Results for “attack modeling” — Page 59

NASA NTRS1987

Experimental study of three-dimensional separated flow surrounding a hemisphere-cylinder at incidence

Detailed experimental surface pressure coefficient measurements, obtained for a hemisphere-cylinder-flare model in a low supersonic flow (freestream M = 1.2) at various angles of attack (0 deg to 27.5 deg), have been analyzed. The pressure values for each angle of attack were smoothed and checked against their respective oil-flow photographs. The smoothed results were then used to validate a theory which relates the number and type of singular points observed in the oil-flow patterns with local surface-pressure extrema.

Meade, Andrew J., Jr.↗

NASA NTRS1988

Application of transonic slender body theory to bodies of varying complexity

Transonic flows over several bodies of varying complexity including a simple wing-body combination and the Shuttle Orbiter have been modeled with the use of slender body theory. Flows with various angles of attack have been considered. Also some off-body flow calculations were conducted. For attached flows, the calculations predicted the pressure results with good accuracy.

Rajagopal, Karunamurthy↗

NASA NTRS1989

Static aeroelastic analysis and tailoring of missile control fins

A concept for enhancing the design of control fins for supersonic tactical missiles is described. The concept makes use of aeroelastic tailoring to create fin designs (for given planforms) that limit the variations in hinge moments that can occur during maneuvers involving high load factors and high angles of attack. It combines supersonic nonlinear aerodynamic load calculations with finite-element structural modeling, static and dynamic structural analysis, and optimization. The problem definition is illustrated. The fin is at least partly made up of a composite material. The layup is fixed, and the orientations of the material principal axes are allowed to vary; these are the design variables. The objective is the magnitude of the difference between the chordwise location of the center of pressure and its desired location, calculated for a given flight condition. Three types of constraints can be imposed: upper bounds on static displacements for a given set of load conditions, lower bounds on specified natural frequencies, and upper bounds on the critical flutter damping parameter at a given set of flight speeds and altitudes. The idea is to seek designs that reduce variations in hinge moments that would otherwise occur. The block diagram describes the operation of the computer program that accomplishes these tasks. There is an option for a single analysis in addition to the optimization.

Mcintosh, S. C., Jr.↗

NASA NTRS1989

Analysis of flight data from a High-Incidence Research Model by system identification methods

Data partitioning and modified stepwise regression were applied to recorded flight data from a Royal Aerospace Establishment high incidence research model. An aerodynamic model structure and corresponding stability and control derivatives were determined for angles of attack between 18 and 30 deg. Several nonlinearities in angles of attack and sideslip as well as a unique roll-dominated set of lateral modes were found. All flight estimated values were compared to available wind tunnel measurements.

Batterson, James G.↗

NASA NTRS1991

Stirling engine: Available tools for long-life assessment

A review is presented for the durability approaches applicable to long-time life assessment of Stirling engine hot-section components. The crucial elements are experimental techniques for generating long-time materials property data (both monotonic and cyclic flow and failure properties); analytic representations of slow strain rate material stress-strain response characteristics (monotonic and cyclic constitutive relations) at high temperatures and low stresses and strains; analytic creep-fatigue-environmental interaction life prediction methods applicable to long lifetimes at high temperatures and small stresses and strains; and experimental verification of life predictions. Long-lifetime design criteria for materials of interest are woefully lacking. Designing against failures due to creep, creep-rupture, fatigue, environmental attack, and creep-fatigue-environmental interaction will require considerable extrapolation. Viscoplastic constitutive models and time-temperature parameters will have to be calibrated for the hot-section materials of interest. Analysis combined with limited verification testing in a short-time regime will be required to build confidence in long-lifetime durability models.

Halford, Gary R.↗

NASA NTRS1991

Stirling engine - Available tools for long-life assessment

A review is presented for the durability approaches applicable to long-time life assessment of Stirling engine hot-section components. The crucial elements are experimental techniques for generating long-time materials property data (both monotonic and cyclic flow and failure properties); analytic representations of slow strain rate material stress-strain response characteristics (monotonic and cyclic constitutive relations) at high temperatures and low stresses and strains; analytic creep-fatigue-environmental interaction life prediction methods applicable to long lifetimes at high temperatures and small stresses and strains; and experimental verification of life predictions. Long-lifetime design criteria for materials of interest are woefully lacking. Designing against failures due to creep, creep-rupture, fatigue, environmental attack, and creep-failure-environmental interaction will require considerable extrapolation. Viscoplastic constitutive models and time-temperature parameters will have to be calibrated for the hot-section materials of interest. Analysis combined with limited verification testing in a short-time regime will be required to build confidence in long-lifetime durability models.

Halford, Gary R.↗

NASA NTRS1956

An Air-flow-direction Pickup Suitable for Telemetering Use on Pilotless Aircraft

A vane-type air-flow-direction pickup is described which is suitable for telemetering angle-of-attack and angle-of-sideslip data from rocket-propelled pilotless aircraft models. Test results which are presented show that the device performs well under high accelerations and is stable throughout a Mach number rage from subsonic to above a Mach number of 2.5.

INSTRUMENTS, FLIGHT↗

NASA NTRS1996

Piloted Parameter Identification Flight Test Maneuvers for Closed Loop Modeling of the F-18 High Alpha Research Vehicle (HARV)

Flight test maneuvers are specified for the F-18 High Alpha Research Vehicle (HARV). The maneuvers were designed for closed loop parameter identification purposes, specifically for longitudinal and lateral linear model parameter estimation at 5, 20, 30, 45, and 60 degrees angle of attack, using the NASA 1A control law. Each maneuver is to be realized by the pilot applying square wave inputs to specific pilot station controls. Maneuver descriptions and complete specifications of the time/amplitude points defining each input are included, along with plots of the input time histories.

Morelli, Eugene A.↗

NASA NTRS1996

Parameter Identification Flight Test Maneuvers for Closed Loop Modeling of the F-18 High Alpha Research Vehicle (HARV)

Flight test maneuvers are specified for the F-18 High Alpha Research Vehicle (HARV). The maneuvers were designed for closed loop parameter identification purposes, specifically for longitudinal and lateral linear model parameter estimation at 5,20,30,45, and 60 degrees angle of attack, using the Actuated Nose Strakes for Enhanced Rolling (ANSER) control law in Thrust Vectoring (TV) mode. Each maneuver is to be realized by applying square wave inputs to specific pilot station controls using the On-Board Excitation System (OBES). Maneuver descriptions and complete specifications of the time / amplitude points defining each input are included, along with plots of the input time histories.

Batterson, James G.↗

NASA NTRS1954

Effects of Fuselage Modifications on the Drag Characteristics of a 1/20-Scale Model of the Convair F-102 Airplane at Transonic Speeds

An investigation has been conducted in the Langley 8-foot transonic tunnel to determine the effects of several fuselage modifications on the transonic drag-rise characteristics of a 1/20-scale model of the Convair F-102 airplane. Tests covered an angle-of-attack range from 0deg to about 10deg and a Mach number range from 0.60 to 1.14. Results indicated that the transonic drag rise .for the basic F-102 airplane could be substantially reduced by extending the fuselage after-body approximately 8 percent of the fuselage length. Tests of other bodies indicated that a shorter (4-percent) afterbody extension may have a similar effect on the drag rise. Further improvement of the axial cross-sectional-area distribution of the 8-percent extended configuration through the addition of fuselage volume resulted in additional reductions in the drag rise at a Mach number of 1.0 and caused no or only slight drag penalties at the higher Mach numbers. The results of the present tests generally substantiate the area-rule concept with respect to the prediction of the transonic drag rise through the use of an equivalent-area body of revolution for a practical delta-wing airplane configuration.

Kelly, Thomas C.↗

NASA NTRS2005

A Comparative Study Using CFD to Predict Iced Airfoil Aerodynamics

WIND, Fluent, and PowerFLOW were used to predict the lift, drag, and moment coefficients of a business-jet airfoil with a rime ice (rough and jagged, but no protruding horns) and with a glaze ice (rough and jagged end has two or more protruding horns) for angles of attack from zero to and after stall. The performance of the following turbulence models were examined by comparing predictions with available experimental data. Spalart-Allmaras (S-A), RNG k-epsilon, shear-stress transport, v(sup 2)-f, and a differential Reynolds stress model with and without non-equilibrium wall functions. For steady RANS simulations, WIND and FLUENT were found to give nearly identical results if the grid about the iced airfoil, the turbulence model, and the order of accuracy of the numerical schemes used are the same. The use of wall functions was found to be acceptable for the rime ice configuration and the flow conditions examined. For rime ice, the S-A model was found to predict accurately until near the stall angle. For glaze ice, the CFD predictions were much less satisfactory for all turbulence models and codes investigated because of the large separated region produced by the horns. For unsteady RANS, WIND and FLUENT did not provide better results. PowerFLOW, based on the Lattice Boltzmann method, gave excellent results for the lift coefficient at and near stall for the rime ice, where the flow is inherently unsteady.

Chi, x.↗

NASA NTRS2011

Sonic Boom Computations for a Mach 1.6 Cruise Low Boom Configuration and Comparisons with Wind Tunnel Data

Accurate analysis of sonic boom pressure signatures using computational fluid dynamics techniques remains quite challenging. Although CFD shows accurate predictions of flow around complex configurations, generating grids that can resolve the sonic boom signature far away from the body is a challenge. The test case chosen for this study corresponds to an experimental wind-tunnel test that was conducted to measure the sonic boom pressure signature of a low boom configuration designed by Gulfstream Aerospace Corporation. Two widely used NASA codes, USM3D and AERO, are examined for their ability to accurately capture sonic boom signature. Numerical simulations are conducted for a free-stream Mach number of 1.6, angle of attack of 0.3 and Reynolds number of 3.85x10(exp 6) based on model reference length. Flow around the low boom configuration in free air and inside the Langley Unitary plan wind tunnel are computed. Results from the numerical simulations are compared with wind tunnel data. The effects of viscous and turbulence modeling along with tunnel walls on the computed sonic boom signature are presented and discussed.

Elmiligui, Alaa A.↗

NASA NTRSDate not supplied

Fluid-Structure Interaction Simulations of the ASPIRE SR03 Supersonic Parachute Flight Test

Research into parachute performance continues to be a source of significant investment from the National Aeronautics and Space Administration to mitigate risks and to enable a variety of exploration missions, including landing on Mars as well as returning to Earth. The cost of flight tests to certify any changes to the current state-of-the-art parachute designs limits the development of next generation parachute systems. Fluid-structure interaction simulations could help accelerate this process once validated. The Launch, Ascent, and Vehicle Aerodynamics team is developing the capability to perform such fluid-structure interaction simulations by coupling a higher-order Cartesian immersed boundary computational fluid dynamics solver with adaptive mesh refinement to a finite element structural dynamics solver in space and time. We continue the effort to validate this tool with the Advanced Supersonic Parachute Inflation Research Experiments SR03 flight test featuring a strengthened parachute akin to the Mars 2020 mission that landed the Perseverance rover on Mars, and a higher freestream dynamic pressure prior to inflation. The effect of the flow conditions’ angle of attack and of the initial parachute shape are quantified. The impact of relaxing modeling assumptions with regards to radial stiffeners on the parachute canopy is also investigated. Results demonstrate improvements in agreement with the pull force recorded during the SR03 flight test as the initial conditions of the flow and parachute are brought closer to those experienced in flight, and further improved when the radial stiffener modeling assumptions are relaxed.

ESM↗

NASA NTRSDate not supplied

Fluid-Structure Interaction Simulations of the ASPIRE SR03 Supersonic Parachute Flight Test

Research into parachute performance continues to be a source of significant investment from the National Aeronautics and Space Administration to mitigate risks and to enable a variety of exploration missions, including landing on Mars as well as returning to Earth. The cost of flight tests to certify any changes to the current state-of-the-art parachute designs limits the development of next generation parachute systems. Fluid-structure interaction simulations could help accelerate this process once validated. The Launch, Ascent, and Vehicle Aerodynamics team is developing the capability to perform such fluid-structure interaction simulations by coupling a higher-order Cartesian immersed boundary computational fluid dynamics solver with adaptive mesh refinement to a finite element structural dynamics solver in space and time. We continue the effort to validate this tool with the Advanced Supersonic Parachute Inflation Research Experiments SR03 flight test featuring a strengthened parachute akin to the Mars 2020 mission that landed the Perseverance rover on Mars, and a higher freestream dynamic pressure prior to inflation. The effect of the flow conditions’ angle of attack and of the initial parachute shape are quantified. The impact of relaxing modeling assumptions with regards to radial stiffeners on the parachute canopy is also investigated. Results demonstrate improvements in agreement with the pull force recorded during the SR03 flight test as the initial conditions of the flow and parachute are brought closer to those experienced in flight, and further improved when the radial stiffener modeling assumptions are relaxed.

ESM↗

NASA NTRSDate not supplied

Broadband Predictions of Optimized Proprotors in Axial Forward Flight

A low-fidelity toolchain was used to predict broadband self-noise trends for three proprotors in axial flight: a baseline (C24ND) and two acoustically constrained proprotors (OPT-III and COPR-3). Rotor loads were predicted with the ANOPP Propeller Analysis System (PAS) and self-noise was predicted with the semiempirical method of Brooks, Pope, and Marcolini (BPM) implemented in the ANOPP2 Self-Noise Internal Functional Module (ASNIFM). Comparisons to experimental data revealed that trends for turbulent boundary layer trailing edge (TBL-TE) noise could be modeled across several flight conditions by increasing the boundary layer thicknesses via the trip setting. Since the Mach number range of the BPM method is exceeded in these predictions, a dependence of boundary layer displacement thickness on blade station Mach number was suggested as a possible reason for needing to model thicker boundary layers, suggesting that the TBL-TE model needs to be developed further. Bluntness vortex shedding noise (BVS) predictions required tuning the trailing edge thickness and trailing edge closure angle for each flight condition to match experimental trends, demonstrating that the BVS noise model is incomplete and that BVS noise may vary with the angle of attack. This study indicates that the BPM self-noise method needs to be improved, which will lead to more accurate broadband predictions.

Broadband self-noise↗

NASA NTRSDate not supplied

Broadband Predictions of Optimized Proprotors in Axial Forward Flight

A low-fidelity toolchain was used to predict broadband self-noise trends for three proprotors in axial flight: a baseline (C24ND) and two acoustically constrained proprotors (OPT-III and COPR-3). Rotor loads were predicted with the ANOPP Propeller Analysis System (PAS) and self-noise was predicted with the semiempirical method of Brooks, Pope, and Marcolini (BPM) implemented in the ANOPP2 Self-Noise Internal Functional Module (ASNIFM). Comparisons to experimental data revealed that trends for turbulent boundary layer trailing edge (TBL-TE) noise could be modeled across several flight conditions by increasing the boundary layer thicknesses via the trip setting. Since the Mach number range of the BPM method is exceeded in these predictions, a dependence of boundary layer displacement thickness on blade station Mach number was suggested as a possible reason for needing to model thicker boundary layers, suggesting that the TBL-TE model needs to be developed further. Bluntness vortex shedding noise (BVS) predictions required tuning the trailing edge thickness and trailing edge closure angle for each flight condition to match experimental trends, demonstrating that the BVS noise model is incomplete and that BVS noise may vary with the angle of attack. This study indicates that the BPM self-noise method needs to be improved, which will lead to more accurate broadband predictions.

Broadband self-noise↗

NASA NTRS1994

An Experimental Study of the Flowfield on a Semispan Rectangular Wing with a Simulated Glaze Ice Accretion

Wind tunnel experiments were conducted in order to study the effect of a simulated glaze ice accretion on the flowfield of a semispan, reflection-plane, rectangular wing at Re = 1.5 million and M = 0.12. A laser Doppler velocimeter was used to map the flowfield on the upper surface of the model in both the clean and iced configurations at alpha = 0, 4, and 8 degrees angle of attack. At low angles of attack, the massive separation bubble aft of the leading edge ice horn was found to behave in a manner similar to laminar separation bubbles. At alpha = 0 and 4 degrees, the locations of transition and reattachment, as deduced from momentum thickness distributions, were found to be in good agreement with transition and reattachment locations in laminar separation bubbles. These values at y/b = 0.470, the centerline measurement location, matched well with data obtained on a similar but two dimensional model. The measured velocity profiles on the iced wing compared reasonably with the predicted profiles from Navier-Stokes computations. The iced-induced separation bubble was also found to have features similar to the recirculating region aft of rearward-facing steps. At alpha = 0 degrees and 4 degrees, reverse flow magnitudes and turbulence intensity levels were typical of those found in the recirculating region aft of rearward-facing steps. The calculated separation streamline aft of the ice horn at alpha = 4 degrees, y/b = 0.470 coincided with the locus of the maximum Reynolds normal stress. The maximum Reynolds normal stress peaked at two locations along the separation streamline. The location of the first peak-value coincided with the transition location, as deduced from the momentum thickness distributions. The location of the second peak was just upstream of reattachment, in good agreement with measurements of flows over similar obstacles. The intermittency factor in the vicinity of reattachment at alpha = 4 degrees, y/b = 0.470, revealed the time-dependent nature of the reattachment process. The size and extent of the separation bubble were found to be a function of angle of attack and the spanwise location. Three dimensional effects were found to be strongest at alpha = 8 degrees. The calculated separation and stagnation streamlines were found to vary little with spanwise location at alpha = 0 degrees. The calculated separation streamlines at alpha = 4 degrees revealed that the bubble was largest near the centerline measurement plane, whereas the tip-induced vortex flow and the model root-tunnel wall boundary-layer interaction reduced the size of the bubble. These effects were found to be most dramatic at alpha = 8 degrees.

Khodadoust, Abdollah↗

NASA NTRS1962

Effects of Booster Transition Section and Rudder Deflection on the Low-Angle-of-Attack Static Stability of a Winged Reentry Vehicle at Mach Numbers of 10.8 and 17.8 in Helium

An investigation has been carried out to obtain aerodynamic stability and control data on a model of a winged reentry vehicle at Mach numbers of 10.8 and 17.8 in helium. The effects of a booster transition section on the static stability were obtained at angles of attack from -5 deg to 15 deg and at angles of sideslip from -5 deg to 10 deg at an angle of attack of 0 deg. Directional control data were also obtained at an angle of attack of 0 deg for sideslip angles from -5 deg to 10 deg. No detailed analysis of the data has been made.

Ladson, Charles L.↗