Results for “attack modeling” — Page 47

NASA NTRS2008

Robustness Analysis and Reliable Flight Regime Estimation of an Integrated Resilent Control System for a Transport Aircraft

Formal robustness analysis of aircraft control upset prevention and recovery systems could play an important role in their validation and ultimate certification. As a part of the validation process, this paper describes an analysis method for determining a reliable flight regime in the flight envelope within which an integrated resilent control system can achieve the desired performance of tracking command signals and detecting additive faults in the presence of parameter uncertainty and unmodeled dynamics. To calculate a reliable flight regime, a structured singular value analysis method is applied to analyze the closed-loop system over the entire flight envelope. To use the structured singular value analysis method, a linear fractional transform (LFT) model of a transport aircraft longitudinal dynamics is developed over the flight envelope by using a preliminary LFT modeling software tool developed at the NASA Langley Research Center, which utilizes a matrix-based computational approach. The developed LFT model can capture original nonlinear dynamics over the flight envelope with the ! block which contains key varying parameters: angle of attack and velocity, and real parameter uncertainty: aerodynamic coefficient uncertainty and moment of inertia uncertainty. Using the developed LFT model and a formal robustness analysis method, a reliable flight regime is calculated for a transport aircraft closed-loop system.

Shin, Jong-Yeob↗

NASA NTRS1995

Hypersonic Aerodynamic Characteristics of a Proposed Single-Stage-To-Orbit Vehicle

The hypersonic aerodynamic characteristics of a winged body concept representing a candidate single- stage-to-orbit vehicle which features wing tip fin controllers and elevon/body flap control surfa'Fs are predicted at points along a nominal trajectory for Mach numbers from 5 to 27 and angles of attack from 19 to 32 degrees. Predictions are derived from surface properties based on flow solvers for inviscid and viscous, laminar flows acting as a perfect gas, as a gas in chemical equilibrium and as a gas in chemical non- equilibrium. At a Mach number of 22, the lateral aerodynamic characteristics of the vehicle are determined based on an inviscid analysis at side slip angles of 2 and 4 degrees and 32 degrees angle of attack; a viscous analysis was carried out to determine the effect of gas chemistry model on surface pressure and to determine the incremental aerodynamics for control surface deflections. The results show that the longitudinal pitch characteristics of the baseline configuration, i.e., zero control surface deflections, are significantly altered by real gas chemistry at angles of attack greater than 30 degrees and Mach numbers greater than 9; and, that aerodynamics derived from inviscid solutions are of sufficient accuracy for preliminary analysis. Also, it is shown that a Mach number of 22, the choice of gas chemistry model has a large impact on surface pressure levels at highly localized regions on the vehicle and that the vehicle can be trimmed at control surface deflections less than 11 degrees.

Weilmuenster, K. James↗

NASA NTRS1968

Wind-tunnel measurements of transition on the face of a blunt entry capsule at angle of attack.

Boundary layer transition point on forward portion of blunt Apollo type entry capsule model measured in high Reynolds number wind tunnel

Jones, R. A.↗

NASA NTRS1974

The effect of canard and vertical tails on the aerodynamic characteristics of a model with a 59 deg sweptback wing at a Mach number of 0.30

An investigation has been conducted to determine the effects of canard, canard location, vertical tails, and vertical-tail location on the aerodynamic characteristics of a model having a 59 deg sweptback wing. The investigation was conducted at a Mach number of 0.30, at angles of attack up to 22 deg and at sideslip angles of 0 deg and plus or minus 5 deg. The results of the study indicate that adding the canard to the model had only a slight effect on the lift at the lower angles of attack. At the higher angles of attack there is a significant effect of canard height on lift, canard in the high location (above the wing chord plane) resulting in the highest lifts. The lift drag characteristics are predicted well for the configuration with the mid or high canard locations by combining a potential flow solution on the canard with a potential plus vortex solution on the wing. Variations in the height significantly affect the pitching-moment characteristics of the configuration; the configuration with the low or mid canard location exhibits an increase in stability at the higher lift coefficients, whereas the configuration with the high canard exhibits pitch-up. Adding the vertical tails in the outboard location caused a significant loss in lift at the higher angles of attack; this lift loss was eliminated by moving the vertical tails inboard.

Henderson, W. P.↗

NASA NTRS1992

A LEX blowing technique for post-stall lateral control of trapezoidal wings

An exploratory low-speed wind-tunnel investigation of a pneumatic technique for LEX-induced lift augmetation on a low-aspect-ratio trapezoidal wing near C sub L, MAX is reported. The technique involved lateral sheet ejection from LEX edge-slots, with emphasis on nonsymmetrical blowing to obtain incremental roll control at high angles of attack when the aileron power is declining. Preliminary experiments on a pressure instrumented semispan wing model followed by force/moment measurements on a full-span LEX-wing configuration confirmed the effectiveness of LEX blowing for lateral control improvement up to an angle of attack of 45 deg and with a momentum coefficient of less than 02.

Rao, Dhanvada M.↗

NASA NTRS1983

Magnus effects at high angles of attack and critical Reynolds numbers

The Magnus force and moment experienced by a yawed, spinning cylinder were studied experimentally in low speed and subsonic flows at high angles of attack and critical Reynolds numbers. Flow-field visualization aided in describing a flow model that divides the Magnus phenomenon into a subcritical region, where reverse Magnus loads are experienced, and a supercritical region where these loads are not encountered. The roles of the spin rate, angle of attack, and crossflow Reynolds number in determining the boundaries of the subcritical region and the variations of the Magnus loads were studied.

Seginer, A.↗

NASA NTRS1956

An Experimental Investigation of Sting-Support Effects on Drag and a Comparison with Jet Effects at Transonic Speeds

This paper presents the results of an investigation of sting-support interference on afterbody drag at transonic speeds. Stings with varying diameter, cone angle, and cylindrical length were tested at the rea r of a model with various afterbody shapes. The data were obtained at an angle of attack of O deg. and at Mach numbers from 0.80 to 1.10. It was found that, in general, the addition of a sting caused a drag reduction. A method is presented whereby approximate sting-interferen ce corrections may be made to models with afterbodies and sting suppo rts of similar size and scale to those of this paper provided the bou ndary layer is turbulent at the model base and the Reynolds numbers a re of the same order of magnitude. Reynolds number of the tests prese nted varied from 15.0 x 10 (exp 6) to 17.4 x 10 (exp 6) based on body length. Sting effects from this investigation are compared with data of jet effects on the same afterbodies. The results of this comparis on indicate that for the more gradually contoured afterbodies, a stin g shape can be found which will duplicate the jet effects, but that f or blunt afterbodies no solid sting shape will duplicate the jet effe cts.

Cahn, Maurice S.↗

NASA NTRS2021

Pretest Computational Assessment of Boundary Layer Transition in the NASA Juncture Flow Model with an NACA 0015-Based Wing

The first two phases of the NASA Juncture Flow experiment were carried out on a DLR-F6 swept-wing model and were designed to provide “CFD validation-quality” data toward the assessment and improvement of existing CFD turbulence models in predicting onset and extent of three-dimensional separated flow near the wing-juncture trailing-edge region. The next phase of experiments will involve an NACA 0015-based swept wing, as prior risk reduction experiments had indicated that this wing shape resulted in reduced separation near the juncture region than the DLR-F6 wing, thus providing a better option to evaluate the ability of CFD models to predict incipient turbulent separation. The NACA 0015 measurements will also include IR thermography to infer the variation of transition front with respect to an increasing angle of attack. The primary objective of this work is to computationally make a preliminary assessment of the transition front on both surfaces of the NACA 0015 wing at a crank-chord-based Reynolds number of 2.4 x 106 for four different angles of attack, (0°, 2.5°, 5°, and 7.5°) and to determine the dominant mechanisms responsible for transition. This assessment includes both RANS-based transition models from NASA’s OVERFLOW 2.3b flow solver and linear parabolized stability equations (PSE) stability analysis based on the Langley Stability and Transition Analysis code, LASTRAC. Linear PSE results indicate that the upper surface of the wing is dominated by Tollmien- Schlichting (TS) instabilities, and that the laminar flow region shrinks from about 50% chord to a very small region just downstream of the attachment line as the angle of attack is increased from 0° to 7.5°. Consequently, the transition fronts predicted by the Spalart- Allmaras-based amplification factor transport (AFT-2017b) equation model (which accounts for the TS instabilities alone) and the Menter’s shear-stress transport equation (SST2003)- based Langtry-Menter transition model with ability to account for both TS and crossflow effects (LM2015) compare well with those predicted using linear PSE. On the lower surface of the wing, stationary crossflow (CF) instabilities begin to appear on the inboard portion of the wing in addition to the TS-instabilities for the larger angles of attack (5° and 7.5°), further reducing the laminar flow extent within the inboard region. The LM2015 model that accounts for CF effects is able to replicate this trend but appears to predict a slightly earlier transition. The outcome of this effort will inform the experiment and, when the actual experimental data become available, provide further opportunity to assess and improve the various transition models.

CFD modeling↗

NASA NTRS2021

Pretest Computational Assessment of Boundary Layer Transition in the NASA Juncture Flow Model with an NACA 0015-Based Wing

The first two phases of the NASA Juncture Flow experiment were carried out on a DLR-F6 swept-wing model and were designed to provide “CFD validation-quality” data toward the assessment and improvement of existing CFD turbulence models in predicting onset and extent of three-dimensional separated flow near the wing-juncture trailing-edge region. The next phase of experiments will involve an NACA 0015-based swept wing, as prior risk reduction experiments had indicated that this wing shape resulted in reduced separation near the juncture region than the DLR-F6 wing, thus providing a better option to evaluate the ability of CFD models to predict incipient turbulent separation. The NACA 0015 measurements will also include IR thermography to infer the variation of transition front with respect to an increasing angle of attack. The primary objective of this work is to computationally make a preliminary assessment of the transition front on both surfaces of the NACA 0015 wing at a crank-chord-based Reynolds number of 2.4 x 106 for four different angles of attack, (0°, 2.5°, 5°, and 7.5°) and to determine the dominant mechanisms responsible for transition. This assessment includes both RANS-based transition models from NASA’s OVERFLOW 2.3b flow solver and linear parabolized stability equations (PSE) stability analysis based on the Langley Stability and Transition Analysis code, LASTRAC. Linear PSE results indicate that the upper surface of the wing is dominated by Tollmien- Schlichting (TS) instabilities, and that the laminar flow region shrinks from about 50% chord to a very small region just downstream of the attachment line as the angle of attack is increased from 0° to 7.5°. Consequently, the transition fronts predicted by the Spalart- Allmaras-based amplification factor transport (AFT-2017b) equation model (which accounts for the TS instabilities alone) and the Menter’s shear-stress transport equation (SST2003)- based Langtry-Menter transition model with ability to account for both TS and crossflow effects (LM2015) compare well with those predicted using linear PSE. On the lower surface of the wing, stationary crossflow (CF) instabilities begin to appear on the inboard portion of the wing in addition to the TS-instabilities for the larger angles of attack (5° and 7.5°), further reducing the laminar flow extent within the inboard region. The LM2015 model that accounts for CF effects is able to replicate this trend but appears to predict a slightly earlier transition. The outcome of this effort will inform the experiment and, when the actual experimental data become available, provide further opportunity to assess and improve the various transition models.

CFD modeling↗

NASA NTRS1983

Experimental investigation of forebody and wing leading-edge vortex interactions at high angles of attack

An experimental investigation was conducted to assess the vortex flow-field interactions on an advanced, twin-jet fighter aircraft configuration at high angles of attack. Flow-field surveys were conducted on a small-scale model in the Northrop 0.41 - by 0.60-meter water tunnel and, where appropriate, the qualitative observations were correlated with low-speed wind tunnel data trends obtained on a large-scale model of the advanced fighter in the NASA Langley Research Center 30- by 60-foot (9.1- by 18.3-meter) facility. Emphasis was placed on understanding the interactions of the forebody and LEX-wing vortical flows, defining the effects on rolling moment variation with sideslip, and identifying modifications to control or regulate the vortex interactions at high angles of attack. The water tunnel flow visualization results and wind tunnel data trend analysis revealed the potential for strong interactions between the forebody and LEX vortices at high angles of attack. In particular, the forebody flow development near the nose could be controlled by means of carefully-positioned radome strakes. The resultant strake-induced flow-field changes were amplified downstream by the more powerful LEX vortical motions with subsequent large effects on wing flow separation characteristics.

Erickson, G. E.↗

NASA NTRS1980

Experimental and computational study of transonic flow about swept wings

An experimental investigation of NACA 0010 and 10% circular arc wing models, swept at 45 deg, spanning a channel, and at zero angle of attack is described. Measurements include chordwise and spanwise surface pressure distributions and oil-flow patterns for a range of transonic Mach numbers and Reynolds numbers. Calculations using a new three-dimensional Navier-Stokes code and a two-equation turbulence model are included for the circular-arc wing flow. Reasonable agreement between measurements and computations is obtained.

Bertelrud, A.↗

NASA NTRS2011

Development of the Transport Class Model (TCM) Aircraft Simulation From a Sub-Scale Generic Transport Model (GTM) Simulation

A six degree-of-freedom, flat-earth dynamics, non-linear, and non-proprietary aircraft simulation was developed that is representative of a generic mid-sized twin-jet transport aircraft. The simulation was developed from a non-proprietary, publicly available, subscale twin-jet transport aircraft simulation using scaling relationships and a modified aerodynamic database. The simulation has an extended aerodynamics database with aero data outside the normal transport-operating envelope (large angle-of-attack and sideslip values). The simulation has representative transport aircraft surface actuator models with variable rate-limits and generally fixed position limits. The simulation contains a generic 40,000 lb sea level thrust engine model. The engine model is a first order dynamic model with a variable time constant that changes according to simulation conditions. The simulation provides a means for interfacing a flight control system to use the simulation sensor variables and to command the surface actuators and throttle position of the engine model.

Hueschen, Richard M.↗

NASA NTRS1972

Space shuttle: Heat transfer investigation of the McDonnell-Douglas delta wing orbiter at a nominal Mach number of 10.5

Heat transfer tests for the delta wing orbiter were conducted in a hypervelocity wind tunnel. A 1.1 percent scale model was tested at a Mach number of approximately 10.5 over an angle of attack range from 10 to 60 degrees over a length Reynolds number range from 5 times 10 to the 6th power to 24 times 10 to the 6th power. Heat transfer results were obtained from model surface heat gage measurements and thermographic phosphor paint. Limited pressure measurements were obtained.

Eaves, R. H.↗

NASA NTRS1975

Results of investigations on a 0.004-scale 140C modified configuration space shuttle vehicle orbiter model (74-0) in the NASA/Langley Research Center hypersonic helium tunnel

Data obtained during a wind tunnel test of a 0.004-scale 140C modified configuration SSV orbiter are documented. The test was conducted during August 1974 with 80 occupancy hours charged, and all runs were conducted at a nominal Mach number of 20 and at Reynolds numbers of 0.7, 1.0, 1.8, and 1,100,000 based on body length. The complete -140C modified model was tested with various elevon settings at angles of attack from 10 to 50 degrees at zero yaw and from angles of sideslip of -10 to +10 at 35 deg angle of attack. The purpose of this test was to obtain high hypersonic longitudinal and lateral-directional stability and control characteristics of the updated SSV configuration.

Hawthorne, P. J.↗

NASA NTRS1975

Design and testing of new center inlet and S-duct for B-727 airplane with refanned JT8D engines

The work described in this paper was part of the NASA refan program. The airflow requirement of the refanned JT8D (-100 series) engine increased about 50% above that of the basic JT8D. This required a redesign of the center inlet and S-duct of the Boeing 727 airplane. The paper describes the design constraints for the S-duct and the analytic method used to define the lines of the new duct. Model tests that were conducted at static, angle-of-attack, and crosswind conditions are described with a variety of flow control devices. Test results showed that the new inlet and S-duct have a pressure recovery comparable to that of the existing inlet. By employing corotating vortex generators, less flow distortion was obtained for the core region than the existing duct has with its counter rotating vortex generators.

Ting, C. T.↗

NASA NTRS1977

Aerodynamic characteristics at Mach 6 of a hypersonic research airplane concept having a 70 deg swept delta wing

The hypersonic aerodynamic characteristics of an air-launched, delta-wing research aircraft concept were investigated at Mach 6. The effect of various components such as nose shape, wing camber, wing location, center vertical tail, wing tip fins, forward delta wing, engine nacelle, and speed brakes was also studied. Tests were conducted with a 0.021 scale model at a Reynolds number, based on model length, of 10.5 million and over an angel of attack range from -4 deg to 20 deg. Results show that most configurations with a center vertical tail have static longitudinal stability at trim, static directional stability at angles of attack up to 12 deg, and static lateral stability throughout the angle of attack range. Configurations with wing tip fins generally have static longitudinal stability at trim, have lateral stability at angles of attack above 8 deg, and are directionally unstable over the angle of attack range.

Clark, L. E.↗

NASA NTRS1978

Space Shuttle Orbiter trimmed center-of-gravity extension study. Volume 4: Effects of configuration modifications on the aerodynamic characteristics of the 139B orbiter at Mach 20.3

Force tests were conducted at Mach 20.3 to determine the effect of several forebody, wing-fillet, and canard modifications on the hypersonic trim capability of a 139B Space Shuttle Orbiter model. Force and moment data were obtained at angles of attack of 10 deg to 54 deg at zero sideslip angle and at a Reynolds number of 1,900,000 based on body length. The results indicated that wing-fillet and canard modifications would increase the allowable forward trimmed center-of-gravity capability by as much as 3.0 percent of the body length.

Scallion, W. I.↗

NASA NTRS1980

Biological and aerodynamic problems with the flight of animals

Biological and aerodynamic considerations related to birds and insects are discussed. A wide field is open for comparative biological, physiological, and aerodynamic investigations. Considerable mathematics related to the flight of animals is presented, including 20 equations. The 15 figures included depict the design of bird and insect wings, diagrams of propulsion efficiency, thrust, lift, and angles of attack and photographs of flapping wing free flying wing only models which were built and flown.

Holst, E. V.↗