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At least 343 records · Page 19
All sourcesNASA NTRSDOE OSTI
NASA NTRS2013

Analysis and Results from a Flush Airdata Sensing System in Close Proximity to Firing Rocket Nozzles

This paper presents information regarding the nosecap Flush Airdata Sensing (FADS) system on Orion’s Pad Abort 1 (PA-1) vehicle. The purpose of the nosecap FADS system was to test whether or not useful data could be obtained from a FADS system if it was placed in close proximity to firing rocket nozzles like the Attitude Control Motor (ACM) nozzles on the PA-1 Launch Abort System. The nosecap FADS system used pressure measurements from a series of pressure ports which were arranged in a cruciform pattern and flush with the surface of the vehicle to estimate values of angle of attack, angle of sideslip, Mach number, impact pressure, and freestream static pressure. This paper will present the algorithms employed by the FADS system along with the development of the calibration datasets and a comparison of the final results to the Best Estimated Trajectory (BET) data for PA-1. Also presented in this paper is a Computational Fluid Dynamics (CFD) study to explore the impact of the ACM on the nosecap FADS system. The comparison of the nosecap FADS system results to the BET and the CFD study showed that more investigation is needed to quantify the impact of the firing rocket motors on the FADS system.

Flush Air Data Sensing System↗
NASA NTRS2013

Analysis and Results from a Flush Airdata Sensing (FADS) System in Close Proximity to Firing Rocket Nozzles

This presentation presents information regarding the nose-cap flush airdata sensing (FADS) system on Orion's Pad Abort 1 (PA-1) vehicle. The purpose of the nose-cap FADS system was to test whether or not useful data could be obtained from a FADS system if it was placed in close proximity to firing rockets nozzles like the attitude control motor (ACM) nozzles on the PA-1 launch abort system (LAS). The nose-cap FADS systems use pressure measurements from a series of pressure ports which are arranged in a cruciform pattern and flush with the surface of the vehicle to estimate values of angle of attack, angle of side-slip, Mach number, impact pressure and free-stream static pressure.

Pad Abort 1↗
NASA NTRSDate not supplied

X-59 Air Data Probe Calibration Wind Tunnel Test

NASA's aeronautical innovators are working with Lockheed Martin to design and build a supersonic X-plane, the X-59, under the the Low Boom Flight Demonstrator (LBFD) Project, that produces a gentle thump rather than a loud sonic boom. The plane will be used to collect human response data over select U.S. communities. The data from the X-59 test flights will be provided to U.S. and international regulators to potentially allow supersonic flight over land, drastically reducing travel time within the U.S. and around the world. Two wind tunnel tests were conducted in the NASA Glenn Research Center (GRC) 8- by 6-Foot Supersonic Wind Tunnel to calibrate the nose probe for the X-59. The probe was successfully tested and calibrated at 19 Mach numbers between Mach 0.25 and Mach 1.7. Most data were collected using continuous roll angle sweeps at a set Mach number and pitch angle. About 725 high quality data runs were recorded during testing to be used in the calibration of the probe. A backup nose probe was also run through the same test matrix as the primary probe. The probe will be the primary instrument for the flight test of the X-59 for determining angle of attack, angle of sideslip, airspeed, and pressure altitude of the aircraft in flight. Knowing the aircraft's speed and attitude is critical for flight safety and is also critical data for the Sonic Boom mission. Areas of interest for the test were sea level takeoff and landing at Mach 0.2, subsonic cruise at Mach 0.9, and supersonic cruise at Mach 1.4. During Phase I flight testing, the air data nose probe and air data system as a whole will be further calibrated in flight to account for additional installed effects of the airframe on the air data system.

X-59↗
NASA NTRSDate not supplied

X-59 Air Data Probe Calibration Wind Tunnel Test

NASA's aeronautical innovators are working with Lockheed Martin to design and build a supersonic X-plane, the X-59, under the the Low Boom Flight Demonstrator (LBFD) Project, that produces a gentle thump rather than a loud sonic boom. The plane will be used to collect human response data over select U.S. communities. The data from the X-59 test flights will be provided to U.S. and international regulators to potentially allow supersonic flight over land, drastically reducing travel time within the U.S. and around the world. Two wind tunnel tests were conducted in the NASA Glenn Research Center (GRC) 8- by 6-Foot Supersonic Wind Tunnel to calibrate the nose probe for the X-59. The probe was successfully tested and calibrated at 19 Mach numbers between Mach 0.25 and Mach 1.7. Most data were collected using continuous roll angle sweeps at a set Mach number and pitch angle. About 725 high quality data runs were recorded during testing to be used in the calibration of the probe. A backup nose probe was also run through the same test matrix as the primary probe. The probe will be the primary instrument for the flight test of the X-59 for determining angle of attack, angle of sideslip, airspeed, and pressure altitude of the aircraft in flight. Knowing the aircraft's speed and attitude is critical for flight safety and is also critical data for the Sonic Boom mission. Areas of interest for the test were sea level takeoff and landing at Mach 0.2, subsonic cruise at Mach 0.9, and supersonic cruise at Mach 1.4. During Phase I flight testing, the air data nose probe and air data system as a whole will be further calibrated in flight to account for additional installed effects of the airframe on the air data system.

X-59↗
NASA NTRS1986

A computational/experimental study of the flow around a body of revolution at angle of attack

The incompressible Navier-Stokes equations are numerically solved for steady flow around an ogive-cylinder (fineness ration 4.5) at angle of attack. The three-dimensional vortical flow is investigated with emphasis on the tip and the near wake region. The implicit, finite-difference computation is performed on the CRAY X-MP computer using the method of pseudo-compressibility. Comparisons of computational results with results of a companion towing tank experiment are presented for two symmetric leeside flow cases of moderate angles of attack. The topology of the flow is discussed and conclusions are drawn concerning the growth and stability of the primary vortices.

Zilliac, Gregory G.↗
NASA NTRS1982

High angle-of-attack characteristics of a forward-swept wing fighter configuration

The NASA-Langley Research Center is currently investigating the high angle-of-attack characteristics of a forward-swept wing fighter technology demonstrator in a cooperative program with the Defense Advanced Research Projects Agency. The program includes static and dynamic wind-tunnel force tests, free-flight model tests, spin-tunnel tests, and piloted simulation. Particular emphasis is placed on identifying the contributions of the forward-swept wing to the configuration aerodynamics at high angles of attack. Results are presented to illustrate the wing contributions in terms of aerodynamic stability, flow visualization, and observation of free-flight characteristics.

Grafton, S. B.↗
NASA NTRS1986

Low Aspect Ratio Wings at High Angles of Attack

A comprehensive evaluation is made of experimental data compiled to date for the flowfields and aerodynamic forces that occur at high angles of attack for low aspect ratio wings with delta, rectangular, clipped delta, and strake/wing planform geometries. Attention is given to wing leading edge-generated vortex breakdown, aspect ratio and compressibility effects, and strake vortex effects on main wing areas. Although the nonlinear effects created by a wing-body combination significantly alter wing-alone aerodynamics, the wing-alone data presented are vital to the development of prediction methodologies for large angle of attack aerodynamics.

R L Stallings Jr.↗
NASA NTRS1990

Near-field noise of a single-rotation propfan at an angle of attack

The near field noise characteristics of a propfan operating at an angle of attack are examined utilizing the unsteady pressure field obtained from a 3-D Euler simulation of the propfan flowfield. The near field noise is calculated employing three different procedures: a direct computation method in which the noise field is extracted directly from the Euler solution, and two acoustic-analogy-based frequency domain methods which utilize the computed unsteady pressure distribution on the propfan blades as the source term. The inflow angles considered are -0.4, 1.6, and 4.6 degrees. The results of the direct computation method and one of the frequency domain methods show qualitative agreement with measurements. They show that an increase in the inflow angle is accompanied by an increase in the sound pressure level at the outboard wing boom locations and a decrease in the sound pressure level at the (inboard) fuselage locations. The trends in the computed azimuthal directivities of the noise field also conform to the measured and expected results.

Nallasamy, M.↗
NASA NTRS1992

An investigation of the effects of aft blowing on a 3.0 caliber tangent ogive body at high angles of attack

An experimental investigation of the effects of aft blowing on the asymmetric vortex flow of a slender, axisymmetric body at high angles of attack was conducted. A 3.0 caliber tangent ogive body fitted with a cylindrical afterbody was tested in a wind tunnel under subsonic, laminar flow test conditions. Asymmetric blowing from both a single nozzle and a double nozzle configuration, positioned near the body apex, was investigated. Aft blowing was observed to alter the vortex asymmetry by moving the blowing-side vortex closer to the body surface while moving the non-blowing-side vortex further away from the body. The effect of increasing the blowing coefficient was to move the blowing-side vortex closer to the body surface at a more upstream location. The data also showed that blowing was more effective in altering the initial vortex asymmetry at the higher angles of attack than at the lower. The effects of changing the nozzle exit geometry were investigated and it was observed that blowing from a nozzle with a low, broad exit geometry was more effective in reducing the vortex asymmetry than blowing from a high, narrow exit geometry.

Gittner, Nathan M.↗
NASA NTRS1972

Analysis of flight data on boundary layer transition at high angles of attack

Boundary layer transition data were obtained on the flight of two cones which reentered at velocities of about 7.0 km/sec. One cone reentered at a nominal zero degree angle of attack and the other, due to an anomaly above the earth atmosphere, reentered at local angles of attack up to 7.0 km/sec. The transition data were obtained from on-board acoustic and electrostatic sensors. A description of the design, calibration, and method used to detect transition from the sensors is included. The flow field calculation used to obtain the local flow properties on the cones is described. Finally, the transition data found from both cone flights is correlated.

Haigh, W. W.↗
NASA NTRS1981

A numerical simulation of hypersonic viscous flow over arbitrary geometries at angle of attack

An implicit conservative, noniterative, finite-difference algorithm that predicts the supersonic, laminar or turbulent viscous flow about arbitrary geometries at large angles of attack is presented. The three-dimensional parabolized form of the thin-layer Navier-Stokes equations are written in generalized coordinates. These equations are solved using the delta form of the Beam-Warming implicit algorithm. Flow field simulations have been obtained for a blunt biconic with windward and leeward cuts and an X-24C lifting body for both laminar and turbulent flow at various Mach numbers and angles of attack. When compared with experiment or with previous theories, these computational predictions show good agreement.

Chaussee, D. S.↗
NASA NTRS1978

Wake vortex measurements of bodies at high angle of attack

Three-dimensional laser velocimeter measurements have been made of the wake vortices of a slender tangent-ogive body which had nose and body fineness ratios of 3.5 and 12, respectively. Data were obtained for an angle of attack to seminose angle ratio of 2.3 at a free-stream Mach number of 0.6 and unit Reynolds number of 2 million/ft. Details of the mean flow field are presented and features of the turbulent and unsteady nature of the vortex flow field are discussed. Problems associated with obtaining meaningful vortex measurements in high-speed flows are addressed.

Owen, F. K.↗