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At least 1,423 records · Page 79
All sourcesNASA NTRSDOE OSTI
NASA NTRS2002

Artificial Immune System Approaches for Aerospace Applications

Artificial Immune Systems (AIS) combine a priori knowledge with the adapting capabilities of biological immune system to provide a powerful alternative to currently available techniques for pattern recognition, modeling, design, and control. Immunology is the science of built-in defense mechanisms that are present in all living beings to protect against external attacks. A biological immune system can be thought of as a robust, adaptive system that is capable of dealing with an enormous variety of disturbances and uncertainties. Biological immune systems use a finite number of discrete "building blocks" to achieve this adaptiveness. These building blocks can be thought of as pieces of a puzzle which must be put together in a specific way-to neutralize, remove, or destroy each unique disturbance the system encounters. In this paper, we outline AIS models that are immediately applicable to aerospace problems and identify application areas that need further investigation.

KrishnaKumar, Kalmanje↗
NASA NTRS1993

Analysis of a high speed civil transport configuration at subsonic flow conditions using a Navier-Stokes solver

Computations of three dimensional vortical flows over a generic High Speed Civil Transport (HSCT) configuration with an aspect ratio of 3.04 are performed using a thin-layer Navier-Stokes solver. The HSCT cruise configuration is modeled without leading or trailing edge flap deflections and without engine nacelles. The flow conditions, which correspond to tests done in the NASA Langley 8-Foot Transonic Pressure Tunnel (TPT), are a subsonic Mach number of 0.3 and Reynolds number of 4.4 million for a range-of-attack (-.23 deg to 17.78 deg). The effects of the farfield boundary location with respect to the body are investigated. The boundary layer is assumed turbulent and simulated using an algebraic turbulence model. The key features of the vortices and their interactions are captured. Grid distribution in the vortex regions is critical for predicting the correct induced lift. Computed forces and surface pressures compare reasonably well with the experimental TPT data.

Lessard, Victor R.↗
NASA NTRS1943

Jet-boundary corrections for reflection-plane models in rectangular wind tunnel

A detailed method for determining the jet-boundary corrections for reflection-plane models in rectangular wind tunnels is presented. The method includes the determination of the tunnel span local distribution and the derivation of equations for the corrections to the angle of attack, the lift and drag coefficients, and the pitching-, rolling-, yawing-, and hinge-moment coefficients. The principle effects of aerodynamic induction and of the boundary-induced curvature of the streamlines have been considered. An example is included to illustrate the method. Numerical values of the more important corrections for reflection-plane models in 7 by 10-foot closed wind tunnels are presented.

Swanson, Robert S↗
NASA NTRS1993

Symmetry breaking in vortical flows over cones - Theory and numerical experiments

A stability analysis suggests that inviscid incompressible flow, independent from angle of attack and regardless whether attached or separated, over slender cones is only marginally stable in regions of decelerating circumferential flow. Reducing slenderness or surface curvature lowers the energy level of harmonic perturbations, and, thus, reduces their impact on the overall stability of flows over slender cones. Associating the notion of instabilities in such flows with the onset of vortex asymmetries provides a model for explaining a variety of flow phenomena in Navier-Stokes simulations of laminar incompressible flows over three right circular cones at moderate to high angles of attack.

Hartwich, Peter M.↗
NASA NTRS1994

Symmetry breaking in vortical flows over cones -- theory and numerical experiments

A stability analysis suggests that inviscid incompressible flow, independent from angle of attack and regardless whether attached or separated, over slender cones is only marginally stable in regions of decelerating circumferential flow. Reducing slenderness or surface curvature lowers the frequency spectrum of the harmonic perturbations and, thus, reduces their impact on the overall stability of flows over slender cones. Associating the notion of instabilities in such flows with the onset of vortex asymmetries provides a model for explaining a variety of flow phenomena in Navier-Stokes simulations of laminar incompressible flows over three right circular cones at moderate to high angles of attack.

Hartwich, Peter M.↗
NASA NTRS1974

Reentry aerodynamic characteristics of a space shuttle solid rocket booster model 449 tested in MSFC 14 by 14 inch TWT (SA26F)

Force tests of a 0.563 percent scale space shuttle solid rocket booster (SRB) model, MSFC Model 449, were conducted at the Marshall Space Flight Center 14 x 14 inch Trisonic Wind Tunnel. There were a total of 134 runs (pitch polars) made. Test Mach numbers were 0.6, 0.9, 1.2, 1.96, 2.74, 3.48, 4.00, 4.45, and 4.96; test angles of attack ranged from minus 10 degrees to 190 degrees; test Reynolds numbers ranged from 4.9 million per foot to 7.1 million per foot; and test roll angles were 0, 45, 90, and 135 degrees. The model was tested with three different engine nozzle/skirts. Two of these engine configurations differed from each other in the magnitude of the volume inside the nozzle and skirt. The third engine configuration had part of the nozzle removed. The model was tested with an electrical tunnel in combination with separation rockets of two different heights.

Johnson, J. D.↗
NASA NTRS1987

Comparison of wind tunnel and flight test afterbody and nozzle pressures for a twin-jet fighter aircraft at transonic speeds

Afterbody and nozzle pressures measured on a 1/12-scale model and in flight on a twin-jet fighter aircraft were compared as Mach number varied from 0.6 to 1.2, Reynolds number from 17.5 million to 302.5 million, and angle of attack from 1 to 7 deg. At Mach 0.6 and 0.8, nozzle pressure coefficient distributions and nozzle axial force coefficients agreed and showed good recompression. At Mach 0.9 and 1.2, flow complexity caused a loss in recompression for both flight and wind tunnel nozzle data. The flight data exhibited less negative values of pressure coefficient and lower axial force coefficients than did the wind tunnel data. Reynolds number effects were noted only at these Mach numbers. Jet temperature and mass flux ratio did not affect the comparisons of nozzle axial flow coefficient. At subsonic speeds, the levels of pressure coefficient distributions on the upper fuselage and lower nacelle surfaces for flight were less negative than those for the model. The model boundary layer thickness at the aft rake station exceeded that for the forward rake station and increased with increasing angle of attack. The flight boundary layer thickness at the aft rake station was less than that for the forward rake station and decreased with increasing angle of attack.

Nugent, Jack↗
NASA NTRS2009

Numerical Solutions for the CAWAPI Configuration on Structured Grids at NASA LaRC, United States

In this chapter numerical simulations of the flow around F-16XL are performed as a contribution to the Cranked Arrow Wing Aerodynamic Project International (CAWAPI) using the PAB3D CFD code. Two turbulence models are used in the calculations: a standard k-epsilon model, and the Shih-Zhu-Lumley (SZL) algebraic stress model. Seven flight conditions are simulated for the flow around the F-16XL where the free stream Mach number varies from 0.242 to 0.97. The range of angles of attack varies from 0 deg to 20 deg. Computational results, surface static pressure, boundary layer velocity profiles, and skin friction are presented and compared with flight data. Numerical results are generally in good agreement with flight data, considering that only one grid resolution is utilized for the different flight conditions simulated in this study. The Algebraic Stress Model (ASM) results are closer to the flight data than the k-epsilon model results. The ASM predicted a stronger primary vortex, however, the origin of the vortex and footprint is approximately the same as in the k-epsilon predictions.

Elmiligui, Alaa A.↗
NASA NTRS2017

Computational Fluid Dynamic Analyses for the High-Lift Common Research Model Using the USM3D and FUN3D Flow Solvers

Two Navier-Stokes codes were used to compute flow over the High-Lift Common Research Model (HL-CRM) in preparation for a wind tunnel test to be performed at the NASA Langley Research Center 14-by-22-Foot Subsonic Tunnel in fiscal year 2018. Both flight and wind tunnel conditions were simulated by the two codes at set Mach numbers and Reynolds numbers over a full angle-of-attack range for three configurations: cruise, landing and takeoff. Force curves, drag polars and surface pressure contour comparisons are shown for the two codes. The lift and drag curves compare well for the cruise configuration up to 10deg angle of attack but not as well for the other two configurations. The drag polars compare reasonably well for all three configurations. The surface pressure contours compare well for some of the conditions modeled but not as well for others.

Rivers, Melissa↗
NASA NTRS1988

Surface flow visualization of separated flows on the forebody of an F-18 aircraft and wind-tunnel model

A method of in-flight surface flow visualization similar to wind-tunnel-model oil flows is described for cases where photo-chase planes or onboard photography are not practical. This method, used on an F-18 aircraft in flight at high angles of attack, clearly showed surface flow streamlines in the fuselage forebody. Vortex separation and reattachment lines were identified with this method and documented using postflight photography. Surface flow angles measured at the 90 and 270 degrees meridians show excellent agreement with the wind tunnel data for a pointed tangent ogive with an aspect ratio of 3.5. The separation and reattachment line locations were qualitatively similar to the F-18 wind-tunnel-model oil flows but neither the laminar separation bubble nor the boundary-layer transition on the wind tunnel model were evident in the flight surface flows. The separation and reattachment line locations were in fair agreement with the wind tunnel data for the 3.5 ogive. The elliptical forebody shape of the F-18 caused the primary separation lines to move toward the leeward meridian. Little effect of angle of attack on the separation locations was noted for the range reported.

Fisher, David F.↗
NASA NTRS1991

A computational study of the F-5A forebody emphasizing directional stability

CFD has been used to study the F-5A forebody flowfield at high angle-of-attack combined with sideslip. The classic wind tunnel experiment demonstrating the dominant contribution of the F-5A forebody to directional stability at high angle-of-attack has been simulated computationally over an angle-of-attack range from 10 deg to 45 deg. The key wind tunnel trend for C (n-beta) was obtained computationally using the code known as cfl3d to solve the Reynolds' equations using a modified form of the Baldwin-Lomax turbulence model. The computations provide detailed and fascinating insights into the physics of flowfield. The results of the investigation show that CFD has reached a level of development where computational methods can be used for high angle-of-attack aerodynamic design.

Mason, William H.↗
NASA NTRS1972

Acoustic characteristics of large-scale STOL model at forward speed

Wind-tunnel investigations of the acoustic characteristics of the externally blown jet flap (EBF) and augmentor wing STOL concepts are dicussed. The large-scale EBF model was equipped with a triple-slotted blown by four JT15D turbofan engines with circular, coannular exhaust nozzles. The large-scale augmentor wing model was equipped with an unlined augmentor blown by a slot primary nozzle. The effects of airspeed and angle of attack on the acoustics of the EBF were small. At a forward speed of 60 knots, the impingement noise of the landing flap was approximately 2 db lower than in the static tests. Angle of attack increased the impingement noise approximately 0.1 decibels per degree. Flap deflection had a greater effect on the acoustics of the augmentor wing than did airspeed. For a nozzle pressure ratio of 1.9, the peak perceived noise level of the landing flap was 3 to 5 PNdb higher than that of the takeoff flap. The total sound power was also significantly higher for landing indicating that turning in the augmentor generated acoustic energy. Airspeed produced a small aft shift in acoustic directivity with no significant change in the peak perceived noise levels or sound power levels.

Falarski, M. D.↗
NASA NTRS1973

Aerodynamic investigations on a 0.004 scale model MCR 0074 baseline space shuttle launch vehicle at Mach numbers between 0.6 and 4.96

A test of a 0.004-scale MCR 0074 Baseline Launch Configuration Space Shuttle model was conducted in the NASA-MSFC 14 x 14-inch Trisonic Wind Tunnel (MSFC TWT 566). The objective of the test was to determine the effects of model parametric variations on aerodynamic static stability characteristics over a Mach number range from 0.6 to 4.96. Angles-of-attack from minus 10 deg to plus 10 deg at 0 deg sideslip and angles-of-sideslip from minus 10 deg to plus 10 deg at minus 5 deg, 0 deg, and plus 5 deg angle-of-attack were investigated. The basic configuration investigated was the integrated vehicle consisting of the orbiter, and external tank, and two solid rocket boosters. It was designated 03T9S3.

Ramsey, P.↗
NASA NTRS1972

Acoustic characteristics of large-scale STOL models at forward speed

Wind-tunnel investigations of the acoustic characteristics of the externally blown jet flap (EBF) and augmentor wing STOL concepts are discussed. The large-scale EBF model was equipped with a triple-slotted flap blown by four JT15D turbofan engines with circular, coannular exhaust nozzles. The large-scale augmentor wing model was equipped with an unlined augmentor blown by a slot primary nozzle. The effects of airspeed and angle of attack on the acoustics of the EBF were small. Flap deflection had a greater effect on the acoustics of the augmentor wing than did airspeed. The total sound power was also significantly higher for landing indicating that turning in the augmentor generated acoustic energy. Airspeed produced a small aft shift in acoustic directivity with no significant change in the peak perceived noise levels or sound power levels. Small-scale research of the acoustics for the augmentor wing has shown that by blowing an acoustically treated augmentor with a lobed primary nozzle, the 95-PNdb noise level goal can be achieved or surpassed.

Falarski, M. D.↗
NASA NTRS1973

Heat transfer phase change paint test (OH-42) of a Rockwell International SSV orbiter in the NASA/LRC Mach 8 variable density wind tunnel

Phase change paint tests of a Rockwell International .00593-scale space shuttle orbiter were conducted in the Langley Research Center's Variable Density Wind Tunnel. The test objectives were to determine the effects of various wing/underbody configurations on the aerodynamic heating rates and boundary layer transition during simulated entry conditions. Several models were constructed. Each varied from the other in either wing cuff radius, airfoil thickness, or wing-fuselage underbody blending. Two ventral fins were glued to the fuselage underside of one model to test the interference heating effects. Simulated Mach 8 entry data were obtained for each configuration at angles of attack ranging from 25 to 40 deg, and a Reynolds number variation of one million to eight million. Elevon, bodyflap, and rudder flare deflections were tested. Oil flow visualization and Schlieren photographs were obtained to aid in reducing the phase change paint data as well as to observe the flow patterns peculiar to each configuration.

Jones, R.↗
NASA NTRS1974

Results of investigations (OA20A) on a 0.015-scale 140A/B configuration space shuttle vehicle orbiter model in the NASA/Langley Research Center unitary plan wind tunnel

Data obtained from a wind tunnel test of a 0.015-scale 140A/B configuration SSV orbiter model are presented. This test was conducted with runs at Mach numbers of 2.5, 3.9, and 4.6 for a constant Reynolds number of 2.5 x 1,000,000/foot. Only one model configuration, the complete 140A/B orbiter vehicle, was investigated; various control-surface settings were run through angles-of-attack from -4 to +42 degrees at 0 and +3 degrees of yaw and through angles-of-sideslip from -4 to +6 degrees at 0, +10, +20, and +30 degrees pitch. The purpose of this test was to establish and verify longitudinal and lateral-directional stability and control characteristics for the updated SSV configuration.

Nichols, M. E.↗
NASA NTRS1974

Stability and control characteristics for the inner mold line configuration of the space shuttle orbiter (OA110)

Experimental aerodynamic investigations were conducted on a sting mounted 0.0405-scale representation of the -140A/B inner mold line (IML) space shuttle orbiter in 7.75 x 11 foot low speed wind tunnel, during the time period from 18 March 1974 to 20 March 1974. The primary test objectives were to establish basic longitudinal and lateral-directional stability and control characteristics for the IML orbiter. Additional configurations investigated were sealed elevon hingeline gaps, sealed rudder split line and hingeline gaps, larger radius leading edge on the vertical tail, and sealed speedbrake base. Aerodynamic force and moment data for the orbiter were measured in the body-axis system by an internally mounted, six-component strain gage balance. The model was sting mounted with the center of rotation located at approximately the wing trailing edge. The nominal angle of attack range was from -4 to +30 degrees. Yaw polars were recorded over a nominal yaw angle range from -14 to +14 degrees at constant angles of attack of 0, + or - 5, 10, 15 and 20 degrees.

Hughes, T.↗