretization of the vortex sheet into vortex segments, verified the symmetrical and asymmetrical roll-up process of the trailing vortices. Also, the effect of wing wake on tail planes is calculated. It is concluded that at high lift the assumption of flat wake for longitudinal and lateral-directional characteristics should be reexamined.

  9. Comparison of aerodynamic characteristics of pentagonal and hexagonal shaped bridge decks

    NASA Astrophysics Data System (ADS)

    Haque, Md. Naimul; Katsuchi, Hiroshi; Yamada, Hitoshi; Nishio, Mayuko

    2016-07-01

    Aerodynamics of the long-span bridge deck should be well understood for an efficient design of the bridge system. For practical bridges various deck shapes are being recommended and adopted, yet not all of their aerodynamic behaviors are well interpreted. In the present study, a numerical investigation was carried out to explore the aerodynamic characteristics of pentagonal and hexagonal shaped bridge decks. A relative comparison of steady state aerodynamic responses was made and the flow field was critically analyzed for better understanding the aerodynamic responses. It was found that the hexagonal shaped bridge deck has better aerodynamic characteristics as compared to the pentagonal shaped bridge deck.

  10. Aerodynamic Characteristics of Two Waverider-Derived Hypersonic Cruise Configurations

    NASA Technical Reports Server (NTRS)

    Cockrell, Charles E., Jr.; Huebner, Lawrence D.; Finley, Dennis B.

    1996-01-01

    An evaluation was made on the effects of integrating the required aircraft components with hypersonic high-lift configurations known as waveriders to create hypersonic cruise vehicles. Previous studies suggest that waveriders offer advantages in aerodynamic performance and propulsion/airframe integration (PAI) characteristics over conventional non-waverider hypersonic shapes. A wind-tunnel model was developed that integrates vehicle components, including canopies, engine components, and control surfaces, with two pure waverider shapes, both conical-flow-derived waveriders for a design Mach number of 4.0. Experimental data and limited computational fluid dynamics (CFD) solutions were obtained over a Mach number range of 1.6 to 4.63. The experimental data show the component build-up effects and the aerodynamic characteristics of the fully integrated configurations, including control surface effectiveness. The aerodynamic performance of the fully integrated configurations is not comparable to that of the pure waverider shapes, but is comparable to previously tested hypersonic models. Both configurations exhibit good lateral-directional stability characteristics.

  11. Wind Tunnel Tests on Aerodynamic Characteristics of Advanced Solid Rocket

    NASA Astrophysics Data System (ADS)

    Kitamura, Keiichi; Fujimoto, Keiichiro; Nonaka, Satoshi; Irikado, Tomoko; Fukuzoe, Moriyasu; Shima, Eiji

    The Advanced Solid Rocket is being developed by JAXA (Japan Aerospace Exploration Agency). Since its configuration has been changed very recently, its aerodynamic characteristics are of great interest of the JAXA Advanced Solid Rocket Team. In this study, we carried out wind tunnel tests on the aerodynamic characteristics of the present configuration for Mach 1.5. Six test cases were conducted with different body configurations, attack angles, and roll angles. A six component balance, oilflow visualization, Schlieren images were used throughout the experiments. It was found that, at zero angle-of-attack, the flow around the body were perturbed and its drag (axial force) characteristics were significantly influenced by protruding body components such as flanges, cable ducts, and attitude control units of SMSJ (Solid Motor Side Jet), while the nozzle had a minor role. With angle-of-attack of five degree, normal force of CNα = 3.50±0.03 was measured along with complex flow features observed in the full-component model; whereas no crossflow separations were induced around the no-protuberance model with CNα = 2.58±0.10. These values were almost constant with respect to the angle-of-attack in both of the cases. Furthermore, presence of roll angle made the flow more complicated, involving interactions of separation vortices. These data provide us with fundamental and important aerodynamic insights of the Advanced Solid Rocket, and they will be utilized as reference data for the corresponding numerical analysis.

  12. Supersonic Aerodynamic Design Improvements of an Arrow-Wing HSCT Configuration Using Nonlinear Point Design Methods

    NASA Technical Reports Server (NTRS)

    Unger, Eric R.; Hager, James O.; Agrawal, Shreekant

    1999-01-01

    This paper is a discussion of the supersonic nonlinear point design optimization efforts at McDonnell Douglas Aerospace under the High-Speed Research (HSR) program. The baseline for these optimization efforts has been the M2.4-7A configuration which represents an arrow-wing technology for the High-Speed Civil Transport (HSCT). Optimization work on this configuration began in early 1994 and continued into 1996. Initial work focused on optimization of the wing camber and twist on a wing/body configuration and reductions of 3.5 drag counts (Euler) were realized. The next phase of the optimization effort included fuselage camber along with the wing and a drag reduction of 5.0 counts was achieved. Including the effects of the nacelles and diverters into the optimization problem became the next focus where a reduction of 6.6 counts (Euler W/B/N/D) was eventually realized. The final two phases of the effort included a large set of constraints designed to make the final optimized configuration more realistic and they were successful albeit with a loss of performance.

  13. Aerodynamic Characteristics of High Speed Trains under Cross Wind Conditions

    NASA Astrophysics Data System (ADS)

    Chen, W.; Wu, S. P.; Zhang, Y.

    2011-09-01

    Numerical simulation for the two models in cross-wind was carried out in this paper. The three-dimensional compressible Reynolds-averaged Navier-Stokes equations(RANS), combined with the standard k-ɛ turbulence model, were solved on multi-block hybrid grids by second order upwind finite volume technique. The impact of fairing on aerodynamic characteristics of the train models was analyzed. It is shown that, the flow separates on the fairing and a strong vortex is generated, the pressure on the upper middle car decreases dramatically, which leads to a large lift force. The fairing changes the basic patterns around the trains. In addition, formulas of the coefficient of aerodynamic force at small yaw angles up to 24° were expressed.

  14. Predicting aerodynamic characteristic of typical wind turbine airfoils using CFD

    SciTech Connect

    Wolfe, W.P.; Ochs, S.S.

    1997-09-01

    An investigation was conducted into the capabilities and accuracy of a representative computational fluid dynamics code to predict the flow field and aerodynamic characteristics of typical wind-turbine airfoils. Comparisons of the computed pressure and aerodynamic coefficients were made with wind tunnel data. This work highlights two areas in CFD that require further investigation and development in order to enable accurate numerical simulations of flow about current generation wind-turbine airfoils: transition prediction and turbulence modeling. The results show that the laminar-to turbulent transition point must be modeled correctly to get accurate simulations for attached flow. Calculations also show that the standard turbulence model used in most commercial CFD codes, the k-e model, is not appropriate at angles of attack with flow separation. 14 refs., 28 figs., 4 tabs.

  15. Integrating aerodynamics and structures in the minimum weight design of a supersonic transport wing

    NASA Technical Reports Server (NTRS)

    Barthelemy, Jean-Francois M.; Wrenn, Gregory A.; Dovi, Augustine R.; Coen, Peter G.; Hall, Laura E.

    1992-01-01

    An approach is presented for determining the minimum weight design of aircraft wing models which takes into consideration aerodynamics-structure coupling when calculating both zeroth order information needed for analysis and first order information needed for optimization. When performing sensitivity analysis, coupling is accounted for by using a generalized sensitivity formulation. The results presented show that the aeroelastic effects are calculated properly and noticeably reduce constraint approximation errors. However, for the particular example selected, the error introduced by ignoring aeroelastic effects are not sufficient to significantly affect the convergence of the optimization process. Trade studies are reported that consider different structural materials, internal spar layouts, and panel buckling lengths. For the formulation, model and materials used in this study, an advanced aluminum material produced the lightest design while satisfying the problem constraints. Also, shorter panel buckling lengths resulted in lower weights by permitting smaller panel thicknesses and generally, by unloading the wing skins and loading the spar caps. Finally, straight spars required slightly lower wing weights than angled spars.

  16. Analysis of preflutter and postflutter characteristics with motion-matched aerodynamic forces

    NASA Technical Reports Server (NTRS)

    Cunningham, H. J.

    1978-01-01

    The development of the equations of dynamic equilibrium for a lifting surface from Lagrange's equation is reviewed and restated for general exponential growing and decaying oscillatory motion. Aerodynamic forces for this motion are obtained from the three-dimensional supersonic kernel function that is newly generalized to complex reduced frequencies. Illustrative calculations were made for two flutter models at supersonic Mach numbers. Preflutter and postflutter motion isodecrement curves were obtained. This type of analysis can be used to predict preflutter behavior during flutter testing and to predict postflutter behavior for use in the design of flutter suppression systems.

  17. Wind tunnel investigation of the interaction and breakdown characteristics of slender wing vortices at subsonic, transonic, and supersonic speeds

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.

    1991-01-01

    The vortex dominated aerodynamic characteristics of a generic 65 degree cropped delta wing model were studied in a wind tunnel at subsonic through supersonic speeds. The lee-side flow fields over the wing-alone configuration and the wing with leading edge extension (LEX) added were observed at M (infinity) equals 0.40 to 1.60 using a laser vapor screen technique. These results were correlated with surface streamline patterns, upper surface static pressure distributions, and six-component forces and moments. The wing-alone exhibited vortex breakdown and asymmetry of the breakdown location at the subsonic and transonic speeds. An earlier onset of vortex breakdown over the wing occurred at transonic speeds due to the interaction of the leading edge vortex with the normal shock wave. The development of a shock wave between the vortex and wing surface caused an early separation of the secondary boundary layer. With the LEX installed, wing vortex breakdown asymmetry did not occur up to the maximum angle of attack in the present test of 24 degrees. The favorable interaction of the LEX vortex with the wing flow field reduced the effects of shock waves on the wing primary and secondary vortical flows. The direct interaction of the wing and LEX vortex cores diminished with increasing Mach number. The maximum attainable vortex-induced pressure signatures were constrained by the vacuum pressure limit at the transonic and supersonic speeds.

  18. Study of spectral characteristics of radiation from a thermal wake of a pulsating optical discharge in a supersonic air flow

    SciTech Connect

    Malov, A N; Orishich, A M; Terent'eva, Ya S

    2015-10-31

    The spectral characteristics of the thermal wake of a pulsating optical discharge (POD) in a supersonic air flow are studied. The POD is stimulated by radiation of a mechanically Q-switched, repetitively pulsed CO{sub 2} laser with a pulse repetition rate of 7 – 150 kHz and a power up to 4.5 kW. The flow is produced by means of the supersonic aerodynamic MAU-M setup having a conic nozzle with a critical cross-section size of 50 mm, the Mach number being 1.3 – 1.6. We describe in detail the system of optical diagnostics that allows the detection of the spectrum of the weak thermal wake glow against the background of high-power POD radiation. The glow of the thermal wake is due to the emission of light by atoms and ions of nitrogen and oxygen, carried by the flow in the form of hot low-density gas clouds (caverns). The wavelengths of the thermal wake emission and the data on the transitions, corresponding to the spectral lines are presented. (laser applications and other topics in quantum electronics)

  19. Aerodynamic design and analysis of the AST-204, AST-205, and AST-206 blended wing-fuse large supersonic transport configuration concepts

    NASA Technical Reports Server (NTRS)

    Martin, G. L.; Walkley, K. B.

    1980-01-01

    The aerodynamic design and analysis of three blended wing-fuselage supersonic cruise configurations providing four, five, and six abreast seating was conducted using a previously designed supersonic cruise configuration as the baseline. The five abreast configuration was optimized for wave drag at a Mach number of 2.7. The four and six abreast configurations were also optimized at Mach 2.7, but with the added constraint that the majority of their structure be common with the five abreast configuration. Analysis of the three configurations indicated an improvement of 6.0, 7.5, and 7.7 percent in cruise lift-to-drag ratio over the baseline configuration for the four, five, and six abreast configurations, respectively.

  20. Development of a pulsed uniform supersonic gas expansion system based on an aerodynamic chopper for gas phase reaction kinetic studies at ultra-low temperatures

    NASA Astrophysics Data System (ADS)

    Jiménez, E.; Ballesteros, B.; Canosa, A.; Townsend, T. M.; Maigler, F. J.; Napal, V.; Rowe, B. R.; Albaladejo, J.

    2015-04-01

    A detailed description of a new pulsed supersonic uniform gas expansion system is presented together with the experimental validation of the setup by applying the CRESU (French acronym for Cinétique de Réaction en Ecoulement Supersonique Uniforme or Reaction Kinetics in a Uniform Supersonic Flow) technique to the gas-phase reaction of OH radicals with 1-butene at ca. 23 K and 0.63 millibars of helium (carrier gas). The carrier gas flow, containing negligible mixing ratios of OH-precursor and 1-butene, is expanded from a high pressure reservoir (337 millibars) to a low pressure region (0.63 millibars) through a convergent-divergent nozzle (Laval type). The novelty of this experimental setup is that the uniform supersonic flow is pulsed by means of a Teflon-coated aerodynamic chopper provided with two symmetrical apertures. Under these operational conditions, the designed Laval nozzle achieves a temperature of (22.4 ± 1.4) K in the gas jet. The spatial characterization of the temperature and the total gas density within the pulsed uniform supersonic flow has also been performed by both aerodynamical and spectroscopic methods. The gas consumption with this technique is considerably reduced with respect to a continuous CRESU system. The kinetics of the OH+1-butene reaction was investigated by the pulsed laser photolysis/laser induced fluorescence technique. The rotation speed of the disk is temporally synchronized with the exit of the photolysis and the probe lasers. The rate coefficient (kOH) for the reaction under investigation was then obtained and compared with the only available data at this temperature.

  1. Hydrodynamic Characteristics of an Aerodynamically Refined Planing-Tail Hull

    NASA Technical Reports Server (NTRS)

    McKann, Robert; Suydam, Henry B.

    1948-01-01

    The hydrodynamic characteristics of an aerodynamically refined planing-tail hull were determined from dynamic model tests in Langley tank no. 2. Stable take-off could be made for a wide range of locations of the center of gravity. The lower porpoising limit peak was high, but no upper limit was encountered. Resistance was high, being about the same as that of float seaplanes. A reasonable range of trims for stable landings was available only in the aft range of center-of-gravity locations.

  2. Steady, Oscillatory, and Unsteady Subsonic and Supersonic Aerodynamics, production version (SOUSSA-P 1.1). Volume 1: Theoretical manual. [Green function

    NASA Technical Reports Server (NTRS)

    Morino, L.

    1980-01-01

    Recent developments of the Green's function method and the computer program SOUSSA (Steady, Oscillatory, and Unsteady Subsonic and Supersonic Aerodynamics) are reviewed and summarized. Applying the Green's function method to the fully unsteady (transient) potential equation yields an integro-differential-delay equation. With spatial discretization by the finite-element method, this equation is approximated by a set of differential-delay equations in time. Time solution by Laplace transform yields a matrix relating the velocity potential to the normal wash. Premultiplying and postmultiplying by the matrices relating generalized forces to the potential and the normal wash to the generalized coordinates one obtains the matrix of the generalized aerodynamic forces. The frequency and mode-shape dependence of this matrix makes the program SOUSSA useful for multiple frequency and repeated mode-shape evaluations.

  3. Propulsion and airframe aerodynamic interactions of supersonic V/STOL configurations. Volume 2: Wind tunnel test force and moment data report

    NASA Technical Reports Server (NTRS)

    Zilz, D. E.

    1985-01-01

    A wind tunnel model of a supersonic V/STOL fighter configuration has been tested to measure the aerodynamic interaction effects which can result from geometrically close-coupled propulsion system/airframe components. The approach was to configure the model to represent two different test techniques. One was a conventional test technique composed of two test modes. In the Flow-Through mode, absolute configuration aerodynamics are measured, including inlet/airframe interactions. In the Jet-Effects mode, incremental nozzle/airframe interactions are measured. The other test technique is a propulsion simulator approach, where a sub-scale, externally powered engine is mounted in the model. This allows proper measurement of inlet/airframe and nozzle/airframe interactions simultaneously. This is Volume 2 of 2: Wind Tunnel Test Force and Moment Data Report.

  4. Aerodynamic characteristics of flying fish in gliding flight.

    PubMed

    Park, Hyungmin; Choi, Haecheon

    2010-10-01

    The flying fish (family Exocoetidae) is an exceptional marine flying vertebrate, utilizing the advantages of moving in two different media, i.e. swimming in water and flying in air. Despite some physical limitations by moving in both water and air, the flying fish has evolved to have good aerodynamic designs (such as the hypertrophied fins and cylindrical body with a ventrally flattened surface) for proficient gliding flight. Hence, the morphological and behavioral adaptations of flying fish to aerial locomotion have attracted great interest from various fields including biology and aerodynamics. Several aspects of the flight of flying fish have been determined or conjectured from previous field observations and measurements of morphometric parameters. However, the detailed measurement of wing performance associated with its morphometry for identifying the characteristics of flight in flying fish has not been performed yet. Therefore, in the present study, we directly measure the aerodynamic forces and moment on darkedged-wing flying fish (Cypselurus hiraii) models and correlated them with morphological characteristics of wing (fin). The model configurations considered are: (1) both the pectoral and pelvic fins spread out, (2) only the pectoral fins spread with the pelvic fins folded, and (3) both fins folded. The role of the pelvic fins was found to increase the lift force and lift-to-drag ratio, which is confirmed by the jet-like flow structure existing between the pectoral and pelvic fins. With both the pectoral and pelvic fins spread, the longitudinal static stability is also more enhanced than that with the pelvic fins folded. For cases 1 and 2, the lift-to-drag ratio was maximum at attack angles of around 0 deg, where the attack angle is the angle between the longitudinal body axis and the flying direction. The lift coefficient is largest at attack angles around 30∼35 deg, at which the flying fish is observed to emerge from the sea surface. From glide polar

  5. Theoretical characteristics in supersonic flow of two types of control surfaces on triangular wings

    NASA Technical Reports Server (NTRS)

    Tucker, Warren A; Nelson, Robert L

    1949-01-01

    Methods based on the linearized theory for supersonic flow were used to find the characteristics of two types of control surfaces on thin triangular wings. The first type, the constant-chord partial-span flap, was considered to extend either outboard from the center of the wing or inboard from the wing tip. The second type, the full-triangular-tip flap, was treated only for the case in which the Mach number component normal to the leading edge is supersonic. For each type, expressions were found for the lift, rolling-moment, pitching-moment, and hinge-moment characteristics.

  6. Deep-Stall Aerodynamic Characteristics of T-Tail Aircraft

    NASA Technical Reports Server (NTRS)

    Taylor, Robert T.; Ray, Edward J.

    1965-01-01

    A wind-tunnel research program has been under-taken by the NASA to study the aerodynamic characteristics of T-tail aircraft at high angles of attack. The program was designed to show the effects on longitudinal stability and control of several configuration variables. The results to date do not allow the formulation of general design rules, but the effects of several configuration variables have been noted to have a prime influence on the post-stall characteristics. An increase in tail size, changes in the location of fuselage-mounted engine nacelles, and reduced fuselage-forebody lift were all found to have a beneficial effect on static longitudinal stability at high angles of attack.

  7. Study of aerodynamic technology for single-cruise-engine V/STOL fighter/attack aircraft

    NASA Technical Reports Server (NTRS)

    Hess, J. R.; Bear, R. L.

    1982-01-01

    A viable, single engine, supersonic V/STOL fighter/attack aircraft concept was defined. This vectored thrust, canard wing configuration utilizes an advanced technology separated flow engine with fan stream burning. The aerodynamic characteristics of this configuration were estimated and performance evaluated. Significant aerodynamic and aerodynamic propulsion interaction uncertainties requiring additional investigation were identified. A wind tunnel model concept and test program to resolve these uncertainties and validate the aerodynamic prediction methods were defined.

  8. Aerodynamic characteristics of Lockheed delta-body orbiter and stage-and-one-half launch vehicle

    NASA Technical Reports Server (NTRS)

    Velligan, F. A.; Svendsen, H. O.

    1971-01-01

    An experimental wind tunnel test program was conducted to investigate the subsonic through high supersonic aerodynamic characteristics of the Lockheed delta lifting body orbiter and stage-and-one-half launch vehicle. Analyses and results of these data are presented. A 0.01-scale model of the LS 200-5 system was designed and fabricated for testing in wind tunnels. Orbiter and launch configurations were tested over a speed range of Mach 0.6 to 2.0, whereas only the orbiter was tested over a speed range of Mach 2.3 to 4.6. Six-component force and moment data, base pressures, and schlieren photos were obtained at various angles-of-attack and sideslip. A 0.03-scale model of the orbiter was also designed, fabricated, and tested in a wind tunnel. Six-component force and moment data, base pressure, and a limited amount of tuft flow visualization data were obtained on a variety of configuration combinations.

  9. Model aerodynamic test results for two variable cycle engine coannular exhaust systems at simulated takeoff and cruise conditions. [Lewis 8 by 6-foot supersonic wind tunnel tests

    NASA Technical Reports Server (NTRS)

    Nelson, D. P.

    1980-01-01

    Wind tunnel tests were conducted to evaluate the aerodynamic performance of a coannular exhaust nozzle for a proposed variable stream control supersonic propulsion system. Tests were conducted with two simulated configurations differing primarily in the fan duct flowpaths: a short flap mechanism for fan stream control with an isentropic contoured flow splitter, and an iris fan nozzle with a conical flow splitter. Both designs feature a translating primary plug and an auxiliary inlet ejector. Tests were conducted at takeoff and simulated cruise conditions. Data were acquired at Mach numbers of 0, 0.36, 0.9, and 2.0 for a wide range of nozzle operating conditions. At simulated supersonic cruise, both configurations demonstrated good performance, comparable to levels assumed in earlier advanced supersonic propulsion studies. However, at subsonic cruise, both configurations exhibited performance that was 6 to 7.5 percent less than the study assumptions. At take off conditions, the iris configuration performance approached the assumed levels, while the short flap design was 4 to 6 percent less.

  10. Experimental and theoretical study of aerodynamic characteristics of some lifting bodies at angles of attack from -10 degrees to 53 degrees at Mach numbers from 2.30 to 4.62

    NASA Technical Reports Server (NTRS)

    Spearman, M. Leroy; Torres, Abel O.

    1994-01-01

    Lifting bodies are of interest for possible use as space transportation vehicles because they have the volume required for significant payloads and the aerodynamic capability to negotiate the transition from high angles of attack to lower angles of attack (for cruise flight) and thus safely reenter the atmosphere and perform conventional horizontal landings. Results are presented for an experimental and theoretical study of the aerodynamic characteristics at supersonic speeds for a series of lifting bodies with 75 deg delta planforms, rounded noses, and various upper and lower surface cambers. The camber shapes varied in thickness and in maximum thickness location, and hence in body volume. The experimental results were obtained in the Langley Unitary Plan Wind Tunnel for both the longitudinal and the lateral aerodynamic characteristics. Selected experimental results are compared with calculated results obtained through the use of the Hypersonic Arbitrary-Body Aerodynamic Computer Program.

  11. Effects of Reynolds number and model support on the supersonic aerodynamic chacteristics of a 140 deg-included-angle cone

    NASA Technical Reports Server (NTRS)

    Trescot, C. D., Jr.; Brown, C. A., Jr.; Howell, D. T.

    1974-01-01

    An investigation has been made in the Langley Unitary Plan wind tunnel to determine the effects of Reynolds number and sting-support interference on the static aerodynamic characteristics of a 140 deg-included-angle cone. Base pressures and forces and moments of the model were measured at Mach numbers of 1.50, 2.00, 2.94, and 4.00 for ratios of sting diameter to model diameter that varied from 0.125 to 0.500 through an angle-of-attack range from about minus 4 deg to 13 deg. The Reynolds number, based on model diameter 4.80 in. was varied from 161,000 to 415,000.

  12. X-31 aerodynamic characteristics determined from flight data

    NASA Technical Reports Server (NTRS)

    Kokolios, Alex

    1993-01-01

    The lateral aerodynamic characteristics of the X-31 were determined at angles of attack ranging from 20 to 45 deg. Estimates of the lateral stability and control parameters were obtained by applying two parameter estimation techniques, linear regression, and the extended Kalman filter to flight test data. An attempt to apply maximum likelihood to extract parameters from the flight data was also made but failed for the reasons presented. An overview of the System Identification process is given. The overview includes a listing of the more important properties of all three estimation techniques that were applied to the data. A comparison is given of results obtained from flight test data and wind tunnel data for four important lateral parameters. Finally, future research to be conducted in this area is discussed.

  13. An analytical procedure for evaluating shuttle abort staging aerodynamic characteristics

    NASA Technical Reports Server (NTRS)

    Meyer, R.

    1973-01-01

    An engineering analysis and computer code (AERSEP) for predicting Space Shuttle Orbiter - HO Tank longitudinal aerodynamic characteristics during abort separation has been developed. Computed results are applicable at Mach numbers above 2 for angle-of-attack between plus or minus 10 degrees. No practical restrictions on orbiter-tank relative positioning are indicated for tank-under-orbiter configurations. Input data requirements and computer running times are minimal facilitating program use for parametric studies, test planning, and trajectory analysis. In a majority of cases AERSEP Orbiter-Tank interference predictions are as accurate as state-of-the-art estimates for interference-free or isolated-vehicle configurations. AERSEP isolated-orbiter predictions also show excellent correlation with data.

  14. Aerodynamic characteristics of proposed assured crew return capability (ACRC) configurations

    NASA Astrophysics Data System (ADS)

    Ware, George M.; Spencer, Bernard, Jr.; Micol, John R.

    1989-07-01

    The aerodynamic characteristics of seven reentry configurations suggested as possible candidate vehicles to return crew members from the U.S. Space Station Freedom to earth has been reviewed. The shapes varied from those capable of purely ballistic entry to those capable of gliding entry and fromk parachute landing to conventional landing. Data were obtained from existing (published and unpublished) sources and from recent wind tunnel tests. The lifting concepts are more versatile and satisfy all the mission requirements. Two of the lifting shapes studied appear promising - a lifting body and a deployable wing concept. The choice of an ACRC concept, however, will be made after all factors involving transportation from earth to orbit and back to earth again have been weighed.

  15. Experimental investigation on the characteristics of supersonic fuel spray and configurations of induced shock waves

    PubMed Central

    Wang, Yong; Yu, Yu-song; Li, Guo-xiu; Jia, Tao-ming

    2017-01-01

    The macro characteristics and configurations of induced shock waves of the supersonic sprays are investigated by experimental methods. Visualization study of spray shape is carried out with the high-speed camera. The macro characteristics including spray tip penetration, velocity of spray tip and spray angle are analyzed. The configurations of shock waves are investigated by Schlieren technique. For supersonic sprays, the concept of spray front angle is presented. Effects of Mach number of spray on the spray front angle are investigated. The results show that the shape of spray tip is similar to blunt body when fuel spray is at transonic region. If spray entered the supersonic region, the oblique shock waves are induced instead of normal shock wave. With the velocity of spray increasing, the spray front angle and shock wave angle are increased. The tip region of the supersonic fuel spray is commonly formed a cone. Mean droplet diameter of fuel spray is measured using Malvern’s Spraytec. Then the mean droplet diameter results are compared with three popular empirical models (Hiroyasu’s, Varde’s and Merrigton’s model). It is found that the Merrigton’s model shows a relative good correlation between models and experimental results. Finally, exponent of injection velocity in the Merrigton’s model is fitted with experimental results. PMID:28054555

  16. Experimental investigation on the characteristics of supersonic fuel spray and configurations of induced shock waves

    NASA Astrophysics Data System (ADS)

    Wang, Yong; Yu, Yu-Song; Li, Guo-Xiu; Jia, Tao-Ming

    2017-01-01

    The macro characteristics and configurations of induced shock waves of the supersonic sprays are investigated by experimental methods. Visualization study of spray shape is carried out with the high-speed camera. The macro characteristics including spray tip penetration, velocity of spray tip and spray angle are analyzed. The configurations of shock waves are investigated by Schlieren technique. For supersonic sprays, the concept of spray front angle is presented. Effects of Mach number of spray on the spray front angle are investigated. The results show that the shape of spray tip is similar to blunt body when fuel spray is at transonic region. If spray entered the supersonic region, the oblique shock waves are induced instead of normal shock wave. With the velocity of spray increasing, the spray front angle and shock wave angle are increased. The tip region of the supersonic fuel spray is commonly formed a cone. Mean droplet diameter of fuel spray is measured using Malvern’s Spraytec. Then the mean droplet diameter results are compared with three popular empirical models (Hiroyasu’s, Varde’s and Merrigton’s model). It is found that the Merrigton’s model shows a relative good correlation between models and experimental results. Finally, exponent of injection velocity in the Merrigton’s model is fitted with experimental results.

  17. Investigation on flow and mixing characteristics of supersonic mixing layer induced by forced vibration of cantilever

    NASA Astrophysics Data System (ADS)

    Zhang, Dongdong; Tan, Jianguo; Lv, Liang

    2015-12-01

    The mixing process has been an important issue for the design of supersonic combustion ramjet engine, and the mixing efficiency plays a crucial role in the improvement of the combustion efficiency. In the present study, nanoparticle-based planar laser scattering (NPLS), particle image velocimetry (PIV) and large eddy simulation (LES) are employed to investigate the flow and mixing characteristics of supersonic mixing layer under different forced vibration conditions. The indexes of fractal dimension, mixing layer thickness, momentum thickness and scalar mixing level are applied to describe the mixing process. Results show that different from the development and evolution of supersonic mixing layer without vibration, the flow under forced vibration is more likely to present the characteristics of three-dimensionality. The laminar flow region of mixing layer under forced vibration is greatly shortened and the scales of rolled up Kelvin-Helmholtz vortices become larger, which promote the mixing process remarkably. The fractal dimension distribution reveals that comparing with the flow without vibration, the turbulent fluctuation of supersonic mixing layer under forced vibration is more intense. Besides, the distribution of mixing layer thickness, momentum thickness and scalar mixing level are strongly influenced by forced vibration. Especially, when the forcing frequency is 4000 Hz, the mixing layer thickness and momentum thickness are 0.0391 m and 0.0222 m at the far field of 0.16 m, 83% and 131% higher than that without vibration at the same position, respectively.

  18. Computational modeling of aerodynamic characteristics in sprayed and spiraled precalciner

    NASA Astrophysics Data System (ADS)

    Li, Xiangguo; Ma, Baoguo; Hu, Zhenwu

    2008-08-01

    Based on the structural and work characteristics of a spiraled and sprayed precalciner, the RNG k- ɛ model and the SIMPLE method were used to simulate the aerodynamic characteristics in a sprayed and spiraled precalciner. The simulation results demonstrate that the flow area of airflow was increased abruptly due to the reduced part of the bottom of precalciners, which attributed to a sprayed effect. With the mix of the tertiary air with the swirl flow and secondary air, a high-speed zone was formed in the opposite side of the inlet of tertiary air, in which the highest speed was 32.97 m/s. Moreover, the inlet of raw meal designed in the high-speed zone can be propitious to the decentralization of the raw meal. A back-flow zone was formed near the side of the inlet of tertiary air, in which the velocity was negative. From the analysis of the results, the flow field of the precalciner is composed of a sprayed zone, a high-speed zone, a back-flow zone and cylinder zone; moreover, the simulation results agree with those of the engineering compared to the in situ results. The results also showed that the CFD method can be used to give the basis for optimizing the geometrical design and flow parameters of a precalciner.

  19. Mathematical modeling of the aerodynamic characteristics in flight dynamics

    NASA Technical Reports Server (NTRS)

    Tobak, M.; Chapman, G. T.; Schiff, L. B.

    1984-01-01

    Basic concepts involved in the mathematical modeling of the aerodynamic response of an aircraft to arbitrary maneuvers are reviewed. The original formulation of an aerodynamic response in terms of nonlinear functionals is shown to be compatible with a derivation based on the use of nonlinear functional expansions. Extensions of the analysis through its natural connection with ideas from bifurcation theory are indicated.

  20. Modeling of aircraft unsteady aerodynamic characteristics. Part 1: Postulated models

    NASA Technical Reports Server (NTRS)

    Klein, Vladislav; Noderer, Keith D.

    1994-01-01

    A short theoretical study of aircraft aerodynamic model equations with unsteady effects is presented. The aerodynamic forces and moments are expressed in terms of indicial functions or internal state variables. The first representation leads to aircraft integro-differential equations of motion; the second preserves the state-space form of the model equations. The formulations of unsteady aerodynamics is applied in two examples. The first example deals with a one-degree-of-freedom harmonic motion about one of the aircraft body axes. In the second example, the equations for longitudinal short-period motion are developed. In these examples, only linear aerodynamic terms are considered. The indicial functions are postulated as simple exponentials and the internal state variables are governed by linear, time-invariant, first-order differential equations. It is shown that both approaches to the modeling of unsteady aerodynamics lead to identical models.

  1. Aerodynamic Characteristics, Database Development and Flight Simulation of the X-34 Vehicle

    NASA Technical Reports Server (NTRS)

    Pamadi, Bandu N.; Brauckmann, Gregory J.; Ruth, Michael J.; Fuhrmann, Henri D.

    2000-01-01

    An overview of the aerodynamic characteristics, development of the preflight aerodynamic database and flight simulation of the NASA/Orbital X-34 vehicle is presented in this paper. To develop the aerodynamic database, wind tunnel tests from subsonic to hypersonic Mach numbers including ground effect tests at low subsonic speeds were conducted in various facilities at the NASA Langley Research Center. Where wind tunnel test data was not available, engineering level analysis is used to fill the gaps in the database. Using this aerodynamic data, simulations have been performed for typical design reference missions of the X-34 vehicle.

  2. Aerodynamic characteristics of wheelchairs. [Langley V/STOL wind tunnel tests for human factors engineering

    NASA Technical Reports Server (NTRS)

    Coe, P. L., Jr.

    1979-01-01

    The overall aerodynamic drag characteristics of a conventional wheelchair were defined and the individual drag contributions of its components were determined. The results show that a fiftieth percentile man sitting in the complete wheelchair would experience an aerodynamic drag coefficient on the order of 1.4.

  3. The impact of emerging technologies on an advanced supersonic transport

    NASA Technical Reports Server (NTRS)

    Driver, C.; Maglieri, D. J.

    1986-01-01

    The effects of advances in propulsion systems, structure and materials, aerodynamics, and systems on the design and development of supersonic transport aircraft are analyzed. Efficient propulsion systems with variable-cycle engines provide the basis for improved propulsion systems; the propulsion efficienies of supersonic and subsonic engines are compared. Material advances consist of long-life damage-tolerant structures, advanced material development, aeroelastic tailoring, and low-cost fabrication. Improvements in the areas of aerodynamics and systems are examined. The environmental problems caused by engine emissions, airport noise, and sonic boom are studied. The characteristics of the aircraft designed to include these technical advances are described.

  4. An Interactive Method of Characteristics Java Applet to Design and Analyze Supersonic Aircraft Nozzles

    NASA Technical Reports Server (NTRS)

    Benson, Thomas J.

    2014-01-01

    The Method of Characteristics (MOC) is a classic technique for designing supersonic nozzles. An interactive computer program using MOC has been developed to allow engineers to design and analyze supersonic nozzle flow fields. The program calculates the internal flow for many classic designs, such as a supersonic wind tunnel nozzle, an ideal 2D or axisymmetric nozzle, or a variety of plug nozzles. The program also calculates the plume flow produced by the nozzle and the external flow leading to the nozzle exit. The program can be used to assess the interactions between the internal, external and plume flows. By proper design and operation of the nozzle, it may be possible to lessen the strength of the sonic boom produced at the rear of supersonic aircraft. The program can also calculate non-ideal nozzles, such as simple cone flows, to determine flow divergence and nonuniformities at the exit, and its effect on the plume shape. The computer program is written in Java and is provided as free-ware from the NASA Glenn central software server.

  5. Wind-tunnel studies of the effects of simulated damage on the aerodynamic characteristics of airplanes and missiles

    NASA Technical Reports Server (NTRS)

    Spearman, M. L.

    1979-01-01

    In order to assess the effects on static aerodynamic characteristics of battle damage to an aircraft or missile, wind tunnel studies were performed on models from which all or parts of the wing or horizontal or vertical tail had been removed. The effects of damage on the lift, longitudinal stability, lateral stability and directional stability of a swept-wing fighter are presented, along with the effects of wing removal on the control requirements of a delta-wing fighter. Results indicate that the loss of a major part of the vertical tail will probably result in the loss of the aircraft at any speed, while the loss of major parts of the horizontal tail generally results in catastrophic instability at subsonic speeds but, at low supersonic speeds, may allow the aircraft to return to friendly territory before pilot ejection. Major damage to the wing may be sustained without the loss of aircraft or pilot. The loss of some of the aerodynamic surfaces of cruise or surface-to-air missiles may result in catastrophic instability or may permit a ballistic trajectory to be maintained, depending upon the location of the lost surface with respect to the center of gravity of the missile.

  6. Aerodynamic Characteristics of Caliber .22 Long Rifle Match Ammunition

    DTIC Science & Technology

    1990-11-01

    range. The Aerodynamic-s Range is an enclosed, cimate -controlled range, instrumented with spark-photography stations to record the motion of the...slant of groups in the wind from a right-hand twist of rifling is due to aerodynamic jump, which is an effective change in the vertical angle of...Director, USAHEL ATTN: SLCHE-IS, Mr. B. Corona Mr. P. Ellis Mr. J. Torre 64 USER EVALUArION SHEET/ CHANGE OF ADDRIESS9 This Uboray underk a c-effor- to imo t

  7. Characteristics of Pressure Sensitive Paint Intrusiveness Effects on Aerodynamic Data

    NASA Technical Reports Server (NTRS)

    Amer, Tahani R.; Liu, Tianshu; Oglesby, Donald M.

    2001-01-01

    One effect of using pressure sensitive paint (PSP) is the potential intrusiveness to the aerodynamic characteristics of the model. The paint thickness and roughness may affect the pressure distribution, and therefore, the forces and moments on the wind tunnel model. A study of these potential intrusive effects was carried out at NASA Langley Research Center where a series of wind tunnel tests were conducted using the Modem Design of Experiments (MDOE) test approach. The PSP effects on the integrated forces were measured on two different models at different test conditions in both the Low Turbulence Pressure Tunnel (LTPT) and the Unitary Plan Wind Tunnel (UPWT) at Langley. The paint effect was found to be very small over a range of Reynolds numbers, Mach numbers and angles of attack. This is due to the very low surface roughness of the painted surface. The surface roughness, after applying the NASA Langley developed PSP, was lower than that of the clean wing. However, the PSP coating had a localized effects on the pressure taps, which leads to an appreciable decrease in the pressure tap reading.

  8. Acoustic and aerodynamic characteristics of ejectives in Amharic

    NASA Astrophysics Data System (ADS)

    Demolin, Didier

    2004-05-01

    This paper invetsigates the main phonetic characteristics that distinguishes ejectives from pulmonic sounds in Amharic. In this language, there are five ejectives that can be phonemically singleton or geminate. Duration measurements have been made in intervocalic position for pulmonic stops and for each type of ejective, taking into account the overall duration and VOT. Results show that ejective stops have a higher amplitude burst than pulmonic stops. The duration of the noise is shorter for ejective fricatives compared to pulmonic fricatives. At the end of ejective fricatives, there is a 30-ms glottal lag that is not present in pulmonic fricatives. Geminate ejectives are realized by delaying the elevation of the larynx. This can be observed on the spectrographic data by an increase of the noise at the end of the geminate ejectives. Aerodynamic data have been collected in synchronization with the acoustic recordings. The main observations are that pharyngeal pressures values are much higher than what is usually assumed (up to 40 CmH2O for velars) and that the delayed command in the elevation of the larynx of geminate ejectives is shown by two phases in the rise of pharyngeal pressure.

  9. Comparison of Theoretical and Experimental Unsteady Aerodynamics of Linear Oscillating Cascade With Supersonic Leading-Edge Locus. Video

    NASA Technical Reports Server (NTRS)

    Ramsey, John K.; Erwin, Dan

    2004-01-01

    The movie file contained in the DVD contains footage of the NASA/Ohio State Supersonic Oscillating Cascade Facility wind tunnel run with an oscillating mechanism. Links to view the movie can also be found in figures 4 and 8 in the online PDF version that this record references, CASI ID 20040082334.

  10. Characteristics and measurement of supersonic projectile shock waves by a 32-microphone ring array

    NASA Astrophysics Data System (ADS)

    Chang, Ho; Wu, Yan-Chyuan; Tsung, Tsing-Tshih

    2011-08-01

    This paper discusses about the characteristics of supersonic projectile shock wave in muzzle region during firing of high explosive anti-tank (HEAT) and high explosive (HE) projectiles. HEAT projectiles are fired horizontally at a muzzle velocity of Mach 3.5 from a medium caliber tank gun equipped with a newly designed multi-perforated muzzle brake, whereas HE projectiles are fired at elevation angles at a muzzle velocity of Mach 2 from a large caliber howitzer equipped with a newly designed double-baffle muzzle brake. In the near field, pressure signatures of the N-wave generated from projectiles are measured by 32-microphone ring array wrapped by cotton sheath. Records measured by the microphone array are used to demonstrate several key characteristics of the shock wave of supersonic projectile. All measurements made in this study can be a significant reference for developing guns, tanks, or the chassis of fighting vehicles.

  11. Theoretical face pressure and drag characteristics of forward-facing steps in supersonic turbulent boundary layers

    NASA Technical Reports Server (NTRS)

    Patel, D. K.; Czarnecki, K. R.

    1975-01-01

    A theoretical investigation of the pressure distributions and drag characteristics was made for forward facing steps in turbulent flow at supersonic speeds. An approximate solution technique proposed by Uebelhack has been modified and extended to obtain a more consistent numerical procedure. A comparison of theoretical calculations with experimental data generally indicated good agreement over the experimentally available range of ratios of step height to boundary layer thickness from 7 to 0.05.

  12. Aerodynamics of the Viggen 37 aircraft. Part 1: General characteristics at low speed

    NASA Technical Reports Server (NTRS)

    Karling, K.

    1986-01-01

    A description of the aerodynamics of the Viggen 37 and its performances, especially at low speeds is presented. The aerodynamic requirements for the design of the Viggen 37 aircraft are given, including the basic design, performance requirement, and aerodynamic characteristics, static and dynamic load test results and flight test results. The Viggen 37 aircraft is designed to be used for air attack, surveillance, pursuit, and training applications. It is shown that this aircraft is suitable for short runways, and has good maneuvering, acceleration, and climbing characteristics. The design objectives for this aircraft were met by utilizing the effect produced by the interference between two triangular wings, positioned in tandem.

  13. Design and aerodynamic characteristics of a span morphing wing

    NASA Astrophysics Data System (ADS)

    Yu, Yuemin; Liu, Yanju; Leng, Jinsong

    2009-03-01

    Flight vehicles are often designed to function around a primary operating point such as an efficient cruise or a high maneuverability mode. Performance and efficiency deteriorate rapidly as the airplane moves towards other portions of the flight envelope. One solution to this quandary is to radically change the shape of the aircraft. This yields both improved efficiency and a larger flight envelope. This global shape change is an example of morphing aircraft . One concept of morphing is the span morphing wing in which the wingspan is varied to accommodate multiple flight regimes. This type of design allows for at least two discreet modes of the aircraft. The original configuration, in which the extensible portion of the wing is fully retracted, yields a high speed dash mode. Fully extending the wing provides the aircraft with a low speed mode tailored for fine tracking and loiter tasks. This paper discusses the design of a span morphing wing that permits a change in the aspect ratio while simultaneously supporting structural wing loads. The wing cross section is maintained by NACA 4412 rib sections . The span morphing wing was investigated in different configurations. The wing area and the aspect ratio of the span morphing wing increase as the wings pan increases. Computational aerodynamics are used to estimate the performance and dynamic characteristics of each wing shape of this span morphing wing as its wingspan is changed. Results show that in order to obtain the same lift, the conventional wing requires a larger angle of attach(AOA) than that of the span morphing wing.The lift of the span morphing wing increases as the wing span ,Mach number and AOA increases.

  14. Aerodynamic characteristics and respiratory deposition of fungal fragments

    NASA Astrophysics Data System (ADS)

    Cho, Seung-Hyun; Seo, Sung-Chul; Schmechel, Detlef; Grinshpun, Sergey A.; Reponen, Tiina

    The purpose of this study was to investigate the aerodynamic characteristics of fungal fragments and to estimate their respiratory deposition. Fragments and spores of three different fungal species ( Aspergillus versicolor, Penicillium melinii, and Stachybotrys chartarum) were aerosolized by the fungal spore source strength tester (FSSST). An electrical low-pressure impactor (ELPI) measured the size distribution in real-time and collected the aerosolized fungal particles simultaneously onto 12 impactor stages in the size range of 0.3-10 μm utilizing water-soluble ZEF-X10 coating of the impaction stages to prevent spore bounce. For S. chartarum, the average concentration of released fungal fragments was 380 particles cm -3, which was about 514 times higher than that of spores. A. versicolor was found to release comparable amount of spores and fragments. Microscopic analysis confirmed that S. chartarum and A. versicolor did not show any significant spore bounce, whereas the size distribution of P. melinii fragments was masked by spore bounce. Respiratory deposition was calculated using a computer-based model, LUDEP 2.07, for an adult male and a 3-month-old infant utilizing the database on the concentration and size distribution of S. chartarum and A. versicolor aerosols measured by the ELPI. Total deposition fractions for fragments and spores were 27-46% and 84-95%, respectively, showing slightly higher values in an infant than in an adult. For S. chartarum, fragments demonstrated 230-250 fold higher respiratory deposition than spores, while the number of deposited fragments and spores of A. versicolor were comparable. It was revealed that the deposition ratio (the number of deposited fragments divided by that of deposited spores) in the lower airways for an infant was 4-5 times higher than that for an adult. As fungal fragments have been shown to contain mycotoxins and antigens, further exposure assessment should include the measurement of fungal fragments for

  15. Numerical analysis of supersonic gas-dynamic characteristic in laser cutting

    NASA Astrophysics Data System (ADS)

    Guo, Shaogang; Jun, Hu; Lei, Luo; Yao, Zhenqiang

    2009-01-01

    The influence of the processing parameters on the dynamic characteristic of supersonic impinging jet in laser cutting is studied numerically. The numerical modeling of a supersonic jet impinging on a plate with a hole is presented to analyze the gas jet-workpiece interaction. The model is able to make quantitative predictions of the effect of the standoff distance and exit Mach number on the mass flow rate and the axial thrust. The numerical results show that the suitable cutting range is slightly different for different exit Mach number, but the optimal cutting parameter for certain exit total pressure is nearly changeless. So the better cut quality and capacity can be obtained mainly by setting the suitable standoff distance for a certain nozzle pressure.

  16. On-orbit free molecular flow aerodynamic characteristics of a proposal space operations center configuration

    NASA Technical Reports Server (NTRS)

    Romere, P. O.

    1982-01-01

    A proposed configuration for a Space Operations Center is presented in its eight stages of buildup. The on orbit aerodynamic force and moment characteristics were calculated for each stage based upon free molecular flow theory. Calculation of the aerodynamic characteristics was accomplished through the use of an orbital aerodynamic computer program, and the computation method is described with respect to the free molecular theory used. The aerodynamic characteristics are presented in tabulated form for each buildup stage at angles of attack from 0 to 360 degrees and roll angles from -60 to +60 degrees. The reference altitude is 490 kilometers, however, the data should be applicable for altitudes below 490 kilometers down to approximately 185 kilometers.

  17. Nozzle and wing geometry effects on OTW aerodynamic characteristics

    NASA Technical Reports Server (NTRS)

    Vonglahn, U.; Groesbeck, D.

    1976-01-01

    The effects of nozzle geometry and wing size on the aerodynamic performance of several 5:1 aspect ratio slot nozzles are presented for over-the-wing (OTW) configurations. Nozzle geometry variables include roof angle, sidewall cutback, and nozzle chordwise location. Wing variables include chord size, and flap deflection. Several external deflectors also were included for comparison. The data indicate that good flow turning may not necessarily provide the best aerodynamic performance. The results suggest that a variable exhaust nozzle geometry offers the best solution for a viable OTW configuration.

  18. A parametric study of planform and aeroelastic effects on aerodynamic center, alpha- and q- stability derivatives. Appendix A: A computer program for calculating alpha- and q- stability derivatives and induced drag for thin elastic aeroplanes at subsonic and supersonic speeds

    NASA Technical Reports Server (NTRS)

    Roskam, J.; Lan, C.; Mehrotra, S.

    1972-01-01

    The computer program used to determine the rigid and elastic stability derivatives presented in the summary report is listed in this appendix along with instructions for its use, sample input data and answers. This program represents the airplane at subsonic and supersonic speeds as (a) thin surface(s) (without dihedral) composed of discrete panels of constant pressure according to the method of Woodward for the aerodynamic effects and slender beam(s) for the structural effects. Given a set of input data, the computer program calculates an aerodynamic influence coefficient matrix and a structural influence coefficient matrix.

  19. Supersonic Cruise Research 1979, part 1

    NASA Technical Reports Server (NTRS)

    1980-01-01

    Aerodynamics, stability and control, propulsion, and environmental factors of the supersonic cruise aircraft are discussed. Other topics include airframe structures and materials, systems integration, and economics.

  20. Characteristics of laser supersonic heating method for producing micro metallic particles

    NASA Astrophysics Data System (ADS)

    Lin, Shih-Lung; Lin, Jehnming

    2005-10-01

    In this article, the authors analyzed the process characteristics of laser supersonic heating method for producing metallic particles and predicted the in-flight tracks and shapes of micro-particles. A pulse Nd-YAG laser was used to heat the carbon steel target placed within an air nozzle. The high-pressure air with supersonic velocity was used to carry out carbon steel particles in the nozzle. The shock wave structures at the nozzle exit were visualized by the shadowgraph method. The carbon steel particles produced by laser supersonic heating method were grabbed and the spraying angles of the particle tracks were visualized. The velocity of the in-flight particles was measured by a photodiode sensor and compared with the numerical result. The solidification of carbon steel particles with diameters of 1-50 μm in compressible flow fields were investigated. The result shows that there is no significant difference in the dimension of solid carbon steel particles produced at shock wave fields under various entrance pressures (3-7 bar) with a constant laser energy radiation.

  1. Transonic aerodynamic characteristics of a proposed wing-body reusable launch vehicle concept

    NASA Technical Reports Server (NTRS)

    Springer, A. M.

    1995-01-01

    A proposed wing-body reusable launch vehicle was tested in the NASA Marshall Space Flight Center's 14 x 14-inch trisonic wind tunnel during the winter of 1994. This test resulted in the vehicle's subsonic and transonic, Mach 0.3 to 1.96, longitudinal and lateral aerodynamic characteristics. The effects of control surface deflections on the basic vehicle's aerodynamics, including a body flap, elevons, ailerons, and tip fins, are presented.

  2. Transonic Aerodynamic Characteristics of a Proposed Wing Body Reusable Launch Vehicle Concept

    NASA Technical Reports Server (NTRS)

    Springer, Anthony M.

    1996-01-01

    A proposed wing body reusable launch vehicle was tested in the Marshall Space Flight Center (MSFC) 14 x 14 inch trisonic wind tunnel during the winter of 1994. this test resulted in the vehicle's subsonic and transonic (Mach 03 to 1.96) longitudinal and lateral aerodynamic characteristics. The effects of control surface deflections on the basic vehicle aerodynamics including a body flap, elevons, ailerons, and tip fins are presented.

  3. Theoretical and Experimental Unsteady Aerodynamics Compared for a Linear Oscillating Cascade With a Supersonic Leading-Edge Locus

    NASA Technical Reports Server (NTRS)

    Ramsey, John K.; Erwin, Dan

    2005-01-01

    Experimental data were obtained to help validate analytical and computational fluid dynamics (CFD) codes used to compute unsteady cascade aerodynamics in a supersonicaxial- flow regime. Results from two analytical codes and one CFD code were compared with experimental data. One analytical code did not account for airfoil thickness or camber; another, using piston theory (piston code), accounted for thickness and camber upstream of the first shockwave/airfoil impingement locations. The Euler CFD code accounted fully for airfoil shape.

  4. Aerodynamic characteristics of NACA 4412 airfoil sction with flap

    NASA Astrophysics Data System (ADS)

    Ockfen, Alex E.; Matveev, Konstantin I.

    2009-09-01

    Wing-in-Ground vehicles and aerodynamically assisted boats take advantage of increased lift and reduced drag of wing sections in the ground proximity. At relatively low speeds or heavy payloads of these craft, a flap at the wing trailing-ground-effect flow id numerically investigated in this study. The computational method consists of a steady-state, incompressible, finite volume method utilizing the Spalart-Allmaras turbulence model. Grid generation and solution of the Navier-Stokes equations are completed flow with a flap, as well as ground-effect motion without a flap. Aerodynamic forces are plain flap. Changes in the flow introduced with the flap addition are also discussed. Overall, the use of a flap on wings with small attack angles is found to be beneficial for small flap deflections up to 5% of the chord, where the contribution of lift augmentation exceeds the drag increase, yielding an augmented lift-to-drag ratio

  5. Performance characteristics of aerodynamically optimum turbines for wind energy generators

    NASA Technical Reports Server (NTRS)

    Rohrbach, C.; Worobel, R.

    1975-01-01

    This paper presents a brief discussion of the aerodynamic methodology for wind energy generator turbines, an approach to the design of aerodynamically optimum wind turbines covering a broad range of design parameters, some insight on the effect on performance of nonoptimum blade shapes which may represent lower fabrication costs, the annual wind turbine energy for a family of optimum wind turbines, and areas of needed research. On the basis of the investigation, it is concluded that optimum wind turbines show high performance over a wide range of design velocity ratios; that structural requirements impose constraints on blade geometry; that variable pitch wind turbines provide excellent power regulation and that annual energy output is insensitive to design rpm and solidity of optimum wind turbines.

  6. Aerodynamic characteristics of sixteen electric, hybrid, and subcompact vehicles

    NASA Technical Reports Server (NTRS)

    Kurtz, D. W.

    1979-01-01

    An elementary electric and hybrid vehicle aerodynamic data base was developed using data obtained on sixteen electric, hybrid, and sub-compact production vehicles tested in the Lockheed-Georgia low-speed wind tunnel. Zero-yaw drag coefficients ranged from a high of 0.58 for a boxey delivery van and an open roadster to a low of about 0.34 for a current four-passenger proto-type automobile which was designed with aerodynamics as an integrated parameter. Vehicles were tested at yaw angles up to 40 degrees and a wing weighting analysis is presented which yields a vehicle's effective drag coefficient as a function of wing velocity and driving cycle. Other parameters investigated included the effects of windows open and closed, radiators open and sealed, and pop-up headlights. Complete six-component force and moment data are presented in both tabular and graphical formats. Only limited commentary is offered since, by its very nature, a data base should consist of unrefined reference material. A justification for pursuing efficient aerodynamic design of EHVs is presented.

  7. Parametric experimental studies on mixing characteristics within a low area ratio rectangular supersonic gaseous ejector

    NASA Astrophysics Data System (ADS)

    Karthick, S. K.; Rao, Srisha M. V.; Jagadeesh, G.; Reddy, K. P. J.

    2016-07-01

    We use the rectangular gaseous supersonic ejector as a platform to study the mixing characteristics of a confined supersonic jet. The entrainment ratio (ER) of the ejector, the non-mixed length (LNM), and potential core length (LPC) of the primary supersonic jet are measures to characterize mixing within the supersonic ejector. Experiments are carried out on a low area ratio rectangular supersonic ejector with air as the working fluid in both primary and secondary flows. The design Mach number of the nozzle (MPD = 1.5-3.0) and primary flow stagnation pressure (Pop = 4.89-9.89 bars) are the parameters that are varied during experimentation. Wall static pressure measurements are carried out to understand the performance of the ejector as well as to estimate the LNM (the spatial resolution is limited by the placement of pressure transducers). Well-resolved flow images (with a spatial resolution of 50 μm/pixel and temporal resolution of 1.25 ms) obtained through Planar Laser Mie Scattering (PLMS) show the flow dynamics within the ejector with clarity. The primary flow and secondary flow are seeded separately with acetone that makes the LNM and LPC clearly visible in the flow images. These parameters are extracted from the flow images using in-house image processing routines. A significant development in this work is the definition of new scaling parameters within the ejector. LNM, non-dimensionalized with respect to the fully expanded jet height hJ, is found to be a linear function of the Mach number ratio (Mach number ratio is defined as the ratio of design Mach number (MPD) and fully expanded Mach number (MPJ) of the primary jet). This definition also provides a clear demarcation of under-expanded and over-expanded regimes of operation according to [MPD/MPJ] > 1 and [MPD/MPJ] < 1, respectively. It is observed that the ER increased in over-expanded mode (to 120%) and decreased in under-expanded mode (to 68%). Similarly, LNM decreased (to 21.8%) in over-expanded mode

  8. Aerodynamic characteristics at Mach numbers of 1.5, 1.8, and 2.0 of a blended wing-body configuration with and without integral canards

    NASA Technical Reports Server (NTRS)

    Robins, A. W.; Lamb, M.; Miller, D. S.

    1979-01-01

    An exploratory, experimental, and theoretical investigation was made of a cambered, twisted, and blended wing-body concept with and without integral canard surfaces. Theoretical calculations of the static longitudinal and lateral aerodynamic characteristics of the wing-body configurations were compared with the characteristics obtained from tests of a model in the Langley Unitary Plan wind tunnel. Mach numbers of 1.5, 1.8, and 2.0 and a Reynolds number per meter of 6.56 million were used in the calculations and tests. Overall results suggest that planform selection is extremely important and that the supplemental application of new calculation techniques should provide a process for the design of supersonic wings in which spanwise distribution of upwash and leading-edge thrust might be rationally controlled and exploited.

  9. Theoretical evaluation of high speed aerodynamics for arrow wing configurations

    NASA Technical Reports Server (NTRS)

    Dollyhigh, S. M.

    1978-01-01

    A limited study in the use of theoretical methods to calculate the high speed aerodynamics of arrow wing supersonic cruise configurations was conducted. The study consisted of correlations with existing wind tunnel data at Mach numbers from 0.8 to 2.7, using theoretical methods to extrapolate the wind tunnel data to full scale flight conditions, and presentation of a typical supersonic data package for an advanced supersonic transport application prepared using the theoretical methods. A brief description of the methods and their application was given. In general, all three methods had excellent correlation with wind tunnel data at supersonic speeds for drag and lift characteristics and fair to poor agreement with pitching moment characteristics. The VORLAX program had excellent correlation with wind tunnel data at subsonic speeds for lift and pitching moment characteristics and fair agreement in drag characteristics.

  10. Comparison of the Aerodynamic Characteristics of Similar Models in Two Size Wind Tunnels at Transonic Speeds

    NASA Technical Reports Server (NTRS)

    Springer, Anthony M.

    1998-01-01

    The aerodynamic characteristics of two similar models of a lifting body configuration were run in two transonic wind tunnels, one a 16 foot the other a 14-inch and are compared. The 16 foot test used a 2% model while the 14-inch test used a 0.7% scale model. The wind tunnel model configurations varied only in vertical tail size and an aft sting shroud. The results from these two tests compare the effect of tunnel size, Reynolds number, dynamic pressure and blockage on the longitudinal aerodynamic characteristics of the vehicle. The data accuracy and uncertainty are also presented. It was concluded from these tests that the data resultant from a small wind tunnel compares very well to that of a much larger wind tunnel in relation to total vehicle aerodynamic characteristics.

  11. Investigation of aerodynamic characteristics of a hypersonic flow around bodies of revolution with a permeable tip

    NASA Astrophysics Data System (ADS)

    Sidnyaev, N. I.

    2007-03-01

    Results of experimental investigations of aerodynamic characteristics of models of high-velocity flying vehicles consisting of a combination of a blunt cone, a cylinder, and a conical tail fin are presented. The model forebody is cooled by porous blowing. The choice of such a configuration is determined by the necessity of optimizing the arrangement of high-velocity flying vehicles on the launcher and their aerodynamic characteristics under conditions of intense surface mass transfer (decrease in drag and heat transfer and increase in static and dynamic stability).

  12. Space shuttle plume/simulation application: Results and math model supersonic data

    NASA Technical Reports Server (NTRS)

    Boyle, W.; Conine, B.; Bell, G.

    1979-01-01

    The analysis of pressure and gage wind tunnel data from space shuttle wind tunnel test IA138 was performed to define the aerodynamic influence of the main propulsion system and solid rocket booster plumes on the total vehicles, elements, and components of the space shuttle vehicle during the supersonic portion of ascent flight. A math model of the plume induced aerodynamic characteristics was developed for a range of Mach numbers to match the forebody aerodynamic math model. The base aerodynamic characteristics are presented in terms of forces and moments versus attitude. Total vehicle base and forebody aerodynamic characteristics are presented in terms of aerodynamic coefficients for Mach numbers from 1.55 to 2.5.

  13. Effects of Leading Edge Defect on the Aerodynamic and Flow Characteristics of an S809 Airfoil

    PubMed Central

    Wang, Yan; Zheng, Xiaojing; Hu, Ruifeng; Wang, Ping

    2016-01-01

    Background Unexpected performance degradation occurs in wind turbine blades due to leading edge defect when suffering from continuous impacts with rain drops, hails, insects, or solid particles during its operation life. To assess this issue, this paper numerically investigates the steady and dynamic stall characteristics of an S809 airfoil with various leading edge defects. More leading edge defect sizes and much closer to practical parameters are investigated in the paper. Methodology Numerical computation is conducted using the SST k-ω turbulence model, and the method has been validated by comparison with existed published data. In order to ensure the calculation convergence, the residuals for the continuity equation are set to be less than 10−7 and 10−6 in steady state and dynamic stall cases. The simulations are conducted with the software ANSYS Fluent 13.0. Results It is found that the characteristics of aerodynamic coefficients and flow fields are sensitive to leading edge defect both in steady and dynamic conditions. For airfoils with the defect thickness of 6%tc, leading edge defect has a relative small influence on the aerodynamics of S809 airfoil. For other investigated defect thicknesses, leading edge defect has much greater influence on the flow field structures, pressure coefficients and aerodynamic characteristics of airfoil at relative small defect lengths. For example, the lift coefficients decrease and drag coefficients increase sharply after the appearance of leading edge defect. However, the aerodynamic characteristics could reach a constant value when the defect length is large enough. The flow field, pressure coefficient distribution and aerodynamic coefficients do not change a lot when the defect lengths reach to 0.5%c,1%c, 2%c and 3%c with defect thicknesses of 6%tc, 12%tc,18%tc and 25%tc, respectively. In addition, the results also show that the critical defect length/thickness ratio is 0.5, beyond which the aerodynamic characteristics

  14. Aerodynamic Characteristics of Two Rotary Wing UAV Designs

    NASA Technical Reports Server (NTRS)

    Jones, Henry E.; Wong, Oliver D.; Noonan, Kevin W.; Reis, Deane G.; Malovrh, Brendon D.

    2006-01-01

    This paper presents the results of an experimental investigation of two rotary-wing UAV designs. The primary goal of the investigation was to provide a set of interactional aerodynamic data for an emerging class of rotorcraft. The present paper provides an overview of the test and an introduction to the test articles, and instrumentation. Sample data in the form of a parametric study of fixed system lift and drag coefficient response to changes in configuration and flight condition for both rotor off and on conditions are presented. The presence of the rotor is seen to greatly affect both the character and magnitude of the response. The affect of scaled stores on body drag is observed to be dependent on body shape.

  15. AERODYNAMIC CHARACTERISTICS OF TWO ROTARY WING UAV DESIGNS

    NASA Technical Reports Server (NTRS)

    Jones, Henry E.; Wong, Oliver D.; Noonan, Kevin W.; Reis, Deane G.; Malovrh, Brendon D.

    2006-01-01

    This paper presents the results of an experimental investigation of two rotary-wing UAV designs. The primary goal of the investigation was to provide a set of interactional aerodynamic data for an emerging class of rotorcraft. The present paper provides an overview of the test and an introduction to the test articles, and instrumentation. Sample data in the form of a parametric study of fixed system lift and drag coefficient response to changes in configuration and flight condition for both rotor off and on conditions are presented. The presence of the rotor is seen to greatly affect both the character and magnitude of the response. The affect of scaled stores on body drag is observed to be dependent on body shape.

  16. A preliminary study of the performance and characteristics of a supersonic executive aircraft

    NASA Technical Reports Server (NTRS)

    Mascitti, V. R.

    1977-01-01

    The impact of advanced supersonic technologies on the performance and characteristics of a supersonic executive aircraft was studied in four configurations with different engine locations and wing/body blending and an advanced nonafterburning turbojet or variable cycle engine. An M 2.2 design Douglas scaled arrow-wing was used with Learjet 35 accommodations. All four configurations with turbojet engines meet the performance goals of 5926 km (3200 n.mi.) range, 1981 meters (6500 feet) takeoff field length, and 77 meters per second (150 knots) approach speed. The noise levels of of turbojet configurations studied are excessive. However, a turbojet with mechanical suppressor was not studied. The variable cycle engine configuration is deficient in range by 555 km (300 n.mi) but nearly meets subsonic noise rules (FAR 36 1977 edition), if coannular noise relief is assumed. All configurations are in the 33566 to 36287 kg (74,000 to 80,000 lbm) takeoff gross weight class when incorporating current titanium manufacturing technology.

  17. Supersonic aerodynamic characteristics of a circular body Earth-to-Orbit vehicle

    NASA Technical Reports Server (NTRS)

    Ware, George M.; Engelund, Walter C.; Macconochie, Ian O.

    1994-01-01

    The circular body configuration is a generic single- or multi-stage reusable Earth-to-orbit transport. A thick clipped-delta wing is the major lifting surface. For directional control, three different vertical fin arrangements were investigated: a conventional aft-mounted center fin, wingtip fins, and a nose-mounted fin. The tests were conducted in the Langley Unitary Plan Wind Tunnel. The configuration is longitudinally stable about the estimated center of gravity of 0.72 body length up to a Mach number of about 3.0. Above Mach 3.0, the model is longitudinally unstable at low angles of attack but has a stable secondary trim point at angles of attack above 30 deg. The model has sufficient pitch control authority with elevator and body flap to produce stable trim over the test range. The model with the center fin is directionally stable at low angles of attack up to a Mach number of 3.90. The rudder-like surfaces on the tip fins and the all-movable nose fin are designed as active controls to produce artificial directional stability and are effective in producing yawing moment. The wing trailing-edge aileron surfaces are effective in producing rolling moment, but they also produce large adverse yawing moment.

  18. Effects of stores on longitudinal aerodynamic characteristics of a fighter at supersonic speeds

    NASA Technical Reports Server (NTRS)

    Dollyhigh, S. M.; Sangiorgio, G.; Monta, W. J.

    1978-01-01

    Experimental investigations of single and twin stores representative of advanced, elliptical cross section missile concepts were made at Mach numbers from 1.60 to 2.16 to substantiate theoretically predicted results. The stores were mounted on the fuselage of a model representing a fighter configuration. Store base closure effects in the carriage condition were also obtained through tests with and without base closure fairings.

  19. Effects of reaction control system jet flow field interactions on the aerodynamic characteristics of a 0.010-scale space shuttle orbiter model in the Langley Research Center 31 inch CFHT (OA85)

    NASA Technical Reports Server (NTRS)

    Daileda, J. J.; Marroquin, J.

    1974-01-01

    An experimental investigation was conducted to obtain detailed effects on supersonic vehicle hypersonic aerodynamic and stability and control characteristics of reaction control system jet flow field interactions with the local vehicle flow field. A 0.010-scale model was used. Six-component force data and wing, elevon, and body flap surface pressure data were obtained through an angle-of-attack range of -10 to +35 degrees with 0 deg angle of sideslip. The test was conducted with yaw, pitch and roll jet simulation at a free-stream Mach number of 10.3 and reaction control system plume simulation of flight dynamic pressures of 5, 10 and 20 PSF.

  20. Aerodynamic Characteristics of Missile Configurations with Wings of Low Aspect Ratio for Various Combinations of Forebodies, Afterbodies, and Nose Shapes for Combined Angles of Attack and Sideslip at a Mach Number of 2.01

    NASA Technical Reports Server (NTRS)

    Robinson, Ross B

    1957-01-01

    An investigation has been made in the Langley 4-by-4-foot supersonic pressure tunnel to determine the aerodynamic characteristics of a series of missile configurations having low-aspect-ratio wings at a Mach number of 2.01. The effects of wing plan form and size, length-diameter ratio, forebody and afterbody length, boattailed and flared afterbodies, and component force and moment data are presented for combined angles of attack and sideslip to about 28 degrees. No analysis of the data was made in this report.

  1. Aerodynamic Characteristics of Parachutes at Mach Numbers from 1.6 to 3

    NASA Technical Reports Server (NTRS)

    1961-01-01

    Aerodynamic Characteristics of Parachutes at Mach Numbers from 1.6 to 3. A wind-tunnel investigation was conducted to determine the parameters affecting the aerodynamic performance of drogue parachutes in the Mach number range from 1.6 to 3. Flow studies of both rigid and flexible-parachute models were made by means of high-speed schlieren motion pictures and drag coefficients of the flexible-parachute models were measured at simulated altitudes from about 50,000 to 120,000 feet. [Entire movie available on DVD from CASI as Doc ID 20070030970. Contact help@sti.nasa.gov

  2. Transonic aerodynamic characteristics of a supercritical-wing transport model with trailing-edge controls

    NASA Technical Reports Server (NTRS)

    Mann, M. J.; Langhans, R. A.

    1977-01-01

    The effects of wing trailing-edge control surfaces on the static transonic aerodynamic characteristics of a transport configuration with a supercritical wing were studied. The configuration was tested with both an area-ruled fuselage and a cylindrical fuselage. The Mach number range was from 0.80 to 0.96 and the angle of attack range was from -1 deg to 12 deg. The Reynolds number was 1,580,000 based on the mean aerodynamic chord. Tabular data are presented.

  3. Development of a morphing flap using shape memory alloy actuators: the aerodynamic characteristics of a morphing flap

    NASA Astrophysics Data System (ADS)

    Ko, Seung-Hee; Bae, Jae-Sung; Rho, Jin-Ho

    2014-07-01

    The discontinuous contour of a wing with conventional flaps diminishes the aerodynamic performance of an aircraft. A wing with a continuous contour does not experience extreme flow stream fluctuations during flight, and consequently has good aerodynamic characteristics. In this study, a morphing flap using shape memory alloy actuators is proposed, designed and fabricated, and its aerodynamic characteristics are investigated using aerodynamic analyses and wind tunnel tests. The ribs of the morphing flap are designed and fabricated with multiple elements joined together in a way that allows relative rotations of adjacent elements and forms a smooth contour of the morphing flap. The aerodynamic analyses of this multiple-element morphing-flap wing are performed using XFLR pro; its aerodynamic performance is compared with that of a mechanical-flap wing, and is measured through wind-tunnel tests.

  4. Overview of Selected Measurement Techniques for Aerodynamics Testing in the NASA Langley Unitary Plan Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.

    2000-01-01

    An overview is given of selected measurement techniques used in the NASA Langley Research Center (LaRC) Unitary Plan Wind Tunnel (UPWT) to determine the aerodynamic characteristics of aerospace vehicles operating at supersonic speeds. A broad definition of a measurement technique is adopted in this paper and is any qualitative or quantitative experimental approach that provides information leading to the improved understanding of the supersonic aerodynamic characteristics. On surface and off-surface measurement techniques used to obtain discrete (point) and global (field) measurements and planar and global flow visualizations are described, and examples of all methods are included. The discussion is limited to recent experiences in the UPWT and is. therefore, not an exhaustive review of existing experimental techniques. The diversity and high quality of the measurement techniques and the resultant data illustrate the capabilities of a around-based experimental facility and the key role that it plays in the advancement of our understanding, prediction, and control of supersonic aerodynamics.

  5. Noise and economic characteristics of an advanced blended supersonic transport concept

    NASA Technical Reports Server (NTRS)

    Molloy, J. K.; Grantham, W. D.; Neubauer, M. J., Jr.

    1982-01-01

    Noise and economic characteristics were obtained for an advanced supersonic transport concept that utilized wing body blending, a double bypass variable cycle engine, superplastically formed and diffusion bonded titanium in both the primary and secondary structures, and an alternative interior arrangement that provides increased seating capacity. The configuration has a cruise Mach number of 2.62, provisions for 290 passengers, a mission range of 8.19 Mm (4423 n.mi.), and an average operating cruise lift drag ratio of 9.23. Advanced operating procedures, which have the potential to reduce airport community noise, were explored by using a simulator. Traded jet noise levels of 105.7 and 103.4 EPNdB were obtained by using standard and advanced takeoff operational procedures, respectively. A new method for predicting lateral attenuation was utilized in obtaining these jet noise levels.

  6. Flow characteristics and micro-scale metallic particle formation in the laser supersonic heating technique

    NASA Astrophysics Data System (ADS)

    Lin, Shih-Lung; Lin, Jehnming

    2007-02-01

    The characteristics of the supersonic flow of the laser heating technique for producing micro-scale metallic particles were investigated in this study. A numerical model was established to predict the flow fields and particle trajectories leaving a spray nozzle with shock wave effects. The compressible flow of the shock waves and the trajectories of particles in diameters of 1-20 μm were simulated and compared with the flow visualization. In the experiment, a pulsed Nd-YAG laser was used as heat source on a carbon steel target within the nozzle, and the carbon steel particles were ejected by high-pressure air. The result shows that the shock wave structures were generated at various entrance pressures, and there is a significant increase in the amount of carbon steel particles and the spraying angles by increasing the entrance air pressure.

  7. Comparison of Theoretical and Experimental Heat-Transfer Characteristics of Bodies of Revolution at Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Scherrer, Richard

    1951-01-01

    An investigation of the three important factors that determine convective heat-transfer characteristics at supersonic speeds, location boundary-layer transition, recovery factor, and heat-transfer parameter has been performed at Mach numbers from 1.49 to 1.18. The bodies of revolution that were tested had, in most cases, laminar boundary layers, and the test results have been compared with available theory. Boundary-layer transition was found to be affected by heat transfer. Adding heat to a laminar boundary layer caused transition to move forward on the test body, while removing heat caused transition to move rearward. These experimental results and the implications of boundary-layer-stability theory are in qualitative agreement.

  8. Numerical and experimental study on flame structure characteristics in a supersonic combustor with dual-cavity

    NASA Astrophysics Data System (ADS)

    Yang, Yixin; Wang, Zhenguo; Sun, Mingbo; Wang, Hongbo; Li, Li

    2015-12-01

    Combined numerical and experimental approaches have been implemented to investigate the quasi-steady flame characteristics of supersonic combustion in tandem and parallel dual-cavity. In simulation, a hybrid Large Eddy Simulation (LES)/assumed sub-grid Probability Density Function (PDF) closure model was carried out. Comparison of calculation and experiment as well as comparison of the two configurations are qualitatively and quantitatively performed regarding the flame structure and other flowfield features. Simulation shows a good level of agreement with experimental observation and measurement in terms of instantaneous and time-averaged results. Given the same fuel equivalence ratio, the parallel dual-cavity with the two opposite injections gathers the major combustion around the cavities, thus leading to the concentrated heat release, the greatly extended recirculation zones and the converging-diverging core flow path. Only intermittent stray flame packets can be found in the downstream region. Flame in the combustor with tandem dual-cavity appears to be stabilized by the upstream cavity shear layer and grows gradually to the second cavity, peaking its most intensity in the middle section between the two cavities. For both dual-cavity configurations, the strongest reaction takes place in near chemistry stoichiometric region around the flame edge, and is mainly confined in the supersonic region supported by the inner subsonic combustion. The coexistence of three parts plays a vital role in flame stabilization in the parallel and tandem dual-cavity: a reacting reservoir transferring hot products and activated radicals within the cavity recirculation zone, the hydrogen-rich premixed flame in the jet mixing region, and the downstream diffusion flames supported by the upstream premixed combustion region. In addition, for the parallel dual-cavity under the given condition, significant reaction are present near jet exit upstream the cavity leading edge.

  9. Numerical simulation of inducing characteristics of high energy electron beam plasma for aerodynamics applications

    NASA Astrophysics Data System (ADS)

    Yongfeng, DENG; Jian, JIANG; Xianwei, HAN; Chang, TAN; Jianguo, WEI

    2017-04-01

    The problem of flow active control by low temperature plasma is considered to be one of the most flourishing fields of aerodynamics due to its practical advantages. Compared with other means, the electron beam plasma is a potential flow control method for large scale flow. In this paper, a computational fluid dynamics model coupled with a multi-fluid plasma model is established to investigate the aerodynamic characteristics induced by electron beam plasma. The results demonstrate that the electron beam strongly influences the flow properties, not only in the boundary layers, but also in the main flow. A weak shockwave is induced at the electron beam injection position and develops to the other side of the wind tunnel behind the beam. It brings additional energy into air, and the inducing characteristics are closely related to the beam power and increase nonlinearly with it. The injection angles also influence the flow properties to some extent. Based on this research, we demonstrate that the high energy electron beam air plasma has three attractive advantages in aerodynamic applications, i.e. the high energy density, wide action range and excellent action effect. Due to the rapid development of near space hypersonic vehicles and atmospheric fighters, by optimizing the parameters, the electron beam can be used as an alternative means in aerodynamic steering in these applications.

  10. Unsteady Aerodynamic Effects on the Flight Characteristics of an F-16XL Configuration

    NASA Technical Reports Server (NTRS)

    Wang, Zhongjun; Lan, C. Edward; Brandon, Jay M.

    2000-01-01

    Unsteady aerodynamic models based on windtunnel forced oscillation test data and analyzed with a fuzzy logic algorithm arc incorporated into an F-16XL flight simulation code. The reduced frequency needed in the unsteady models is numerically calculated by using a limited prior time history of state variables in a least-square sense. Numerical examples arc presented to show the accuracy of the calculated reduced frequency. Oscillatory control inputs are employed to demonstrate the differences in the flight characteristics based on unsteady and quasi-steady aerodynamic models. Application of the unsteady aerodynamic models is also presented and the results are compared with one set of F16XIL longitudinal maneuver flight data. It is shown that the main differences in dynamic response are in the lateral-directional characteristics, with the quasi-steady model being more stable than the flight vehicle, while the unsteady model being more unstable. Similar conclusions can also be made in a simulated rapid sideslipping roll. To improve unsteady aerodynamic modeling, it is recommended to acquire test data with coupled motions in pitch, roll and yaw.

  11. Aerodynamic characteristics of the National Launch System (NLS) 1 1/2 stage launch vehicle

    NASA Technical Reports Server (NTRS)

    Springer, A. M.; Pokora, D. C.

    1994-01-01

    The National Aeronautics and Space Administration (NASA) is studying ways of assuring more reliable and cost effective means to space. One launch system studied was the NLS which included the l l/2 stage vehicle. This document encompasses the aerodynamic characteristics of the 1 l/2 stage vehicle. To support the detailed configuration definition two wind tunnel tests were conducted in the NASA Marshall Space Flight Center's 14x14-Inch Trisonic Wind Tunnel during 1992. The tests were a static stability and a pressure test, each utilizing 0.004 scale models. The static stability test resulted in the forces and moments acting on the vehicle. The aerodynamics for the reference configuration with and without feedlines and an evaluation of three proposed engine shroud configurations were also determined. The pressure test resulted in pressure distributions over the reference vehicle with and without feedlines including the reference engine shrouds. These pressure distributions were integrated and balanced to the static stability coefficients resulting in distributed aerodynamic loads on the vehicle. The wind tunnel tests covered a Mach range of 0.60 to 4.96. These ascent flight aerodynamic characteristics provide the basis for trajectory and performance analysis, loads determination, and guidance and control evaluation.

  12. Characteristic boundary conditions for three-dimensional transonic unsteady aerodynamics

    NASA Technical Reports Server (NTRS)

    Whitlow, W., Jr.

    1984-01-01

    Characteristic far-field boundary conditions for the three-dimensional unsteady transonic small disturbance potential equation have been developed. The boundary conditions were implemented in the XTRAN3S finite difference code and tested for a flat plate rectangular wing with a pulse in angle of attack; the freestream Mach number was 0.85. The calculated force response shows that the characteristic boundary conditions reduce disturbances that are reflected from the computational boundaries.

  13. Study of aerodynamic technology for VSTOL fighter/attack aircraft: Horizontal attitude concept

    NASA Technical Reports Server (NTRS)

    Brown, S. H.

    1978-01-01

    A horizontal attitude VSTOL (HAVSTOL) supersonic fighter attack aircraft powered by RALS turbofan propulsion system is analyzed. Reaction control for subaerodynamic flight is obtained in pitch and yaw from the RALS and roll from wingtip jets powered by bleed air from the RALS duct. Emphasis is placed on the development of aerodynamic characteristics and the identification of aerodynamic uncertainties. A wind tunnel program is shown to resolve some of the uncertainties. Aerodynamic data developed are static characteristics about all axes, control effectiveness, drag, propulsion induced effects and reaction control characteristics.

  14. Longitudinal aerodynamic characteristics of an externally blown flap powered lift model with several propulsive system simulators

    NASA Technical Reports Server (NTRS)

    Hoad, D. R.

    1974-01-01

    An investigation of a four-engine externally blown flap (EBF) powered-lift transport was conducted in the Langley V/STOL tunnel to determine the effect of different engine configurations on the longitudinal aerodynamic characteristics. The different engine configurations were simulated by five different sets of propulsion simulators on a single aircraft model. Longitudinal aerodynamic data were obtained for each simulator on each flap deflection corresponding to cruise, take-off, and landing at a range of angles of attack and various thrust coefficients. The bypass ratio (BPR) 6.2 engine simulator provided the best lift and drag characteristics of the five simulators tested in the take-off and landing configurations. The poor performance of the BPR 10.0 and 3.2 engine simulators can be attributed to a mismatch of engine-model sizes or poor engine location and orientation. Isolated engine wake surveys indicated that a reasonable assessment of the aerodynamic characteristics of an engine-wing-flap configuration could be made if qualitative information were available which defined the engine wake characteristics. All configurations could be trimmed easily with relatively small horizontal-tail incidence angles; however, the take-off landing configurations required a high-lift tail.

  15. Subsonic aerodynamic and flutter characteristics of several wings calculated by the SOUSSA P1.1 panel method

    NASA Technical Reports Server (NTRS)

    Yates, E. C., Jr.; Cunningham, H. J.; Desmarais, R. N.; Silva, W. A.; Drobenko, B.

    1982-01-01

    The SOUSSA (steady, oscillatory, and unsteady subsonic and supersonic aerodynamics) program is the computational implementation of a general potential flow analysis (by the Green's function method) that can generate pressure distributions on complete aircraft having arbitrary shapes, motions and deformations. Some applications of the initial release version of this program to several wings in steady and oscillatory motion, including flutter are presented. The results are validated by comparisons with other calculations and experiments. Experiences in using the program as well as some recent improvements are described.

  16. The Aerodynamic Characteristics of a Slotted Clark Y Wing as Affected by the Auxiliary Airfoil Position

    NASA Technical Reports Server (NTRS)

    Wenzinger, Carl J; Shortal, Joseph A

    1932-01-01

    Aerodynamic force tests on a slotted Clark Y wing were conducted in a vertical wind tunnel to determine the best position for a given auxiliary airfoil with respect to the main wing. A systematic series of 100 changes in location of the auxiliary airfoil were made to cover all the probable useful ranges of slot gap, slot width, and slot depth. The results of the investigation may be applied to the design of automatic or controlled slots on wings with geometric characteristics similar to the wing tested. The best positions of the auxiliary airfoil were covered by the range of the tests, and the position for desired aerodynamic characteristics may easily be obtained from charts prepared especially for the purpose.

  17. Aerodynamic characteristics at Mach 6 of a wing-body concept for a hypersonic research airplane

    NASA Technical Reports Server (NTRS)

    Dillon, J. L.; Pittman, J. L.

    1978-01-01

    The static aerodynamic characteristics of a 1/30 scale model of a wing-body concept for a high speed research airplane were investigated in the Langley 20 inch Mach six tunnel. The investigation consisted of configuration buildup from the basic body by adding a wing, center vertical tail, three-module scramjet, and six-module scramjet engine. The test Mach number was six at a Reynolds number, based on model fuselage length, of about 13,700,000. The test angle-of-attack range was 4 to 20 D at constant angles of sideslip of 0, 2, and 4 deg. The elevons were deflected from 10 to -15 D for pitch control. Roll and yaw control were investigated. Experimental aerodynamic characteristics are compared with analytical elements.

  18. Calculation of aerodynamic characteristics of airplane configurations at high angles of attack

    NASA Technical Reports Server (NTRS)

    Tseng, J. B.; Lan, C. Edward

    1988-01-01

    Calculation of longitudinal and lateral directional aerodynamic characteristics of airplanes by the VORSTAB code is examined. The numerical predictions are based on the potential flow theory with corrections of high angle of attack phenomena; namely, vortex flow and boundary layer separation effects. To account for the vortex flow effect, vortex lift, vortex action point, augmented vortex lift and vortex breakdown effect through the method of suction analogy are included. The effect of boundary layer separation is obtained by matching the nonlinear section data with the three dimensional lift characteristics iteratively. Through correlation with results for nine fighter configurations, it is concluded that reasonably accurate prediction of longitudinal and static lateral directional aerodynamics can be obtained with the VORSTAB code up to an angle of attack at which wake interference and forebody vortex effect are not important. Possible reasons for discrepancy at higher angles of attack are discussed.

  19. Study of aerodynamic technology for single-cruise-engine V/STOL fighter/attack aircraft

    NASA Technical Reports Server (NTRS)

    Mark, L.

    1982-01-01

    Conceptual designs and analyses were conducted on two V/STOL supersonic fighter/attack aircraft. These aircraft feature low footprint temperature and pressure thrust augmenting ejectors in the wings for vertical lift, combined with a low wing loading, low wave drag airframe for outstanding cruise and supersonic performance. Aerodynamic, propulsion, performance, and mass properties were determined and are presented for each aircraft. Aerodynamic and Aero/Propulsion characteristics having the most significant effect on the success of the up and away flight mode were identified, and the certainty with which they could be predicted was defined. A wind tunnel model and test program are recommended to resolve the identified uncertainties.

  20. Configuration Aerodynamics: Past - Present - Future

    NASA Technical Reports Server (NTRS)

    Wood, Richard M.; Agrawal, Shreekant; Bencze, Daniel P.; Kulfan, Robert M.; Wilson, Douglas L.

    1999-01-01

    The Configuration Aerodynamics (CA) element of the High Speed Research (HSR) program is managed by a joint NASA and Industry team, referred to as the Technology Integration Development (ITD) team. This team is responsible for the development of a broad range of technologies for improved aerodynamic performance and stability and control characteristics at subsonic to supersonic flight conditions. These objectives are pursued through the aggressive use of advanced experimental test techniques and state of the art computational methods. As the HSR program matures and transitions into the next phase the objectives of the Configuration Aerodynamics ITD are being refined to address the drag reduction needs and stability and control requirements of High Speed Civil Transport (HSCT) aircraft. In addition, the experimental and computational tools are being refined and improved to meet these challenges. The presentation will review the work performed within the Configuration Aerodynamics element in 1994 and 1995 and then discuss the plans for the 1996-1998 time period. The final portion of the presentation will review several observations of the HSR program and the design activity within Configuration Aerodynamics.

  1. Prediction of static aerodynamic characteristics for slender bodies alone and with lifting surfaces to very high angles of attack

    NASA Technical Reports Server (NTRS)

    Jorgensen, L. H.

    1977-01-01

    An engineering-type method is presented for computing normal-force and pitching-moment coefficients for slender bodies of circular and noncircular cross section alone and with lifting surfaces. In this method, a semi-empirical term representing viscous-separation crossflow is added to a term representing potential-theory crossflow. For many bodies of revolution, computed aerodynamic characteristics are shown to agree with measured results for investigated free-stream Mach numbers from 0.6 to 2.9. The angles of attack extend from 0 deg to 180 deg for M = 2.9 from 0 deg to 60 deg for M = 0.6 to 2.0. For several bodies of elliptic cross section, measured results are also predicted reasonably well over the investigated Mach number range from 0.6 to 2.0 and at angles of attack from 0 deg to 60 deg. As for the bodies of revolution, the predictions are best for supersonic Mach numbers. For body-wing and body-wing-tail configurations with wings of aspect ratios 3 and 4, measured normal-force coefficients and centers are predicted reasonably well at the upper test Mach number of 2.0. Vapor-screen and oil-flow pictures are shown for many body, body-wing and body-wing-tail configurations. When spearation and vortex patterns are asymmetric, undesirable side forces are measured for the models even at zero sideslip angle. Generally, the side-force coefficients decrease or vanish with the following: increase in Mach number, decrease in nose fineness ratio, change from sharp to blunt nose, and flattening of body cross section (particularly the body nose).

  2. Prediction of static aerodynamic characteristics for slender bodies alone and with lifting surfaces to very high angles of attack

    NASA Technical Reports Server (NTRS)

    Jorgensen, L. H.

    1976-01-01

    An engineering-type method is presented for computing normal-force and pitching-moment coefficients for slender bodies of circular and noncircular cross section alone and with lifting surfaces. In this method, a semi-empirical term representing viscous-separation crossflow is added to a term representing potential-theory crossflow. For many bodies of revolution, computed aerodynamic characteristics are shown to agree with measured results for investigated free-stream Mach numbers from 0.6 to 2.9. For several bodies of elliptic cross section, measured results are also predicted reasonably well over the investigated Mach number range from 0.6 to 2.0 and at angles of attack from 0 to 60 deg. As for the bodies of revolution, the predictions are best for supersonic Mach numbers. For body-wing and body-wing-tail configurations with wings of aspect ratios 3 and 4, measured normal-force coefficients and centers are predicted reasonably well at the upper test Mach number of 2.0. However, with a decrease in Mach number to 0.6, the agreement for C sub N rapidly deteriorates, although the normal-force centers remain in close agreement. Vapor-screen and oil-flow pictures are shown for many body, body-wing, and body-wing-tail configurations. When separation and vortex patterns are asymmetric, undesirable side forces are measured for the models even at zero sideslip angle. Generally, the side-force coefficients decrease or vanish with the following: increase in Mach number, decrease in nose fineness ratio, change from sharp to blunt nose, and flattening of body cross section (particularly the body nose).

  3. Aerodynamic characteristics of the standard dynamics model in coning motion at Mach 0.6

    NASA Technical Reports Server (NTRS)

    Jermey, C.; Schiff, L. B.

    1985-01-01

    A wind tunnel test was conducted on the Standard Dynamics Model (a simplified generic fighter aircraft shape) undergoing coning motion at Mach 0.6. Six component force and moment data are presented for a range of angle of attack, sideslip, and coning rates. At the relatively low non-dimensional coning rate employed (omega b/2V less than or equal to 0.04), the lateral aerodynamic characteristics generally show a linear variation with coning rate.

  4. Application of NASTRAN to large deflection supersonic flutter of panels

    NASA Technical Reports Server (NTRS)

    Mei, C.; Rogers, J. L., Jr.

    1976-01-01

    Flat panel flutter at high supersonic Mach number is analyzed using NASTRAN Level 16.0 by means of modifications to the code. Two-dimensional plate theory and quasi-steady aerodynamic theory are employed. The finite element formulation and solution procedure are presented. Modifications to the NASTRAN code are discussed. Convergence characteristics of the iteration processes are also briefly discussed. Effects of aerodynamic damping, boundary support condition and applied in-plane loading are included. Comparison of nonlinear vibration and linear flutter results with analytical solutions demonstrate that excellent accuracy is obtained with NASTRAN.

  5. Transonic Aerodynamic Characteristics of a Wing-Body Combination having a 52.5 deg Sweptback Wing of Aspect Ratio 3 with Conical Camber and Designed for a Mach Number of the Square Root of 2

    NASA Technical Reports Server (NTRS)

    Igoe, William B.; Re, Richard J.; Cassetti, Marlowe

    1961-01-01

    An investigation has been made of the effects of conical wing camber and supersonic body indentation on the aerodynamic characteristics of a wing-body configuration at transonic speeds. Wing aspect ratio was 3.0, taper ratio was 0.1, and quarter-chord line sweepback was 52.5 deg with airfoil sections of 0.03 thickness ratio. The tests were conducted in the Langley 16-foot transonic tunnel at various Mach numbers from 0.80 to 1.05 at angles of attack from -4 deg to 14 deg. The cambered-wing configuration achieved higher lift-drag ratios than a similar plane-wing configuration. The camber also reduced the effects of wing-tip flow separation on the aerodynamic characteristics. In general, no stability or trim changes below wing-tip flow separation resulted from the use of camber. The use of supersonic body indentation improved the lift-drag ratios at Mach numbers from 0.96 to 1.05.

  6. Theoretical evaluation of high-speed aerodynamics for arrow-wing configurations

    NASA Technical Reports Server (NTRS)

    Dollyhigh, S. M.

    1979-01-01

    The use of the theoretical methods to calculate the high-speed aerodynamic characteristics of arrow-wing supersonic cruise configurations was studied. Included are correlations of theoretical predictions with wind-tunnel data at Mach numbers from 0.8 to 2.7, examples of the use of theoretical methods to extrapolate the wind-tunnel data to full-scale flight condition, and presentation of a typical supersonic data package for an advanced supersonic transport application. A brief description of the methods and their application is given.

  7. The Effects of Streamwise-Deflected Wing Tips on the Aerodynamic Characteristics of an Aspect Ratio-2 Triangular Wing, Body, and Tail Combination

    NASA Technical Reports Server (NTRS)

    Peterson, Victor L.

    1959-01-01

    An investigation has been conducted on a triangular wing and body combination to determine the effects on the aerodynamic characteristics resulting from deflecting portions of the wing near the tips 900 to the wing surface about streamwise hinge lines. Experimental data were obtained for Mach numbers of 0.70, 1.30, 1.70, and 2.22 and for angles of attack ranging from -5 deg to +18 deg at sideslip angles of 0 deg and 5 deg. The results showed that the aerodynamic center shift experienced by the triangular wing and body combination as the Mach number was increased from subsonic to supersonic could be reduced by about 40 percent by deflecting the outboard 4 percent of the total area of each wing panel. Deflection about the same hinge line of additional inboard surfaces consisting of 2 percent of the total area of each wing panel resulted in a further reduction of the aerodynamic center travel of 10 percent. The resulting reductions in the stability were accompanied by increases in the drag due to lift and, for the case of the configuration with all surfaces deflected, in the minimum drag. The combined effects of reduced stability and increased drag of the untrimmed configuration on the trimmed lift-drag ratios were estimated from an analysis of the cases in which the wing-body combination with or without tips deflected was assumed to be controlled by a canard. The configurations with deflected surfaces had higher trimmed lift-drag ratios than the model with undeflected surfaces at Mach numbers up to about 1.70. Deflecting either the outboard surfaces or all of the surfaces caused the directional stability to be increased by increments that were approximately constant with increasing angle of attack at each Mach number. The effective dihedral was decreased at all angles of attack and Mach numbers when the surfaces were deflected.

  8. The method of characteristics for the determination of supersonic flow over bodies of revolution at small angles of attack

    NASA Technical Reports Server (NTRS)

    Ferri, Antonio

    1951-01-01

    The method of characteristics has been applied for the determination of the supersonic-flow properties around bodies of revolution at a small angle of attack. The system developed considers the effect of the variation of entropy due to the curved shock and determines a flow that exactly satisfies the boundary conditions in the limits of the simplifications assumed. Two practical methods for numerical calculations are given. (author)

  9. Mixing characteristics of a moderate aspect ratio screeching supersonic rectangular jet

    NASA Astrophysics Data System (ADS)

    Valentich, Griffin; Upadhyay, Puja; Kumar, Rajan

    2016-05-01

    Flow field characteristics of a moderate aspect ratio supersonic rectangular jet were examined at two overexpanded, a perfectly expanded, and an underexpanded jet conditions. The underexpanded and one overexpanded operating condition were of maximum screech, while the second overexpanded condition was of minimum screech intensity. Streamwise particle image velocimetry was performed along both major and minor axes of the jet and the measurements were made up to 30 nozzle heights, h, where h is the small dimension of the nozzle. Select cross planes were examined using stereoscopic particle image velocimetry to investigate the jet development and the role streamwise vortices play in jet spreading at each operating condition. The results show that streamwise vortices present at the nozzle corners along with vortices excited by screech tones play a major role in the jet evolution. All cases except for the perfectly expanded operating condition exhibited axis switching at streamwise locations ranging from 11 to 16 nozzle heights downstream of the exit. The overexpanded condition of maximum screech showed the most upstream switch over, while the underexpanded case showed the farthest downstream. Both of the maximum screeching cases developed into a diamond cross-sectional profile far downstream of the exit, while the ideally expanded case maintained a rectangular shape. The overexpanded minimum screeching case eventually decayed into an oblong profile.

  10. Flight-measured lift and drag characteristics of a large, flexible, high supersonic cruise airplane

    NASA Technical Reports Server (NTRS)

    Arnaiz, H. H.

    1977-01-01

    Flight measurements of lift, drag, and angle of attack were obtained for the XB-70 airplane, a large, flexible, high supersonic cruise airplane. This airplane had a length of over 57 meters, a takeoff gross mass of over 226,800 kilograms, and a design cruise speed of Mach 3 at an altitude of 21,340 meters. The performance measurements were made at Mach numbers from 0.72 to 3.07 and altitudes from approximately 7620 meters to 21,340 meters. The measurements were made to provide data for evaluating the techniques presently being used to design and predict the performance of aircraft in this category. Such performance characteristics as drag polars, lift-curve slopes, and maximum lift-to-drag ratios were derived from the flight data. The base drag of the airplane, changes in airplane drag with changes in engine power setting at transonic speeds, and the magnitude of the drag components of the propulsion system are also discussed.

  11. Pressure recovery, drag, and subcritical stability characteristics of conical supersonic diffusers with boundary-layer removal

    NASA Technical Reports Server (NTRS)

    Obey, Leonard T; Englert, Gerald W; Nussdorfer, Theodore J , Jr

    1952-01-01

    A study of two 20 degrees half-angle, low mass-flow ratio conical supersonic inlets with cone boundary-layer bleed was made on a 16-inch ram-jet engine in the Lewis 8- by 6-foot supersonic wind tunnel. A greater stable subcritical range of operation was obtained with the bleed inlets than with the corresponding inlet without boundary-layer bleed. The drag added by the bleed system was small.

  12. Evaluation of VSAERO in prediction of aerodynamic characteristics of helicopter hub fairings

    NASA Technical Reports Server (NTRS)

    Louie, Alexander

    1989-01-01

    A low-order panel code, VSAERO, was used to predict the aerodynamic characteristics of helicopter hub fairings. Since the simulation of this kind of bluff body by VSAERO was not documented before, the VSAERO solutions were correlated with experimental data to establish their validity. The validation process revealed that simulation of the aerodynamic environment around a hub fairing was sensitive to several modeling parameters. Some of these parameters are body and wake panels arrangement, streamwise and spanwise separation location, and the most prominent one-the wake modeling. Three wake models were used: regular wake, separated wake, and jet model. The regular wake is a wake with negligible thickness (thin wake). It is represented by a single vortex sheet. The separated wake and the jet model in the present application are wakes with finite thickness (thick wake). They consist of a vortex sheet enclosing a region of low-energy flow. The results obtained with the reqular wake were marginally acceptable for sharp-edged hub fairings. For all other cases under consideration, the jet model results correlated slightly better. The separated wake, which seemed to be the most appropriate model, caused the solution to diverge. While the regular wake was straight-forward to apply in simulations, the jet model was not. It requires the user to provide information about the doublet strength gradient on wake panels by guessing the efflux velocities at the wake shedding location. In summary, VSAERO neither predicts accurately the aerodynamic characteristics of helicopter hub fairings nor was cost effective.

  13. Numerical Study of Aerodynamic Characteristics of a Symmetric NACA Section with Simulated Ice Shapes

    NASA Astrophysics Data System (ADS)

    Tabatabaei, N.; Cervantes, M. J.; Trivedi, C.; Aidanpää, Jan-Olof

    2016-09-01

    To develop a numerical model of icing on wind turbine blades, a CFD simulation was conducted to investigate the effect of critical ice accretions on the aerodynamic characteristics of a 0.610 m chord NACA 0011 airfoil section. Aerodynamic performance coefficients and pressure profile were calculated and compared with the available measurements for a chord Reynolds number of 1.83x106. Ice shapes were simulated with flat plates (spoiler-ice) extending along the span of the wing. Lift, drag, and pressure coefficients were calculated in zero angle of attack through the steady state and transient simulations. Different approaches of numerical studies have been applied to investigate the icing conditions on the blades. The simulated separated flow over the sharp spoilers is challenging and can be seen as a worst test case for validation. It allows determining a reliable strategy to simulate real ice shapes [1] for which the detailed validation cannot easily be provided.

  14. TAD- THEORETICAL AERODYNAMICS PROGRAM

    NASA Technical Reports Server (NTRS)

    Barrowman, J.

    1994-01-01

    This theoretical aerodynamics program, TAD, was developed to predict the aerodynamic characteristics of vehicles with sounding rocket configurations. These slender, axisymmetric finned vehicle configurations have a wide range of aeronautical applications from rockets to high speed armament. Over a given range of Mach numbers, TAD will compute the normal force coefficient derivative, the center-of-pressure, the roll forcing moment coefficient derivative, the roll damping moment coefficient derivative, and the pitch damping moment coefficient derivative of a sounding rocket configured vehicle. The vehicle may consist of a sharp pointed nose of cone or tangent ogive shape, up to nine other body divisions of conical shoulder, conical boattail, or circular cylinder shape, and fins of trapezoid planform shape with constant cross section and either three or four fins per fin set. The characteristics computed by TAD have been shown to be accurate to within ten percent of experimental data in the supersonic region. The TAD program calculates the characteristics of separate portions of the vehicle, calculates the interference between separate portions of the vehicle, and then combines the results to form a total vehicle solution. Also, TAD can be used to calculate the characteristics of the body or fins separately as an aid in the design process. Input to the TAD program consists of simple descriptions of the body and fin geometries and the Mach range of interest. Output includes the aerodynamic characteristics of the total vehicle, or user-selected portions, at specified points over the mach range. The TAD program is written in FORTRAN IV for batch execution and has been implemented on an IBM 360 computer with a central memory requirement of approximately 123K of 8 bit bytes. The TAD program was originally developed in 1967 and last updated in 1972.

  15. Mixing characteristics of a transverse jet injection into supersonic crossflows through an expansion wall

    NASA Astrophysics Data System (ADS)

    Liu, Chaoyang; Wang, Zhenguo; Wang, Hongbo; Sun, Mingbo

    2016-12-01

    Mixing characteristics of a transverse jet injection into supersonic crossflows through an expansion plate are investigated using large eddy simulation (LES), where the expansion effects on the mixing are analyzed emphatically by comparing to the flat-plate counterpart. An adaptive central-upwind weighted essentially non-oscillatory (WENO) scheme along with multi-threaded and multi-process MPI/OpenMP parallel is adopted to improve the accuracy and efficiency of the calculations. Progressive mesh refinement study is performed to assess the grid resolution and solution convergence. Statistic results obtained are compared to the experimental data and recently performed classical numerical simulation, which validates the reliability of the present LES codes. Firstly, the jet mixing mechanisms in the flowfield with expansion plate are revealed. It indicates that the large-scale vortices in the windward side of jet plume induced by Kelvin-Helmholtz (K-H) instability contribute to the mixing in the near-field, while the entrainment by the counter-rotating vortices and molecular diffusion dominate the mixing process in the far-field. Furthermore, the effects of wall expansion on the flow and mixing characteristics are discussed. The boundary layer across the expansion corner is relaminarized and the profiles of streamwise velocity are distinctly changed. Then the separation region ahead of jet plume is more close to the wall, and the breaking process of large-scale vortices in the windward side of jet plume starts earlier. However, the favorable pressure gradient generated by wall expansion reduces the mixing efficiency and brings a greater total pressure loss.

  16. A program to compute three-dimensional subsonic unsteady aerodynamic characteristics using the doublet lattice method, L216 (DUBFLEX). Volume 2: Supplemental system design and maintenance document

    NASA Technical Reports Server (NTRS)

    Harrison, B. A.; Richard, M.

    1979-01-01

    The information necessary for execution of the digital computer program L216 on the CDC 6600 is described. L216 characteristics are based on the doublet lattice method. Arbitrary aerodynamic configurations may be represented with combinations of nonplanar lifting surfaces composed of finite constant pressure panel elements, and axially summetric slender bodies composed of constant pressure line elements. Program input consists of configuration geometry, aerodynamic parameters, and modal data; output includes element geometry, pressure difference distributions, integrated aerodynamic coefficients, stability derivatives, generalized aerodynamic forces, and aerodynamic influence coefficient matrices. Optionally, modal data may be input on magnetic field (tape or disk), and certain geometric and aerodynamic output may be saved for subsequent use.

  17. Nonlinear aerodynamic wing design

    NASA Technical Reports Server (NTRS)

    Bonner, Ellwood

    1985-01-01

    The applicability of new nonlinear theoretical techniques is demonstrated for supersonic wing design. The new technology was utilized to define outboard panels for an existing advanced tactical fighter model. Mach 1.6 maneuver point design and multi-operating point compromise surfaces were developed and tested. High aerodynamic efficiency was achieved at the design conditions. A corollary result was that only modest supersonic penalties were incurred to meet multiple aerodynamic requirements. The nonlinear potential analysis of a practical configuration arrangement correlated well with experimental data.

  18. Measured and Computed Hypersonic Aerodynamic/Aeroheating Characteristics for an Elliptically Blunted Flared Cylinder

    NASA Technical Reports Server (NTRS)

    Greene, Francis A.; Buck, Gregory M.; Wood, William A.

    2001-01-01

    Computational and experimental hypersonic aerodynamic forces and moments and aeroheating levels for Kistler Aerospace Corporation's baseline orbiter vehicle at incidence are presented. Experimental data were measured in ground-based facilities at the Langley Research Center and predictions were performed using the Langley Aerothermodynamic Upwind Relaxation Algorithm code. The test parameters were incidence (-4 to 24 degrees), freestream Mach number (6 to 10), freestream ratio of specific heats (1.2 to 1.4), and freestream Reynolds number (0.5 to 8.0 million per foot). The effects of these parameters on aerodynamic characteristics, as well as the effects of Reynolds number on measured heating levels are discussed. Good agreement between computational and experimental aerodynamic and aeroheating values were observed over the wide range of test parameters examined. Reynolds number and ratio of specific heats were observed to significantly alter the trim L/D value. At Mach 6, laminar flow, was observed along the entire windward centerline up to the flare for all angles and Reynolds numbers tested. Flow over the flare transitioned front laminar to transitional turbulent between 4 and 8 million per foot at 8 and 12 degrees angle of attack, and near 4 million per foot at 16 degrees angle of attack.

  19. Measured and Computed Hypersonic Aerodynamic/Aeroheating Characteristics for an Elliptically Blunted Flared Cylinder

    NASA Technical Reports Server (NTRS)

    Greene, Francis A.; Buck, Gregory M.; Wood, William A.

    2001-01-01

    Computational and experimental hypersonic aerodynamic forces and moments and aeroheating levels for Kistler Aerospace Corporation's baseline orbiter vehicle at incidence are presented. Experimental data were measured in ground-based facilities at the Langley Research Center and predictions were performed using the Langley Aerothermodynamic Upwind Relaxation Algorithm code. The test parameters were incidence (-4 to 24 degrees), freestream Mach number (6 to 10),freestream ratio o specific heats (1.2 to 1.4), and freestream Reynolds number (0.5 to 8.0 million per foot). The effects of these parameters on aerodynamic characteristics, as well as the effects of Reynolds number on measured heating levels are discussed. Good agreement between computational and experimental aerodynamic and aeroheating values were observed over the wide range of test parameters examined. Reynolds number and ratio of specific heats were observed to significantly alter the trim L/D value. At Mach 6, laminar flow was observed along the entire windward centerline tip to the flare for all angles and Reynolds numbers tested. Flow over the flare transitioned from laminar to transitional/turbulent between 4 and 8 million per foot at 8 and 12 degrees angle of attack, and near 4 million per foot at 16 degrees angle of attack.

  20. Supersonic dynamic stability characteristics of the test technique demonstrator NASP configuration

    NASA Technical Reports Server (NTRS)

    Dress, David A.; Boyden, Richmond P.; Cruz, Christopher I.

    1992-01-01

    Wind tunnel tests of a National Aero-Space Plane (NASP) configuration were conducted in both test sections of the Langley Unitary Plan Wind Tunnel. The model used is a Langley designed blended body NASP configuration. Dynamic stability characteristics were measured on this configuration at Mach numbers of 2.0, 2.5, 3.5, and 4.5. In addition to tests of the baseline configuration, component buildup tests were conducted. The test results show that the baseline configuration generally has positive damping about all three axes with only isolated exceptions. In addition, there was generally good agreement between the in-pulse dynamic parameters and the corresponding static data which were measured during another series of tests in the Unitary Plan Wind Tunnel. Also included are comparisons of the experimental damping parameters with results from the engineering predictive code APAS (Aerodynamic Preliminary Analysis System). These comparisons show good agreement at low angles of attack; however, the comparisons are generally not as good at the higher angles of attack.

  1. Theoretical aerodynamic characteristics of a family of slender wing-tail-body combinations

    NASA Technical Reports Server (NTRS)

    Lomax, Harvard; Byrd, Paul F

    1951-01-01

    The aerodynamic characteristics of an airplane configuration composed of a swept-back, nearly constant chord wing and a triangular tail mounted on a cylindrical body are presented. The analysis is based on the assumption that the free-stream Mach number is near unity or that the configuration is slender. The calculations for the tail are made on the assumption that the vortex system trailing back from the wing is either a sheet lying entirely in the plane of the flat tail surface or has completely "rolled up" into two point vortices that lie either in, above, or below the plane of the tail surface.

  2. Prediction of Aerodynamic Characteristics of Fighter Wings at High Angles of Attack.

    DTIC Science & Technology

    1984-03-01

    method coupled with iterative routines for wake location, viscous effects and vortex flows. Applications of the techniques to a number of...AD-A145 1@7 PREDICTION OF AERODYNAMIC CHARACTERISTICS OF FIGHTER i/2 WIINGS AT HIGH ANGLES OF ATTACK(U) ANALYTICAL METHODS INC REDMOND WA B MASKEW ET...ATTACK I B. !4askew T.S. Vaidyanathan J.K. Nathman F.A. Dvorak Analytical Methods , Inc. 2047 - 152nd Avenue N.E. Redmond, Washington 98052 CONTRACT

  3. The characteristics of the ground vortex and its effect on the aerodynamics of the STOL configuration

    NASA Technical Reports Server (NTRS)

    Stewart, Vearle R.

    1988-01-01

    The interaction of the free stream velocity on the wall jet formed by the impingement of deflected engine thrust results in a rolled up vortex which exerts sizable forces on a short takeoff (STOL) airplane configuration. Some data suggest that the boundary layer under the free stream ahead of the configuration may be important in determining the extent of the travel of the wall jet into the oncoming stream. Here, early studies of the ground vortex are examined, and those results are compared to some later data obtained with moving a model over a fixed ground board. The effect of the ground vortex on the aerodynamic characteristics are discussed.

  4. Cylinder wake influence on the tonal noise and aerodynamic characteristics of a NACA0018 airfoil

    NASA Astrophysics Data System (ADS)

    Takagi, Y.; Fujisawa, N.; Nakano, T.; Nashimoto, A.

    2006-11-01

    The influence of cylinder wake on discrete tonal noise and aerodynamic characteristics of a NACA0018 airfoil is studied experimentally in a uniform flow at a moderate Reynolds number. The experiments are carried out by measuring sound pressure levels and spectrum, separation and the reattachment points, pressure distribution, fluid forces, mean-flow and turbulence characteristics around the airfoil with and without the cylinder wake. Present results indicate that the tonal noise from the airfoil is suppressed by the influence of the cylinder wake and the aerodynamic characteristics are improved in comparison with the case without the cylinder wake. These are mainly due to the separation control of boundary layers over the airfoil caused by the wake-induced transition, which is observed by surface flow visualization with liquid- crystal coating. The PIV measurements of the flow field around the airfoil confirm that highly turbulent velocity fluctuation of the cylinder wake induces the transition of the boundary layers and produces an attached boundary layer over the airfoil. Then, the vortex shedding phenomenon near the trailing edge of pressure surface is removed by the influence of the wake and results in the suppression of tonal noise.

  5. Supersonic axial-flow fan flutter

    NASA Technical Reports Server (NTRS)

    Ramsey, John K.

    1988-01-01

    Lane's (1957) analytical formulation of the unsteady pressure distribution on an oscillating two-dimensional flat plate cascade in supersonic axial flow has been developed into a computer code. This unsteady aerodynamic code has shown good agreement with other published data. This code has also been incorporated into an existing aeroelastic code to analyze the NASA Lewis supersonic through-flow fan design.

  6. Progress in supersonic cruise technology

    NASA Technical Reports Server (NTRS)

    Driver, C.

    1983-01-01

    The Supersonic Cruise Research (SCR) program identified significant improvements in the technology areas of aerodynamics, structures, propulsion, noise reduction, takeoff and landing procedures, and advanced configuration concepts. These improvements, when combined in a large supersonic cruise vehicle, offer a far greater technology advance than generally realized. They offer the promise of an advanced commercial family of aircraft which are environmentally acceptable, have flexible range-payload capability, and are economically viable. These same areas of technology have direct application to smaller advanced military aircraft and to supersonic executive aircraft. Several possible applications will be addressed.

  7. Supersonic Chordwise Bending Flutter in Cascades

    DTIC Science & Technology

    1975-05-31

    such a flutter boundary can be made by utilizing the trend lines predicted from a supersonic analysis based on supersonic cascade theory (Appendix I...bonding agent was injected via hypodermic needles after the blade tabs were properly inserted, The integrity and repeatability of the mounting of the indi...in conjunction with NASTRAN predictions and supersonic cascade aerodynamic computa- tions. Comparisons between theory and experiment are discussed. DD

  8. Large-scale aerodynamic characteristics of airfoils as tested in the variable density wind tunnel

    NASA Technical Reports Server (NTRS)

    Jacobs, Eastman N; Anderson, Raymond F

    1931-01-01

    In order to give the large-scale characteristics of a variety of airfoils in a form which will be of maximum value, both for airplane design and for the study of airfoil characteristics, a collection has been made of the results of airfoil tests made at full-scale values of the reynolds number in the variable density wind tunnel of the National Advisory Committee for Aeronautics. They have been corrected for tunnel wall interference and are presented not only in the conventional form but also in a form which facilitates the comparison of airfoils and from which corrections may be easily made to any aspect ratio. An example showing the method of correcting the results to a desired aspect ratio has been given for the convenience of designers. In addition, the data have been analyzed with a view to finding the variation of the aerodynamic characteristics of airfoils with their thickness and camber.

  9. Aerodynamic characteristics of the HL-20 and HL-20A lifting-body configurations

    NASA Technical Reports Server (NTRS)

    Ware, George M.; Spencer, Bernard, Jr.; Micol, John R.

    1991-01-01

    The data show that the HL-20 is longitudinally and laterally stable over the test range from Mach 10 to 0.2. At hypersonic speeds it has a trimmed lift/drag ratio of 1.4. This values gives the vehicle a cross range capability similar to that of the Space Shuttle. At subsonic speeds, the HL-20 has a trimmed lift/drag ratio of about 3.6. Replacing the flat plate outboard fins with fins having an airfoil shape, increased the maximum trimmed L/D to 4.3. Preliminary evaluation of configuration modifications (the HL-20A series), indicates that trim at higher values of lift at hypersonic speeds could be achieved with an L/D of about 1.0. In the supersonic range, the lift and directional stability characteristics were improved. The untrimmed subsonic L/D was increased to 5.8 with airfoil fins.

  10. Aerodynamic preliminary analysis system 2. Part 1: Theory

    NASA Technical Reports Server (NTRS)

    Bonner, E.; Clever, W.; Dunn, K.

    1991-01-01

    An aerodynamic analysis system based on potential theory at subsonic and/or supersonic speeds and impact type finite element solutions at hypersonic conditions is described. Three dimensional configurations having multiple nonplanar surfaces of arbitrary planform and bodies of noncircular contour may be analyzed. Static, rotary, and control longitudinal and lateral directional characteristics may be generated. The analysis was implemented on a time sharing system in conjunction with an input tablet digitizer and an interactive graphics input/output display and editing terminal to maximize its responsiveness to the preliminary analysis problem. The program provides an efficient analysis for systematically performing various aerodynamic configuration tradeoff and evaluation studies.

  11. Low-Disturbance Flow Characteristics of the NASA-Ames Laminar Flow Supersonic Wind Tunnel

    NASA Technical Reports Server (NTRS)

    Wolf, Stephen W. D.; Laub, James A.; Davis, Sanford S. (Technical Monitor)

    1994-01-01

    A unique, low-disturbance (quiet) supersonic wind tunnel has been commissioned at the NASA-Ames Fluid Mechanics Laboratory (FML) to support Supersonic Laminar Flow Control (SLFC) research. Known as the Laminar Flow Supersonic Wind Tunnel (LFSWT), this tunnel is designed to operate at potential cruise Mach numbers and unit Reynolds numbers (Re) of the High Speed Civil Transport (HSCT). The need to better understand the receptivity of the transition phenomena on swept (HSCT) wings to attachment-line contamination and cross-flows has provided the impetus for building the LFSWT. Low-disturbance or "quiet" wind tunnels are known to be an essential part of any meaningful boundary layer transition research. In particular, the receptivity of supersonic boundary layers to wind tunnel disturbances can significantly alter the transition phenomena under investigation on a test model. Consequently, considerable effort has gone into the design of the LFSWT to provide quiet flow. The paper describes efforts to quantify the low-disturbance flows in the LFSWT operating at Mach 1.6, as a precursor to transition research on wing models. The research includes: (1) Flow measurements in both the test section and settling chamber of the LFSWT, using a full range of measurement techniques; (2) Study of the state of the test section boundary layer so far by using a single hot-wire mounted above the floor centerline, with and without boundary layer trips fitted at the test section entrance; (3) The effect of flow quality of unsteady supersonic diffuser flow, joint steps and gaps, and wall vibration.

  12. Aerodynamic characteristics and pressure distributions for an executive-jet baseline airfoil section

    NASA Technical Reports Server (NTRS)

    Allison, Dennis O.; Mineck, Raymond E.

    1993-01-01

    A wind tunnel test of an executive-jet baseline airfoil model was conducted in the adaptive-wall test section of the NASA Langley 0.3-Meter Transonic Cryogenic Tunnel. The primary goal of the test was to measure airfoil aerodynamic characteristics over a wide range of flow conditions that encompass two design points. The two design Mach numbers were 0.654 and 0.735 with corresponding Reynolds numbers of 4.5 x 10(exp 6) and 8.9 x 10(exp 6) based on chord, respectively, and normal-force coefficients of 0.98 and 0.51, respectively. The tests were conducted over a Mach number range from 0.250 to 0.780 and a chord Reynolds number range from 3 x 10(exp 6) to 18 x 10(exp 6). The angle of attack was varied from -2 deg to a maximum below 10 deg with one exception in which the maximum was 14 deg for a Mach number of 0.250 at a chord Reynolds number of 4.5 x 10(exp 6). Boundary-layer transition was fixed at 5 percent of chord on both the upper and lower surfaces of the model for most of the test. The adaptive-wall test section had flexible top and bottom walls and rigid sidewalls. Wall interference was minimized by the movement of the adaptive walls, and the airfoil aerodynamic characteristics were corrected for any residual top and bottom wall interference.

  13. Calculation of static longitudinal aerodynamic characteristics of STOL aircraft with upper surface blown flaps

    NASA Technical Reports Server (NTRS)

    Mendenhall, M. R.; Perkin, S. C., Jr.; Goodwin, F. K.; Spangler, S. B.

    1975-01-01

    An existing prediction method developed for EBF aircraft configurations was applied to USB configurations to determine its potential utility in predicting USB aerodynamic characteristics. An existing wing-flap vortex-lattice computer program was modified to handle multiple spanwise flap segments at different flap angles. A potential flow turbofan wake model developed for circular cross-section jets was used to model a rectangular cross-section jet wake by placing a number of circular jets side by side. The calculation procedure was evaluated by comparison of measured and predicted aerodynamic characteristics on a variety of USB configurations. The method is limited to the case where the flow and geometry of the configuration are symmetric about a vertical plane containing the wing root chord. Comparison of predicted and measured lift and pitching moment coefficients were made on swept wings with one and two engines per wing panel, various flap deflection angles, and a range of thrust coefficients. The results indicate satisfactory prediction of lift for flap deflections up to 55 and thrust coefficients less than 2. The applicability of the prediction procedure to USB configurations is evaluated, and specific recommendations for improvements are discussed.

  14. Semi-Empirical Prediction of Aircraft Low-Speed Aerodynamic Characteristics

    NASA Technical Reports Server (NTRS)

    Olson, Erik D.

    2015-01-01

    This paper lays out a comprehensive methodology for computing a low-speed, high-lift polar, without requiring additional details about the aircraft design beyond what is typically available at the conceptual design stage. Introducing low-order, physics-based aerodynamic analyses allows the methodology to be more applicable to unconventional aircraft concepts than traditional, fully-empirical methods. The methodology uses empirical relationships for flap lift effectiveness, chord extension, drag-coefficient increment and maximum lift coefficient of various types of flap systems as a function of flap deflection, and combines these increments with the characteristics of the unflapped airfoils. Once the aerodynamic characteristics of the flapped sections are known, a vortex-lattice analysis calculates the three-dimensional lift, drag and moment coefficients of the whole aircraft configuration. This paper details the results of two validation cases: a supercritical airfoil model with several types of flaps; and a 12-foot, full-span aircraft model with slats and double-slotted flaps.

  15. A numerical investigation into the aerodynamic characteristics and aeroelastic stability of a footbridge

    NASA Astrophysics Data System (ADS)

    Taylor, I. J.; Vezza, M.

    2009-01-01

    The results of a numerical investigation into the aerodynamic characteristics and aeroelastic stability of a proposed footbridge across a highway in the north of England are presented. The longer than usual span, along with the unusual nature of the pedestrian barriers, indicated that the deck configuration was likely to be beyond the reliable limits of the British design code BD 49/01. The calculations were performed using the discrete vortex method, DIVEX, developed at the Universities of Glasgow and Strathclyde. DIVEX has been successfully validated on a wide range of problems, including the aeroelastic response of bridge deck sections. In particular, the investigation focussed on the effects of non-standard pedestrian barriers on the structural integrity of the bridge. The proposed deck configuration incorporated a barrier comprised of angled flat plates, and the bridge was found to be unstable at low wind speeds, with the plates having a strong turning effect on the flow at the leading edge of the deck. These effects are highlighted in both a static and dynamic analysis of the bridge deck, along with modifications to the design that aim to improve the aeroelastic stability of the deck. Proper orthogonal decomposition (POD) was also used to investigate the unsteady pressure field on the upper surface of the static bridge deck. The results of the flutter investigation and the POD analysis highlight the strong influence of the pedestrian barriers on the overall aerodynamic characteristics and aeroelastic stability of the bridge.

  16. Hypersonic aerodynamic characteristics of an all-body research aircraft configuration

    NASA Technical Reports Server (NTRS)

    Clark, L. E.

    1973-01-01

    An experimental investigation was conducted at Mach 6 to determine the hypersonic aerodynamic characteristics of an all-body, delta-planform, hypersonic research aircraft (HYFAC configuration). The aerodynamic characteristics were obtained at Reynolds numbers based on model length of 2.84 million and 10.5 million and over an angle-of-attack range from minus 4 deg to 20 deg. The experimental results show that the HYFAC configuration is longitudinally stable and can be trimmed over the range of test conditions. The configuration had a small degree of directional stability over the angle-of-attack range and positive effective dihedral at angles of attack greater than 2 deg. Addition of canards caused a decrease in longitudinal stability and an increase in directional stability. Oil-flow studies revealed extensive areas of separated and vortex flow on the fuselage lee surface. A limited comparison of wind-tunnel data with several hypersonic approximations indicated that, except for the directional stability, the tangent-cone method gave adequate agreement at control settings between 5 deg and minus 5 deg and positive lift coefficient. A limited comparison indicated that the HYFAC configuration had greater longitudinal stability than an elliptical-cross-section configuration, but a lower maximum lift-drag ratio.

  17. Numerical and Experimental Study on Aerodynamic Characteristics of Basic Airfoils at Low Reynolds Numbers

    NASA Astrophysics Data System (ADS)

    Hirata, Katsuya; Kawakita, Masatoshi; Iijima, Takayoshi; Koga, Mitsuhiro; Kihira, Mitsuhiko; Funaki, Jiro

    The aerodynamic characteristics of airfoils have been researched in higher Reynolds-number ranges more than 106, in a historic context closely related with the developments of airplanes and fluid machineries in the last century. However, our knowledge is not enough at low and middle Reynolds-number ranges. So, in the present study, we investigate such basic airfoils as a NACA0015, a flat plate and the flat plates with modified fore-face and after-face geometries at Reynolds number Re < 1.0×105, using two- and three-dimensional computations together with wind-tunnel and water-tank experiments. As a result, we have revealed the effect of the Reynolds number Re upon the minimum drag coefficient CDmin. Besides, we have shown the effects of attack angle α upon various aerodynamic characteristics such as the lift coefficient CL, the drag coefficient CD and the lift-to-drag ratio CL/CD at Re = 1.0×102, discussing those effects on the basis of both near-flow-field information and surface-pressure profiles. Such results suggest the importance of sharp leading edges, which implies the possibility of an inversed NACA0015. Furthermore, concerning the flat-plate airfoil, we investigate the influences of fore-face and after-face geometries upon such effects.

  18. Effect of symmetrical vortex shedding on the longitudinal aerodynamic characteristics of wing-body-tail combinations

    NASA Technical Reports Server (NTRS)

    Mendenhall, M. R.; Nielsen, J. N.

    1975-01-01

    An engineering prediction method for determining the longitudinal aerodynamic characteristics of wing-body-tail combinations is developed. The method includes the effects of nonlinear aerodynamics of components and the interference between components. Nonlinearities associated with symmetrical vortex shedding from the nose of the body are considered as well as the nonlinearities associated with the separation vortices from the leading edges and side edges of the lifting surfaces. The wing and tail characteristics are calculated using lifting surface theories which include effects of incidence, camber, twist, and induced velocities from external sources of disturbance such as bodies and vortices. The lifting surface theories calculate the distribution of leading edge and side edge suction which is converted to vortex lift using the Polhamus suction analogy. Correlation curves are developed to determine the fraction of the theoretical suction force which is converted into vortex lift. The prediction method is compared with experimental data on a variety of aircraft configurations to assess the accuracy and limitations of the method.

  19. Transonic Aerodynamic Loading Characteristics of a Wing-Body-Tail Combination Having a 52.5 deg. Sweptback Wing of Aspect Ratio 3 With Conical Wing Camber and Body Indentation for a Design Mach Number of Square Root of 2

    NASA Technical Reports Server (NTRS)

    Cassetti, Marlowe D.; Re, Richard J.; Igoe, William B.

    1961-01-01

    An investigation has been made of the effects of conical wing camber and body indentation according to the supersonic area rule on the aerodynamic wing loading characteristics of a wing-body-tail configuration at transonic speeds. The wing aspect ratio was 3, taper ratio was 0.1, and quarter-chord-line sweepback was 52.5 deg. with 3-percent-thick airfoil sections. The tests were conducted in the Langley 16-foot transonic tunnel at Mach numbers from 0.80 to 1.05 and at angles of attack from 0 deg. to 14 deg., with Reynolds numbers based on mean aerodynamic chord varying from 7 x 10(exp 6) to 8 x 10(exp 6). Conical camber delayed wing-tip stall and reduced the severity of the accompanying longitudinal instability but did not appreciably affect the spanwise load distribution at angles of attack below tip stall. Body indentation reduced the transonic chordwise center-of-pressure travel from about 8 percent to 5 percent of the mean aerodynamic chord.

  20. Effect of wing design on the longitudinal aerodynamic characteristics of a wing-body model at subsonic speeds

    NASA Technical Reports Server (NTRS)

    Henderson, W. P.; Huffman, J. K.

    1972-01-01

    An investigation has been conducted to determine the effects of wing camber and twist on the longitudinal aerodynamic characteristics of a wingbody configuration. Three wings were used each having the same planform (aspect ratio of 2.5 and leading-edge sweep angle of 44 deg.) but differing in amounts of camber and twist (wing design lift coefficient). The wing design lift coefficients were 0, 0.35, and 0.70. The investigation was conducted over a Mach number range from 0.20 to 0.70 at angles of attack up to about 22 deg. The effect of wing strakes on the aerodynamic characteristics of the cambered wings was also studied. A comparison of the experimentally determined aerodynamic characteristics with theoretical estimates is also included.

  1. Effect of the Surface Condition of a Wing on the Aerodynamic Characteristics of an Airplane

    NASA Technical Reports Server (NTRS)

    Defrance, S J

    1934-01-01

    In order to determine the effect of the surface conditions of a wing on the aerodynamic characteristics of an airplane, tests were conducted in the N.A.C.A. full-scale wind tunnel on the Fairchild F-22 airplane first with normal commercial finish of wing surface and later with the same wing polished. Comparison of the characteristics of the airplane with the two surface conditions shows that the polish caused a negligible change in the lift curve, but reduced the minimum drag coefficient by 0.001. This reduction in drag if applied to an airplane with a given speed of 200 miles per hour and a minimum drag coefficient of 0.025 would increase the speed only 2.9 miles per hour, but if the speed remained the same, the power would be reduced 4 percent.

  2. FLPP IXV Re-Entry Vehicle, Supersonic Charectisation Based on DNW SST Wind Tunnel Tests and CFD

    NASA Astrophysics Data System (ADS)

    Kapteijn, C.; Maseland, H.; Chiarelli, C.; Mareschi, V.; Tribot, J.-P.; Binetti, P.; Walloscheck, T.

    2009-01-01

    The European Space Agency ESA, has engaged in 2004, the IXV project (Intermediate eXperimental Vehicle) which is part of the FLPP (Future Launcher Preparatory Programme) aiming at answering to critical technological issues for controlled re-entry, while supporting the future generation launchers and to improve in general European capabilities in the strategic field of atmospheric re-entry for future space transportation, exploration and scientific applications. The IXV key mission and system objectives are the design, development, manufacturing, assembling and on- ground to in-flight verification of an autonomous European lifting and aerodynamically controlled re- entry system, integrating the critical re- entry technologies at the system level. In particular, the IXV shall demonstrate system integrated key technologies such as lifting flight control by means of aerodynamic surfaces that are one of the main primary objectives of the experimental investigation. Lifting and aerodynamic controlled re-entry represents a significant capability advancement with respect to the ballistic re-entry of capsules like the ARD. Since hypersonic aerodynamics is essentially different from supersonic aerodynamics, the current mission is to perform an atmospheric re-entry in combination with a safe recovery the in supersonic flight regime. However, mission extension to trimmed transonic flight is under consideration based on a preliminary analysis of the aerodynamic characteristics of the IXV configuration. Since the beginning of the IXV project, an aerodynamic data base (AEDB) has been built up and continuously updated integrating the additional information mainly provided by means of CFD (ie: Euler and Navier-Stokes) and lately also by means of WTTs. This AEDB serves for flying qualities analysis and for re-entry simulations. During the development phase B2/C1, the effectiveness of the control surfaces and their impact on te vehicle's aerodynamic forces in the supersonic regime is

  3. Method and apparatus for starting supersonic compressors

    DOEpatents

    Lawlor, Shawn P [Bellevue, WA

    2012-04-10

    A supersonic gas compressor. The compressor includes aerodynamic duct(s) situated on a rotor journaled in a casing. The aerodynamic duct(s) generate a plurality of oblique shock waves for efficiently compressing a gas at supersonic conditions. The convergent inlet is adjacent to a bleed air collector, and during acceleration of the rotor, bypass gas is removed from the convergent inlet via a collector to enable supersonic shock stabilization. Once the oblique shocks are stabilized at a selected inlet relative Mach number and pressure ratio, the bleed of bypass gas from the convergent inlet via the bypass gas collectors is eliminated.

  4. Aerodynamic Characteristics of a Circular Cylinder at Mach Number of 6.86 and Angles of Attack up to 90 Degrees

    NASA Technical Reports Server (NTRS)

    Penland, Jim A

    1954-01-01

    Pressure-distribution and force tests of a circular cylinder have been made in the Langley 11-inch hypersonic tunnel at a Mach number of 6.86, a Reynolds number of 129,000 based on diameter, and angles of attack up to 90 degrees. The results are compared with the hypersonic approximation of Grimminger, Williams, and Young and with a simple modification of the Newtonian flow theory. The comparison of experimental results shows that either theory gives adequate general aerodynamic characteristics but that the modified Newtonian theory gives a more accurate prediction of the pressure distribution. The calculated crossflow drag coefficients plotted as a function of crossflow Mach number were found to be in reasonable agreement with similar results obtained from other investigations at lower supersonic Mach numbers. Comparison of the results of this investigation with data obtained at a lower Mach number indicates that the drag coefficient of a cylinder normal to the flow is relatively constant for Mach numbers above about 4.

  5. Aerodynamic Characteristics of Three Deep-Stepped Planing-Tail Flying-Boat Hulls

    NASA Technical Reports Server (NTRS)

    Riebe, John M.; Naeseth, Rodger L.

    1947-01-01

    An investigation was made in the Langley 300 MPH 7- by 10-foot tunnel to determine the aerodynamic characteristics of three deep-stepped planing-tail flying-boat hulls differing only in the amount of step fairing. The hulls were derived by increasing the unfaired step depth of a planing-tail hull of a previous aerodynamic investigation to a depth about 92 percent of the hull beam. Tests were also made on a transverse-stepped hull with an extended afterbody for the purpose of comparison and in order to extend and verify the results of a previous investigation. The investigation indicated that the extended afterbody hull had a minimum drag coefficient about the same as a conventional hull, 0.0066, and an angle-of-attack range for minimum drag coefficient of 0.0057 which was 14 percent less than the transverse stepped hull with extended afterbody; the hulls with step fairing had up to 44 percent less minimum drag coefficient than the transverse-stepped hull, or slightly more drag than a streamlined body having approximately the same length and volume. Longitudinal and lateral instability varied little with step fairing and was about the same as a conventional hull.

  6. Steady-state and transitional aerodynamic characteristics of a wing in simulated heavy rain

    NASA Technical Reports Server (NTRS)

    Campbell, Bryan A.; Bezos, Gaudy M.

    1989-01-01

    The steady-state and transient effects of simulated heavy rain on the subsonic aerodynamic characteristics of a wing model were determined in the Langley 14- by 22-Foot Subsonic Tunnel. The 1.29 foot chord wing was comprised of a NACA 23015 airfoil and had an aspect ratio of 6.10. Data were obtained while test variables of liquid water content, angle of attack, and trailing edge flap angle were parametrically varied at dynamic pressures of 10, 30, and 50 psf (i.e., Reynolds numbers of .76x10(6), 1.31x10(6), and 1.69x10(6)). The experimental results showed reductions in lift and increases in drag when in the simulated rain environment. Accompanying this was a reduction of the stall angle of attack by approximately 4 deg. The transient aerodynamic performance during transition from dry to wet steady-state conditions varied between a linear and a nonlinear transition.

  7. Wind Tunnel Aerodynamic Characteristics of a Transport-type Airfoil in a Simulated Heavy Rain Environment

    NASA Technical Reports Server (NTRS)

    Bezos, Gaudy M.; Dunham, R. Earl, Jr.; Gentry, Garl L., Jr.; Melson, W. Edward, Jr.

    1992-01-01

    The effects of simulated heavy rain on the aerodynamic characteristics of an NACA 64-210 airfoil section equipped with leading-and trailing-edge high-lift devices were investigated in the Langley 14- by 22-Foot Subsonic Tunnel. The model had a chord of 2.5 ft, a span of 8 ft, and was mounted on the tunnel centerline between two large endplates. Aerodynamic measurements in and out of the simulated rain environment were obtained for dynamic pressures of 30 and 50 psf and an angle-of-attack range of 0 to 20 degrees for the cruise configuration. The rain intensity was varied to produce liquid water contents ranging from 16 to 46 gm/cu m. The results obtained for various rain intensity levels and tunnel speeds showed significant losses in maximum lift capability and increases in drag for a given lift as the liquid water content was increased. The results obtained on the landing configuration also indicate a progressive decrease in the angle of attack at which maximum lift occurred and an increase in the slope of the pitching-moment curve as the liquid water content was increased. The sensitivity of test results to the effects of the water surface tension was also investigated. A chemical was introduced into the rain environment that reduced the surface tension of water by a factor of 2. The reduction in the surface tension of water did not significantly alter the level of performance losses for the landing configuration.

  8. Experimental and theoretical aerodynamic characteristics of a high-lift semispan wing model

    NASA Technical Reports Server (NTRS)

    Applin, Zachary T.; Gentry, Garl L., Jr.

    1990-01-01

    Experimental and theoretical aerodynamic characteristics were compared for a high-lift, semispan wing configuration that incorporated a slightly modified version of the NASA Advanced Laminar Flow Control airfoil section. The experimental investigation was conducted in the Langley 14- by 22-Foot Subsonic Tunnel at chord Reynolds numbers of 2.36 and 3.33 million. A two-dimensional airfoil code and a three-dimensional panel code were used to obtain aerodynamic predictions. Two-dimensional data were corrected for three-dimensional effects. Comparisons between predicted and measured values were made for the cruise configuration and for various high-lift configurations. Both codes predicted lift and pitching moment coefficients that agreed well with experiment for the cruise configuration. These parameters were overpredicted for all high-lift configurations. Drag coefficient was underpredicted for all cases. Corrected two-dimensional pressure distributions typically agreed well with experiment, while the panel code overpredicted the leading-edge suction peak on the wing. One important feature missing from both of these codes was a capability for separated flow analysis. The major cause of disparity between the measured data and predictions presented herein was attributed to separated flow conditions.

  9. Flow characteristic of in-flight particles in supersonic plasma spraying process

    NASA Astrophysics Data System (ADS)

    Wei, Pei; Wei, Zhengying; Zhao, Guangxi; Du, Jun; Bai, Y.

    2016-09-01

    In this paper, a computational model based on supersonic plasma spraying (SAPS) is developed to describe the plasma jet coupled with the injection of carrier gas and particles for SAPS. Based on a high-efficiency supersonic spraying gun, the 3D computational model of spraying gun was built to study the features of plasma jet and its interactions with the sprayed particles. Further the velocity and temperature of in-flight particles were measured by Spray Watch 2i, the shape of in-flight particles was observed by scanning electron microscope. Numerical results were compared with the experimental measurements and a good agreement has been achieved. The flight process of particles in plasma jet consists of three stages: accelerated stage, constant speed stage and decelerated stage. Numerical and experimental indicates that the H2 volume fraction in mixture gas of Ar + H2 should keep in the range of 23-26 %, and the distance of 100 mm is the optimal spraying distance in Supersonic atmosphere plasma spraying. Particles were melted and broken into small child particles by plasma jet and the diameters of most child particles were less than 30 μm. In general, increasing the particles impacting velocity and surface temperature can decrease the coating porosity.

  10. Prediction of vortex shedding from circular and noncircular bodies in supersonic flow

    NASA Technical Reports Server (NTRS)

    Mendenhall, M. R.; Perkins, S. C., Jr.

    1984-01-01

    An engineering prediction method and associated computer code NOZVTX to predict nose vortex shedding from circular and noncircular bodies in supersonic flow at angles of attack and roll are presented. The body is represented by either a supersonic panel method for noncircular cross sections or line sources and doublets for circular cross sections, and the lee side vortex wake is modeled by discrete vortices in crossflow planes. The three-dimensional steady flow problem is reduced to a two-dimensional, unsteady, separated flow problem for solution. Comparison of measured and predicted surface pressure distributions, flow field surveys, and aerodynamic characteristics is presented for bodies with circular and noncircular cross-sectional shapes.

  11. Advanced structures technology applied to a supersonic cruise arrow-wing configuration

    NASA Technical Reports Server (NTRS)

    Sakata, I. F.; Davis, G. W.

    1976-01-01

    The application of advanced technology to a promising aerodynamic configuration was explored to investigate the improved payload range characteristics over the configuration postulated during the National SST Program. The results of an analytical study performed to determine the best structural approach for design of a Mach number 2.7 arrow-wing supersonic cruise aircraft are highlighted. The data conducted under the auspices of the Structures Directorate of the National Aeronautics and Space Administration, Langley Research Center, established firm technical bases from which further trend studies were conducted to quantitatively assess the benefits and feasibility of using advanced structures technology to arrive at a viable advanced supersonic cruise aircraft.

  12. The effect of winglets on the static aerodynamic stability characteristics of a representative second generation jet transport model

    NASA Technical Reports Server (NTRS)

    Jacobs, P. F.; Flechner, S. G.

    1976-01-01

    A baseline wing and a version of the same wing fitted with winglets were tested. The longitudinal aerodynamic characteristics were determined through an angle-of-attack range from -1 deg to 10 deg at an angle of sideslip of 0 deg for Mach numbers of 0.750, 0.800, and 0.825. The lateral aerodynamic characteristics were determined through the same angle-of-attack range at fixed sideslip angles of 2.5 deg and 5 deg. Both configurations were investigated at Reynolds numbers of 13,000,000, per meter (4,000,000 per foot) and approximately 20,000,000 per meter (6,000,000 per foot). The winglet configuration showed slight increases over the baseline wing in static longitudinal and lateral aerodynamic stability throughout the test Mach number range for a model design lift coefficient of 0.53. Reynolds number variation had very little effect on stability.

  13. Aerodynamic Characteristics at Mach Numbers of 1.41 and 2.01 of a Series of Cranked Wings Ranging in Aspect Ratio from 4.00 to 1.74 in Combination with a Body

    NASA Technical Reports Server (NTRS)

    Sevier, John R., Jr.

    1960-01-01

    A program has been conducted in the Langley 4- by 4-foot supersonic pressure tunnel to determine the effects of certain wing plan-form variations on the aerodynamic characteristics of wing-body combinations at supersonic speeds. The present report deals with the results of tests of a family of cranked wing plan forms in combination with an ogive-cylinder body of revolution. Tests were made at Mach numbers of 1.41 and 2.01 at corresponding values of Reynolds number per foot of 3.0 x 10(exp 6) and 2.5 x 10(exp 6). Results of the tests indicate that the best overall characteristics were obtained with the low-aspect-ratio wings. Plan-form changes which involved decreasing the aspect ratio resulted in higher values of maximum lift-drag ratio, in addition to large increases in wing volume. Indications are that this trend would have continued to exist at aspect ratios even lower than the lowest considered in the present tests. Increases in the maximum lift-drag ratio of about 15 percent over the basic wing were achieved with practically no increase in drag. The severe longitudinal stability associated with the basic cranked wing was no longer present (within the limits of the present tests) on the wings of lower aspect ratio formed by sweeping forward the inboard portion of the trailing edge.

  14. Multifidelity Analysis and Optimization for Supersonic Design

    NASA Technical Reports Server (NTRS)

    Kroo, Ilan; Willcox, Karen; March, Andrew; Haas, Alex; Rajnarayan, Dev; Kays, Cory

    2010-01-01

    Supersonic aircraft design is a computationally expensive optimization problem and multifidelity approaches over a significant opportunity to reduce design time and computational cost. This report presents tools developed to improve supersonic aircraft design capabilities including: aerodynamic tools for supersonic aircraft configurations; a systematic way to manage model uncertainty; and multifidelity model management concepts that incorporate uncertainty. The aerodynamic analysis tools developed are appropriate for use in a multifidelity optimization framework, and include four analysis routines to estimate the lift and drag of a supersonic airfoil, a multifidelity supersonic drag code that estimates the drag of aircraft configurations with three different methods: an area rule method, a panel method, and an Euler solver. In addition, five multifidelity optimization methods are developed, which include local and global methods as well as gradient-based and gradient-free techniques.

  15. Wake structure and aerodynamic characteristics of an auto-propelled pitching airfoil

    NASA Astrophysics Data System (ADS)

    Hanchi, S.; Benkherouf, T.; Mekadem, M.; Oualli, H.; Keirsbulck, L.; Labraga, L.

    2013-05-01

    In the present study, we investigate the wake configuration as well as the flow aerodynamic and propulsive characteristics of a system equipped with a nature-inspired propulsion system. The study focuses on the effect of a set of pitching frequency and amplitude values on the flow behavior for a symmetric foil performing pitching sinusoidal rolling oscillations. The viscous, non-stationary flow around the pitching foil is simulated using ANSYS FLUENT 13. The foil movement is reproduced using the dynamic mesh technique and an in-house developed UDF (User Define Function). Our results show the influence of the pitching frequency and the amplitude on the wake. We provide the mechanisms relating the system behavior to the applied forces. The frequency varies from 1 to 400Hz and the considered amplitudes are 18%, 24%, 30%, 37%, 53%, 82% and 114% of the foil chord.

  16. Low-speed longitudinal and lateral-directional aerodynamic characteristics of the X-31 configuration

    NASA Technical Reports Server (NTRS)

    Banks, Daniel W.; Gatlin, Gregory M.; Paulson, John W., Jr.

    1992-01-01

    An experimental investigation of a 19 pct. scale model of the X-31 configuration was completed in the Langley 14 x 22 Foot Subsonic Tunnel. This study was performed to determine the static low speed aerodynamic characteristics of the basic configuration over a large range of angle of attack and sideslip and to study the effects of strakes, leading-edge extensions (wing-body strakes), nose booms, speed-brake deployment, and inlet configurations. The ultimate purpose was to optimize the configuration for high angle of attack and maneuvering-flight conditions. The model was tested at angles of attack from -5 to 67 deg and at sideslip angles from -16 to 16 deg for speeds up to 190 knots (dynamic pressure of 120 psf).

  17. Aerodynamic stability and control characteristics of TBC shuttle booster AR-11981-3

    NASA Technical Reports Server (NTRS)

    Phelps, E. R.; Watts, L. L.; Ainsworth, R. W.

    1972-01-01

    A scale model of the Boeing Company space shuttle booster configuration 3 was tested in the MSFC 14-inch trisonic wind tunnel. This test was proposed to fill-in the original test run schedule as well as to investigate the aerodynamic stability and control characteristics of the booster with three wing configurations not previously tested. The configurations tested included: (1) a cylindrical booster body with an axisymmetric nose, (2) clipped delta canards that had variable incidence from 0 deg to -60 deg, (3) different aft body mounted wing configurations, (4) two vertical fin configurations, and (5) a Grumman G-3 orbiter configuration. Tests were conducted over a Mach range from 0.6 to 5.0.

  18. An experimental investigation of the aerodynamic characteristics of slanted base ogive cylinders using magnetic suspension technology

    NASA Technical Reports Server (NTRS)

    Alcorn, Charles W.; Britcher, Colin

    1988-01-01

    An experimental investigation is reported on slanted base ogive cylinders at zero incidence. The Mach number range is 0.05 to 0.3. All flow disturbances associated with wind tunnel supports are eliminated in this investigation by magnetically suspending the wind tunnel models. The sudden and drastic changes in the lift, pitching moment, and drag for a slight change in base slant angle are reported. Flow visualization with liquid crystals and oil is used to observe base flow patterns, which are responsible for the sudden changes in aerodynamic characteristics. Hysteretic effects in base flow pattern changes are present in this investigation and are reported. The effect of a wire support attachment on the 0 deg slanted base model is studied. Computational drag and transition location results using VSAERO and SANDRAG are presented and compared with experimental results. Base pressure measurements over the slanted bases are made with an onboard pressure transducer using remote data telemetry.

  19. Aerodynamic characteristics of the ventilated design for flapping wing micro air vehicle.

    PubMed

    Zhang, G Q; Yu, S C M

    2014-01-01

    Inspired by superior flight performance of natural flight masters like birds and insects and based on the ventilating flaps that can be opened and closed by the changing air pressure around the wing, a new flapping wing type has been proposed. It is known that the net lift force generated by a solid wing in a flapping cycle is nearly zero. However, for the case of the ventilated wing, results for the net lift force are positive which is due to the effect created by the "ventilation" in reducing negative lift force during the upstroke. The presence of moving flaps can serve as the variable in which, through careful control of the areas, a correlation with the decrease in negative lift can be generated. The corresponding aerodynamic characteristics have been investigated numerically by using different flapping frequencies and forward flight speeds.

  20. Aerodynamic Characteristics of the Ventilated Design for Flapping Wing Micro Air Vehicle

    PubMed Central

    Zhang, G. Q.; Yu, S. C. M.

    2014-01-01

    Inspired by superior flight performance of natural flight masters like birds and insects and based on the ventilating flaps that can be opened and closed by the changing air pressure around the wing, a new flapping wing type has been proposed. It is known that the net lift force generated by a solid wing in a flapping cycle is nearly zero. However, for the case of the ventilated wing, results for the net lift force are positive which is due to the effect created by the “ventilation” in reducing negative lift force during the upstroke. The presence of moving flaps can serve as the variable in which, through careful control of the areas, a correlation with the decrease in negative lift can be generated. The corresponding aerodynamic characteristics have been investigated numerically by using different flapping frequencies and forward flight speeds. PMID:24683339

  1. Calculation of the longitudinal aerodynamic characteristics of upper-surface-blown wing-flap configurations

    NASA Technical Reports Server (NTRS)

    Mendenhall, M. R.; Spangler, S. B.

    1979-01-01

    An investigation has been carried out to develop an engineering method for predicting the longitudinal aerodynamic characteristics of wing-flap configurations with upper surface blown (USB) high lift devices. Potential flow models of the lifting surfaces and the jet wakes are combined to calculate the induced interference of the engine wakes on the wing and flaps. The wing may have an arbitrary planform with camber and twist and multiple trailing edge flaps. The jet wake model has a rectangular cross section over its entire length and it is positioned such that the wake is tangent to the upper surfaces of the wing and flaps. Comparisons of measured and predicted pressure distributions, spanload distributions, and total lift and pitching-moment coefficients on swept and unswept USB configurations are presented for a wide range of thrust coefficients and flap deflection angles.

  2. Prediction of longitudinal aerodynamic characteristics of STOL configurations with externally blown high lift devices

    NASA Technical Reports Server (NTRS)

    Mendenhall, M. R.; Spangler, S. B.

    1976-01-01

    A theoretical method has been developed to predict the longitudinal aerodynamic characteristics of engine-wing-flap combinations with externally blown flaps (EBF) and upper surface blowing (USB) high lift devices. Potential flow models of the lifting surfaces and the jet wake are combined to calculate the induced interference of the engine wakes on the lifting surfaces. The engine wakes may be circular, elliptic, or rectangular cross-sectional jets, and the lifting surfaces are comprised of a wing with multiple-slotted trailing-edge flaps or a deflected trailing-edge Coanda surface. Results are presented showing comparisons of measured and predicted forces, pitching moments, span-load distributions, and flow fields.

  3. Subsonic longitudinal and lateral aerodynamic characteristics for a systematic series of strake-wing configurations

    NASA Technical Reports Server (NTRS)

    Luckring, J. M.

    1979-01-01

    A systematic wind tunnel study was conducted in the Langley 7 by 10 foot high speed tunnel to help establish a parametric data base of the longitudinal and lateral aerodynamic characteristics for configurations incorporating strake-wing geometries indicative of current and proposed maneuvering aircraft. The configurations employed combinations of strakes with reflexed planforms having exposed spans of 10%, 20%, and 30% of the reference wing span and wings with trapezoidal planforms having leading edge sweep angles of approximately 30, 40, 44, 50, and 60 deg. Tests were conducted at Mach numbers ranging from 0.3 to 0.8 and at angles of attack from approximately -4 to 48 deg at zero sideslip.

  4. Low speed aerodynamic characteristics of a 17 percent thick airfoil section designed for general aviation applications

    NASA Technical Reports Server (NTRS)

    Mcghee, R. J.; Beasley, W. D.

    1973-01-01

    Wind-tunnel tests have been conducted to determine the low-speed two-dimensional aerodynamic characteristics of a 17-percent-thick airfoil designed for general aviation applications (GA(W)-1). The results were compared with predictions based on a theoretical method for calculating the viscous flow about the airfoil. The tests were conducted over a Mach number range from 0.10 to 0.28. Reynolds numbers based on airfoil chord varied from 2.0 million to 20.0 million. Maximum section lift coefficients greater than 2.0 were obtained and section lift-drag ratio at a lift coefficient of 1.0 (climb condition) varied from about 65 to 85 as the Reynolds number increased from about 2.0 million to 6.0 million.

  5. Calculation of the longitudinal aerodynamic characteristics of wing-flap configurations with externally blown flaps

    NASA Technical Reports Server (NTRS)

    Mendenhall, M. R.

    1976-01-01

    An analytical method for predicting the longitudinal aerodynamic characteristics of externally blown flap configurations is described. Two potential flow models make up the prediction method: a wing and flap lifting-surface model and a turbofan engine wake model. A vortex-lattice lifting-surface method is used to represent the wing and multiple-slotted trailing-edge flaps. The jet wake is represented by a series of closely spaced vortex rings normal to a centerline which is free to move to conform to the local flow field. The two potential models are combined in an iterative fashion to predict the jet wake interference effects on a typical EBF configuration. Comparisons of measured and predicted span-load distributions, individual surface forces, forces and moments on the complete configuration, and flow fields are included.

  6. Investigations on the Aerodynamic Characteristics and Blade Excitations of the Radial Turbine with Pulsating Inlet Flow

    NASA Astrophysics Data System (ADS)

    Liu, Yixiong; Yang, Ce; Yang, Dengfeng; Zhang, Rui

    2016-04-01

    The aerodynamic performance, detailed unsteady flow and time-based excitations acting on blade surfaces of a radial flow turbine have been investigated with pulsation flow condition. The results show that the turbine instantaneous performance under pulsation flow condition deviates from the quasi-steady value significantly and forms obvious hysteretic loops around the quasi-steady conditions. The detailed analysis of unsteady flow shows that the characteristic of pulsation flow field in radial turbine is highly influenced by the pulsation inlet condition. The blade torque, power and loading fluctuate with the inlet pulsation wave in a pulse period. For the blade excitations, the maximum and the minimum blade excitations conform to the wave crest and wave trough of the inlet pulsation, respectively, in time-based scale. And toward blade chord direction, the maximum loading distributes along the blade leading edge until 20% chord position and decreases from the leading to trailing edge.

  7. Aerodynamic characteristics of a powered tilt-proprotor wind tunnel model

    NASA Technical Reports Server (NTRS)

    Wilson, J. C.; Mineck, R. E.; Freeman, C. E.

    1976-01-01

    An investigation was conducted in the Langley V/STOL tunnel to determine the performance, stability and control, and rotor-wake interaction effects of a powered tilt-proprotor aircraft model with gimbal-hub rotors. Tests were conducted at representative flight conditions for hover, helicopter, transition, and airplane flight. Force and moment data were obtained for the complete model and for each of the two rotors. In addition to wind-speed variation, the angle of attack, angle of sideslip, rotor speed, rotor collective pitch, longitudinal cyclic pitch, rotor pylon angle, and configuration geometry were varied. The results, presented in graphical form, are available in tabular form to facilitate the validation of analytical methods of defining the aerodynamic characteristics of tilt-proprotor configurations.

  8. Supersonic laser propulsion.

    PubMed

    Rezunkov, Yurii; Schmidt, Alexander

    2014-11-01

    To produce supersonic laser propulsion, a new technique based on the interaction of a laser-ablated jet with supersonic gas flow in a nozzle is proposed. It is shown that such parameters of the jet, such as gas-plasma pressure and temperature in the ablation region as well as the mass consumption rate of the ablated solid propellant, are characteristic in this respect. The results of numerical simulations of the supersonic laser propulsion are presented for two types of nozzle configuration. The feasibility to achieve the momentum coupling coefficient of C(m)∼10(-3) N/W is shown.

  9. A simulator investigation of the influence of engine response characteristics on the approach and landing for an externally blown flap aircraft. Part 2: Aerodynamic model

    NASA Technical Reports Server (NTRS)

    Ciffone, D. L.; Robinson, G. H.

    1973-01-01

    An analysis of the influence of engine response characteristics on the approach and landing of an externally blown flap aircraft was conducted using flight simulator facilities. The configuration of the aerodynamic model is described. The aerodynamic characteristics as a function of angle of attack, thrust coefficient, and flap deflection are presented in tabular form and as graphs.

  10. A Numerical Comparison of Symmetric and Asymmetric Supersonic Wind Tunnels

    NASA Astrophysics Data System (ADS)

    Clark, Kylen D.

    Supersonic wind tunnels are a vital aspect to the aerospace industry. Both the design and testing processes of different aerospace components often include and depend upon utilization of supersonic test facilities. Engine inlets, wing shapes, and body aerodynamics, to name a few, are aspects of aircraft that are frequently subjected to supersonic conditions in use, and thus often require supersonic wind tunnel testing. There is a need for reliable and repeatable supersonic test facilities in order to help create these vital components. The option of building and using asymmetric supersonic converging-diverging nozzles may be appealing due in part to lower construction costs. There is a need, however, to investigate the differences, if any, in the flow characteristics and performance of asymmetric type supersonic wind tunnels in comparison to symmetric due to the fact that asymmetric configurations of CD nozzle are not as common. A computational fluid dynamics (CFD) study has been conducted on an existing University of Michigan (UM) asymmetric supersonic wind tunnel geometry in order to study the effects of asymmetry on supersonic wind tunnel performance. Simulations were made on both the existing asymmetrical tunnel geometry and two axisymmetric reflections (of differing aspect ratio) of that original tunnel geometry. The Reynolds Averaged Navier Stokes equations are solved via NASAs OVERFLOW code to model flow through these configurations. In this way, information has been gleaned on the effects of asymmetry on supersonic wind tunnel performance. Shock boundary layer interactions are paid particular attention since the test section integrity is greatly dependent upon these interactions. Boundary layer and overall flow characteristics are studied. The RANS study presented in this document shows that the UM asymmetric wind tunnel/nozzle configuration is not as well suited to producing uniform test section flow as that of a symmetric configuration, specifically one

  11. Study of aerodynamic technology for VSTOL fighter/attack aircraft: Vertical attitude concept

    NASA Technical Reports Server (NTRS)

    Gerhardt, H. A.; Chen, W. S.

    1978-01-01

    The aerodynamic technology for a vertical attitude VSTOL (VATOL) supersonic fighter/attack aircraft was studied. The selected configuration features a tailless clipped delta wing with leading-edge extension (LEX), maneuvering flaps, top-side inlet, twin dry engines and vectoring nozzles. A relaxed static stability is employed in conjunction with the maneuvering flaps to optimize transonic performance and minimize supersonic trim drag. Control for subaerodynamic flight is obtained by gimballing the nozzles in combination with wing tip jets. Emphasis is placed on the development of aerodynamic characteristics and the identification of aerodynamic uncertainties. A wind tunnel test program is proposed to resolve these uncertainties and ascertain the feasibility of the conceptual design. Ship interface, flight control integration, crew station concepts, advanced weapons, avionics, and materials are discussed.

  12. Study of aerodynamic technology for VSTOL fighter attack aircraft

    NASA Technical Reports Server (NTRS)

    Burhans, W., Jr.; Crafta, V. J., Jr.; Dannenhoffer, N.; Dellamura, F. A.; Krepski, R. E.

    1978-01-01

    Vertical short takeoff aircraft capability, supersonic dash capability, and transonic agility were investigated for the development of Fighter/attack aircraft to be accommodated on ships smaller than present aircraft carriers. Topics covered include: (1) description of viable V/STOL fighter/attack configuration (a high wing, close-coupled canard, twin-engine, control configured aircraft) which meets or exceeds specified levels of vehicle performance; (2) estimates of vehicle aerodynamic characteristics and the methodology utilized to generate them; (3) description of propulsion system characteristics and vehicle mass properties; (4) identification of areas of aerodynamic uncertainty; and (5) a test program to investigate the areas of aerodynamic uncertainty in the conventional flight mode.

  13. Characteristics of heat exchange in the region of injection into a supersonic high-temperature flow

    NASA Astrophysics Data System (ADS)

    Bakirov, F. G.; Shaykhutdinov, Z. G.

    1985-04-01

    An experimental investigation of the local heat transfer coefficient distribution during gas injection into the supersonic-flow portion of a Laval nozzle is discussed. The controlling dimensionless parameters of the investigated process are presented in terms of a generalized relation for the maximum value of the heat transfer coefficient in the nozzle cross section behind the injection hole. Data on the heat transfer coefficient variation along the nozzle length as a function of gas injection rate are also presented, along with the heat transfer coefficient distribution over a cross section of the nozzle.

  14. Supersonic dynamic stability characteristics of a space shuttle orbiter. [wind tunnel tests of scale models

    NASA Technical Reports Server (NTRS)

    Freeman, D. C., Jr.; Boyden, R. P.; Davenport, E. E.

    1976-01-01

    Supersonic forced-oscillation tests of a 0.0165-scale model of a modified 089B Rockwell International shuttle orbiter were conducted in a wind tunnel for several configurations over a Mach range from 1.6 to 4.63. The tests covered angles of attack up to 30 deg. The period and damping of the basic unaugmented vehicle were calculated along the entry trajectory using the measured damping results. Some parameter analysis was made with the measured dynamic derivatives. Photographs of the test configurations and test equipment are shown.

  15. Effect of Length-Beam Ratio on the Aerodynamic Characteristics of Flying-Boat Hulls without Wing Interference

    NASA Technical Reports Server (NTRS)

    Lowry, John G.; Riebe, John M.

    1948-01-01

    Contains experimental results of an investigation of the aerodynamic characteristics of a family of flying boat hulls of length beam ratios 6, 9, 12, and 15 without wing interference. The results are compared with those taken on the same family of hulls in the presence of a wing.

  16. Aerodynamic Characteristics of a Revised Target Drone Vehicle at Mach Numbers from 1.60 to 2.86

    NASA Technical Reports Server (NTRS)

    Blair, A. B., Jr.; Babb, C. Donald

    1968-01-01

    An investigation has been conducted in the Langley Unitary Plan wind tunnel to determine the aerodynamic characteristics of a revised target drone vehicle through a Mach number range from 1.60 to 2.86. The vehicle had canard surfaces and a swept clipped-delta wing with twin tip-mounted vertical tails.

  17. Aerodynamic characteristics of a wing with Fowler flaps including flap loads, downwash, and calculated effect on take-off

    NASA Technical Reports Server (NTRS)

    Platt, Robert C

    1936-01-01

    This report presents the results of wind tunnel tests of a wing in combination with each of three sizes of Fowler flap. The purpose of the investigation was to determine the aerodynamic characteristics as affected by flap chord and position, the air loads on the flaps, and the effect of flaps on the downwash.

  18. Flowfield characteristics of a transverse jet into supersonic flow with pseudo-shock wave

    NASA Astrophysics Data System (ADS)

    Yamauchi, H.; Choi, B.; Takae, K.; Kouchi, T.; Masuya, G.

    2012-11-01

    We performed an experimental investigation of the flowfield of a transverse jet into supersonic flow with a pseudo-shock wave (PSW). In this study, we injected compressed air as the injectant, simulating hydrocarbon fuel. A back pressure control valve generated PSW into Mach 2.5 supersonic flow and controlled its position. The positions of PSW were set at nondimensional distance from the injector by the duct height ( x/ H) of -1.0, -2.5, and -4.0. Particle image velocimetry (PIV) gave us the velocity of the flowfield. Mie scattering of oil mist only with the jet was used to measure the spread of the injectant. Furthermore, gas sampling measurements at the exit of the test section were carried out to determine the injectant mole fraction distributions. Gas sampling data qualitatively matched the intensity of Mie scattering. PIV measurements indicated that far-upstream PSW decelerated the flow speed of the main stream and developed the boundary layer on the wall of the test section. The flow speed deceleration at the corner of the test section was remarkable. The PSW produced nonuniformity in the main stream and reduced the momentum flux of the main stream in front of the injector. The blowing ratio, defined as the square root of the momentum flux ratio, of the jet and the main stream considering the effect of the boundary layer thickness was shown to be a useful parameter to explain the jet behavior.

  19. Aerodynamic Characteristics of a Canard and an Outboard-Tail Airplane Model at High Subsonic Speeds

    NASA Technical Reports Server (NTRS)

    Fournier, Paul G.

    1961-01-01

    An investigation has been made in the Langley high-speed 7- by 10-foot tunnel through a range of Mach numbers from 0.60 to 0.95 of the static longitudinal and lateral stability and control characteristics of a canard airplane configuration and an outboard-tail configuration. The canard model had a twisted wing with approximately 67 deg of sweepback and an aspect ratio of 2.91 and was tested with three trapezoidal canard surfaces having ratios of exposed area to wing area of 0.032, 0.076, and 0.121. The canard model had a single body-mounted vertical tail. The outboard-tail model had its horizontal- and vertical-tail surfaces mounted on slender bodies attached to the wing tips and located to the rear and outboard of the 67 deg sweptback wing of aspect ratio 1.00. The data, which are presented with limited analysis, provide information at high subsonic speeds on these two types of high-speed airplanes which have previously been tested at supersonic speeds and reported in NACA RM L58BO7 and NACA RM L58E20.

  20. Aerodynamic characteristics of a distinct wing-body configuration at Mach 6: Experiment, theory, and the hypersonic isolation principle

    NASA Technical Reports Server (NTRS)

    Penland, J. A.; Pittman, J. L.

    1985-01-01

    An experimental investigation has been conducted to determine the effect of wing leading edge sweep and wing translation on the aerodynamic characteristics of a wing body configuration at a free stream Mach number of about 6 and Reynolds number (based on body length) of 17.9 x 10 to the 6th power. Seven wings with leading edge sweep angles from -20 deg to 60 deg were tested on a common body over an angle of attack range from -12 deg to 10 deg. All wings had a common span, aspect ratio, taper ratio, planform area, and thickness ratio. Wings were translated longitudinally on the body to make tests possible with the total and exposed mean aerodynamic chords located at a fixed body station. Aerodynamic forces were found to be independent of wing sweep and translation, and pitching moments were constant when the exposed wing mean aerodynamic chord was located at a fixed body station. Thus, the Hypersonic Isolation Principle was verified. Theory applied with tangent wedge pressures on the wing and tangent cone pressures on the body provided excellent predictions of aerodynamic force coefficients but poor estimates of moment coefficients.

  1. Effect of posture on the aerodynamic characteristics during take-off in ski jumping.

    PubMed

    Yamamoto, Keizo; Tsubokura, Makoto; Ikeda, Jun; Onishi, Keiji; Baleriola, Sophie

    2016-11-07

    The purpose of this study was to investigate the effects of posture of a ski jumper on aerodynamic characteristics during the take-off using computational fluid dynamics (CFD). The CFD method adopted for this study was based on Large-Eddy Simulation. Body surface data were obtained by 3-D laser scanning of an active ski jumper. Based on video analysis of the actual take-off movement, two sets of motion data were generated (world-class jumper A and less-experienced jumper B). The inlet flow velocity that corresponds to the in-run velocity in actual ski jumping was set to 23.23m/s in the CFD. The aerodynamic force, flow velocity, and vortexes for each model were compared between models. The total drag force acting upon jumper A was lower than that acting upon jumper B through the whole movement. Regarding the total lift force, although jumper A׳s total lift force was less in the in-run posture, it became greater than that of jumper B at the end of the movement. In the latter half of the movement, low air-speed domain expansion was observed at the model׳s back. This domain of jumper B was larger. There were two symmetric vortexes in the wake of jumper A, but the disordered vortexes were observed behind the jumper B. In the case of jumper A, these two distinct vortexes generated by the arms produced a downwash flow in the wake. It is considered that the positioning of the arms in a very low position strongly influences the flow structure.

  2. Experimental Evaluation of the Effect of Angle-of-attack on the External Aerodynamics and Mass Capture of a Symmetric Three-engine Air-breathing Launch Vehicle Configuration at Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Kim, Hyun D.; Frate, Franco C.

    2001-01-01

    A subscale aerodynamic model of the GTX air-breathing launch vehicle was tested at NASA Glenn Research Center's 10- by 10-Foot Supersonic Wind Tunnel from Mach 2.0 to 3.5 at various angles-of-attack. The objective of the test was to investigate the effect of angle-of-attack on inlet mass capture, inlet diverter effectiveness, and the flowfield at the cowl lip plane. The flow-through inlets were tested with and without boundary-layer diverters. Quantitative measurements such as inlet mass flow rates and pitot-pressure distributions in the cowl lip plane are presented. At a 3deg angle-of-attack, the flow rates for the top and side inlets were within 8 percent of the zero angle-of-attack value, and little distortion was evident at the cowl lip plane. Surface oil flow patterns showing the shock/boundary-layer interaction caused by the inlet spikes are shown. In addition to inlet data, vehicle forebody static pressure distributions, boundary-layer profiles, and temperature-sensitive paint images to evaluate the boundary-layer transition are presented. Three-dimensional parabolized Navier-Stokes computational fluid dynamics calculations of the forebody flowfield are presented and show good agreement with the experimental static pressure distributions and boundary-layer profiles. With the boundary-layer diverters installed, no adverse aerodynamic phenomena were found that would prevent the inlets from operating at the required angles-of-attack. We recommend that phase 2 of the test program be initiated, where inlet contraction ratio and diverter geometry variations will be tested.

  3. Low Speed Aerodynamic Characteristics of Wings of Aspect Ratios 3 and 4 Equipped with High Lift Systems

    DTIC Science & Technology

    1980-05-01

    Trailing Edge in CCW Configuration without Tip Fence ................. ... 47 16 - Effect of a Nonround Coanda Trailing Edge on an Aspect Ratio 3 Wing in...fence installed. Figure 16 summarizes the effect of the noncircular Coanda surface on the lift characteristics. The aerodynamic characteristics of the...that of the round Coanda trailing edge depending on the value of a and C Figure 17 is a crossplot of all of the CCW data showing the effect of a wing tip

  4. Mathematical description of nonstationary aerodynamic characteristics of a passenger aircraft model in longitudinal motion at large angles of attack

    NASA Astrophysics Data System (ADS)

    Petoshin, V. I.; Chasovnikov, E. A.

    2011-05-01

    Aerodynamic loads in problems of flight dynamics of passenger aircraft in stalled flow regimes are described using a mathematical model that includes an ordinary linear first-order differential equation. A procedure for determining the parameters of the mathematical model is proposed which is based on approximating experimental frequency characteristics with the frequency characteristics of the linearized mathematical model. The mathematical model was verified by tests of a modern passenger aircraft model in a wind tunnel.

  5. Prediction of dynamic and aerodynamic characteristics of the centrifugal fan with forward curved blades

    NASA Astrophysics Data System (ADS)

    Polanský, Jiří; Kalmár, László; Gášpár, Roman

    2013-12-01

    The main aim of this paper is determine the centrifugal fan with forward curved blades aerodynamic characteristics based on numerical modeling. Three variants of geometry were investigated. The first, basic "A" variant contains 12 blades. The geometry of second "B" variant contains 12 blades and 12 semi-blades with optimal length [1]. The third, control variant "C" contains 24 blades without semi-blades. Numerical calculations were performed by CFD Ansys. Another aim of this paper is to compare results of the numerical simulation with results of approximate numerical procedure. Applied approximate numerical procedure [2] is designated to determine characteristics of the turbulent flow in the bladed space of a centrifugal-flow fan impeller. This numerical method is an extension of the hydro-dynamical cascade theory for incompressible and inviscid fluid flow. Paper also partially compares results from the numerical simulation and results from the experimental investigation. Acoustic phenomena observed during experiment, during numerical simulation manifested as deterioration of the calculation stability, residuals oscillation and thus also as a flow field oscillation. Pressure pulsations are evaluated by using frequency analysis for each variant and working condition.

  6. Operating characteristics of a hydrogen-argon plasma torch for supersonic combustion applications

    SciTech Connect

    Barbi, E.; Mahan, J.R.; O'brien, W.F.; Wagner, T.C.

    1989-04-01

    The residence time of the combustible mixture in the combustion chamber of a scramjet engine is much less than the time normally required for complete combustion. Hydrogen and hydrocarbon fuels require an ignition source under conditions typically found in a scramjet combustor. Analytical studies indicate that the presence of hydrogen atoms should greatly reduce the ignition delay in this environment. Because hydrogen plasmas are prolific sources of hydrogen atoms, a low-power, uncooled hydrogen plasma torch has been built and tested to evaluate its potential as a possible flame holder for supersonic combustion. The torch was found to be unstable when operated on pure hydrogen; however, stable operation could be obtained by using argon as a body gas and mixing in the desired amount of hydrogen. The stability limits of the torch are delineated and its electrical and thermal behavior documented. An average torch thermal efficiency of around 88 percent is demonstrated. 10 references.

  7. Operating characteristics of a hydrogen-argon plasma torch for supersonic combustion applications

    NASA Technical Reports Server (NTRS)

    Barbi, E.; Mahan, J. R.; O'Brien, W. F.; Wagner, T. C.

    1989-01-01

    The residence time of the combustible mixture in the combustion chamber of a scramjet engine is much less than the time normally required for complete combustion. Hydrogen and hydrocarbon fuels require an ignition source under conditions typically found in a scramjet combustor. Analytical studies indicate that the presence of hydrogen atoms should greatly reduce the ignition delay in this environment. Because hydrogen plasmas are prolific sources of hydrogen atoms, a low-power, uncooled hydrogen plasma torch has been built and tested to evaluate its potential as a possible flame holder for supersonic combustion. The torch was found to be unstable when operated on pure hydrogen; however, stable operation could be obtained by using argon as a body gas and mixing in the desired amount of hydrogen. The stability limits of the torch are delineated and its electrical and thermal behavior documented. An average torch thermal efficiency of around 88 percent is demonstrated.

  8. Low-speed longitudinal aerodynamic characteristics of a flat-plate planform model of an advanced fighter configuration

    NASA Technical Reports Server (NTRS)

    Mcgrath, Brian E.; Neuhart, Dan H.; Gatlin, Gregory M.; Oneil, Pat

    1994-01-01

    A flat-plate wind tunnel model of an advanced fighter configuration was tested in the NASA LaRC Subsonic Basic Research Tunnel and the 16- by 24-inch Water Tunnel. The test objectives were to obtain and evaluate the low-speed longitudinal aerodynamic characteristics of a candidate configuration for the integration of several new innovative wing designs. The flat plate test allowed for the initial evaluation of the candidate planform and was designated as the baseline planform for the innovative wing design study. Low-speed longitudinal aerodynamic data were obtained over a range of freestream dynamic pressures from 7.5 psf to 30 psf (M = 0.07 to M = 0.14) and angles-of-attack from 0 to 40 deg. The aerodynamic data are presented in coefficient form for the lift, induced drag, and pitching moment. Flow-visualization results obtained were photographs of the flow pattern over the flat plate model in the water tunnel for angles-of-attack from 10 to 40 deg. The force and moment coefficients and the flow-visualization photographs showed the linear and nonlinear aerodynamic characteristics due to attached flow and vortical flow over the flat plate model. Comparison between experiment and linear theory showed good agreement for the lift and induced drag; however, the agreement was poor for the pitching moment.

  9. Measurements of Aerodynamic Heat Transfer and Boundary-Layer Transition on a 10 deg Cone in Free Flight at Supersonic Mach Numbers up to 5.9

    NASA Technical Reports Server (NTRS)

    Rumsey, Charles B.; Lee, Dorothy B.

    1961-01-01

    Measurements of aerodynamic heat transfer have been made at six stations on the 40-inch-long 10 deg. total-angle conical nose of a rocket- propelled model which was flight tested at Mach numbers up to 5.9. are presented for a range of local Mach number just outside the bound- ary layer on the cone from 1.57 to 5.50, and a range of local Reynolds number from 6.6 x 10(exp 6) to 55.2 x 10(exp 6) based on length from the nose tip.

  10. Minimum energy, liquid hydrogen supersonic cruise vehicle study

    NASA Technical Reports Server (NTRS)

    Brewer, G. D.; Morris, R. E.

    1975-01-01

    The potential was examined of hydrogen-fueled supersonic vehicles designed for cruise at Mach 2.7 and at Mach 2.2. The aerodynamic, weight, and propulsion characteristics of a previously established design of a LH2 fueled, Mach 2.7 supersonic cruise vehicle (SCV) were critically reviewed and updated. The design of a Mach 2.2 SCV was established on a corresponding basis. These baseline designs were then studied to determine the potential of minimizing energy expenditure in performing their design mission, and to explore the effect of fuel price and noise restriction on their design and operating performance. The baseline designs of LH2 fueled aircraft were than compared with equivalent designs of jet A (conventional hydrocarbon) fueled SCV's. Use of liquid hydrogen for fuel for the subject aircraft provides significant advantages in performance, cost, noise, pollution, sonic boom, and energy utilization.

  11. Off-Design Reynolds Number Effects for a Supersonic Transport

    NASA Technical Reports Server (NTRS)

    Owens, Lewis R.; Wahls, Richard A.; Rivers, S. Melissa

    2005-01-01

    A high Reynolds number wind tunnel test was conducted to assess Reynolds number effects on the aerodynamic performance characteristics of a realistic, second-generation supersonic transport concept. The tests included longitudinal studies at transonic and low-speed, high-lift conditions across a range of chord Reynolds numbers (8 million to 120 million). Results presented focus on Reynolds number and static aeroelastic sensitivities at Mach 0.30 and 0.90 for a configuration without a tail. Static aeroelastic effects, which mask Reynolds number effects, were observed. Reynolds number effects were generally small and the drag data followed established trends of skin friction as a function of Reynolds number. A more nose-down pitching moment was produced as Reynolds number increased because of an outward movement of the inboard leading-edge separation at constant angles of attack. This study extends the existing Reynolds number database for supersonic transports operating at off-design conditions.

  12. Supersonic second order analysis and optimization program user's manual

    NASA Technical Reports Server (NTRS)

    Clever, W. C.

    1984-01-01

    Approximate nonlinear inviscid theoretical techniques for predicting aerodynamic characteristics and surface pressures for relatively slender vehicles at supersonic and moderate hypersonic speeds were developed. Emphasis was placed on approaches that would be responsive to conceptual configuration design level of effort. Second order small disturbance theory was utilized to meet this objective. Numerical codes were developed for analysis and design of relatively general three dimensional geometries. Results from the computations indicate good agreement with experimental results for a variety of wing, body, and wing-body shapes. Case computational time of one minute on a CDC 176 are typical for practical aircraft arrangement.

  13. Aerodynamic characteristics of a rotorcraft airfoil designed for the tip region of a main rotor blade

    NASA Technical Reports Server (NTRS)

    Noonan, Kevin W.

    1991-01-01

    A wind tunnel investigation was conducted to determine the 2-D aerodynamic characteristics of a new rotorcraft airfoil designed for application to the tip region (stations outboard of 85 pct. radius) of a helicopter main rotor blade. The new airfoil, the RC(6)-08, and a baseline airfoil, the RC(3)-08, were investigated in the Langley 6- by 28-inch transonic tunnel at Mach numbers from 0.37 to 0.90. The Reynolds number varied from 5.2 x 10(exp 6) at the lowest Mach number to 9.6 x 10(exp 6) at the highest Mach number. Some comparisons were made of the experimental data for the new airfoil and the predictions of a transonic, viscous analysis code. The results of the investigation indicate that the RC(6)-08 airfoil met the design goals of attaining higher maximum lift coefficients than the baseline airfoil while maintaining drag divergence characteristics at low lift and pitching moment characteristics nearly the same as those of the baseline airfoil. The maximum lift coefficients of the RC(6)-08 varied from 1.07 at M=0.37 to 0.94 at M=0.52 while those of the RC(3)-08 varied from 0.91 to 0.85 over the same Mach number range. At lift coefficients of -0.1 and 0, the drag divergence Mach number of both the RC(6)-08 and the RC(3)-08 was 0.86. The pitching moment coefficients of the RC(6)-08 were less negative than those of the RC(3)-08 for Mach numbers and lift coefficients typical of those that would occur on a main rotor blade tip at high forward speeds on the advancing side of the rotor disk.

  14. Characteristics of the NASA-Ames Laminar Flow Supersonic Wind Tunnel for Unique Mach 1.6 Transition Studies

    NASA Technical Reports Server (NTRS)

    Wolf, Stephen W. D.; Laub, James A.

    1997-01-01

    Flow quality measurements have been performed in the unique Laminar Flow Supersonic Wind Tunnel (LFSWT) to examine both mean and dynamic characteristics. The intent was to provide the necessary flow information about this ground test facility, to support meaningful transition research at Mach 1.6 and flight unit Reynolds numbers. This paper is intended to assist other experimentalists with similar goals of characterizing low-supersonic test environments. An array of instrumentation has been used to highlight the importance of proper selection of pressure instruments and data acquisition procedures. We conclude that the test section is low-disturbance (based on classical standards of pressure disturbances less than 0.1% with no specified data bandwidth), and has uniform flow. This is confirmation that the quiet design features of the LFSWT are effective. However, characterization of the test section flow over a 0.25k-5Ok bandwidth shows that the disturbance levels can be greater than classical standards particularly for stagnation pressures less than 9.5 psia (0.65 bar) with low stagnation temperatures. Variability of the flow disturbances in the settling chamber and test section is contained in a narrow frequency bandwidth below 5k Hz, which is associated with resonant frequencies from the pressure reduction system. So far, these disturbances have not impacted transition along the tunnel walls or a 10 degrees cone. However, continual vigilance is required to maintain a known low-disturbance environment for transition research in the LFSWT. Furthermore, the formation of standards for flow quality measurements is strongly recommended, so that transition research can be better isolated from tunnel disturbances.

  15. Improvement of aerodynamic characteristics of a thick airfoil with a vortex cell in sub- and transonic flow

    NASA Astrophysics Data System (ADS)

    Isaev, Sergey; Baranov, Paul; Popov, Igor; Sudakov, Alexander; Usachov, Alexander

    2017-03-01

    The modified SST model (2005) is verified using Rodi- Leschziner-Isaev's approach and the multiblock computational technologies are validated in the VP2/3 code on different-structure overlapping grids by comparing the numerical predictions with the experimental data on transonic flow around an NACA0012 airfoil at an angle of attack of 4o for M=0.7 and Re=4×106. It is proved that the aerodynamic characteristics of a thick (20% of the chord) MQ airfoil mounted at an angle of attack of 2o for Re=107 and over the Mach number range 0.3-0.55 are significantly improved because an almost circular small-size (0.12) vortex cell with a defined volumetric flow rate coefficient of 0.007 during slot suction has been located on the upper airfoil section and an intense trapped vortex has been formed in it. A detailed analysis of buffeting within the self-oscillatory regime of flow around the MQ airfoil with a vortex cell has demonstrated the periodic changes in local and integral characteristics; the lift and the aerodynamic efficiency remain quite high, but inferior to the similar characteristics at M=0.55. It is found that the vortex cell at M=0.7 is inactive, and the aerodynamic characteristics of the MQ airfoil with a vortex cell are close to those of a smooth airfoil without a cell.

  16. Program VSAERO theory document: A computer program for calculating nonlinear aerodynamic characteristics of arbitrary configurations

    NASA Technical Reports Server (NTRS)

    Maskew, Brian

    1987-01-01

    The VSAERO low order panel method formulation is described for the calculation of subsonic aerodynamic characteristics of general configurations. The method is based on piecewise constant doublet and source singularities. Two forms of the internal Dirichlet boundary condition are discussed and the source distribution is determined by the external Neumann boundary condition. A number of basic test cases are examined. Calculations are compared with higher order solutions for a number of cases. It is demonstrated that for comparable density of control points where the boundary conditions are satisfied, the low order method gives comparable accuracy to the higher order solutions. It is also shown that problems associated with some earlier low order panel methods, e.g., leakage in internal flows and junctions and also poor trailing edge solutions, do not appear for the present method. Further, the application of the Kutta conditions is extremely simple; no extra equation or trailing edge velocity point is required. The method has very low computing costs and this has made it practical for application to nonlinear problems requiring iterative solutions for wake shape and surface boundary layer effects.

  17. An experimental study of the aerodynamic characteristics of planar and non-planar outboard wing planforms

    NASA Technical Reports Server (NTRS)

    Naik, D. A.; Ostowari, C.

    1987-01-01

    A series of wind tunnel experiments have been conducted to investigate the aerodynamic characteristics of several planar and nonplanar wingtip planforms. Seven different configurations: base-line rectangular, elliptical, swept and tapered, swept and tapered with dihedral, swept and tapered with anhedral, rising arc, and drooping arc, were investigated for two different spans. The data are available in terms of coefficient plots of force data, flow visualization photographs, and velocity and pressure flowfield surveys. All planforms, particularly the nonplanar, have some advantages over the baseline rectangular planform. Span efficiencies up to 20-percent greater than baseline are a possibility. However, it is suggested that the span efficiency concept might need refinement for nonplanar wings. Flow survey data show the change in effective span with vortex roll-up. The flow visualization shows the occurrence of mushroom-cell-separation flow patterns at angles of attack corresponding to stall. These grow with an increase in post-stall angle of attack. For the larger aspect ratios, the cells are observed to split into sub-cells at the higher angles of attack. For all angles of attack, some amount of secondary vortex flow is observed for the planar and nonplanar out-board planforms with sweep and taper.

  18. High angle-of-attack aerodynamic characteristics of crescent and elliptic wings

    NASA Technical Reports Server (NTRS)

    Vandam, C. P.

    1989-01-01

    Static longitudinal and lateral-directional forces and moments were measured for elliptic- and crescent-wing models at high angles-of-attack in the NASA Langley 14 by 22-Ft Subsonic Tunnel. The forces and moments were obtained for an angle-of-attack range including stall and post-stall conditions at a Reynolds number based on the average wing chord of about 1.8 million. Flow-visualization photographs using a mixture of oil and titanium-dioxide were also taken for several incidence angles. The force and moment data and the flow-visualization results indicated that the crescent wing model with its highly swept tips produced much better high angle-of-attack aerodynamic characteristics than the elliptic model. Leading-edge separation-induced vortex flow over the highly swept tips of the crescent wing is thought to produce this improved behavior at high angles-of-attack. The unique planform design could result in safer and more efficient low-speed airplanes.

  19. The aerodynamic characteristics of vortex ingestion for the F/A-18 inlet duct

    NASA Technical Reports Server (NTRS)

    Anderson, Bernhard H.

    1991-01-01

    A Reduced Navier-Stokes (RNS) solution technique was successfully combined with the concept of partitioned geometry and mesh generation to form a very efficient 3D RNS code aimed at the analysis-design engineering environment. Partitioned geometry and mesh generation is a pre-processor to augment existing geometry and grid generation programs which allows the solver to (1) recluster an existing gridlife mesh lattice, and (2) perturb an existing gridfile definition to alter the cross-sectional shape and inlet duct centerline distribution without returning to the external geometry and grid generator. The present results provide a quantitative validation of the initial value space marching 3D RNS procedure and demonstrates accurate predictions of the engine face flow field, with a separation present in the inlet duct as well as when vortex generators are installed to supress flow separation. The present results also demonstrate the ability of the 3D RNS procedure to analyze the flow physics associated with vortex ingestion in general geometry ducts such as the F/A-18 inlet. At the conditions investigated, these interactions are basically inviscid like, i.e., the dominant aerodynamic characteristics have their origin in inviscid flow theory.

  20. Aerodynamic characteristics of the Toroidal Accelerator Rotor Platform (TARP) wind energy conversion system

    SciTech Connect

    Not Available

    1980-02-01

    This report describes an analytical and experimental research program that has been conducted at Rensselaer Polytechnic Institute for the purpose of evaluating the aerodynamic characteristics of the Toroidal Accelerator Rotor Platform (TARP) wind energy conversion system. The TARP is an obstruction type flow concentrator and accelerator which converts ambient winds into low pressure, high kinetic energy zones in the immediate proximity of a wind energy conversion unit. A TARP may be described as being substantially the shape of an inner section of a hollow toroid. A twin rotor system of any kind may be mounted within the peripheral flow channel about a TARP structure such that each rotor is situated in the optimum accelerated flow velocity region for best energy recovery. In a series of preliminary experimental tests, the pressure distribution about the basic TARP configuration was obtained at Reynolds numbers based on the TARP's minimum diameter ranging from about 1.1 x 10/sup 5/ to 9.0 x 10/sup 5/.

  1. Aerodynamic characteristics of a propeller-powered high-lift semispan wing

    NASA Technical Reports Server (NTRS)

    Gentry, Garl L., Jr.; Takallu, M. A.; Applin, Zachary T.

    1994-01-01

    A small-scale semispan high-lift wing-flap system equipped under the wing with a turboprop engine assembly was tested in the LaRC 14- by 22-Foot Subsonic Tunnel. Experimental data were obtained for various propeller rotational speeds, nacelle locations, and nacelle inclinations. To isolate the effects of the high lift system, data were obtained with and without the flaps and leading-edge device. The effects of the propeller slipstream on the overall longitudinal aerodynamic characteristics of the wing-propeller assembly were examined. Test results indicated that the lift coefficient of the wing could be increased by the propeller slipstream when the rotational speed was increased and high-lift devices were deployed. Decreasing the nacelle inclination (increased pitch down) enhanced the lift performance of the system much more than varying the vertical or horizontal location of the nacelle. Furthermore, decreasing the nacelle inclination led to higher lift curve slope values, which indicated that the powered wing could sustain higher angles of attack near maximum lift performance. Any lift augmentation was accompanied by a drag penalty due to the increased wing lift.

  2. Aerodynamic characteristics of a 1/6-scale powered model of the rotor systems research aircraft

    NASA Technical Reports Server (NTRS)

    Mineck, R. E.; Freeman, C. E.

    1977-01-01

    A wind-tunnel investigation was conducted to determine the effects of the main-rotor wake on the aerodynamic characteristics of the rotor systems research aircraft (RSRA). For the investigation, a 1/6-scale model with a four-blade articulated main rotor was used. Tests were conducted with and without the main rotor. Both the helicopter and the compound helicopter were tested. The latter configuration included the auxiliary thrust engines and the variable-incidence wing. Data were obtained over ranges of angle of attack, angle of sideslip, and main-rotor collective pitch angle at several main-rotor advance ratios. Results are presented for the total loads on the airframe as well as the loads on the rotor, the wing, and the tail. The results indicated that without the effect of the rotor wake, the RSRA had static longitudinal and directional stability and positive effective dihedral. With the effect of the main rotor and its wake, the RSRA exhibited longitudinal instability but retained static directional stability and positive effective dihedral.

  3. Calculation of the longitudinal aerodynamic characteristics of upper-surface-blown wing-flap configurations

    NASA Technical Reports Server (NTRS)

    Mendenhall, M. R.; Spangler, S. B.

    1978-01-01

    An engineering method for predicting the longitudinal aerodynamic characteristics of wing-flap configurations with upper surface blowing (USB) was developed. Potential flow models were incorporated into the prediction method: a wing and flap lifting surface model and a jet wake model. The wing-flap model used a vortex-lattice to represent the wing and flaps. The wing had an arbitrary planform and camber and twist, and the flap system was made up of a Coanda flap and other flap segments of arbitrary size. The jet wake model consisted of a series of closely spaced rectangular vortex rings. The wake was positioned such that it was tangent to the upper surface of the wing and flap between the exhaust nozzle and the flap trailing edge. It was specified such that the mass, momentum, and spreading rates were similar to actual USB jet wakes. Comparisons of measured and predicted pressure distributions, span load distributions, and total lift and pitching-moment coefficients on swept and unswept USB configurations are included. A wide range of thrust coefficients and flap deflection angles were considered at angles of attack up to the onset of stall.

  4. Numerical Simulations of the Steady and Unsteady Aerodynamic Characteristics of a Circulation Control Wing Airfoil

    NASA Technical Reports Server (NTRS)

    Liu, Yi; Sankar, Lakshmi N.; Englar, Robert J.; Ahuja, Krishan K.

    2003-01-01

    The aerodynamic characteristics of a Circulation Control Wing (CCW) airfoil have been numerically investigated, and comparisons with experimental data have been made. The configuration chosen was a supercritical airfoil with a 30 degree dual-radius CCW flap. Steady and pulsed jet calculations were performed. It was found that the use of steady jets, even at very small mass flow rates, yielded a lift coefficient that is comparable or superior to conventional high-lift systems. The attached flow over the flap also gave rise to lower drag coefficients, and high L/D ratios. Pulsed jets with a 50% duty cycle were also studied. It was found that they were effective in generating lift at lower reduced mass flow rates compared to a steady jet, provided the pulse frequency was sufficiently high. This benefit was attributable to the fact that the momentum coefficient of the pulsed jet, during the portions of the cycle when the jet was on, was typically twice as much as that of a steady jet.

  5. Investigation of Aerodynamic and Icing Characteristics of a Flush Alternate Inlet Induction System Air Scoop

    NASA Technical Reports Server (NTRS)

    Lewis, James P.

    1953-01-01

    An investigation has been made in the NACA Lewis icing research tunnel to determine the aerodynamic and icing characteristics of a full-scale induction-system air-scoop assembly incorporating a flush alternate inlet. The flush inlet was located immediately downstream of the offset ram inlet and included a 180 deg reversal and a 90 deg elbow in the ducting between inlet and carburetor top deck. The model also had a preheat-air inlet. The investigation was made over a range of mass-air- flow ratios of 0 to 0.8, angles of attack of 0 and 4 deg airspeeds of 150 to 270 miles per hour, air temperatures of 0 and 25 F various liquid-water contents, and droplet sizes. The ram inlet gave good pressure recovery in both clear air and icing but rapid blockage of the top-deck screen occurred during icing. The flush alternate inlet had poor pressure recovery in both clear air and icing. The greatest decreases in the alternate-inlet pressure recovery were obtained at icing conditions of low air temperature and high liquid-water content. No serious screen icing was observed with the alternate inlet. Pressure and temperature distributions on the carburetor top deck were determined using the preheat-air supply with the preheat- and alternate-inlet doors in various positions. No screen icing occurred when the preheat-air system was operated in combination with alternate-inlet air flow.

  6. Low-speed, high-lift aerodynamic characteristics of slender, hypersonic accelerator-type configurations

    NASA Technical Reports Server (NTRS)

    Gatlin, Gregory M.

    1989-01-01

    Two investigations were conducted in the Langley 14 by 22 Foot Subsonic Tunnel to determine the low-speed aerodynamic characteristics of a generic hypersonic accelerator-type configuration. The model was a delta wing configuration incorporating a conical forebody, a simulated wrap-around engine package, and a truncated conical aftbody. Six-component force and moment data were obtained over a range of attack from -4 to 30 degrees and for a sideslip range of + or - 20 degrees. In addition to tests of the basic configuration, component build-up tests were conducted; and the effects of power, forebody nose geometry, canard surfaces, fuselage strakes, and engines on the lower surface alone were also determined. Control power available from deflections of wing flaps and aftbody flaps was also investigated and found to be significantly increased during power-on conditions. Large yawing moments resulted from asymmetric flow fields exhibited by the forebody as revealed by both surface pressure data and flow visualization. Increasing nose bluntness reduced the yawing-moment asymmetry, and the addition of a canard eliminated the yawing-moment asymmetry.

  7. Analysis of some aerodynamic characteristics due to wing-jet interaction

    NASA Technical Reports Server (NTRS)

    Fillman, G. L.; Lan, C. E.

    1979-01-01

    The results of two separate theoretical investigations are presented. A program was used which is capable of predicting the aerodynamic characteristics of both upper-surface blowing (USB) and over-wing blowing (OWB) configurations. A theoretical analysis of the effects of over-wing blowing jets on the induced drag of a 50 deg sweep back wing was developed. Experiments showed net drag reductions associated with the well known lift enhancement due to over-wing blowing. The mechanisms through which this drag reduction is brought about are presented. Both jet entrainment and the so called wing-jet interaction play important roles in this process. The effects of a rectangular upper-surface blowing jet were examined for a wide variety of planforms. The isolated effects of wing taper, sweep, and aspect ratio variations on the incremental lift due to blowing are presented. The effects of wing taper ratio and sweep angle were found to be especially important parameters when considering the relative levels of incremental lift produced by an upper-surface blowing configuration.

  8. Research on a two-dimensional inlet for a supersonic V/STOL propulsion system. Appendix A

    NASA Technical Reports Server (NTRS)

    Mark, J. L.; Mcgarry, M. A.; Reagan, P. V.

    1984-01-01

    The inlet system performance requirements associated with supersonic V/STOL aircraft place extreme demands on the inlet designer. The present effort makes maximum use of flow improvement techniques, proven for high subsonic maneuvering flight and adapts them to the critical static and low speed/high angle-of-attack flight regime of the supersonic V/STOL aircraft. A description of the aerodynamic design, model characteristics, data analysis, discussion, and conclusions concerning the most promising inlet design approaches are contained. The appendix contains the reduced wind tunnel data plots and pressure distribution.

  9. Investigation of Aerodynamic and Icing Characteristics of Recessed Fuel-Vent Configurations

    NASA Technical Reports Server (NTRS)

    Ruggeri, Robert S.; VonGlahn, Uwe H.; Rollins, Vern G.

    1949-01-01

    An investigation has been conducted in the NACA Cleveland icing research tunnel to determine the aerodynamic and icing characteristics of several recessed fuel-vent configurations. The vents were investigated aerodynamically to obtain vent-tube pressures and pressure distributions on the ramp surface as functions of tunnel-air velocity and angle of attack. Icing investigations were made to determine the vent-tube pressure losses for several icing conditions at tunnel-air velocities ranging from 220 to 440 feet per second. In general, under nonicing conditions, the configurations with diverging ramp walls maintained, vent-tube pressures greater than the required marginal value of 2 inches of water positive pressure differential between the fuel cell and the compartment containing the fuel cell for a range of angles of attack from 0 to 14deg at a tunnel-air velocity of approximately 240 feet per second. A configuration haying divergIng ramp sldewalls, a 7deg ramp angle; and vent tubes manifold,ed to a common plenum chamber opening through a slot In the ramp floor gave the greatest vent-tube pressures for all the configurations investigated. The use of the plenum chamber resulted in uniform pressures in all vent tubes. In a cloud-icing condition, roughness caused by ice formations on the airfoil surface ahead of the vent ramp, rather than icing of the vent configuration, caused a rapid loss in vent-tube pressures during the first few minutes of an icing period. Only the configuration having diverging ramp sidewalls, a 7 ramp angle, and a common plenum chamber maintained the required vent-tube pressures throughout a 60-minute icing period, although the ice formations on this configuration were more severe than those observed for the other configurations. No complete closure of vent-tube openings occurred for the configurations investigated. A simulated freezing-rain condition caused a greater and more rapid vent-tube pressure loss than was observed for a cloud

  10. Advanced Supersonic Technology concept AST-100 characteristics developed in a baseline-update study

    NASA Technical Reports Server (NTRS)

    Baber, H. T., Jr.; Swanson, E. E.

    1976-01-01

    The advanced supersonic technology configuration, AST-100, is described. The combination of wing thickness reduction, nacelle recontouring for minimum drag at cruise, and the use of the horizontal tail to produce lift during climb and cruise resulted in an increase in maximum lift-to-drag ratio. Lighter engines and lower fuel weight associated with this resizing result in a six percent reduction in takeoff gross weight. The AST-100 takeoff maximum effective perceived noise at the runway centerline and sideline measurement stations was 114.4 decibels. Since 1.5-decibels tradeoff is available from the approach noise, the required engine noise supression is 4.9 decibels. The AST-100 largest maximum overpressure would occur during transonic climb acceleration when the aircraft was at relatively low altitude. Calculated standard +8 C day range of the AST-100, with a 292 passenger payload, is 7348 km (3968 n.mi). Fuel price is the largest contributor to direct operating cost. However, if the AST-100 were flown subsonically (M = 0.9), direct operating costs would increase approximately 50 percent because of time related costs.

  11. Experimental Investigation of the Low-Speed Aerodynamic Characteristics of a 5.8-Percent Scale Hybrid Wing Body Configuration

    NASA Technical Reports Server (NTRS)

    Gatlin, Gregory M.; Vicroy, Dan D.; Carter, Melissa B.

    2012-01-01

    A low-speed experimental investigation has been conducted on a 5.8-percent scale Hybrid Wing Body configuration in the NASA Langley 14- by 22-Foot Subsonic Tunnel. This Hybrid Wing Body (HWB) configuration was designed with specific intention to support the NASA Environmentally Responsible Aviation (ERA) Project goals of reduced noise, emissions, and fuel burn. This HWB configuration incorporates twin, podded nacelles mounted on the vehicle upper surface between twin vertical tails. Low-speed aerodynamic characteristics were assessed through the acquisition of force and moment, surface pressure, and flow visualization data. Longitudinal and lateral-directional characteristics were investigated on this multi-component model. The effects of a drooped leading edge, longitudinal flow-through nacelle location, vertical tail shape and position, elevon deflection, and rudder deflection have been studied. The basic configuration aerodynamics, as well as the effects of these configuration variations, are presented in this paper.

  12. Aerodynamic characteristics of a high-wing transport configuration with a over-the-wing nacelle-pylon arrangement

    NASA Technical Reports Server (NTRS)

    Henderson, W. P.; Abeyounis, W. K.

    1985-01-01

    An investigation has been conducted in the Langley 16-Foot Transonic Tunnel to determine the effects on the aerodynamic characteristics of a high-wing transport configuration of installing an over-the-wing nacelle-pylon arrangement. The tests are conducted at Mach numbers from 0.70 to 0.82 and at angles of attack from -2 deg to 4 deg. The configurational variables under study include symmetrical and contoured nacelles and pylons, pylon size, and wing leading-edge extensions. The symmetrical nacelles and pylons reduce the lift coefficient, increase the drag coefficient, and cause a nose-up pitching-moment coefficient. The contoured nacelles significantly reduce the interference drag, though it is still excessive. Increasing the pylon size reduces the drag, whereas adding wing leading-edge extension does not affect the aerodynamic characteristics significantly.

  13. Test data from solid propellant plume aerodynamics test program in Ames 6 x 6 foot supersonic wind tunnel (shuttle test FA7) (Ames test 033-66)

    NASA Technical Reports Server (NTRS)

    Hair, L. M.

    1975-01-01

    The aerodynamic effects of plumes from hot combustion gases in the presence of a transonic external flow field were measured to advance plumes simulation technology, extend a previously acquired data base, and provide data to compare with the effects observed using cold gas plumes. A variety of underexpanded plumes issuing from the base of a strut-mounted ogive-cylinder body were produced by combusting solid propellant gas generators. The gas generator fired in a short-duration mode (200 to 300 msec). Propellants containing 16 percent and 2 percent A1 were used, with chamber pressures from 400 to 1800 psia. Conical nozzles of 15 deg half-angle were tested with area ratios of 4 and 8. Pressures were measured in the gas generator combustion chamber, along the nozzle wall, on the base, and along the body rear exterior. Schlieren photographs were taken for all tests. Test data are presented along with a description of the test setup and procedures.

  14. Steady and Oscillatory, Subsonic and Supersonic, Aerodynamic Pressure and Generalized Forces for Complex Aircraft Configurations and Applications to Flutter. M.S. Thesis

    NASA Technical Reports Server (NTRS)

    Chen, L. T.

    1975-01-01

    A general method for analyzing aerodynamic flows around complex configurations is presented. By applying the Green function method, a linear integral equation relating the unknown, small perturbation potential on the surface of the body, to the known downwash is obtained. The surfaces of the aircraft, wake and diaphragm (if necessary) are divided into small quadrilateral elements which are approximated with hyperboloidal surfaces. The potential and its normal derivative are assumed to be constant within each element. This yields a set of linear algebraic equations and the coefficients are evaluated analytically. By using Gaussian elimination method, equations are solved for the potentials at the centroids of elements. The pressure coefficient is evaluated by the finite different method; the lift and moment coefficients are evaluated by numerical integration. Numerical results are presented, and applications to flutter are also included.

  15. Supersonic combustion engine testbed, heat lightning

    NASA Technical Reports Server (NTRS)

    Hoying, D.; Kelble, C.; Langenbahn, A.; Stahl, M.; Tincher, M.; Walsh, M.; Wisler, S.

    1990-01-01

    The design of a supersonic combustion engine testbed (SCET) aircraft is presented. The hypersonic waverider will utilize both supersonic combustion ramjet (SCRAMjet) and turbofan-ramjet engines. The waverider concept, system integration, electrical power, weight analysis, cockpit, landing skids, and configuration modeling are addressed in the configuration considerations. The subsonic, supersonic and hypersonic aerodynamics are presented along with the aerodynamic stability and landing analysis of the aircraft. The propulsion design considerations include: engine selection, turbofan ramjet inlets, SCRAMjet inlets and the SCRAMjet diffuser. The cooling requirements and system are covered along with the topics of materials and the hydrogen fuel tanks and insulation system. A cost analysis is presented and the appendices include: information about the subsonic wind tunnel test, shock expansion calculations, and an aerodynamic heat flux program.

  16. Effects of thickness on the aerodynamic characteristics of an initial low-speed family of airfoils for general aviation applications

    NASA Technical Reports Server (NTRS)

    Mcghee, R. J.; Beasley, W. D.

    1976-01-01

    Wind tunnel tests were conducted to determine the effects of airfoil thickness-ratio on the low speed aerodynamic characteristics of an initial family of airfoils. The results were compared with theoretical predictions obtained from a subsonic viscous method. The tests were conducted over a Mach number range from 0.10 to 0.28. Chord Reynolds numbers varied from about 2.0 x 1 million to 9.0 x 1 million.

  17. Tests of Four Full-scale Propellers to Determine the Effect of Trailing-edge Extensions on Propeller Aerodynamic Characteristics

    NASA Technical Reports Server (NTRS)

    Maynard, Julian D.; Evans, Albert J

    1945-01-01

    Propellers with trailing-edge extensions were studied to determine aerodynamic characteristics. Trailing-edge extension increased power absorbed by propeller with little loss in efficiency. Power coefficient for maximum efficiency was greater for 20% camber type extension than for 20% straight type extension over range of advance ratio of 1.0 to 2.5 although camber type was less efficient. Efficiency was about the same for cruising and high-speed at a high power coefficient for propeller with extension.

  18. Longitudinal aerodynamics of a low-wing lift-fan transport including hover characteristics in and out of ground effect

    NASA Technical Reports Server (NTRS)

    Hoad, D. R.; Gentry, G. L., Jr.

    1975-01-01

    A wind-tunnel investigation was conducted in the Langley V/STOL tunnel to determine the longitudinal aerodynamic characteristics of a six-fan, tip-driven (remote) lift-fan VTOL transport throughout transition. The large midspan lift-fan pods and cruise fans were removed to determine their influence on the stability and control of the configuration. Data were obtained in the hovering mode for ranges of model height above ground. The data are presented without analysis or discussion.

  19. Analysis of wind tunnel test results for a 9.39-per cent scale model of a VSTOL fighter/attack aircraft. Volume 1: Study overview. [aerodynamic characteristics

    NASA Technical Reports Server (NTRS)

    Lummus, J. R.; Joyce, G. T.; Omalley, C. D.

    1980-01-01

    The ability of current methodologies to accurately predict the aerodynamic characteristics identified as uncertainties was evaluated for two aircraft configurations. The two wind tunnel models studied horizontal altitude takeoff and landing V/STOL fighter aircraft derivatives.

  20. Wind Tunnel Investigation of the Effects of Surface Porosity and Vertical Tail Placement on Slender Wing Vortex Flow Aerodynamics at Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.

    2007-01-01

    A wind tunnel experiment was conducted in the NASA Langley Research Center (LaRC) Unitary Plan Wind Tunnel (UPWT) to determine the effects of passive surface porosity and vertical tail placement on vortex flow development and interactions about a general research fighter configuration at supersonic speeds. Optical flow measurement and flow visualization techniques were used that featured pressure sensitive paint (PSP), laser vapor screen (LVS), and schlieren, These techniques were combined with conventional electronically-scanned pressure (ESP) and six-component force and moment measurements to quantify and to visualize the effects of flow-through porosity applied to a wing leading edge extension (LEX) and the placement of centerline and twin vertical tails on the vortex-dominated flow field of a 65 cropped delta wing model. Test results were obtained at free-stream Mach numbers of 1.6, 1.8, and 2.1 and a Reynolds number per foot of 2.0 million. LEX porosity promoted a wing vortex-dominated flow field as a result of a diffusion and weakening of the LEX vortex. The redistribution of the vortex-induced suction pressures contributed to large nose-down pitching moment increments but did not significantly affect the vortex-induced lift. The trends associated with LEX porosity were unaffected by vertical tail placement. The centerline tail configuration generally provided more stable rolling moments and yawing moments compared to the twin wing-mounted vertical tails. The strength of a complex system of shock waves between the twin tails was reduced by LEX porosity.

  1. Nonlinear flutter analysis of stiffened composite panels in supersonic flow

    NASA Astrophysics Data System (ADS)

    Yuan, Kaihua; Qiu, Zhiping

    2010-02-01

    The flutter instability of stiffened composite panels subjected to aerodynamic forces in the supersonic flow is investigated. Based on Hamilton’s principle, the aeroelastic model of the composite panel is established by using the von Karman large deflection plate theory, piston theory aerodynamics and the quasi-steady thermal stress theory. Then, using the finite element method along with Bogner-Fox-Schmit elements and three-dimensional beam elements, the nonlinear equations of motion are derived. The effect of stiffening scheme on the flutter critical dynamic pressure is demonstrated through the numerical example, and the nonlinear flutter characteristics of stiffened composite panels are also analyzed in the time domain. This will lay the foundation for design of panel structures employed in aerospace vehicles.

  2. Investigation of Aerodynamic and Icing Characteristics of Water-Inertia-Separation Inlets for Turbojet Engines

    NASA Technical Reports Server (NTRS)

    VonGlahn, Uwe; Blatz, R. E.

    1950-01-01

    The results of an investigation of several internal water-inertia-separation inlets consisting of a main duct and an alternate duct designed to prevent automatically the entrance of large quantities of water into a turbojet engine in icing conditions are presented. Total-pressure losses and icing characteristics for a direct-ram inlet and the inertia-separation inlets are compared at similar aerodynamic and simulated icing conditions. Complete ice protection for inlet guide vanes could not be achieved with the inertia-separation inlets investigated. Approximately 8 percent of the volume of water entering the nacelles remained. In the air passing into the compressor inlet. Heavy alternate-duct-elbow ice formations caused by secondary inertia separation resulted in rapid total-pressure losses and decreases in mass flow. The duration in an icing condition for an inertia-separation- inlet, without local surface heating, was increased approximately four times above that for a direct-ram inlet with a compressor-inlet screen. For normal nonicing operation, the inertia-separation- inlet total-pressure losses were comparable to a direct-ram installation. The pressure losses and the circumferential uniformity of the mass flow in all the inlets were relatively independent of angle of attack. Use of an inertia-separation inlet would in most cases require a larger diameter nacelle than a direct-ram inlet in order to obtain an alternate duct sufficiently large to pass the required engine air flow at duct Mach numbers below 1.0 at the minimum area.

  3. Experimental Aerodynamic Characteristics of an Oblique Wing for the F-8 OWRA

    NASA Technical Reports Server (NTRS)

    Kennelly, Robert A., Jr.; Carmichael, Ralph L.; Smith, Stephen C.; Strong, James M.; Kroo, Ilan M.

    1999-01-01

    An experimental investigation was conducted during June-July 1987 in the NASA Ames 11-Foot Transonic Wind Tunnel to study the aerodynamic performance and stability and control characteristics of a 0.087-scale model of an F-8 airplane fitted with an oblique wing. This effort was part of the Oblique Wing Research Aircraft (OWRA) program performed in conjunction with Rockwell International. The Ames-designed, aspect ratio 10.47, tapered wing used specially designed supercritical airfoils with 0.14 thickness/chord ratio at the root and 0.12 at the 85% span location. The wing was tested at two different mounting heights above the fuselage. Performance and longitudinal stability data were obtained at sweep angles of 0deg, 30deg, 45deg, 60deg, and 65deg at Mach numbers ranging from 0.30 to 1.40. Reynolds number varied from 3.1 x 10(exp 6)to 5.2 x 10(exp 6), based on the reference chord length. Angle of attack was varied from -5deg to 18deg. The performance of this wing is compared with that of another oblique wing, designed by Rockwell International, which was tested as part of the same development program. Lateral-directional stability data were obtained for a limited combination of sweep angles and Mach numbers. Sideslip angle was varied from -5deg to +5deg. Landing flap performance was studied, as were the effects of cruise flap deflections to achieve roll trim and tailor wing camber for various flight conditions. Roll-control authority of the flaps and ailerons was measured. A novel, deflected wing tip was evaluated for roll-control authority at high sweep angles.

  4. Experimental Study on Aerodynamic Characteristics of a Stepped-Nose Obstacle

    NASA Astrophysics Data System (ADS)

    Cakrawala, Anang; Umemura, Akira

    Wind tunnel experiments were conducted to obtain a comprehensive understanding of the aerodynamic characteristics of two-dimensional stepped-nose obstacles. In our previous numerical study on the flow around stepped-nose obstacles at a zero angle of attack, the following favorable properties have been found for such step configuration that the flow separating from the front corners reattaches to the leading edges of the side surfaces. (1) The strong vortices trapped in the step region produce the suction forces acting on the step walls to cancel the drag force acting the front surface. (2) The suppression of large-scale flow separation on the obstacle’s sides reduces not only the suction force acting on the back surface, but also the lateral force fluctuation to a great degree. (3) The resulting net drag coefficient is much smaller than that of square/rectangular obstacles. In the present experimental study, the drag coefficient of stepped-nosed obstacles with various step height and length, at zero angle of attack, was measured to identify the optimum step configuration for large drag reduction. The effect of attack angle on drag, lift and moment coefficients was examined to gain insight into the static and dynamic stability of stepped-nose obstacles. The effect of step configuration on the Strouhal number was also examined. It was found that the stepped nose brought about static stability to the obstacle with a rather large step length-to-height ratio, but neither static nor dynamic stability was derived for the optimal step configuration with maximum drag reduction.

  5. Low-speed aerodynamic characteristics of a twin-engine general aviation configuration with aft-fuselage-mounted pusher propellers

    NASA Technical Reports Server (NTRS)

    Dunham, Dana Morris; Gentry, Garl L., Jr.; Manuel, Gregory S.; Applin, Zachary T.; Quinto, P. Frank

    1987-01-01

    An investigation was conducted to determine the aerodynamic characteristics of an advanced turboprop aircraft model with aft-pylon-mounted pusher propellers. Tests were conducted through an angle-of-attack range of -8 to 28 degrees, and an angle-of-sideslip range of -20 to 20 degrees at free-stream conditions corresponding to Reynolds numbers of 0.55 to 2.14 x 10 to the 6th power based on mean aerodynamic chord. Test results show that for the unpowered configurations the maximum lift coefficients for the cruise, takeoff, and landing configurations are 1.45, 1.90, and 2.10, respectively. Nacelle installation results in a drag coefficient increase of 0.01. Increasing propeller thrust results in a significant increase in lift for angles of attack above stall and improves the longitudinal stability. The cruise configuration remains longitudinally stable to an angle of attack 5 degrees beyond the stall angle, the takeoff configuration is stable 4 degrees beyond stall angle, and the landing configuration is stable 3 degrees beyond stall angle. The predominant effect of symmetric thrust on the lateral-directional aerodynamic characteristics is in the post-stall region, where additional rudder control is available with power on.

  6. Supersonic LFC: Challenges and opportunities

    NASA Technical Reports Server (NTRS)

    Powell, Arthur G.

    1992-01-01

    The discussion and viewgraphs on supersonic laminar control are provided. The high fuel fractions required for long range supersonic airplanes give significant leverage to technologies for cruise drag reduction such as laminar flow control (LFC). Fuel burn benefits are further enhanced when sizing effects are considered. These effects may even be powerful enough to reduce airplane production cost over a turbulent baseline. This is an important goal for LFC technology development. The results of aerodynamics studies on the application of LFC technology to the highly swept wings of supersonic airplanes are presented. Important questions of applicability, realistic benefit, and critical application issues, addressed in a NASA-sponsored study conducted by McDonnell Douglas Corporation in 1987-88 are reviewed. Efforts aimed at establishing the feasibility of demonstrating extensive laminarization on the F-16XL-2 airplane are summarized.

  7. Supersonic transport vis-a-vis energy savings

    NASA Technical Reports Server (NTRS)

    Cormery, G.

    1979-01-01

    The energy and economic saving modifications in supersonic transportation are studied. Modifications in the propulsion systems and in the aerodynamic configurations of the Concorde aircraft to reduce noise generation and increase fuel efficiency are discussed. The conversion of supersonic aircraft from fuel oils to synthetic fuels is examined.

  8. Aerodynamic characteristics of a canard-controlled missile at Mach numbers of 1.5 and 2.0.

    NASA Technical Reports Server (NTRS)

    Kassner, D. L.; Wettlaufer, B.

    1977-01-01

    A typical missile model with nose mounted canards and cruciform tail surfaces was tested in the Ames 6- by 6-Foot Wind Tunnel to determine the contributions of the component aerodynamic surfaces to the static aerodynamic characteristics at Mach numbers of 1.5 and 2.0 and Reynolds number of 1 million based on body diameter. Data were obtained at angles of attack ranging from -3 deg to 12 deg for various stages of model build-up (i.e., with and without canard and/or tail surfaces). Results were obtained both with the model unrolled and rolled 45 deg. For the canard and tail arrangements investigated, the model was trimmable at angles of attack up to about 10 deg with canard deflections of 9 deg. Also, the tail arrangements studied provided ample pitch stability. there were no appreciable effects of model roll orientation.

  9. Comparison Between Theory and Experiment for Wings at Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Vincenti, Walter G

    1951-01-01

    This paper presents a critical comparison made between experimental and theoretical results for the aerodynamic characteristics of wings at supersonic flight speeds. As a preliminary, a brief, nonmathematical review is given of the basic assumptions and general findings of supersonic wing theory in two and three dimensions. Published data from two-dimensional pressure-distribution tests are then used to illustrate the effects of fluid viscosity and to assess the accuracy of linear theory as compared with the more exact theories which are available in the two-dimensional case. Finally, an account is presented of an NACA study of the over-all force characteristics of three-dimensional wings at supersonic speed. In this study, the lift, pitching moment, and drag characteristics of several families of wings of varying plan form and section were measured in the wind tunnel and compared with values predicted by the three-dimensional linear theory. The regions of agreement and disagreement between experiment and theory are noted and discussed.

  10. Comparison of theoretically predicted lateral-directional aerodynamic characteristics with full-scale wind tunnel data on the ATLIT airplane

    NASA Technical Reports Server (NTRS)

    Griswold, M.; Roskam, J.

    1980-01-01

    An analytical method is presented for predicting lateral-directional aerodynamic characteristics of light twin engine propeller-driven airplanes. This method is applied to the Advanced Technology Light Twin Engine airplane. The calculated characteristics are correlated against full-scale wind tunnel data. The method predicts the sideslip derivatives fairly well, although angle of attack variations are not well predicted. Spoiler performance was predicted somewhat high but was still reasonable. The rudder derivatives were not well predicted, in particular the effect of angle of attack. The predicted dynamic derivatives could not be correlated due to lack of experimental data.

  11. The Edge supersonic transport

    NASA Technical Reports Server (NTRS)

    Agosta, Roxana; Bilbija, Dushan; Deutsch, Marc; Gallant, David; Rose, Don; Shreve, Gene; Smario, David; Suffredini, Brian

    1992-01-01

    As intercontinental business and tourism volumes continue their rapid expansion, the need to reduce travel times becomes increasingly acute. The Edge Supersonic Transport Aircraft is designed to meet this demand by the year 2015. With a maximum range of 5750 nm, a payload of 294 passengers and a cruising speed of M = 2.4, The Edge will cut current international flight durations in half, while maintaining competitive first class, business class, and economy class comfort levels. Moreover, this transport will render a minimal impact upon the environment, and will meet all Federal Aviation Administration Part 36, Stage III noise requirements. The cornerstone of The Edge's superior flight performance is its aerodynamically efficient, dual-configuration design incorporating variable-geometry wingtips. This arrangement combines the benefits of a high aspect ratio wing at takeoff and low cruising speeds with the high performance of an arrow-wing in supersonic cruise. And while the structural weight concerns relating to swinging wingtips are substantial, The Edge looks to ever-advancing material technologies to further increase its viability. Heeding well the lessons of the past, The Edge design holds economic feasibility as its primary focus. Therefore, in addition to its inherently superior aerodynamic performance, The Edge uses a lightweight, largely windowless configuration, relying on a synthetic vision system for outside viewing by both pilot and passengers. Additionally, a fly-by-light flight control system is incorporated to address aircraft supersonic cruise instability. The Edge will be produced at an estimated volume of 400 aircraft and will be offered to airlines in 2015 at $167 million per transport (1992 dollars).

  12. Supersonic compressor

    DOEpatents

    Lawlor, Shawn P.; Novaresi, Mark A.; Cornelius, Charles C.

    2008-02-26

    A gas compressor based on the use of a driven rotor having an axially oriented compression ramp traveling at a local supersonic inlet velocity (based on the combination of inlet gas velocity and tangential speed of the ramp) which forms a supersonic shockwave axially, between adjacent strakes. In using this method to compress inlet gas, the supersonic compressor efficiently achieves high compression ratios while utilizing a compact, stabilized gasdynamic flow path. Operated at supersonic speeds, the inlet stabilizes an oblique/normal shock system in the gasdyanamic flow path formed between the gas compression ramp on a strake, the shock capture lip on the adjacent strake, and captures the resultant pressure within the stationary external housing while providing a diffuser downstream of the compression ramp.

  13. Navier-Stokes simulations of Orbiter aerodynamic characteristics including pitch trim and bodyflap

    NASA Technical Reports Server (NTRS)

    Weilmuenster, K. James; Gnoffo, Peter A.; Greene, Francis A.

    1994-01-01

    An analysis of the longitudinal aerodynamics of the shuttle orbiter in the hypersonic flight regime is made through the use of computational fluid dynamics. Particular attention is given to establishing the cause of the 'pitching moment anomaly,' which occurred on the orbiter's first flight, and to computing the aerodynamics of a complete orbiter configuration at flight conditions. Data from ground-based facilities as well as orbiter flight data are used to validate the computed results. Analysis shows that the pitching moment anomaly is a real-gas chemistry effect that was not simulated in ground-based facilities, which used air as a test gas. Computed flight aerodynamics for the orbiter are within 5% of the measured flight values and trim bodyflap deflections are predicted to within 10%.

  14. Measurements of Aerodynamic Heat Transfer and Boundary-Layer Transition on a 15 deg. Cone in Free Flight at Supersonic Mach Numbers up to 5.2

    NASA Technical Reports Server (NTRS)

    Rumsey, Charles B.; Lee, Dorothy B.

    1961-01-01

    Measurements of aerodynamic heat transfer have been made at several stations on the 15 deg total-angle conical nose of a rocket-propelled model in free flight at Mach numbers up to 5.2. Data are presented for a range of local Mach number just outside the boundary layer from 1.40 to 4.65 and a range of local Reynolds number from 3.8 x 10(exp 6) to 46.5 x 10(exp 6), based on length from the nose tip to a measurement station. Laminar, transitional, and turbulent heat-transfer coefficients were measured. The laminar data were in agreement with laminar theory for cones, and the turbulent data agreed well with turbulent theory for cones using Reynolds number based on length from the nose tip. At a nearly constant ratio of wall to local static temperature of 1.2 the Reynolds number of transition increased from 14 x 10(exp 6) to 30 x 10(exp 6) as Mach number increased from 1.4 to 2.9 and then decreased to 17 x 10(exp 6) as Mach number increased to 3.7. At Mach numbers near 3.5, transition Reynolds numbers appeared to be independent of skin temperature at skin temperatures very cold with respect to adiabatic wall temperature. The transition Reynolds number was 17.7 x 10(exp 6) at a condition of Mach number and ratio of wall to local static temperature near that for which three-dimensional disturbance theory has been evaluated and has predicted laminar boundary-layer stability to very high Reynolds numbers (approximately 10(exp 12)).

  15. Experimental Hypersonic Aerodynamic Characteristics of the 2001 Mars Surveyor Precision Lander with Flap

    NASA Technical Reports Server (NTRS)

    Horvath, Thomas J.; OConnell, Tod F.; Cheatwood, F. McNeil; Prabhu, Ramadas K.; Alter, Stephen J.

    2002-01-01

    Aerodynamic wind-tunnel screening tests were conducted on a 0.029 scale model of a proposed Mars Surveyor 2001 Precision Lander (70 deg half angle spherically blunted cone with a conical afterbody). The primary experimental objective was to determine the effectiveness of a single flap to trim the vehicle at incidence during a lifting hypersonic planetary entry. The laminar force and moment data, presented in the form of coefficients, and shock patterns from schlieren photography were obtained in the NASA Langley Aerothermodynamic Laboratory for post-normal shock Reynolds numbers (based on forebody diameter) ranging from 2,637 to 92,350, angles of attack ranging from 0 tip to 23 degrees at 0 and 2 degree sideslip, and normal-shock density ratios of 5 and 12. Based upon the proposed entry trajectory of the 2001 Lander, the blunt body heavy gas tests in CF, simulate a Mach number of approximately 12 based upon a normal shock density ratio of 12 in flight at Mars. The results from this experimental study suggest that when traditional means of providing aerodynamic trim for this class of planetary entry vehicle are not possible (e.g. offset c.g.), a single flap can provide similar aerodynamic performance. An assessment of blunt body aerodynamic effects attributed to a real gas were obtained by synergistic testing in Mach 6 ideal-air at a comparable Reynolds number. From an aerodynamic perspective, an appropriately sized flap was found to provide sufficient trim capability at the desired L/D for precision landing. Inviscid hypersonic flow computations using an unstructured grid were made to provide a quick assessment of the Lander aerodynamics. Navier-Stokes computational predictions were found to be in very good agreement with experimental measurement.

  16. Edge localized mode characteristics during edge localized mode mitigation by supersonic molecular beam injection in Korea Superconducting Tokamak Advanced Research

    SciTech Connect

    Lee, H. Y.; Hong, J. H.; Jang, J. H.; Park, J. S.; Choe, Wonho; Hahn, S. H.; Bak, J. G.; Lee, J. H.; Ko, W. H.; Lee, K. D.; Lee, S. H.; Lee, H. H.; Juhn, J.-W.; Kim, H. S.; Yoon, S. W.; Han, H.; Ghim, Y.-C.

    2015-12-15

    It has been reported that supersonic molecular beam injection (SMBI) is an effective means of edge localized mode (ELM) mitigation. This paper newly reports the changes in the ELM, plasma profiles, and fluctuation characteristics during ELM mitigation by SMBI in Korea Superconducting Tokamak Advanced Research. During the mitigated ELM phase, the ELM frequency increased by a factor of 2–3 and the ELM size, which was estimated from the D{sub α} amplitude, the fractional changes in the plasma-stored energy and the line-averaged electron density, and divertor heat flux during an ELM burst, decreased by a factor of 0.34–0.43. Reductions in the electron and ion temperatures rather than in the electron density were observed during the mitigated ELM phase. In the natural ELM phase, frequency chirping of the plasma fluctuations was observed before the ELM bursts; however, the ELM bursts occurred without changes in the plasma fluctuation frequency in the mitigated ELM phase.

  17. Applied computational aerodynamics

    SciTech Connect

    Henne, P.A.

    1990-01-01

    The present volume discusses the original development of the panel method, the mapping solutions and singularity distributions of linear potential schemes, the capabilities of full-potential, Euler, and Navier-Stokes schemes, the use of the grid-generation methodology in applied aerodynamics, subsonic airfoil design, inverse airfoil design for transonic applications, the divergent trailing-edge airfoil innovation in CFD, Euler and potential computational results for selected aerodynamic configurations, and the application of CFD to wing high-lift systems. Also discussed are high-lift wing modifications for an advanced-capability EA-6B aircraft, Navier-Stokes methods for internal and integrated propulsion system flow predictions, the use of zonal techniques for analysis of rotor-stator interaction, CFD applications to complex configurations, CFD applications in component aerodynamic design of the V-22, Navier-Stokes computations of a complete F-16, CFD at supersonic/hypersonic speeds, and future CFD developments.

  18. Longitudinal aerodynamic characteristics of a low-wing lift-fan transport including hover characteristics in and out of ground effect

    NASA Technical Reports Server (NTRS)

    Hoad, D. R.; Gentry, G. L., Jr.

    1977-01-01

    The longitudinal aerodynamic characteristics of a six-fan, tip-driven (remote) lift-fan VTOL transport through transition were determined by an investigation conducted in the Langley V/STOL tunnel. Tests were also made with the large midspan lift-fan pods and lift-cruise fans removed to determine their their influence on the stability and control of the configuration. Data were obtained for a range of model height above ground.

  19. Formulation of aerodynamic prediction techniques for hypersonic configuration design

    NASA Technical Reports Server (NTRS)

    1979-01-01

    An investigation of approximate theoretical techniques for predicting aerodynamic characteristics and surface pressures for relatively slender vehicles at moderate hypersonic speeds was performed. Emphasis was placed on approaches that would be responsive to preliminary configuration design level of effort. Supersonic second order potential theory was examined in detail to meet this objective. Shock layer integral techniques were considered as an alternative means of predicting gross aerodynamic characteristics. Several numerical pilot codes were developed for simple three dimensional geometries to evaluate the capability of the approximate equations of motion considered. Results from the second order computations indicated good agreement with higher order solutions and experimental results for a variety of wing like shapes and values of the hypersonic similarity parameter M delta approaching one.

  20. 1997 NASA High-Speed Research Program Aerodynamic Performance Workshop. Volume 1; Configuration Aerodynamics

    NASA Technical Reports Server (NTRS)

    Baize, Daniel G. (Editor)

    1999-01-01

    The High-Speed Research Program and NASA Langley Research Center sponsored the NASA High-Speed Research Program Aerodynamic Performance Workshop on February 25-28, 1997. The workshop was designed to bring together NASA and industry High-Speed Civil Transport (HSCT) Aerodynamic Performance technology development participants in areas of Configuration Aerodynamics (transonic and supersonic cruise drag prediction and minimization), High-Lift, Flight Controls, Supersonic Laminar Flow Control, and Sonic Boom Prediction. The workshop objectives were to (1) report the progress and status of HSCT aerodynamic performance technology development; (2) disseminate this technology within the appropriate technical communities; and (3) promote synergy among the scientist and engineers working HSCT aerodynamics. In particular, single- and multi-point optimized HSCT configurations, HSCT high-lift system performance predictions, and HSCT Motion Simulator results were presented along with executive summaries for all the Aerodynamic Performance technology areas.

  1. Aerodynamic preliminary analysis system 2. Part 2: User's manual

    NASA Technical Reports Server (NTRS)

    Sova, G.; Divan, P.; Spacht, L.

    1991-01-01

    An aerodynamic analysis system based on potential theory at subsonic and/or supersonic speeds and impact type finite element solutions at hypersonic conditions is described. Three dimensional configurations have multiple nonplanar surfaces of arbitrary planforms and bodies of noncircular contour may be analyzed. Static, rotary, and control longitudinal and lateral-directional characteristics may be generated. The analysis was implemented on a time sharing system in conjunction with an input tablet digitizer and an interactive graphics input/output display and editing terminal to maximize its responsiveness to the preliminary analysis. Computation times on an IBM 3081 are typically less than one minute of CPU/Mach number at subsonic, supersonic, or hypersonic speeds. This is a user manual for the computer programming.

  2. Transonic aerodynamic characteristics associated with variations in the geometry of the forward portion of irregular planform wings

    NASA Technical Reports Server (NTRS)

    Spencer, B., Jr.; Stone, D. R.

    1973-01-01

    The experimental aerodynamic characteristics of three basic wing planforms on a conceptual orbiter fuselage (designated the LO-100) have been obtained in the 8-Foot Transonic Pressure Tunnel. The study included variations in the forward portion (fillet) of each basic wing. Fillet sweeps to 78 deg were investigated while holding the spanwise intersection of the fillet and wing constant. The data were obtained at Mach numbers of 0.35 to 1.2 and at Reynolds number (depending on Mach number) of 1.9 million to 2.11 million per foot. The angle of attack was varied from about minus 2 deg to 22 deg at 0 deg of sideslip.

  3. Longitudinal aerodynamic and propulsion characteristics of a propulsive-wing V/STOL model at high subsonic speeds

    NASA Technical Reports Server (NTRS)

    Salters, L. B., Jr.; Schmeer, J. W.

    1973-01-01

    The aerodynamic and propulsion characteristics of a 1/6-scale propulsive-wing V/STOL air-powered model was investigated over the Mach number range from 0.40 to 0.96 and at angles of attack from -5 deg to 15 deg for several fan rotational speeds. Three fanduct-exit configurations were tested, including two exit areas. The model with 25-percent-thick wing had a drag-rise Mach number of 0.85, which is typical of aircraft with thinner, conventional, unswept wings.

  4. Aerodynamic Characteristics at High Speeds of Related Full-Scale Propellers Having Different Blade-Section Cambers

    NASA Technical Reports Server (NTRS)

    Maynard, Julian D; Salters, Leland B , Jr

    1957-01-01

    Wind-tunnel tests of a full-scale two-blade NACA 10-(10)(08)-03 (high camber) propeller have been made for a range of blade angles from 20 degrees to 55 degrees at airspeeds up to 500 miles per hour. The results of these tests have been compared with results from previous tests of the NACA 10-(3) (08)-03 (low camber) and NACA 10-(5)(08)-03 (medium camber) propellers to evaluate the effects of blade-section camber on propeller aerodynamic characteristics.

  5. Aerodynamic Characteristics of a Slender Cone-cylinder Body of Revolution at a Mach Number of 3.85

    NASA Technical Reports Server (NTRS)

    Jack, John R

    1951-01-01

    An experimental investigation of the aerodynamics of a slender cone-cylinder body of revolution was conducted at a Mach number of 3.85 for angles of attack of 0 degree to 10 degrees and a Reynolds number of 3.85x10(exp 6). Boundary-layer measurements at zero angle of attack are compared with the compressible-flow formulations for predicting laminar boundary-layer characteristics. Comparison of experimental pressure and force values with theoretical values showed relatively good agreement for small angles of attack. The measured mean skin-friction coefficients agreed well with theoretical values obtained for laminar flow over cones.

  6. Aerodynamic characteristics of a series of single-inlet air-breathing missile configurations

    NASA Technical Reports Server (NTRS)

    Hayes, C.

    1983-01-01

    A series of air-breathing missile configurations was investigated to provide a data base for the design of such missiles. The model could be configured with either a single axisymmetric or a two dimensional inlet located at the bottom of the body. Two tail configurations were investigated: a tri-tail and an X-tail. The tail surfaces could be deflected to provide pitch control. A wing could be located above the inlet on the center line of the model. Tests were made at supersonic Mach numbers with the inlet open and internal flow, and at subsonic-transonic Mach numbers with the internal duct closed and no internal flow.

  7. Transonic and Low-Supersonic Aeroelastic Analysis of a Two-Degree Airfoil with a Freeplay Non-Linearity

    NASA Astrophysics Data System (ADS)

    KIM, DONG-HYUN; LEE, IN

    2000-07-01

    A two-degree-of-freedom airfoil with a freeplay non-linearity in the pitch and plunge directions has been analyzed in the transonic and low-supersonic flow region, where aerodynamic non-linearities also exist. The primary purpose of this study is to show aeroelastic characteristics due to freeplay structural non-linearity in the transonic and low-supersonic regions. The unsteady aerodynamic forces on the airfoil were evaluated using two-dimensional unsteady Euler code, and the resulting aeroelastic equations are numerically integrated to obtain the aeroelastic time responses of the airfoil motions and to investigate the dynamic instability. The present model has been considered as a simple aeroelastic model, which is equivalent to the folding fin of an advanced generic missile. From the results of the present study, characteristics of important vibration responses and aeroelastic instabilities can be observed in the transonic and supersonic regions, especially considering the effect of structural non-linearity in the pitch and plunge directions. The regions of limit-cycle oscillation are shown at much lower velocities, especially in the supersonic flow region, than the divergent flutter velocities of the linear structure model. It is also shown that even small freeplay angles can lead to severe dynamic instabilities and dangerous fatigue conditions for the flight vehicle wings and control fins.

  8. Flight Test Determined Aerodynamics Force and Moment Characteristics of the X-43A Research Vehicle at Mach 7.0

    NASA Technical Reports Server (NTRS)

    Davis, Mark C.; White, J. Terry

    2006-01-01

    The second flight of the HYPER-X Program afforded a unique opportunity to determine the aerodynamic force and moment characteristics of an airframe integrated scramjet powered aircraft in hypersonic flight. These data were gathered via a repeated series of pitch, yaw, and roll doublets, frequency sweeps, and pull-up/push-over maneuvers performed throughout the X-43A cowl-closed descent phase. The subject flight research maneuvers were conducted in a Mach number range of 6.8 to 0.95 at altitudes from 92,000 ft to sea level. In this flight regime, the dynamic pressure varied from 1300 psf to 400 psf with angle-of-attack ranging from 0 deg to 14 deg. The flight-extracted aerodynamics were compared with pre-flight predictions based on wind tunnel test data. The X-43A flight-derived axial force was found to be 10 to 15 percent higher than prediction. Under-predictions of similar magnitude were observed for the normal force. For Mach numbers greater than 4, the X-43A flight-derived stability and control characteristics resulted in larger than predicted static margins, with the largest discrepancy approximately 5-inches forward along the X(CG) at Mach 6. This would result in less static margin in pitch. The X-43A predicted lateral-directional stability and control characteristics matched well with flight data when allowance was made for the high uncertainty in angle-of-sideslip.

  9. Flight-Test-Determined Aerodynamic Force and Moment Characteristics of the X-43A at Mach 7.0

    NASA Technical Reports Server (NTRS)

    Davis. Marl C.; White, J. Terry

    2006-01-01

    The second flight of the Hyper-X program afforded a unique opportunity to determine the aerodynamic force and moment characteristics of an airframe-integrated scramjet-powered aircraft in hypersonic flight. These data were gathered via a repeated series of pitch, yaw, and roll doublets; frequency sweeps; and pushover-pullup maneuvers performed throughout the X-43A cowl-closed descent. Maneuvers were conducted at Mach numbers of 6.80 to 0.95 and altitudes from 92,000 ft msl to sea level. The dynamic pressure varied from 1300 psf to 400 psf with the angle of attack ranging from 0 deg to 14 deg. The flight-extracted aerodynamics were compared with preflight predictions based on wind-tunnel-test data. The X-43A flight-derived axial force was found to be 10 percent to 15 percent higher than prediction. Under-predictions of similar magnitude were observed for the normal force. For Mach numbers above 4.0, the flight-derived stability and control characteristics resulted in larger-than-predicted static margins, with the largest discrepancy approximately 5 in. forward along the x-axis center of gravity at Mach 6.0. This condition would result in less static margin in pitch. The predicted lateral-directional stability and control characteristics matched well with flight data when allowance was made for the high uncertainty in angle of sideslip.

  10. X-43A Flight-Test-Determined Aerodynamic Force and Moment Characteristics at Mach 7.0

    NASA Technical Reports Server (NTRS)

    Davis, Mark C.; White, J. Terry

    2008-01-01

    The second flight of the Hyper-X program afforded a unique opportunity to determine the aerodynamic force and moment characteristics of an airframe-integrated scramjet-powered aircraft in hypersonic flight. These data were gathered via a repeated series of pitch, yaw, and roll doublets, frequency sweeps, and pushover-pullup maneuvers performed throughout the X-43A cowl-closed descent. Maneuvers were conducted at Mach numbers of 6.80-0.95 and at altitudes from 92,000 ft mean sea level to sea level. The dynamic pressure varied from 1300 to 400 psf with the angle of attack ranging from 0 to 14 deg. The flight-extracted aerodynamics were compared with preflight predictions based on wind-tunnel test data. The X-43A flight-derived axial force was found to be 10-15%higher than prediction. Underpredictions of similar magnitude were observed for the normal force. For Mach numbers above 4.0, the flight-derived stability and control characteristics resulted in larger-than-predicted static margins, with the largest discrepancy approximately 5 in. forward along the x-axis center of gravity at Mach 6.0. This condition would result in less static margin in pitch. The predicted lateral-directional stability and control characteristics matched well with flight data when allowance was made for the high uncertainty in angle of sideslip.

  11. Aerodynamic Characteristics of an Aerospace Vehicle During a Subsonic Pitch-Over Maneuver

    NASA Technical Reports Server (NTRS)

    Kleb, William L.

    1996-01-01

    Time-dependent CFD has been used to predict aerospace vehicle aerodynamics during a subsonic rotation maneuver. The inviscid 3D3U code is employed to solve the 3-D unsteady flow field using an unstructured grid of tetrahedra. As this application represents a challenge to time-dependent CFD, observations concerning spatial and temporal resolution are included. It is shown that even for a benign rotation rate, unsteady aerodynamic effects are significant during the maneuver. Possibly more significant, however, the rotation maneuver creates ow asymmetries leading to yawing moment, rolling moment, and side force which are not present in the quasi-steady case. A series of steady solutions at discrete points in the maneuver are also computed for comparison with wind tunnel measurements and as a means of quantifying unsteady effects.

  12. Modeling of aircraft unsteady aerodynamic characteristics. Part 2: Parameters estimated from wind tunnel data

    NASA Technical Reports Server (NTRS)

    Klein, Vladislav; Noderer, Keith D.

    1995-01-01

    Aerodynamic equations with unsteady effects were formulated for an aircraft in one-degree-of-freedom, small-amplitude, harmonic motion. These equations were used as a model for aerodynamic parameter estimation from wind tunnel oscillatory data. The estimation algorithm was based on nonlinear least squares and was applied in three examples to the oscillatory data in pitch and roll of 70 deg triangular wing and an X-31 model, and in-sideslip oscillatory data of the High Incidence Research Model 2 (HIRM 2). All three examples indicated that a model using a simple indicial function can explain unsteady effects observed in measured data. The accuracy of the estimated parameters and model verification were strongly influenced by the number of data points with respect to the number of unknown parameters.

  13. Plasma Influence on Characteristics of Aerodynamic Friction and Separation Flow Location

    DTIC Science & Technology

    2007-11-02

    Mirror- galvanometer oscillograph NO43.1 (6); • · Shadow Schlieren device IAB-451 (7-10). After a modernization the oscilloscope “Tektronix TDS...amplifier is transmitted to the mirror- galvanometer oscillograph (6), which in turn records the pressure variation diagram on a Plasma Aerodynamics...balance (4) is used, the signal from which is also transmitted to the 8-ANCh amplifier and then to the mirror- galvanometer oscillograph and is

  14. Experimental Hypersonic Aerodynamic Characteristics of the Space Shuttle Orbiter for a Range of Damage Scenarios

    NASA Technical Reports Server (NTRS)

    Brauckman, Gregory J.; Scallion, William I.

    2003-01-01

    Aerodynamic tests in support of the Columbia accident investigation were conducted in two hypersonic wind tunnels at the NASA Langley Research Center, the 20-Inch Mach 6 Air Tunnel and the 20-Inch Mach 6 CF4 Tunnel. The primary purpose of these tests was to measure the forces and moments generated by a variety of outer mold line alterations (damage scenarios) using 0.0075-scale models of the Space Shuttle Orbiter (approximately 10 inches in length). Simultaneously acquired global heat transfer mappings were obtained for a majority of the configurations tested. Test parameters include angles of attack from 38 to 42 deg, unit Reynolds numbers from 0.26 to 3.0 x10^6 per foot, and normal shock density ratios of 5 (Mach 6 air) and 12 (Mach 6 CF4). The damage scenarios evaluated included asymmetric boundary layer transition, gouges in the windward surface acreage thermal protection system tiles, wing leading edge damage (partially and fully missing reinforced carbon-carbon (RCC) panels), holes through the wing from the windward surface to the leeside, deformation of the wing windward surface, and main landing gear door and/or gear deployment. The aerodynamic data were compared to the magnitudes and directions observed in flight, and the heating images were evaluated in terms of the location of the generated disturbances and how these disturbance might relate to the response of discrete gages on the Columbia Orbiter vehicle during entry. The measured aerodynamic increments were generally small in magnitude, as were the flight-derived values during most of the entry. Asymmetric boundary layer transition (ABLT) results were consistent with the flight-derived Shuttle ABLT model, but not with the observed flight trends for STS-107. The partially missing leading edge panel results best matched both the early aerodynamic and heating trends observed in flight. A progressive damage scenario is presented that qualitatively matches the flight observations for the full entry.

  15. Aerodynamic and Flight Dynamic Characteristics of the New Family of 5. 56mm NATO Ammunition

    DTIC Science & Technology

    1985-10-01

    5 LIST OF TABLES .... 7 I. INTRODUCTION ........... ......, 9..,,.......,,.,.., .-. IeI TEST MATERIEL AND PROCEDURE...transonic and subsonic speeds. All Phase II aeroballistic tests were fired in the BRL Aerodynamics Range, using the same Phase I weapon mounting...shown in Figure 1. - The Phase III testing also used the weapon mounting system of Figure 1, but with the gun moved to one of the three firing positions

  16. The influence of vehicle aerodynamic and control response characteristics on driver-vehicle performance

    NASA Technical Reports Server (NTRS)

    Alexandridis, A. A.; Repa, B. S.; Wierwille, W. W.

    1978-01-01

    The effects of changes in understeer, control sensitivity, and location of the lateral aerodynamic center of pressure (c.p.) of a typical passenger car on the driver's opinion and on the performance of the driver-vehicle system were studied in a moving-base driving simulator. Twelve subjects with no prior experience on the simulator and no special driving skills performed regulation tasks in the presence of both random and step wind gusts.

  17. Combustion Characteristics of Liquid Normal Alkane Fuels in a Model Combustor of Supersonic Combustion Ramjet Engine

    NASA Astrophysics Data System (ADS)

    今村, 宰; 石川, 雄太; 鈴木, 俊介; 福本, 皓士郎; 西田, 俊介; 氏家, 康成; 津江, 光洋

    Effect of kinds of one-component n-alkane liquid fuels on combustion characteristics was investigated experimentally using a model combustor of scramjet engine. The inlet condition of a model combustor is 2.0 of Mach number, up to 2400K of total temperature, and 0.38MPa of total pressure. Five kinds of n-alkane are tested, of which carbon numbers are 7, 8, 10, 13, and 16. They are more chemically active and less volatile with an increase of alkane carbon number. Fuels are injected to the combustor in the upstream of cavity with barbotage nitrogen gas and self-ignition performance was investigated. The result shows that self-ignition occurs with less equivalence ratio when alkane carbon number is smaller. This indicates that physical characteristic of fuel, namely volatile of fuel, is dominant for self-ignition behavior. Effect on flame-holding performance is also examined with adding pilot hydrogen and combustion is kept after cutting off pilot hydrogen with the least equivalence ratio where alkane carbon number is from 8 to 10. These points are discussed qualitatively from the conflict effect of chemical and physical properties on alkane carbon number.

  18. Method and apparatus for starting supersonic compressors

    DOEpatents

    Lawlor, Shawn P

    2013-08-06

    A supersonic gas compressor with bleed gas collectors, and a method of starting the compressor. The compressor includes aerodynamic duct(s) situated for rotary movement in a casing. The aerodynamic duct(s) generate a plurality of oblique shock waves for efficiently compressing a gas at supersonic conditions. A convergent inlet is provided adjacent to a bleed gas collector, and during startup of the compressor, bypass gas is removed from the convergent inlet via the bleed gas collector, to enable supersonic shock stabilization. Once the oblique shocks are stabilized at a selected inlet relative Mach number and pressure ratio, the bleed of bypass gas from the convergent inlet via the bypass gas collectors is effectively eliminated.

  19. Aerodynamic Characteristics of a Flying-Boat Hull Having a Length-Beam Ratio of 15, TED No. NACA 2206

    NASA Technical Reports Server (NTRS)

    Riebe, John M.; Naeseth, Rodger L.

    1951-01-01

    An investigation was made in the Langley 300 MPH 7- by 10-foot tunnel to determine the aerodynamic characteristics of a flying-boat hull of a length-beam ratio of 15 in the presence of a wing. The investigation was an extension of previous tests made on hulls of length-beam ratios of 6, 9, and 12; these hulls were designed to have approximately the same hydrodynamic performance with respect to spray and resistance characteristics. Comparison with the previous investigation at lower length-beam ratios indicated a reduction in minimum drag coefficients of 0.0006 (10 peroent)with fixed transition when the length-beam ratio was extended from 12 to 15. As with the hulls of lower length-beam ratio, the drag reduction with a length-beam ratio of 15 occurred throughout the range of angle of attack tested and the angle of attack for minimum drag was in the range from 2deg to 3deg. Increasing the length-beam ratio from 12 to 15 reduced the hull longitudinal instability by an mount corresponding to an aerodynamic-center shift of about 1/2 percent of the mean aerodynamic chord of the hypothetical flying boat. At an angle of attack of 2deg, the value of the variation of yawing-moment coefficient with angle of yaw for a length-beam ratio of 15 was 0.00144, which was 0.00007 larger than the value for a length-beam ratio of 12.

  20. A Subsonic Wind-Tunnel Study to Determine the Buffet and Static Aerodynamic Characteristics of a Systematic Series of Wings. Phase 1

    NASA Technical Reports Server (NTRS)

    Ray, Edward J.; Taylor, Robert T.

    1968-01-01

    A wind-tunnel investigation has been conducted in the Langley High-Speed 7- by 10-Foot Tunnel to determine the buffet and static aerodynamic characteristics of a systematic wing series at Mach numbers ranging from 0.23 to 0.94. The results have indicated that for a given Mach number, the wings which display superior aerodynamic efficiency characteristics generally display the highest buffet free lift coefficient. The characteristics exhibited by the wings which were considered have indicated that correlations can be made between the onset of buffet and selected divergences in the static aerodynamic characteristics. Axial force has been found to be the most sensitive static component to the onset of buffeting.

  1. Mechanical characteristics of stability-bleed valves for a supersonic inlet. [for the YF-12 aircraft

    NASA Technical Reports Server (NTRS)

    Neiner, G. H.; Dustin, M. O.; Cole, G. L.

    1977-01-01

    Mechanical characteristics of a set of direct-operated relief valves used in a throat-bypass stability-bleed system designed for the YF-12 aircraft inlet are described. A comparison of data taken before and after the windtunnel tests (at room temperature) showed that both the effective spring rate and the piston friction had decreased during the wind tunnel tests. In neither the effective spring rate nor the piston friction was the magnitude of change great enough to cause significant impairment of overall system effectiveness. No major valve mechanical problems were encountered in any of the tests. During high temperature bench tests, piston frictional drag increased. The friction returned to its initial room temperature value when the stability-bleed valve was disassembled and reassembled. The problem might be solved by using a different material for the piston sleeve bearing and the piston rings.

  2. The linearized characteristics method and its application to practical nonlinear supersonic problems

    NASA Technical Reports Server (NTRS)

    Ferri, Antonio

    1952-01-01

    The methods of characteristics has been linearized by assuming that the flow field can be represented as a basic flow field determined by nonlinearized methods and a linearized superposed flow field that accounts for small changes of boundary conditions. The method has been applied to two-dimensional rotational flow where the basic flow is potential flow and to axially symmetric problems where conical flows have been used as the basic flows. In both cases the method allows the determination of the flow field to be simplified and the numerical work to be reduced to a few calculations. The calculations of axially symmetric flow can be simplified if tabulated values of some coefficients of the conical flow are obtained. The method has also been applied to slender bodies without symmetry and to some three-dimensional wing problems where two-dimensional flow can be used as the basic flow. Both problems were unsolved before in the approximation of nonlinear flow.

  3. Aerodynamics for the Mars Phoenix Entry Capsule

    NASA Technical Reports Server (NTRS)

    Edquist, Karl T.; Desai, Prasun N.; Schoenenberger, Mark

    2008-01-01

    Pre-flight aerodynamics data for the Mars Phoenix entry capsule are presented. The aerodynamic coefficients were generated as a function of total angle-of-attack and either Knudsen number, velocity, or Mach number, depending on the flight regime. The database was constructed using continuum flowfield computations and data from the Mars Exploration Rover and Viking programs. Hypersonic and supersonic static coefficients were derived from Navier-Stokes solutions on a pre-flight design trajectory. High-altitude data (free-molecular and transitional regimes) and dynamic pitch damping characteristics were taken from Mars Exploration Rover analysis and testing. Transonic static coefficients from Viking wind tunnel tests were used for capsule aerodynamics under the parachute. Static instabilities were predicted at two points along the reference trajectory and were verified by reconstructed flight data. During the hypersonic instability, the capsule was predicted to trim at angles as high as 2.5 deg with an on-axis center-of-gravity. Trim angles were predicted for off-nominal pitching moment (4.2 deg peak) and a 5 mm off-axis center-ofgravity (4.8 deg peak). Finally, hypersonic static coefficient sensitivities to atmospheric density were predicted to be within uncertainty bounds.

  4. An experimental and theoretical analysis of the aerodynamic characteristics of a biplane-winglet configuration. M.D. Thesis

    NASA Technical Reports Server (NTRS)

    Gall, P. D.

    1984-01-01

    Improving the aerodynamic characteristics of an airplane with respect to maximizing lift and minimizing induced and parasite drag are of primary importance in designing lighter, faster, and more efficient aircraft. Previous research has shown that a properly designed biplane wing system can perform superiorly to an equivalent monoplane system with regard to maximizing the lift-to-drag ratio and efficiency factor. Biplanes offer several potential advantages over equivalent monoplanes, such as a 60-percent reduction in weight, greater structural integrity, and increased roll response. The purpose of this research is to examine, both theoretically and experimentally, the possibility of further improving the aerodynamic characteristics of the biplanes configuration by adding winglets. Theoretical predictions were carried out utilizing vortex-lattice theory, which is a numerical method based on potential flow theory. Experimental data were obtained by testing a model in the Pennsylvania State University's subsonic wind tunnel at a Reynolds number of 510,000. The results showed that the addition of winglets improved the performance of the biplane with respect to increasing the lift-curve slope, increasing the maximum lift coefficient, increasing the efficiency factor, and decreasing the induced drag. A listing of the program is included in the Appendix.

  5. Aerodynamic Design Exploration for Reusable Launch Vehicle Using Genetic Algorithm with Navier Stokes Solver

    NASA Astrophysics Data System (ADS)

    Tatsukawa, Tomoaki; Nonomura, Taku; Oyama, Akira; Fujii, Kozo

    In this study, aerodynamic design exploration for reusable launch vehicle (RLV) is conducted using genetic algorithm with Navier-Stokes solver to understand the aerodynamic characteristics for various body configurations and find design information such as tradeoff information among objectives. The multi-objective aerodynamic design optimization for minimizing zero-lift drag at supersonic condition, maximizing maximum lift-to-drag ratio (L/D) at subsonic condition, maximizing maximum L/D at supersonic condition, and maximizing volume of shape is conducted for bi-conical shape RLV based on computational fluid dynamics (CFD). The total number of evaluation in multi-objective optimization is 400, and it is necessary for evaluating one body configuration to conduct 8 CFD runs. In total, 3200 CFD runs are conducted. The analysis of Pareto-optimal solutions shows that there are various trade-off relations among objectives clearly, and the analysis of flow fields shows that the shape for the minimum drag configuration is almost the same as that of the shape for the maximum L/D configuration at supersonic condition. The shape for the maximum L/D at subsonic condition obtains additional lift at the kink compared with the minimum drag configuration. It leads to enhancement of L/D.

  6. Numerical investigation of the aerodynamic and structural characteristics of a corrugated wing

    NASA Astrophysics Data System (ADS)

    Hord, Kyle

    Previous experimental studies on static, bio-inspired corrugated wings have shown that they produce favorable aerodynamic properties such as delayed stall compared to streamlined wings and flat plates at high Reynolds numbers (Re ≥ 4x104). The majority of studies have been carried out with scaled models of dragonfly forewings from the Aeshna Cyanea in either wind tunnels or water channels. In this thesis, the aerodynamics of a corrugated airfoil was studied using computational fluid dynamics methods at a low Reynolds number of 1000. Structural analysis was also performed using the commercial software SolidWorks 2009. The flow field is described by solving the incompressible Navier-Stokes equations on an overlapping grid using the pressure-Poisson method. The equations are discretized in space with second-order accurate central differences. Time integration is achieved through the second-order Crank-Nicolson implicit method. The complex vortex structures that form in the corrugated airfoil valleys and around the corrugated airfoil are studied in detail. Comparisons are made with experimental measurements from corrugated wings and also with simulations of a flat plate. Contrary to the studies at high Reynolds numbers, our study shows that at low Reynolds numbers the wing corrugation does not provide any aerodynamic benefit compared to a smoothed flat plate. Instead, the corrugated profile generates more pressure drag which is only partially offset by the reduction of friction drag, leading to more total drag than the flat plate. Structural analysis shows that the wing corrugation can increase the resistance to bending moments on the wing structure. A smoothed structure has to be three times thicker to provide the same stiffness. It was concluded the corrugated wing has the structural benefit to provide the same resistance to bending moments with a much reduced weight.

  7. Measured unsteady transonic aerodynamic characteristics of an elastic supercritical wing with an oscillating control surface

    NASA Technical Reports Server (NTRS)

    Seidel, D. A.; Sandford, M. C.; Eckstrom, C. V.

    1985-01-01

    Transonic steady and unsteady aerodynamic data were measured on a large elastic wing in the NASA Langley Transonic Dynamics Tunnel. The wing had a supercritical airfoil shape and a leading-edge sweepback of 28.8 deg. The wing was heavily instrumented to measure both static and dynamic pressures and deflections. A hydraulically driven outboard control surface was oscillated to generate unsteady airloads on the wing. Representative results from the wind tunnel tests are presented and discussed, and the unexpected occurrence of an unusual dynamic wing instability, which was sensitive to angle of attack, is reported.

  8. Aerodynamic characteristics of airfoils VI : continuation of reports nos. 93, 124, 182, 244, and 286

    NASA Technical Reports Server (NTRS)

    1930-01-01

    This collection of data on airfoils has been made from the published reports of a number of the leading aerodynamic laboratories of this country and Europe. The information which was originally expressed according to the different customs of the several laboratories is here presented in a uniform series of charts and tables suitable for use of designing engineers and for purposes of general reference. The authority for the results here presented is given as the name of the laboratory at which the experiments were conducted, with the size of the model, wind velocity, and year of test.

  9. Calculation of the aerodynamic characteristics of tapered wings with partial-span flaps

    NASA Technical Reports Server (NTRS)

    Person, Henry A; Anderson, Raymond F

    1939-01-01

    Factors derived from wing theory are presented. By means of these factors, the angle of zero lift, the lift-curve slope, the pitching moment, the aerodynamic-center position, and the induced drag of tapered wings with partial-span flaps may be calculated. The factors are given for wings of aspect ratios 6 and 10 , of taper ratios from 0.25 to 1.00, and with flaps of various length. An example is presented of the method of application of the factors. Fair agreement with experimental results is shown for two wings of different taper ratio having plain flaps of various spacing.

  10. Error Estimate of the Ares I Vehicle Longitudinal Aerodynamic Characteristics Based on Turbulent Navier-Stokes Analysis

    NASA Technical Reports Server (NTRS)

    Abdol-Hamid, Khaled S.; Ghaffari, Farhad

    2011-01-01

    Numerical predictions of the longitudinal aerodynamic characteristics for the Ares I class of vehicles, along with the associated error estimate derived from an iterative convergence grid refinement, are presented. Computational results are based on the unstructured grid, Reynolds-averaged Navier-Stokes flow solver USM3D, with an assumption that the flow is fully turbulent over the entire vehicle. This effort was designed to complement the prior computational activities conducted over the past five years in support of the Ares I Project with the emphasis on the vehicle s last design cycle designated as the A106 configuration. Due to a lack of flight data for this particular design s outer mold line, the initial vehicle s aerodynamic predictions and the associated error estimates were first assessed and validated against the available experimental data at representative wind tunnel flow conditions pertinent to the ascent phase of the trajectory without including any propulsion effects. Subsequently, the established procedures were then applied to obtain the longitudinal aerodynamic predictions at the selected flight flow conditions. Sample computed results and the correlations with the experimental measurements are presented. In addition, the present analysis includes the relevant data to highlight the balance between the prediction accuracy against the grid size and, thus, the corresponding computer resource requirements for the computations at both wind tunnel and flight flow conditions. NOTE: Some details have been removed from selected plots and figures in compliance with the sensitive but unclassified (SBU) restrictions. However, the content still conveys the merits of the technical approach and the relevant results.

  11. Longitudinal aerodynamic characteristics of a generic fighter model with a wing designed for sustained transonic maneuver conditions

    NASA Technical Reports Server (NTRS)

    Ferris, J. C.

    1986-01-01

    A wind-tunnel investigation was made to determine the longitudinal aerodynamic characteristics of a fixed-wing generic fighter model with a wing designed for sustained transonic maneuver conditions. The airfoil sections on the wing were designed with a two-dimensional nonlinear computer code, and the root and tip section were modified with a three-dimensional code. The wing geometric characteristics were as follows: a leading-edge sweep of 45 degrees, a taper ratio of 0.2141, an aspect ratio of 3.30, and a thickness ratio of 0.044. The model was investigated at Mach numbers from 0.600 to 1.200, at Reynolds numbers, based on the model reference length, from 2,560,000 to 3,970,000, and through a model angle-of-attack range from -5 to +18 degrees.

  12. Predicting aerodynamic characteristics of vortical flows on three-dimensional configurations using a surface-singularity panel method

    NASA Technical Reports Server (NTRS)

    Maskew, B.

    1983-01-01

    A general low-order surface-singularity panel method is used to predict the aerodynamic characteristics of a problem where a wing-tip vortex from one wing closely interacts with an aft mounted wing in a low Reynolds Number flow; i.e., 125,000. Nonlinear effects due to wake roll-up and the influence of the wings on the vortex path are included in the calculation by using a coupled iterative wake relaxation scheme. The interaction also affects the wing pressures and boundary layer characteristics: these effects are also considered using coupled integral boundary layer codes and preliminary calculations using free vortex sheet separation modelling are included. Calculated results are compared with water tunnel experimental data with generally remarkably good agreement.

  13. Experimental study of the effects of Reynolds number on high angle of attack aerodynamic characteristics of forebodies during rotary motion

    NASA Technical Reports Server (NTRS)

    Pauley, H.; Ralston, J.; Dickes, E.

    1995-01-01

    The National Aeronautics and Space Administration and the Defense Research Agency (United Kingdom) have ongoing experimental research programs in rotary-flow aerodynamics. A cooperative effort between the two agencies is currently underway to collect an extensive database for the development of high angle of attack computational methods to predict the effects of Reynolds number on the forebody flowfield at dynamic conditions, as well as to study the use of low Reynolds number data for the evaluation of high Reynolds number characteristics. Rotary balance experiments, including force and moment and surface pressure measurements, were conducted on circular and rectangular aftbodies with hemispherical and ogive noses at the Bedford and Farnborough wind tunnel facilities in the United Kingdom. The bodies were tested at 60 and 90 deg angle of attack for a wide range of Reynolds numbers in order to observe the effects of laminar, transitional, and turbulent flow separation on the forebody characteristics when rolling about the velocity vector.

  14. Low-speed aerodynamic characteristics of a 13 percent thick medium speed airfoil designed for general aviation applications

    NASA Technical Reports Server (NTRS)

    Mcghee, R. J.; Beasley, W. D.

    1979-01-01

    Wind tunnel tests were conducted to determine the low speed, two dimensional aerodynamic characteristics of a 13percent thick medium speed airfoil designed for general aviation applications. The results were compared with data for the 13 percent thick low speed airfoil. The tests were conducted over a Mach number range from 0.10 to 0.32, a chord Reynolds number range from 2.0 x 10 to the 6th power to 12.0 x 10 to the 6th power, and an angle of attack frange from about -8 deg to 10 deg. The objective of retaining good high-lift low speed characteristics for an airfoil designed to have good medium speed cruise performance was achieved.

  15. Low-speed aerodynamic characteristics of a 17-percent-thick medium speed airfoil designed for general aviation applications

    NASA Technical Reports Server (NTRS)

    Mcghee, R. J.; Beaseley, W. D.

    1980-01-01

    Wind tunnel tests were conducted to determine the low speed two dimensional aerodynamic characteristics of a 17 percent thick medium speed airfoil (MS(1)-0317) designed for general aviation applications. The results were compared with data for the 17 percent thick low speed airfoil (LS(1)-0417) and the 13 percent thick medium speed airfoil (MS(1)-0313). Theoretical predictions of the drag rise characteristics of this airfoil are also provided. The tests were conducted in the Langley low turbulence pressure tunnel over a Mach number range from 0.10 to 0.32, a chord Reynolds number range from 2 million to 12 million, and an angle of attack range from about -8 to 20 deg.

  16. QCSEE under-the-wing engine-wing-flap aerodynamic profile characteristics

    NASA Technical Reports Server (NTRS)

    Bloomer, H. E.; Samanich, N. E.

    1982-01-01

    As part of a broad-based NASA program to provide a technology base for future propulsion requirements for powered-lift aircraft, the Quiet, Clean, Short-Haul, Experimental Engine (QCSEE) program was begun by the Lewis Research Center in 1974. The initial buildup of the under-the-wing (UTW) engine was tested by the contractor at his test site. The UTW engine was delivered to Lewis in 1978 for further testing with wing and flap segments simulating an installation on a short-haul transport aircraft. The engine was also tested alone as an aid in identifying the various noise sources and their levels. As part of these tests the aerodynamic profiles at the exhaust nozzle and on the surfaces and in the wake of the wing-flap system were measured. This report documents, in plots and tabular form, the significant results from those tests. The results are presented as tabulations of aerodynamic data for all of the test points and as profiles of pressure, temperature, velocity, and normalized velocity and pressure for selected conditions. One of the main conclusions was that the measured flap surface temperatures were surprisingly low for both approach and takeoff flap settings.

  17. Aerodynamics of cyclist posture, bicycle and helmet characteristics in time trial stage.

    PubMed

    Chabroux, Vincent; Barelle, Caroline; Favier, Daniel

    2012-07-01

    The present work is focused on the aerodynamic study of different parameters, including both the posture of a cyclist's upper limbs and the saddle position, in time trial (TT) stages. The aerodynamic influence of a TT helmet large visor is also quantified as a function of the helmet inclination. Experiments conducted in a wind tunnel on nine professional cyclists provided drag force and frontal area measurements to determine the drag force coefficient. Data statistical analysis clearly shows that the hands positioning on shifters and the elbows joined together are significantly reducing the cyclist drag force. Concerning the saddle position, the drag force is shown to be significantly increased (about 3%) when the saddle is raised. The usual helmet inclination appears to be the inclination value minimizing the drag force. Moreover, the addition of a large visor on the helmet is shown to provide a drag coefficient reduction as a function of the helmet inclination. Present results indicate that variations in the TT cyclist posture, the saddle position and the helmet visor can produce a significant gain in time (up to 2.2%) during stages.

  18. Experimental Hypersonic Aerodynamic Characteristics of the Space Shuttle Orbiter for a Range of Damage Scenarios

    NASA Technical Reports Server (NTRS)

    Brauckmann, Gregory J.; Scallion, William I.

    2004-01-01

    Aerodynamic tests in support of the Columbia accident investigation were conducted in two hypersonic wind tunnels at the NASA Langley Research Center, the 20-Inch Mach 6 Air Tunnel and the 20-Inch CF4 Tunnel. The primary purpose of these tests was to measure the forces and moments generated by a variety of outer mold line alterations (damage scenarios) using 0.0075-scale models of the Space Shuttle Orbiter. Simultaneously acquired global heat transfer mappings were obtained for a majority of the configurations tested. Test parametrics included angles of attack from 38 to 42 deg, unit Reynolds numbers from 0.3 x 10(exp 6) to 3.0 x 10(exp 6) per foot, and normal shock density ratios of 5 (Mach 6 air) and 12 (CF4). The damage scenarios evaluated included asymmetric boundary layer transition, gouges in the windward surface thermal protection system tiles, wing leading edge damage (partially and fully missing reinforced carbon-carbon (RCC) panels), deformation of the wing windward surface, and main landing gear and/or door deployment. The measured aerodynamic increments for the damage scenarios examined were generally small in magnitude, as were the flight-derived values during most of the entry prior to loss of communication. A progressive damage scenario is presented that qualitatively matches the flight observations for the STS-107 entry.

  19. Effects of independent variation of Mach and Reynolds numbers on the low-speed aerodynamic characteristics of the NACA 0012 airfoil section

    NASA Technical Reports Server (NTRS)

    Ladson, Charles L.

    1988-01-01

    A comprehensive data base is given for the low speed aerodynamic characteristics of the NACA 0012 airfoil section. The Langley low-turbulence pressure tunnel is the facility used to obtain the data. Included in the report are the effects of Mach number and Reynolds number and transition fixing on the aerodynamic characteristics. Presented are also comparisons of some of the results with previously published data and with theoretical estimates. The Mach number varied from 0.05 to 0.36. The Reynolds number, based on model chord, varied from 3 x 10 to the 6th to 12 x 10 to the 6th power.

  20. Aerodynamic Characteristics of a Feathered Dinosaur Measured Using Physical Models. Effects of Form on Static Stability and Control Effectiveness

    PubMed Central

    Evangelista, Dennis; Cardona, Griselda; Guenther-Gleason, Eric; Huynh, Tony; Kwong, Austin; Marks, Dylan; Ray, Neil; Tisbe, Adrian; Tse, Kyle; Koehl, Mimi

    2014-01-01

    We report the effects of posture and morphology on the static aerodynamic stability and control effectiveness of physical models based on the feathered dinosaur, Microraptor gui, from the Cretaceous of China. Postures had similar lift and drag coefficients and were broadly similar when simplified metrics of gliding were considered, but they exhibited different stability characteristics depending on the position of the legs and the presence of feathers on the legs and the tail. Both stability and the function of appendages in generating maneuvering forces and torques changed as the glide angle or angle of attack were changed. These are significant because they represent an aerial environment that may have shifted during the evolution of directed aerial descent and other aerial behaviors. Certain movements were particularly effective (symmetric movements of the wings and tail in pitch, asymmetric wing movements, some tail movements). Other appendages altered their function from creating yaws at high angle of attack to rolls at low angle of attack, or reversed their function entirely. While M. gui lived after Archaeopteryx and likely represents a side experiment with feathered morphology, the general patterns of stability and control effectiveness suggested from the manipulations of forelimb, hindlimb and tail morphology here may help understand the evolution of flight control aerodynamics in vertebrates. Though these results rest on a single specimen, as further fossils with different morphologies are tested, the findings here could be applied in a phylogenetic context to reveal biomechanical constraints on extinct flyers arising from the need to maneuver. PMID:24454820

  1. Effects of forebody strakes and Mach number on overall aerodynamic characteristics of configuration with 55 deg cropped delta wing

    NASA Technical Reports Server (NTRS)

    Erickson, Gary E.; Rogers, Lawrence W.

    1992-01-01

    A wind tunnel data base was established for the effects of chine-like forebody strakes and Mach number on the longitudinal and lateral-directional characteristics of a generalized 55 degree cropped delta wing-fuselage-centerline vertical tail configuration. The testing was conducted in the 7- by 10-Foot Transonic Tunnel at the David Taylor Research Center at free-stream Mach numbers of 0.40 to 1.10 and Reynolds numbers based on the wing mean aerodynamic chord of 1.60 x 10(exp 6) to 2.59 x 10(exp 6). The best matrix included angles of attack from 0 degree to a maximum of 28 degree, angles of sidesip of 0, +5, and -5 degrees, and wing leading-edge flat deflection angles of 0 and 30 degrees. Key flow phenomena at subsonic and transonic conditions were identified by measuring off-body flow visualization with a laser screen technique. These phenomena included coexisting and interacting vortex flows and shock waves, vortex breakdown, vortex flow interactions with the vertical tail, and vortices induced by flow separation from the hinge line of the deflected wing flap. The flow mechanisms were correlated with the longitudinal and lateral-directional aerodynamic data trends.

  2. Predicted Aerodynamic Characteristics of a NACA 0015 Airfoil Having a 25% Integral-Type Trailing Edge Flap

    NASA Technical Reports Server (NTRS)

    Hassan, Ahmed

    1999-01-01

    Using the two-dimensional ARC2D Navier-Stokes flow solver analyses were conducted to predict the sectional aerodynamic characteristics of the flapped NACA-0015 airfoil section. To facilitate the analyses and the generation of the computational grids, the airfoil with the deflected trailing edge flap was treated as a single element airfoil with no allowance for a gap between the flap's leading edge and the base of the forward portion of the airfoil. Generation of the O-type computational grids was accomplished using the HYGRID hyperbolic grid generation program. Results were obtained for a wide range of Mach numbers, angles of attack and flap deflections. The predicted sectional lift, drag and pitching moment values for the airfoil were then cast in tabular format (C81) to be used in lifting-line helicopter rotor aerodynamic performance calculations. Similar were also generated for the flap. Mathematical expressions providing the variation of the sectional lift and pitching moment coefficients for the airfoil and for the flap as a function of flap chord length and flap deflection angle were derived within the context of thin airfoil theory. The airfoil's sectional drag coefficient were derived using the ARC2D drag predictions for equivalent two dimensional flow conditions.

  3. 1997 NASA High-Speed Research Program Aerodynamic Performance Workshop. Volume 1; Configuration Aerodynamics

    NASA Technical Reports Server (NTRS)

    Baize, Daniel G. (Editor)

    1999-01-01

    The High-Speed Research Program and NASA Langley Research Center sponsored the NASA High-Speed Research Program Aerodynamic Performance Workshop on February 25-28, 1997. The workshop was designed to bring together NASA and industry High-Speed Civil Transport (HSCT) Aerodynamic Performance technology development participants in area of Configuration Aerodynamics (transonic and supersonic cruise drag prediction and minimization), High-Lift, Flight Controls, Supersonic Laminar Flow Control, and Sonic Boom Prediction. The workshop objectives were to (1) report the progress and status of HSCT aerodyamic performance technology development; (2) disseminate this technology within the appropriate technical communities; and (3) promote synergy among the scientist and engineers working HSCT aerodynamics. In particular, single- and multi-point optimized HSCT configurations, HSCT high-lift system performance predictions, and HSCT Motion Simulator results were presented along with executive summaries for all the Aerodynamic Performance technology areas.

  4. Aerodynamic experimentation with ducted models as applied to hypersonic air-breathing vehicles

    NASA Astrophysics Data System (ADS)

    Goon'ko, Yu. P.

    A methodology of experimentation in high supersonic wind tunnels for studying aerodynamic characteristics of hypersonic flying vehicles powered by air-breathing engines is discussed. Investigations of such total aerodynamic forces as drag, lift and pitching moment at testing the models are implicit when the air flow through the model ducts is accomplished so that to provide the simulation of the external flow around the airplane and flow over the inlets, but the operating engines and, hence, the exhaust jets are not modeled. The methods used for testing such models are based on the measurement of duct stream parameters alongside with the balance measurement of aerodynamic forces acting on the models. In the tests, aerometric tools are used such as narrow metering nozzles (plugs), pitot and static pressure probes, stagnation temperature probes and pressure orifices in walls of the model duct. The aerometric data serve to determine the flow rate and momentum of the stream at the duct exit. The internal non-simulated forces of the model ducts are also determined using the conservation equations for energy, mass flow and momentum, and these forces are eliminated from the aerodynamic test results. The techniques of the said model testing have been well developed as applied to supersonic aircraft, however their application for hypersonic vehicles whose models are tested at high supersonic speeds, Mach number M∞>4, implies some specific features. In the present paper, the results of experimental and theoretical study of these features are discussed. Some experimental data on aerodynamics of hypersonic aircraft models received in methodological tests are also presented. The tunnel experiments have been carried out in the Mach number range M∞=2-6.

  5. Model aerodynamic test results for two variable cycle engine coannular exhaust systems at simulated takeoff and cruise conditions. Comprehensive data report. Volume 3: Graphical data book 1

    NASA Technical Reports Server (NTRS)

    Nelson, D. P.

    1981-01-01

    A graphical presentation of the aerodynamic data acquired during coannular nozzle performance wind tunnel tests is given. The graphical data consist of plots of nozzle gross thrust coefficient, fan nozzle discharge coefficient, and primary nozzle discharge coefficient. Normalized model component static pressure distributions are presented as a function of primary total pressure, fan total pressure, and ambient static pressure for selected operating conditions. In addition, the supersonic cruise configuration data include plots of nozzle efficiency and secondary-to-fan total pressure pumping characteristics. Supersonic and subsonic cruise data are given.

  6. Aerodynamic Properties Analysis of Rapid Prototyped Models Versus Conventional Machined Models

    NASA Technical Reports Server (NTRS)

    Springer, A.; Cooper, K.

    1998-01-01

    Initial studies of the aerodynamic characteristics of proposed launch vehicles can be made more accurately if lower cost, high fidelity aerodynamic models are available for wind tunnel testing early in the design phase. This paper discusses the results of a study undertaken at NASA's Marshall Space Flight Center to determine if four rapid prototyping methods using a variety of materials are suitable for the design and manufacturing of high speed wind tunnel models in direct testing applications. It also gives an analysis of whether these materials and processes are of sufficient strength and fidelity to withstand the testing environment. In addition to test data, costs and turn-around times for the various models are given. Based on the results of this study, it can be concluded that rapid prototyping models show promise in limited direct application for preliminary aerodynamic development studies at subsonic, transonic, and supersonic speeds.

  7. Unsteady aerodynamic analysis of space shuttle vehicles. Part 1: Summary report

    NASA Technical Reports Server (NTRS)

    Ericsson, L. E.; Reding, J. P.

    1973-01-01

    An analysis of the unsteady aerodynamics of space shuttle vehicles was performed. The results show that slender wing theory can be modified to give the potential flow static and dynamic characteristics over a large Mach number range from M = 0 to M 1. A semi-empirical analytic approximation is derived for the loads induced by the leading edge vortex; and it is shown that the developed analytic technique gives good prediction of experimentally determined steady and unsteady delta wing aerodynamics, including the effects of leading edge roundness. At supersonic speeds, attached leading edge flow is established and shock-induced flow separation effects become of concern. Analysis of experimental results for a variety of boost configurations led to a definition of the main features of the flow interference effects between orbiter (delta wing) and booster. The effects of control deflection on the unsteady aerodynamics of the delta-wing orbiter were also evaluated.

  8. Aerodynamic characterisation and trajectory simulations for the Ariane-5 booster recovery system

    NASA Astrophysics Data System (ADS)

    Meiboom, F. P.

    One of the most critical aspects of the early phases of the development of the Ariane-5 booster recovery system was the determination of the behavior of the booster during its atmospheric reentry, since this behavior determines the start conditions for the parachute system elements. A combination of wind-tunnel tests (subsonic and supersonic) and analytical methods was applied to define the aerodynamic characteristics of the booster. This aerodynamic characterization in combination with information of the ascent trajectory, atmospheric properties and booster mass and inertia were used as input for the 6-DOF trajectory simulations of the vehicle. Uncertainties in aerodynamic properties and deviations in atmospheric and booster properties were incorporated to define the range of initial conditions for the parachute system, utilizing stochastic (Monte-Carlo) methods.

  9. On the aerodynamic characteristics of hovering rigid and flexible hawkmoth-like wings

    NASA Astrophysics Data System (ADS)

    Lua, K. B.; Lai, K. C.; Lim, T. T.; Yeo, K. S.

    2010-12-01

    Insect wings are subjected to fluid, inertia and gravitational forces during flapping flight. Owing to their limited rigidity, they bent under the influence of these forces. Numerical study by Hamamoto et al. (Adv Robot 21(1-2):1-21, 2007) showed that a flexible wing is able to generate almost as much lift as a rigid wing during flapping. In this paper, we take a closer look at the relationship between wing flexibility (or stiffness) and aerodynamic force generation in flapping hovering flight. The experimental study was conducted in two stages. The first stage consisted of detailed force measurement and flow visualization of a rigid hawkmoth-like wing undergoing hovering hawkmoth flapping motion and simple harmonic flapping motion, with the aim of establishing a benchmark database for the second stage, which involved hawkmoth-like wing of different flexibility performing the same flapping motions. Hawkmoth motion was conducted at Re = 7,254 and reduced frequency of 0.26, while simple harmonic flapping motion at Re = 7,800 and 11,700, and reduced frequency of 0.25. Results show that aerodynamic force generation on the rigid wing is governed primarily by the combined effect of wing acceleration and leading edge vortex generated on the upper surface of the wing, while the remnants of the wake vortices generated from the previous stroke play only a minor role. Our results from the flexible wing study, while generally supportive of the finding by Hamamoto et al. (Adv Robot 21(1-2):1-21, 2007), also reveal the existence of a critical stiffness constant, below which lift coefficient deteriorates significantly. This finding suggests that although using flexible wing in micro air vehicle application may be beneficial in term of lightweight, too much flexibility can lead to deterioration in flapping performance in terms of aerodynamic force generation. The results further show that wings with stiffness constant above the critical value can deliver mean lift coefficient

  10. Tesseract supersonic business transport

    NASA Technical Reports Server (NTRS)

    Reshotko, Eli; Garbinski, Gary; Fellenstein, James; Botting, Mary; Hooper, Joan; Ryan, Michael; Struk, Peter; Taggart, Ben; Taillon, Maggie; Warzynski, Gary

    1992-01-01

    This year, the senior level Aerospace Design class at Case Western Reserve University developed a conceptual design of a supersonic business transport. Due to the growing trade between Asia and the United States, a transpacific range was chosen for the aircraft. A Mach number of 2.2 was chosen, too, because it provides reasonable block times and allows the use of a large range of materials without a need for active cooling. A payload of 2,500 lbs. was assumed corresponding to a complement of nine passengers and crew, plus some light cargo. With these general requirements set, the class was broken down into three groups. The aerodynamics of the aircraft were the responsibility of the first group. The second developed the propulsion system. The efforts of both the aerodynamics and propulsion groups were monitored and reviewed for weight considerations and structural feasibility by the third group. Integration of the design required considerable interaction between the groups in the final stages. The fuselage length of the final conceptual design was 107.0 ft, while the diameter of the fuselage was 7.6 ft. The delta wing design consisted of an aspect ratio of 1.9 with a wing span of 47.75 ft and mid-chord length of 61.0 ft. A SNECMA MCV 99 variable-cycle engine design was chosen for this aircraft.

  11. Tesseract: Supersonic business transport

    NASA Technical Reports Server (NTRS)

    Reshotko, Eli; Garbinski, Gary

    1992-01-01

    This year, the senior level Aerospace Design class at Case Western Reserve University developed a conceptual design of a supersonic business transport. Due to the growing trade between Asia and the United States, a transpacific range has been chosen for the aircraft. A Mach number of 2.2 was chosen too because it provides reasonable block times and allows the use of a large range of materials without a need for active cooling. A payload of 2500 lbs. has been assumed corresponding to a complement of nine (passengers and crew) plus some light cargo. With these general requirements set, the class was broken down into three groups. The aerodynamics of the aircraft were the responsibility of the first group. The second developed the propulsion system. The efforts of both the aerodynamics and propulsion groups were monitored and reviewed for weight considerations and structural feasibility by the third group. Integration of the design required considerable interaction between the groups in the final stages. The fuselage length of the final conceptual design was 107.0 ft. while the diameter of the fuselage was 7.6 ft. The delta wing design consisted of an aspect ratio of 1.9 with a wing span of 47.75 ft and midcord length of 61.0 ft. A SNEMCA MCV 99 variable-cycle engine design was chosen for this aircraft.

  12. HYSHOT-2 Aerodynamics

    NASA Astrophysics Data System (ADS)

    Cain, T.; Owen, R.; Walton, C.

    2005-02-01

    The scramjet flight test Hyshot-2, flew on the 30 July 2002. The programme, led by the University of Queensland, had the primary objective of obtaining supersonic combustion data in flight for comparison with measurements made in shock tunnels. QinetiQ was one of the sponsors, and also provided aerodynamic data and trajectory predictions for the ballistic re-entry of the spinning sounding rocket. The unconventional missile geometry created by the nose-mounted asymmetric-scramjet in conjunction with the high angle of attack during re-entry makes the problem interesting. This paper presents the wind tunnel measurements and aerodynamic calculations used as input for the trajectory prediction. Indirect comparison is made with data obtained in the Hyshot-2 flight using a 6 degree-of-freedom trajectory simulation.

  13. Study of the supersonic propeller

    NASA Technical Reports Server (NTRS)

    Fabri, Jean; Siestrunck, Raymond

    1953-01-01

    In this paper a propeller having all sections operating at supersonic speeds is designated a supersonic propeller regardless of flight speed. Analyses assume subsonic flight speeds but very high rotational speeds. A very elementary analysis of the efficiency of a jet-propeller system is presented. A propeller analysis based on conventional vortex blade element theory is presented and reduced to a single point method which leads to an expression for optimum advance ratio in terms of hub-tip diameter ratio and airfoil fineness ratio. An expression for propeller efficiency in terms of advance ratio, hub-tip diameter ratio, and airfoil thickness ratio is also presented. Use is made of theoretical airfoil characteristics at supersonic speeds. A study of blade section interference, blade shock and expansion fields, at supersonic section speeds is presented. An example taken indicates that an efficiency of seventy percent can be obtained with a propeller having a tip Mach number of 2.3.

  14. Handbook of Supersonic Aerodynamics. Volume 5

    DTIC Science & Technology

    1953-08-01

    Beattie - Bridgeman equation of state was used instead. It is 2V - / B bB A ( 5 0 1 1 2 T (1- T 1 + -- - - T (1500.111... Equation 1500.111-2) A a Beattie - Bridgeman coefficient (cf. Equation 1500.111-2) b a Beattie - Bridgeman coefficient (cf. Equation 1500.111-2) b ( 1... Bridgeman coefficient (cf. Equation 1500.111-2)0 c a Beattie - Bridgeman coefficient (cf. Equation 1500.111-2) P1 M2 c ( RT (cf.

  15. Handbook of Supersonic Aerodynamics Volume 1

    DTIC Science & Technology

    1950-04-01

    34 1 Q w .Hi TI TU 0 TITU 00 4-1-4-4- it i;tit .§ L_ ’ 4=* i...represents the possible series of states having a constant flow per unit area and a constant stagnation tem- pera tiu?e. Two Fanno lines are shown

  16. Analysis of the effect of engine characteristics on the external aerodynamics of STOL wing propulsion systems

    NASA Technical Reports Server (NTRS)

    Albers, J. A.

    1972-01-01

    The effects of engine presssure ratio, engine size, and engine location on the pressure distribution, lift coefficient, and flow field of a STOL wing propulsion system are presented. The flow variables of the engines are included in the two-dimensional potential flow analysis by considering the effects of mass flow coefficient at the engine inlet and thrust coefficient at the engine exit. A functional relation between these coefficients and engine pressure ratio is given. The results of this study indicate that the effect of engine pressure ratio on the external aerodynamics is a function of engine location. For engines located on the bottom of the wing, the highest pressure ratio engine resulted in the highest lift coefficient. For engines located on the top of the wing, the lowest pressure ratio engine resulted in the highest lift coefficient.

  17. Impact of pulsed blowing jet on aerodynamic characteristics of wind turbine airfoils

    NASA Astrophysics Data System (ADS)

    Bobonea, Andreea

    2012-11-01

    Wind turbine growth in size and weight made it impossible to control turbines passively as they were controlled in the past. Current efforts focus on increasing their aerodynamic efficiency and operational range through active flow control methods. One of the main methods of active flow control is the usage of blowing devices with constant or pulsed jets. By adding stored high-momentum air through slots into the boundary layer, they overcome adverse pressure gradients and postpone separation. Pulsed blowing sends short pulses rather than a continuous jet of fluid into the boundary layer and has been found to be more effective. Through CFD simulations over a 2D wind turbine airfoil, this research highlights the impact of different slot geometries with constant/pulsed blowing, on the effectiveness of this active flow control technique.

  18. Space shuttle: Aerodynamic characteristics of a 162-inch diameter solid rocket booster with and without strakes

    NASA Technical Reports Server (NTRS)

    Johnson, J. D.; Radford, W. D.; Rampy, J. M.

    1973-01-01

    Tests conducted at NASA-Langley have shown that a small flap or strake can generate a significant amount of lift on a circular cylinder with large cross flow. If strakes are placed on the opposite sides and ends on a circular body, a moment will be produced about the center of mass of the body. The purpose of this test was to determine the static-aerodynamic forces and moments of a 162-inch diameter SRB (PRR) with and without strakes. The total angle-of-attack range of the SRB test was from -10 to 190 degrees. Model roll angles were 0, 45, 90, and 135 degrees with some intermediate angles. The Mach range was from 0.6 to 3.48. The 0.00494 scale model was designated as MSFC No. 449.

  19. DSMC method on aerodynamic heating and temperature characteristic of hypersonic rarefied flows

    NASA Astrophysics Data System (ADS)

    Ma, Jing; Bao, Xingdong; Mao, Hongxia; Dong, Yanbing

    2016-10-01

    Aerodynamic heating is one of important factors affecting hypersonic aircraft design. The Direct Simulation Monte Carlo method (DSMC) has evolved years into a powerful numerical technique for the computation of complex, non-equilibrium gas flows. In atmospheric target, non-equilibrium conditions occur at high altitude and in regions of flow fields with small length scales. In this paper, the theoretical basis of the DSMC technique is discussed. In addition, the methods used in DSMC are described for simulation of high temperature, real gas effects and gas-surface interactions. Combined with the solution of heat transfer in material, heat-flux distribution and temperature distribution of the different shape structures was calculated in rarefied conditions.

  20. Low Density Supersonic Decelerator (LDSD) Supersonic Flight Dynamics Test (SFDT) Plume Induced Environment Modelling

    NASA Technical Reports Server (NTRS)

    Mobley, B. L.; Smith, S. D.; Van Norman, J. W.; Muppidi, S.; Clark, I

    2016-01-01

    Provide plume induced heating (radiation & convection) predictions in support of the LDSD thermal design (pre-flight SFDT-1) Predict plume induced aerodynamics in support of flight dynamics, to achieve targeted freestream conditions to test supersonic deceleration technologies (post-flight SFDT-1, pre-flight SFDT-2)

  1. Reynolds Number Effects on a Supersonic Transport at Transonic Conditions

    NASA Technical Reports Server (NTRS)

    Wahls, R. N.; Owens, L. R.; Rivers, S. M. B.

    2001-01-01

    A High Speed Civil Transport configuration was tested in the National Transonic Facility at the NASA Langley Research Center as part of NASA's High Speed Research Program. The primary purposes of the tests were to assess Reynolds number scale effects and the high Reynolds number aerodynamic characteristics of a realistic, second generation supersonic transport while providing data for the assessment of computational methods. The tests included longitudinal and lateral/directional studies at low speed high-lift and transonic conditions across a range of Reynolds numbers from that available in conventional wind tunnels to near flight conditions. Results are presented which focus on both the Reynolds number and static aeroelastic sensitivities of longitudinal characteristics at Mach 0.90 for a configuration without an empennage.

  2. High transonic speed transport aircraft study. [aerodynamic characteristics of single-fuselage, yawed-wing configuration

    NASA Technical Reports Server (NTRS)

    Kulfan, R. M.; Neumann, F. D.; Nisbet, J. W.; Mulally, A. R.; Murakami, J. K.; Noble, E. C.; Mcbarron, J. P.; Stalter, J. L.; Gimmestad, D. W.; Sussman, M. B.

    1973-01-01

    An initial design study of high-transonic-speed transport aircraft has been completed. Five different design concepts were developed. These included fixed swept wing, variable-sweep wing, delta wing, double-fuselage yawed-wing, and single-fuselage yawed-wing aircraft. The boomless supersonic design objectives of range=5560 Km (3000 nmi), payload-18 143 kg (40 000lb), Mach=1.2, and FAR Part 36 aircraft noise levels were achieved by the single-fuselage yawed-wing configuration with a gross weight of 211 828 Kg (467 000 lb). A noise level of 15 EPNdB below FAR Part 36 requirements was obtained with a gross weight increase to 226 796 Kg (500 000 lb). Although wing aeroelastic divergence was a primary design consideration for the yawed-wing concepts, the graphite-epoxy wings of this study were designed by critical gust and maneuver loads rather than by divergence requirements. The transonic nacelle drag is shown to be very sensitive to the nacelle installation. A six-degree-of-freedom dynamic stability analysis indicated that the control coordination and stability augmentation system would require more development than for a symmetrical airplane but is entirely feasible. A three-phase development plan is recommended to establish the full potential of the yawed-wing concept.

  3. Space shuttle orbiter trimmed center-of-gravity extension study. Volume 5: Effects of configuration modifications on the aerodynamic characteristics of the 140A/B orbiter at Mach numbers of 2.5, 3.95 and 4.6

    NASA Technical Reports Server (NTRS)

    Phillips, W. P.; Fournier, R. H.

    1979-01-01

    Supersonic aerodynamic characteristics are presented for the 140A/B space shuttle orbiter configuration (0.010 scale) and for the configuration modified to incorporate geometry changes in the wing planform fillet region. The modifications designed to extend the orbiter's longitudinal trim capability to more forward center-of-gravity locations, included reshaping of the baseline wing planform fillet and adding canards. The investigation was made in the high Mach number test section of the Langley Unitary Plan Wind Tunnel at a Reynolds number of approximately 2.2 million based on fuselage reference length. The angle-of-attack range for the investigation extended from -1 deg to 31 deg. Data were obtained with the elevators and body flap deflected at appropriate negative and positive conditions to assess the trim limits.

  4. The equivalent angle-of-attack method for estimating the nonlinear aerodynamic characteristics of missile wings and control surfaces

    NASA Technical Reports Server (NTRS)

    Hemsch, M. J.; Nielsen, J. N.

    1982-01-01

    A method has been developed for estimating the nonlinear aerodynamic characteristics of missile wing and control surfaces. The method is based on the following assumption: if a fin on a body has the same normal-force coefficient as a wing alone composed of two of the same fins joined together at their root chords, then the other force and moment coefficients of the fin and the wing alone are the same including the nonlinearities. The method can be used for deflected fins at arbitrary bank angles and at high angles of attack. In the paper, a full derivation of the method is given, its accuracy demonstrated and its use in extending missile data bases is shown.

  5. Effect of aileron deflections on the aerodynamic characteristics of a semispan model of a subsonic energy-efficient transport

    NASA Technical Reports Server (NTRS)

    Jacobs, P. F.

    1985-01-01

    An investigation was conducted in the Langley 8 Foot Transonic Pressure Tunnel to determine the effect of aileron deflections on the aerodynamic characteristics of a subsonic energy efficient transport (EET) model. The semispan model had an aspect ratio 10 supercritical wing and was configured with a conventionally located set of ailerons (i.e., a high speed aileron located inboard and a low speed aileron located outboard). Data for the model were taken over a Mach number range from 0.30 to 0.90 and an angle of attack range from approximately -2 deg to 10 deg. The Reynolds number was 2.5 million per foot for Mach number = 0.30 and 4 million per foot for the other Mach numbers. Model force and moment data, aileron effectiveness parameters, aileron hinge moment data, otherwise pressure distributions, and spanwise load data are presented.

  6. PROGRAM VSAERO: A computer program for calculating the non-linear aerodynamic characteristics of arbitrary configurations: User's manual

    NASA Technical Reports Server (NTRS)

    Maskew, B.

    1982-01-01

    VSAERO is a computer program used to predict the nonlinear aerodynamic characteristics of arbitrary three-dimensional configurations in subsonic flow. Nonlinear effects of vortex separation and vortex surface interaction are treated in an iterative wake-shape calculation procedure, while the effects of viscosity are treated in an iterative loop coupling potential-flow and integral boundary-layer calculations. The program employs a surface singularity panel method using quadrilateral panels on which doublet and source singularities are distributed in a piecewise constant form. This user's manual provides a brief overview of the mathematical model, instructions for configuration modeling and a description of the input and output data. A listing of a sample case is included.

  7. A computer program for calculating aerodynamic characteristics of low aspect-ratio wings with partial leading-edge separation

    NASA Technical Reports Server (NTRS)

    Mehrotra, S. C.; Lan, C. E.

    1978-01-01

    The necessary information for using a computer program to predict distributed and total aerodynamic characteristics for low aspect ratio wings with partial leading-edge separation is presented. The flow is assumed to be steady and inviscid. The wing boundary condition is formulated by the Quasi-Vortex-Lattice method. The leading edge separated vortices are represented by discrete free vortex elements which are aligned with the local velocity vector at midpoints to satisfy the force free condition. The wake behind the trailing edge is also force free. The flow tangency boundary condition is satisfied on the wing, including the leading and trailing edges. The program is restricted to delta wings with zero thickness and no camber. It is written in FORTRAN language and runs on CDC 6600 computer.

  8. Two-dimensional aerodynamic characteristics of several rotorcraft airfoils at Mach numbers from 0.35 to 0.90

    NASA Technical Reports Server (NTRS)

    Noonan, K. W.; Bingham, G. J.

    1977-01-01

    An investigation was conducted in the Langley 6- by 28-inch transonic tunnel and the 6- by 19-inch transonic tunnel to determine the two-dimensional aerodynamic characteristics of several rotorcraft airfoils at Mach numbers from 0.35 to 0.90. The airfoils differed in thickness, thickness distribution, and camber. The FX69-H-098, the BHC-540, and the NACA 0012 airfoils were investigated in the 6- by 28-inch tunnel at Reynolds numbers (based on chord) from about 4.7 to 9.3 million at the lowest and highest test Mach numbers respectively. The FX69-H-098, the NLR-1, the BHC-540, and the NACA 23012 airfoils were investigated in the 6- by 19-inch tunnel at Reynolds numbers from about 0.9 to 2.2 million at the lowest and highest test Mach numbers respectively.

  9. Effect of tail size reductions on longitudinal aerodynamic characteristics of a three surface F-15 model with nonaxisymmetric nozzles

    NASA Technical Reports Server (NTRS)

    Frassinelli, Mark C.; Carson, George T., Jr.

    1990-01-01

    An investigation was conducted in the Langley 16-Foot Transonic Tunnel to determine the effects of horizontal and vertical tail size reductions on the longitudinal aerodynamic characteristics of a modified F-15 model with canards and 2-D convergent-divergent nozzles. Quantifying the drag decrease at low angles of attack produced by tail size reductions was the primary focus. The model was tested at Mach numbers of 0.40, 0.90, and 1.20 over an angle of attack of -2 degree to 10 degree. The nozzle exhaust flow was simulated using high pressure air at nozzle pressure ratios varying from 1.0 (jet off) to 7.5. Data were obtained on the baseline configuration with and without tails as well as with reduced horizontal and/or vertical tail sizes that were 75, 50, and 25 percent of the baseline tail areas.

  10. Aerodynamic Characteristics of a 14-Percent-Thick NASA Supercritical Airfoil Designed for a Normal-Force Coefficient of 0.7

    NASA Technical Reports Server (NTRS)

    Harris, C. D.

    1975-01-01

    This report documents the experimental aerodynamic characteristics of a 14 percent thick supercritical airfoil based on an off design sonic pressure plateau criterion. The design normal force coefficient was 0.7. The results are compared with those of the family related 10 percent thick supercritical airfoil 33. Comparisons are also made between experimental and theoretical characteristics and composite drag rise characteristics derived for a full scale Reynolds number of 40 million.

  11. Aerodynamic characteristics of wings designed with a combined-theory method to cruise at a Mach number of 4.5

    NASA Technical Reports Server (NTRS)

    Mack, Robert J.

    1988-01-01

    A wind-tunnel study was conducted to determine the capability of a method combining linear theory and shock-expansion theory to design optimum camber surfaces for wings that will fly at high-supersonic/low-hypersonic speeds. Three force models (a flat-plate reference wing and two cambered and twisted wings) were used to obtain aerodynamic lift, drag, and pitching-moment data. A fourth pressure-orifice model was used to obtain surface-pressure data. All four wing models had the same planform, airfoil section, and centerbody area distribution. The design Mach number was 4.5, but data were also obtained at Mach numbers of 3.5 and 4.0. Results of these tests indicated that the use of airfoil thickness as a theoretical optimum, camber-surface design constraint did not improve the aerodynamic efficiency or performance of a wing as compared with a wing that was designed with a zero-thickness airfoil (linear-theory) constraint.

  12. Longitudinal aerodynamic characteristics of a wing-winglet model designed at M = 0.8, C sub L = 0.4 using linear aerodynamic theory

    NASA Technical Reports Server (NTRS)

    Kuhlman, J. M.

    1983-01-01

    Wind tunnel test results have been presented herein for a subsonic transport type wing fitted with winglets. Wind planform was chosen to be representative of wings used on current jet transport aircraft, while wing and winglet camber surfaces were designed using two different linear aerodynamic design methods. The purpose of the wind tunnel investigation was to determine the effectiveness of these linear aerodynamic design computer codes in designing a non-planar transport configuration which would cruise efficiently. The design lift coefficient was chosen to be 0.4, at a design Mach number of 0.8. Force and limited pressure data were obtained for the basic wing, and for the wing fitted with the two different winglet designs, at Mach numbers of 0.60, 0.70, 0.75 and 0.80 over an angle of attack range of -2 to +6 degrees, at zero sideslip. The data have been presented without analysis to expedite publication.

  13. The rebirth of supersonic transport

    NASA Technical Reports Server (NTRS)

    Rosen, Robert; Williams, Louis J.

    1993-01-01

    A development status evaluation is presented for the aerodynamic and propulsion system technologies that will be required to make a second-generation SST, or High Speed Civil Transport (HSCT) sufficiently (20 dB) quieter near airports and in supersonic cruise (sonic boom), less NO(x)-emitting, and longer-range than Concorde. NASA's Ames, Lewis, and Langley facilities are intensively concerned with the development of such an environmentally benign and economically feasible Mach 2.4 HSCT; efforts have concentrated on the design of a powerplant with the requisite specific fuel consumption, low emissions, and low noise.

  14. Investigation of the radiation brightness of gases around a burning model moving at supersonic velocity

    NASA Astrophysics Data System (ADS)

    Baulin, N. N.; Kuvalkin, D. G.; Piliugin, N. N.; Taganov, O. K.; Tikhomirov, S. G.

    1987-01-01

    Experimental results are presented on the ablation and shape change of burning models made of a pyrotechnic composition moving in air at supersonic velocity. A radiometer was used to measure the radiation brightness at a wavelength of 0.63 micron in the shock layer and wake of the burning models. The glow characteristics are determined as a function of the initial air pressure in the path of motion; and a theoretical model for the motion and ablation of burning bodies flying at supersonic velocity is developed which satisfactorily describes the experimental results. The present study is of interest in connection with the aerodynamic heating of vehicles flying at hypersonic velocity in planetary atmospheres.

  15. A Survey of Supersonic Retropropulsion Technology for Mars Entry, Descent, and Landing

    NASA Technical Reports Server (NTRS)

    Korzun, Ashley M.; Cruz, Juan R.; Braun, Robert D.

    2007-01-01

    This paper presents a literature survey on supersonic retropropulsion technology as it applies to Mars entry, descent, and landing (EDL). The relevance of this technology to the feasibility of Mars EDL is shown to increase with ballistic coefficient to the point that it is likely required for human Mars exploration. The use of retropropulsion to decelerate an entry vehicle from hypersonic or supersonic conditions to a subsonic velocity is the primary focus of this review. Discussed are systems-level studies, general flowfield characteristics, static aerodynamics, vehicle and flowfield stability considerations, and aerothermodynamics. The experimental and computational approaches used to develop retropropulsion technology are also reviewed. Finally, the applicability and limitations of the existing literature and current state-of-the-art computational tools to future missions are discussed in the context of human and robotic Mars exploration.

  16. Numerical Simulation of Laser-driven In-Tube Accelerator on Supersonic Condition

    NASA Astrophysics Data System (ADS)

    Kim, Sukyum; Jeung, In-Seuck; Choi, Jeong-Yeol

    2004-03-01

    Recently, several laser propulsion vehicles have been launched successfully. But these vehicles remained in a very low subsonic flight. Laser-driven In-Tube Accelerator (LITA) is developed as unique laser propulsion system at Tohoku University. In this paper, flow characteristics and momentum coupling coefficients are studied numerically in the supersonic condition with the same configuration of LITA. Because of the aerodynamic drag, the coupling coefficient could not get correctly especially at the low energy input. In this study, the coupling coefficient was calculated using the concept of the effective impulse.

  17. Numerical Simulation of Laser-driven In-Tube Accelerator on Supersonic Condition

    SciTech Connect

    Kim, Sukyum; Jeung, In-Seuck; Choi, Jeong-Yeol

    2004-03-30

    Recently, several laser propulsion vehicles have been launched successfully. But these vehicles remained in a very low subsonic flight. Laser-driven In-Tube Accelerator (LITA) is developed as unique laser propulsion system at Tohoku University. In this paper, flow characteristics and momentum coupling coefficients are studied numerically in the supersonic condition with the same configuration of LITA. Because of the aerodynamic drag, the coupling coefficient could not get correctly especially at the low energy input. In this study, the coupling coefficient was calculated using the concept of the effective impulse.

  18. Wind tunnel investigation of the aerodynamic characteristics of symmetrically deflected ailerons of the F-8C airplane. [conducted in the Langley 8-foot transonic pressure tunnel

    NASA Technical Reports Server (NTRS)

    Gera, J.

    1977-01-01

    A .042-scale model of the F-8C airplane was investigated in a transonic wind tunnel at high subsonic Mach numbers and a range of angles of attack between-3 and 20 degrees. The effect of symmetrically deflected ailerons on the longitudinal aerodynamic characteristics was measured. Some data were also obtained on the lateral control effectiveness of asymmetrically deflected horizontal tail surfaces.

  19. Development of the X-33 Aerodynamic Uncertainty Model

    NASA Technical Reports Server (NTRS)

    Cobleigh, Brent R.

    1998-01-01

    An aerodynamic uncertainty model for the X-33 single-stage-to-orbit demonstrator aircraft has been developed at NASA Dryden Flight Research Center. The model is based on comparisons of historical flight test estimates to preflight wind-tunnel and analysis code predictions of vehicle aerodynamics documented during six lifting-body aircraft and the Space Shuttle Orbiter flight programs. The lifting-body and Orbiter data were used to define an appropriate uncertainty magnitude in the subsonic and supersonic flight regions, and the Orbiter data were used to extend the database to hypersonic Mach numbers. The uncertainty data consist of increments or percentage variations in the important aerodynamic coefficients and derivatives as a function of Mach number along a nominal trajectory. The uncertainty models will be used to perform linear analysis of the X-33 flight control system and Monte Carlo mission simulation studies. Because the X-33 aerodynamic uncertainty model was developed exclusively using historical data rather than X-33 specific characteristics, the model may be useful for other lifting-body studies.

  20. Experimental Investigation on Airfoil Shock Control by Plasma Aerodynamic Actuation

    NASA Astrophysics Data System (ADS)

    Sun, Quan; Cheng, Bangqin; Li, Yinghong; Cui, Wei; Jin, Di; Li, Jun

    2013-11-01

    An experimental investigation on airfoil (NACA64—215) shock control is performed by plasma aerodynamic actuation in a supersonic tunnel (Ma = 2). The results of schlieren and pressure measurement show that when plasma aerodynamic actuation is applied, the position moves forward and the intensity of shock at the head of the airfoil weakens. With the increase in actuating voltage, the total pressure measured at the head of the airfoil increases, which means that the shock intensity decreases and the control effect increases. The best actuation effect is caused by upwind-direction actuation with a magnetic field, and then downwind-direction actuation with a magnetic field, while the control effect of aerodynamic actuation without a magnetic field is the most inconspicuous. The mean intensity of the normal shock at the head of the airfoil is relatively decreased by 16.33%, and the normal shock intensity is relatively reduced by 27.5% when 1000 V actuating voltage and upwind-direction actuation are applied with a magnetic field. This paper theoretically analyzes the Joule heating effect generated by DC discharge and the Lorentz force effect caused by the magnetic field. The discharge characteristics are compared for all kinds of actuation conditions to reveal the mechanism of shock control by plasma aerodynamic actuation.

  1. Applicability of commercial CFD tools for assessment of heavy vehicle aerodynamic characteristics.

    SciTech Connect

    Pointer, W. D.; Sofu, T.; Chang, J.; Weber, D.; Nuclear Engineering Division

    2008-12-01

    In preliminary validation studies, computational predictions from the commercial CFD codes Star-CD were compared with detailed velocity, pressure and force balance data from experiments completed in the 7 ft. by 10 ft. wind tunnel at NASA Ames using a Generic Conventional Model (GCM) that is representative of typical current-generation tractor-trailer geometries. Lessons learned from this validation study were then applied to the prediction of aerodynamic drag impacts associated with various changes to the GCM geometry, including the addition of trailer based drag reduction devices and modifications to the radiator and hood configuration. Add-on device studies have focused on ogive boat tails, with initial results indicating that a seven percent reduction in drag coefficient is easily achievable. Radiator and hood reconfiguration studies have focused on changing only the size of the radiator and angle of the hood components without changes to radii of curvature between the radiator grill and hood components. Initial results indicate that such changes lead to only modest changes in drag coefficient.

  2. Longitudinal aerodynamic characteristics of light, twin-engine, propeller-driven airplanes

    NASA Technical Reports Server (NTRS)

    Wolowicz, C. H.; Yancey, R. B.

    1972-01-01

    Representative state-of-the-art analytical procedures and design data for predicting the longitudinal static and dynamic stability and control characteristics of light, propeller-driven airplanes are presented. Procedures for predicting drag characteristics are also included. The procedures are applied to a twin-engine, propeller-driven airplane in the clean configuration from zero lift to stall conditions. The calculated characteristics are compared with wind-tunnel and flight data. Included in the comparisons are level-flight trim characteristics, period and damping of the short-period oscillatory mode, and windup-turn characteristics. All calculations are documented.

  3. Gas turbine engine with supersonic compressor

    SciTech Connect

    Roberts, II, William Byron; Lawlor, Shawn P.

    2015-10-20

    A gas turbine engine having a compressor section using blades on a rotor to deliver a gas at supersonic conditions to a stator. The stator includes one or more of aerodynamic ducts that have converging and diverging portions for deceleration of the gas to subsonic conditions and to deliver a high pressure gas to combustors. The aerodynamic ducts include structures for changing the effective contraction ratio to enable starting even when designed for high pressure ratios, and structures for boundary layer control. In an embodiment, aerodynamic ducts are provided having an aspect ratio of two to one (2:1) or more, when viewed in cross-section orthogonal to flow direction at an entrance to the aerodynamic duct.

  4. Aerodynamic characteristics of a feathered dinosaur measured using physical models. Effects of form on static stability and control effectiveness.

    PubMed

    Evangelista, Dennis; Cardona, Griselda; Guenther-Gleason, Eric; Huynh, Tony; Kwong, Austin; Marks, Dylan; Ray, Neil; Tisbe, Adrian; Tse, Kyle; Koehl, Mimi

    2014-01-01

    We report the effects of posture and morphology on the static aerodynamic stability and control effectiveness of physical models based on the feathered dinosaur, [Formula: see text]Microraptor gui, from the Cretaceous of China. Postures had similar lift and drag coefficients and were broadly similar when simplified metrics of gliding were considered, but they exhibited different stability characteristics depending on the position of the legs and the presence of feathers on the legs and the tail. Both stability and the function of appendages in generating maneuvering forces and torques changed as the glide angle or angle of attack were changed. These are significant because they represent an aerial environment that may have shifted during the evolution of directed aerial descent and other aerial behaviors. Certain movements were particularly effective (symmetric movements of the wings and tail in pitch, asymmetric wing movements, some tail movements). Other appendages altered their function from creating yaws at high angle of attack to rolls at low angle of attack, or reversed their function entirely. While [Formula: see text]M. gui lived after [Formula: see text]Archaeopteryx and likely represents a side experiment with feathered morphology, the general patterns of stability and control effectiveness suggested from the manipulations of forelimb, hindlimb and tail morphology here may help understand the evolution of flight control aerodynamics in vertebrates. Though these results rest on a single specimen, as further fossils with different morphologies are tested, the findings here could be applied in a phylogenetic context to reveal biomechanical constraints on extinct flyers arising from the need to maneuver.

  5. Transonic Aerodynamic Characteristics of Two Wedge Airfoil Sections Including Unsteady Flow Studies

    NASA Technical Reports Server (NTRS)

    Johnston, Patrick J.

    1959-01-01

    A two-dimensional wind-tunnel investigation has been conducted on a 20-percent-thick single-wedge airfoil section. Steady-state forces and moments were determined from pressure measurements at Mach numbers from 0.70 to about 1.25. Additional information on the flows about the single wedge is provided by means of instantaneous pressure measurements at Mach numbers up to unity. Pressure distributions were also obtained on a symmetrical double-wedge or diamond-shaped profile which had the same leading-edge included angle as the single-wedge airfoil. A comparison of the data on the two profiles to provide information on the effects of the afterbody showed that with the exception of drag, the single-wedge profile proved to be aerodynamically superior to the diamond profile in all respects. The lift effectiveness of the single-wedge airfoil section far exceeded that of conventional thin airfoil sections over the speed range of the investigation. Pitching-moment irregularities, caused by negative loadings near the trailing edge, generally associated with conventional airfoils of equivalent thicknesses were not exhibited by the single-wedge profile. Moderately high pulsating pressures existing over the base of the single-wedge airfoil section were significantly reduced as the Mach number was increased beyond 0.92 and the boundaries of the dead airspace at the base of the model converged to eliminate the vortex street in the wake. Increasing the leading-edge radius from 0 to 1 percent of the chord had a minor effect on the steady-state forces and generally raised the level of pressure pulsations over the forward part of the single-wedge profile.

  6. Aerodynamic Characteristics of SC1095 and SC1094 R8 Airfoils

    DTIC Science & Technology

    2003-12-01

    Development, and Engineering Command Ames Research Center Moffett Field, California December 2003 National Aeronautics and Space Administration Ames...60A ROTOR BLADE AND AIRFOILS ................................................................................... 2 EVALUATION OF SECTION CHARACTERISTICS...Characteristics of SC1095 and SC1094 R8 Airfoils WILLIAM G. BOUSMAN Aeroflightdynamics Directorate U.S. Army Research, Development, and Engineering Command Ames

  7. An aerodynamic tradeoff study of the scissor wing configuration

    NASA Technical Reports Server (NTRS)

    Selberg, Bruce P.; Rokhsaz, Kamran; Housh, Clinton S.

    1990-01-01

    A scissor wing configuration, consisting of two independently sweeping wings was numerically studied. This configuration was also compared with an equivalent fixed wing baseline. Aerodynamic and stability and control characteristics of these geometries were investigated over a wide range of flight Mach numbers. It is demonstrated that in the purely subsonic flight regime, the scissor wing can achieve higher aerodynamic efficiency as the result of slightly higher aspect ratio. In the transonic regime, the lift to drag ratio of the scissor wing is shown to be higher than that of the baseline, for higer values of the lift coefficient. This tends to make the scissor wing more efficient during transonic cruise at high altitudes as well as during air combat at all altitudes. In supersonic flight, where the wings are maintained at maximum sweep angle, the scissor wing is shown to have a decided advantage in terms of reduced wave drag. From the view point of stability and control, the scissor wing is shown to have distinct advantages. It is shown that this geometry can maintain a constant static margin in supersonic as well as subsonic flight, by proper sweep scheduling. Furthermore, it is demonstrated that addition of wing mounted elevons can greatly enhance control authority in pitch and roll.

  8. Aerodynamic beam generator for large particles

    DOEpatents

    Brockmann, John E.; Torczynski, John R.; Dykhuizen, Ronald C.; Neiser, Richard A.; Smith, Mark F.

    2002-01-01

    A new type of aerodynamic particle beam generator is disclosed. This generator produces a tightly focused beam of large material particles at velocities ranging from a few feet per second to supersonic speeds, depending on the exact configuration and operating conditions. Such generators are of particular interest for use in additive fabrication techniques.

  9. Low-speed aerodynamic characteristics of a 16-percent-thick variable-geometry airfoil designed for general aviation applications

    NASA Technical Reports Server (NTRS)

    Barnwell, R. W.; Noonan, K. W.; Mcghee, R. J.

    1978-01-01

    Tests were conducted in the Langley low-turbulence pressure tunnel to determine the aerodynamic characteristics of climb, cruise, and landing configurations. These tests were conducted over a Mach number range from 0.10 to 0.35, a chord Reynolds number range from 2.0 x 1 million to 20.0 x 1 million, and an angle-of-attack range from -8 deg to 20 deg. Results show that the maximum section lift coefficients increased in the Reynolds number range from 2.0 x 1 million to 9.0 x 1 million and reached values of approximately 2.1, 1.8, and 1.5 for the landing, climb, and cruise configurations, respectively. Stall characteristics, although of the trailing-edge type, were abrupt. The section lift-drag ratio of the climb configuration with fixed transition near the leading edge was about 78 at a lift coefficient of 0.9, a Mach number of 0.15, and a Reynolds number of 4.0 x 1 million. Design lift coefficients of 0.9 and 0.4 for the climb and cruise configurations were obtained at the same angle of attack, about 6 deg, as intended. Good agreement was obtained between experimental results and the predictions of a viscous, attached-flow theoretical method.

  10. Effect of Ground Proximity on the Aerodynamic Characteristics of Aspect-Ratio-1 Airfoils With and Without End Plates

    NASA Technical Reports Server (NTRS)

    Carter, Arthur W.

    1961-01-01

    An investigation has been made to determine the effect of ground proximity on the aerodynamic characteristics of aspect-ratio-1 airfoils. The investigation was made with the model moving over the water in a towing tank in order to eliminate the effects of wind-tunnel walls and of boundary layer on ground boards at small ground clearances. The results indicated that, as the ground was approached, the airfoils experienced an increase in lift-curve slope and a reduction in induced drag; thus, lift-drag ratio was increased. As the ground was approached, the profile drag remained essentially constant for each airfoil. Near the ground, the addition of end plates to the airfoil resulted in a large increase in lift-drag ratio. The lift characteristics of the airfoils indicated stability of height at positive angles of attack and instability of height at negative angles; therefore, the operating range of angles of attack would be limited to positive values. At positive angles of attack, the static longitudinal stability was increased as the height above the ground was reduced. Comparison of the experimental data with Wieselsberger's ground-effect theory (NACA Technical Memorandum 77) indicated generally good agreement between experiment and theory for the airfoils without end plates.

  11. Two-dimensional aerodynamic characteristics of several polygon-shaped cross-sectional models applicable to helicopter fuselages

    NASA Technical Reports Server (NTRS)

    Kelley, Henry L.; Crowell, Cynthia A.; Wilson, John C.

    1992-01-01

    A wind-tunnel investigation was conducted to determine 2-D aerodynamic characteristics of nine polygon-shaped models applicable to helicopter fuselages. The models varied from 1/2 to 1/5 scale and were nominally triangular, diamond, and rectangular in shape. Side force and normal force were obtained at increments of angle of flow incidence from -45 to 90 degrees. The data were compared with results from a baseline UH-60 tail-boom cross-section model. The results indicate that the overall shapes of the plots of normal force and side force were similar to the characteristic shape of the baseline data; however, there were important differences in magnitude. At a flow incidence of 0 degrees, larger values of normal force for the polygon models indicate an increase in fuselage down load of 1 to 2.5 percent of main-rotor thrust compared with the baseline value. Also, potential was indicated among some of the configurations to produce high fuselage side forces and yawing moments compared with the baseline model.

  12. Transonic wind tunnel test of a supersonic nozzle installation

    NASA Technical Reports Server (NTRS)

    Yetter, J. A.; Evelyn, G. B.; Mercer, C.

    1982-01-01

    The design of the propulsion system installation affects strongly the total drag and overall performance of an aircraft, and the concept, placement, and integration details of the exhaust nozzle are major considerations in the configuration definition. As part of the NASA Supersonic Cruise Research (SCR) program, a wind tunnel test program has been conducted to investigate exhaust nozzle-airframe interactions at transonic speeds. First phase testing is to establish guidelines for follow-on testing. A summary is provided of the results of first phase testing, taking into account the test approach, the effect of nozzle closure on aircraft aerodynamic characteristics, nozzle installation effects and nacelle interference drag, and an analytical study of the effects of nozzle closure on the aircraft.

  13. Development of a Parachute System for Deceleration of Flying Vehicles in Supersonic Regimes

    NASA Astrophysics Data System (ADS)

    Pilyugin, N. N.; Khlebnikov, V. S.

    2010-09-01

    Aerodynamic problems arising during design and development of braking systems for re-entry vehicles are analyzed. Aerodynamic phenomena and laws valid in a supersonic flow around a pair of bodies having different shapes are studied. Results of this research can be used in solving application problems (arrangement and optimization of experiments; design and development of various braking systems for re-entry vehicles moving with supersonic speeds in the atmosphere).

  14. The aerodynamic characteristics of eight very thick airfoils from tests in the variable density wind tunnel

    NASA Technical Reports Server (NTRS)

    Jacobs, Eastman N

    1932-01-01

    Report presents the results of wind tunnel tests on a group of eight very thick airfoils having sections of the same thickness as those used near the roots of tapered airfoils. The tests were made to study certain discontinuities in the characteristic curves that have been obtained from previous tests of these airfoils, and to compare the characteristics of the different sections at values of the Reynolds number comparable with those attained in flight. The discontinuities were found to disappear as the Reynolds number was increased. The results obtained from the large-scale airfoil, a symmetrical airfoil having a thickness ratio of 21 per cent, has the best general characteristics.

  15. Flight Dynamics of an Aeroshell Using an Attached Inflatable Aerodynamic Decelerator

    NASA Technical Reports Server (NTRS)

    Cruz, Juan R.; Schoenenberger, Mark; Axdahl, Erik; Wilhite, Alan

    2009-01-01

    An aeroelastic analysis of the behavior of an entry vehicle utilizing an attached inflatable aerodynamic decelerator during supersonic flight is presented. The analysis consists of a planar, four degree of freedom simulation. The aeroshell and the IAD are assumed to be separate, rigid bodies connected with a spring-damper at an interface point constraining the relative motion of the two bodies. Aerodynamic forces and moments are modeled using modified Newtonian aerodynamics. The analysis includes the contribution of static aerodynamic forces and moments as well as pitch damping. Two cases are considered in the analysis: constant velocity flight and planar free flight. For the constant velocity and free flight cases with neutral pitch damping, configurations with highly-stiff interfaces exhibit statically stable but dynamically unstable aeroshell angle of attack. Moderately stiff interfaces exhibit static and dynamic stability of aeroshell angle of attack due to damping induced by the pitch angle rate lag between the aeroshell and IAD. For the free-flight case, low values of both the interface stiffness and damping cause divergence of the aeroshell angle of attack due to the offset of the IAD drag force with respect to the aeroshell center of mass. The presence of dynamic aerodynamic moments was found to influence the stability characteristics of the vehicle. The effect of gravity on the aeroshell angle of attack stability characteristics was determined to be negligible for the cases investigated.

  16. Vortex Flows at Supersonic Speeds

    NASA Technical Reports Server (NTRS)

    Wood, Richard M.; Wilcox, Floyd J., Jr.; Bauer, Steven X. S.; Allen, Jerry M.

    2003-01-01

    A review of research conducted at the National Aeronautics and Space Administration (NASA) Langley Research Center (LaRC) into high-speed vortex flows during the 1970s, 1980s, and 1990s is presented. The data are for flat plates, cavities, bodies, missiles, wings, and aircraft with Mach numbers of 1.5 to 4.6. Data are presented to show the types of vortex structures that occur at supersonic speeds and the impact of these flow structures on vehicle performance and control. The data show the presence of both small- and large-scale vortex structures for a variety of vehicles, from missiles to transports. For cavities, the data show very complex multiple vortex structures exist at all combinations of cavity depth to length ratios and Mach number. The data for missiles show the existence of very strong interference effects between body and/or fin vortices. Data are shown that highlight the effect of leading-edge sweep, leading-edge bluntness, wing thickness, location of maximum thickness, and camber on the aerodynamics of and flow over delta wings. Finally, a discussion of a design approach for wings that use vortex flows for improved aerodynamic performance at supersonic speeds is presented.

  17. Aerodynamic Characteristics of NACA 23012 and 23021 Airfoils with 20-Percent-chord External-Airfoil Flaps of NACA 23012 Section

    NASA Technical Reports Server (NTRS)

    Platt, Robert C; Abbott, Ira H

    1937-01-01

    Report presents the results of an investigation of the general aerodynamic characteristics of the NACA 23012 and 23021 airfoils, each equipped with a 0.20c external flap of NACA 23012 section. The tests were made in the NACA 7 by 10-foot and variable-density wind tunnels and covered a range of Reynolds numbers that included values corresponding to those for landing conditions of a wide range of airplanes. Besides a determination of the variation of lift and drag characteristics with position of the flap relative to the main airfoil, complete aerodynamic characteristics of the airfoil-flap combination with a flap hinge axis selected to give small hinge moments were measured in the two tunnels. Some measurements of air loads on the flap itself in the presence of the wing were made in the 7 by 10-foot wind tunnel.

  18. Supersonic Retropropulsion Flight Test Concepts

    NASA Technical Reports Server (NTRS)

    Post, Ethan A.; Dupzyk, Ian C.; Korzun, Ashley M.; Dyakonov, Artem A.; Tanimoto, Rebekah L.; Edquist, Karl T.

    2011-01-01

    NASA's Exploration Technology Development and Demonstration Program has proposed plans for a series of three sub-scale flight tests at Earth for supersonic retropropulsion, a candidate decelerator technology for future, high-mass Mars missions. The first flight test in this series is intended to be a proof-of-concept test, demonstrating successful initiation and operation of supersonic retropropulsion at conditions that replicate the relevant physics of the aerodynamic-propulsive interactions expected in flight. Five sub-scale flight test article concepts, each designed for launch on sounding rockets, have been developed in consideration of this proof-of-concept flight test. Commercial, off-the-shelf components are utilized as much as possible in each concept. The design merits of the concepts are compared along with their predicted performance for a baseline trajectory. The results of a packaging study and performance-based trade studies indicate that a sounding rocket is a viable launch platform for this proof-of-concept test of supersonic retropropulsion.

  19. Supersonic wing design concepts employing nonlinear flows

    NASA Technical Reports Server (NTRS)

    Miller, D. S.

    1984-01-01

    Three nonlinear flow concepts for the design of supersonic wings are reviewed. The specific concepts are: leading-edge thrust, supercritical crossflow, and leading-edge vortex flow. The major results of the experimental-theoretical studies supporting the development of these concepts are presented and discussed. Also, supporting aerodynamic prediction methods are described and example applications are given. Recommendations for further development of each concept are made.

  20. Lateral-directional aerodynamic characteristics of light, twin-engine, propeller driven airplanes

    NASA Technical Reports Server (NTRS)

    Wolowicz, C. H.; Yancey, R. B.

    1972-01-01

    Analytical procedures and design data for predicting the lateral-directional static and dynamic stability and control characteristics of light, twin engine, propeller driven airplanes for propeller-off and power-on conditions are reported. Although the consideration of power effects is limited to twin engine airplanes, the propeller-off considerations are applicable to single engine airplanes as well. The procedures are applied to a twin engine, propeller driven, semi-low-wing airplane in the clean configuration through the linear lift range. The calculated derivative characteristics are compared with wind tunnel and flight data. Included in the calculated characteristics are the spiral mode, roll mode, and Dutch roll mode over the speed range of the airplane.