Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “attack modeling”

Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 721 records · Page 40
All sourcesNASA NTRSDOE OSTI
NASA NTRS1977

Effects of canard location on the aerodynamic characteristics of a blunt-nosed missile at Mach numbers of 1.5 and 2.0

A blunt-nosed 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 and canard-deflection angles from -3 deg to 15 deg for various stages of model build-up (i.e., with and without canard and/or tail surfaces). Results were obtained with the canards at two different nose locations. For the canard and tail arrangements investigated, the model was trimmable at angles of attack up to about 4 deg or 5 deg with canard deflections of 9 deg. For this blunt-nosed model, there was little effect of canard location on trim angle of attack. The tail arrangements studied provided ample pitch stability.

Kassner, D. L.↗
NASA NTRS1975

Aerodynamic testing technique for twin fuselage models at hypersonic speeds

A testing technique for obtaining the static aerodynamic characteristics of twin-fuselage configurations at hypersonic speeds by using a conventional single-balance installation has been evaluated. Data from a triple-fuselage model and a single-fuselage model were summed and then halved to obtain the characteristics for a twin-fuselage model of the same scale. The three related models were evaluated experimentally at Mach 20.3 in helium and Mach 6 in air for an angle-of-attack range from minus 6 deg to 50 deg. The Reynolds numbers, based on model length, were 1.88 million for the Mach 20.3 tests and 2.55 million for the Mach 6 tests.

Ashby, G. C., Jr.↗
NASA NTRS1958

Investigation of the Static Lateral Stability and Aileron Characteristics of a 0.067-Scale Model of the Bell X-2 Airplane at Mach Numbers of 2.29, 2.78, 3.22, and 3.71

Tests were performed in the high. Mach number test section of the Langley Unitary Plan wind tunnel to determine the static lateral stability. and aileron characteristics of a 0.067-scale model of the Bell X-2 airplane at Mach numbers of 2.29, 2. 78, 3.22, and. 3.71. The results of this investigation indicated that the directional stability of the model was low with directional instability occurring at Mach numbers higher than 3.1 and. angles of attack higher than about 5.0 deg (equivalent lift coefficient of about 0.18). The yaw due to aileron deflection was adverse and, with 10 deg of differential aileron deflection, large enough to overbalance the available directional restoring moment at all angles of attack higher than about 5.0 deg (equivalent lift coefficient of about 0.21) and Mach numbers higher than 2. 5. The model also had positive effective dihedral for all test attitudes and. Mach numbers. A combination of the lateral-stability parameters with the aileron characteristics to form a lateral-stability criterion for a maneuver using ailerons alone indicated that the model has characteristics which would. give unstable aperiodic behavior (divergence) over a large part of the test Mach number and angle-of-attack range.

Fournier, Roger H.↗
NASA NTRS2007

On Problems Associated with Modeling Wing-Tail Configurations from Wind Tunnel Data

This paper considers factors that contribute to poor identification of unsteady aerodynamics from wind tunnel data for an airliner configuration. One approach to modeling a wing-tail configuration is considered and applied to both steady and large-amplitude forced pitch oscillation wind tunnel data taken over a wide range of angles of attack but a limited range of amplitude and frequencies. The identified models fit the measured data well but in some cases with inaccurate parameters. Only limited conclusions can be drawn from analysis of the current data set until further experiments can be performed to resolve the identification issues. The analysis of measured and simulated data provides some insights and guidance on how an effective experiment may be designed for wing-tail configurations with nonlinear unsteady aerodynamics.

Murphy, Patrick C.↗
NASA NTRS2018

Computational Component Build-Up for the X-57 Distributed Electric Propulsion Aircraft

A computational study of the wing for the distributed electric propulsion X-57 Maxwell airplane configuration at cruise and takeoff/landing conditions was completed. Three unstructured-mesh, Navier-Stokes computational fluid dynamics methods, FUN3D, USM3D and Kestrel, were used to predict the performance buildup of components to the full X-57 configuration. The goal of the X-57 wing and distributed electric propulsion system design was to meet or exceed the required lift coefficient of 3.95 for a stall speed of 58 knots. The X-57 Maxwell airplane was designed with a small, high aspect ratio cruise wing that was designed for a high cruise lift coefficient of 0.75 at a cruise speed of 150 knots and altitude of 8,000 ft, with an angle of attack of approximately 0deg. The computational data indicates that the X-57 full aircraft drag would meet the cruise drag goal with a 25 count drag margin. The cruise configuration maximum lift coefficient is 2.07 and without including the stabilator is 1.86 at an angle of attack of 14 deg, predicted with the USM3D flow solver using the Spalart-Allmaras turbulence model. The maximum lift coefficient for the high-lift wing (with the 30deg flap deflection) without the stabilator contribution is 2.60 at an angle of attack of 13 deg. For high-lift blowing conditions with 13.7 hp/prop, the maximum lift coefficient excluding the stabilator is 4.426 at (alpha) = 13 deg. Therefore, the lift augmentation from the high-lift propellers is 1.7 and the total lift augmentation from the high-lift system (30 deg flap deflection and the high-lift blowing) is 2.38. The drag for the high-lift wing with 30 deg flap deflection is much higher than the cruise wing configuration, but the high-lift system is used only during a small portion of the flight envelope. The pitching moment is relatively constant for both blown and unblown conditions when the stabilator is excluded. Modeling the full geometry has indicated some adverse effects from the fuselage on the wing and stabilator. At high angles of attack, the solutions with the USM3D flow solver using the Spalart-Allmaras turbulence model indicates large flow separation on the wing upper surface between the two high-lift nacelles near the fuselage, and also a reduction in sectional lift on the stabilator in the first 50 percent of the stabilator semispan. However, the large flow separation near the fuselage is mostly eliminated in the solutions predicted with two codes, USM3D and Kestrel, using Hybrid Reynolds-averaged Navier Stokes/Large Eddy Simulation turbulence models.

Deere, Karen A.↗
NASA NTRS1974

Supersonic tests of an 0.015 scale space shuttle mated vehicle model (67-OTS) in the LaRC UPWT to obtain aerodynamic force data (IA42A/B)

Wind tunnel tests were conducted of the NASA/Rockwell 0.015 scale configuration 4 mated space shuttle vehicle. Data were obtained for a range of Mach numbers from 1.60 to 4.63 and angles of attack from minus 10 degrees to plus 10 degrees. A complete model build-up was performed. Longitudinal and lateral directional stability and control data were obtained for the tank alone, tank plus solid rocket boosters, and mated configuration of tank plus orbiter plus solid rocket boosters. Single component rudder hinge moment data were obtained at rudder deflections of zero degrees and minus 20 degrees for each Mach number tested.

Hardin, R.↗
NASA NTRS1993

A wind model for launch vehicle design

A new vector wind profile model is described based on the properties of the quadrivariate normal distribution which will help establish preliminary design ranges for angle of attack, angle of sideslip, aerodynamic pressure, and other launch vehicle flight variables. The model produces wind profiles that are a reasonable substitute for measured wind profile samples. The model wind profiles produce dispersion in aerodynamic load indicators that cover the dispersion range calculated from an extensive sample of Jimsphere wind profiles. This is accomplished with only 12 model wind profiles compared to 1800 Jimsphere profiles for a selected reference altitude. This represents an opportunity for a considerable reduction of computational effort during design phases that require many iterations to establish the merit of various design philosophies.

Adelfang, S. I.↗
NASA NTRS2010

Validating Human Performance Models of the Future Orion Crew Exploration Vehicle

NASA's Orion Crew Exploration Vehicle (CEV) will provide transportation for crew and cargo to and from destinations in support of the Constellation Architecture Design Reference Missions. Discrete Event Simulation (DES) is one of the design methods NASA employs for crew performance of the CEV. During the early development of the CEV, NASA and its prime Orion contractor Lockheed Martin (LM) strived to seek an effective low-cost method for developing and validating human performance DES models. This paper focuses on the method developed while creating a DES model for the CEV Rendezvous, Proximity Operations, and Docking (RPOD) task to the International Space Station. Our approach to validation was to attack the problem from several fronts. First, we began the development of the model early in the CEV design stage. Second, we adhered strictly to M&S development standards. Third, we involved the stakeholders, NASA astronauts, subject matter experts, and NASA's modeling and simulation development community throughout. Fourth, we applied standard and easy-to-conduct methods to ensure the model's accuracy. Lastly, we reviewed the data from an earlier human-in-the-loop RPOD simulation that had different objectives, which provided us an additional means to estimate the model's confidence level. The results revealed that a majority of the DES model was a reasonable representation of the current CEV design.

Wong, Douglas T.↗
NASA NTRS2000

Dynamics of Active Separation Control at High Reynolds Numbers

A series of active flow control experiments were recently conducted at high Reynolds numbers on a generic separated configuration. The model simulates the upper surface of a 20% thick Glauert-Goldschmied type airfoil at zero angle of attack. The flow is fully turbulent since the tunnel sidewall boundary layer flows over the model. The main motivation for the experiments is to generate a comprehensive data base for validation of unsteady numerical simulation as a first step in the development of a CFD design tool, without which it would not be possible to effectively utilize the great potential of unsteady flow control. This paper focuses on the dynamics of several key features of the baseline as well as the controlled flow. It was found that the thickness of the upstream boundary layer has a negligible effect on the flow dynamics. It is speculated that separation is caused mainly by the highly convex surface while viscous effects are less important. The two-dimensional separated flow contains unsteady waves centered on a reduced frequency of 0.8, while in the three dimensional separated flow, frequencies around a reduced frequency of 0.3 and 1 are active. Several scenarios of resonant wave interaction take place at the separated shear-layer and in the pressure recovery region. The unstable reduced frequency bands for periodic excitation are centered on 1.5 and 5, but these reduced frequencies are based on the length of the baseline bubble that shortens due to the excitation. The conventional swept wing-scaling works well for the coherent wave features. Reproduction of these dynamic effects by a numerical simulation would provide benchmark validation.

Pack, LaTunia G.↗
NASA NTRS2000

Dynamics of Active Separation Control at High Reynolds Numbers

A series of active flow control experiments were recently conducted at high Reynolds numbers on a generic separated configuration. The model simulates the upper surface of a 20% thick Glauert-Goldschmied type airfoil at zero angle of attack. The flow is fully turbulent since the tunnel sidewall boundary layer flows over the model. The main motivation for the experiments is to generate a comprehensive data base for validation of unsteady numerical simulation as a first step in the development of a CFD design tool, without which it would not be possible to effectively utilize the great potential of unsteady flow control. This paper focuses on the dynamics of several key features of the baseline as well as the controlled flow. It was found that the thickness of the upstream boundary layer has a negligible effect on the flow dynamics. It is speculated that separation is caused mainly by the highly convex surface while viscous effects are less important. The two-dimensional separated flow contains unsteady waves centered on a reduced frequency of 0.9, while in the three dimensional separated flow, frequencies around a reduced frequency of 0.3 and 1 are active. Several scenarios of resonant wave interaction take place at the separated shear-layer and in the pressure recovery region. The unstable reduced frequency bands for periodic excitation are centered on 1.5 and 5, but these reduced frequencies are based on the length of the baseline bubble that shortens due to the excitation. The conventional works well for the coherent wave features. Reproduction of these dynamic effects by a numerical simulation would provide benchmark validation.

Pack, LaTunia G.↗
NASA NTRS1987

Low speed wind tunnel test of a propulsive wing/canard concept in the STOL configuration. Volume 1: Test description and discussion of results

A propulsive wing/canard model was tested at STOL operating conditions in the NASA Langley Research Center 4 x 7 meter wind tunnel. Longitudinal and lateral/directional aerodynamic characteristics were measured for various flap deflections, angles of attack and sideslip, and blowing coefficients. Testing was conducted for several model heights to determine ground proximity effects on the aerodynamic characteristics. Flow field surveys of local flow angles and velocities were performed behind both the canard and the wing. This is volume 1 of a 2 volume report. The model, instrumentation, and test procedures are described. An analysis of the data is included.

Stewart, V. R.↗
NASA NTRS1987

Low speed wind tunnel test of a propulsive wing/canard concept in the STOL configuration. Volume 2: Test data

A propulsive wind/canard model was tested at STOL operating conditions in the NASA Langley Research Center 4 x 7 meter wind tunnel. Longitudinal and lateral/directional aerodynamic characteristics were measured for various flap deflections, angles of attack and sideslip, and blowing coefficients. Testing was conducted for several model heights to determine ground proximity effects on the aerodynamic characteristics. Flow field surveys of local flow angles and velocities were performed behind both the canard and the wing. This is volume 2 of a 2 volume report. All of the test data in three appendices are presented. Appendix A presented tabulated six component force and moment data, Appendix B presents tabulated wing pressure coefficients, and Appendix C presents the flow field data.

Stewart, V. R.↗
NASA NTRS1989

Supersonic aerodynamic characteristics of a proposed Assured Crew Return Capability (ACRC) lifting-body configuration

An investigation was conducted in the Langley Unitary Plan Wind Tunnel at Mach numbers from 1.6 to 4.5. The model had a low-aspect-ratio body with a flat undersurface. A center fin and two outboard fins were mounted on the aft portion of the upper body. The outboard fins were rolled outboard 40 deg from the vertical. Elevon surfaces made up the trailing edges of the outboard fins, and body flaps were located on the upper and lower aft fuselage. The center fin pivoted about its midchord for yaw control. The model was longitudinally stable about the design center-of-gravity position at 54 percent of the body length. The configuration with undeflected longitudinal controls trimmed near 0 deg angle of attack at Mach numbers from 1.6 to 3.0 where lift and lift-drag ratio were negative. Longitudinal trim was near the maximum lift-drag ratio (1.4) at Mach 4.5. The model was directionally stable over Mach number range except at angles of attack around 4 deg at M = 2.5. Pitch control deflection of more than -10 deg with either elevons or body flaps is needed to trim the model to angles of attack at which lift becomes positive. With increased control deflection, the lifting-body configuration should perform the assured crew return mission through the supersonic speed range.

Ware, George M.↗
NASA NTRS2017

Static Wind Tunnel Testing of a Legged Venus Lander

A static wind tunnel test was conducted to determine the static aerodynamic coe cients of the surface lander design proposed by the Venus In-Situ Atmospheric and Geochemical Explorer (VISAGE). The baseline design of the lander consists of a spherical pressure vessel with a top-mounted circular drag plate for aerodynamic deceleration, and three xed landing legs. A modular subscale lander model, which allowed testing of several drag plate and landing leg con gurations, was fabricated using rapid-prototyping techniques. The model was tested with four di erent drag plates and three di erent leg con gurations, at a dynamic pressure of 1610 Pa and Mach number of 0.15. Testing was conducted using legs of varying diameter in order to determine the appropriate scaling of the results to Reynolds numbers representative of terminal descent at Venus. The static aerodynamic force (axial, normal, and side), and moment (rolling, pitching, yawing) coe cients for each lander con guration were determined as a function of model orientation. All model con gurations were found to be statically stable in the range of total angles of attack considered (0 to 30 deg). However, the drag performance and degree of static stability of the model were dependent on the dihedral angle of the circular drag plate. The results will enable the creation of a preliminary aerodatabase for the VISAGE lander concept, and allow the e ect of design changes on aerodynamic performance to be evaluated.

Rabinovitch, Jason↗
NASA NTRS1989

Numerical simulation and comparison of symmetrical/supercritical airfoils for the near tip region of a helicopter in forward flight

Aerodynamic loads on a multi-bladed helicopter rotor in forward flight at transonic tip conditions are calculated. The unsteady, three-dimensional, time-accurate compressible Reynolds-averaged thin layer Navier-Stokes equations are solved in a rotating coordinate system on a body-conformed, curvilinear grid of C-H topology. Detailed boundary layer and global numerical comparisons of NACA-0012 symmetrical and CAST7-158 supercritical airfoils are made under identical forward flight conditions. The rotor wake effects are modeled by applying a correction to the geometric angle of attack of the blade. This correction is obtained by computing the local induced downwash velocity with a free wake analysis program. The calculations are performed on the Numerical Aerodynamic Simulation Cray 2 and the VPS32 (a derivative of a Cyber 205 at the Langley Research Center) for a model helicopter rotor in forward flight.

Badavi, F. F.↗
NASA NTRS2013

Hybrid Wing Body (HWB) Slat Noise Analysis

This paper presents an analysis of the slat noise for Hybrid Wing Body (HWB) aircraft, based on a database from a 3% scale wind tunnel test. It is shown that the HWB slats are one of the dominant noise components, characterized by its broad spectral shape with a peak frequency that depends on both the mean flow velocity and the aircraft angle of attack, the former following the conventional Strouhal number scaling and the latter explainable by the dependence of the coherence length of the slat unsteady flows on the aircraft angle of attack. While the overall levels of the slat noise are shown to approximately follow the fifth power law in the flow Mach number, the effects of the Mach number manifest themselves in the noise spectra in both the amplitude and the spectral shape. The slat noise amplitude is shown to also depend on the angle of attack, assuming a minimum in the range of 3 to 5 degrees and increasing when the angle of attack moves away from this range. These features are all modeled and incorporated in slat noise prediction methodologies, extending the prediction capability from conventional aircraft designs to HWB configurations. Comparisons between predictions and data show very good agreements both in various parametric trends and in the absolute levels. The HWB aircraft is designed to operate at angles of attack much higher than those of conventional aircraft. This is shown to significantly increase the HWB slat noise. To further illustrate, the test data are extrapolated to full scale and compared with the slat noise of the Boeing 777 aircraft, showing that the former is higher the latter.

Guo, Yueping↗
NASA NTRS1974

Hypersonic stability and control characteristics of the Rockwell International 139-B space shuttle orbiter

The purpose of this study was to determine the hypersonic aerodynamic performance, longitudinal trim, static stability, and the extent of the center-of-gravity boundaries within the design control limits as well as to examine the effect of Reynolds number on the longitudinal stability of the 139-B space shuttle orbiter. A 0.004 scale model (34-0)was tested in helium over an angle-of-attack range of 18 deg to 54 deg at 0 deg and -minus 5 deg sideslip angles. The total pressure was varied for selected configurations to obtain a Reynolds number range, based on model length, from 0.51 million to 3.6 million at Mach numbers between 17.6 and 21.6 Surface oil-flow studies and electron-beam illuminated-flow studies were also conducted at angles-of-attack of 30 deg, 40 deg, and 50 deg for selected Reynolds numbers.

Stone, D. R.↗
NASA NTRS1960

Investigation of the Subsonic Stability and Control Characteristics of a 1/7-Scale Model of the North American X-15 Airplane with and without Fuselage Forebody Strakes

An investigation of the low-subsonic stability and control characteristics of a l/7-scale free-flying model modified to represent closely the North American X-15 airplane (configuration 3) has been made in the Langley full-scale tunnel. Flight conditions at a relatively low altitude were simulated with the center of gravity at 16.0 percent of the mean aerodynamic chord. The longitudinal stability and control were considered to be satisfactory for all flight conditions tested. The lateral flight behavior was generally satisfactory for angles of attack below about 20 deg. At higher angles, however, the model developed a tendency to fly in a side-slipped attitude because of static directional instability at small sideslip angles. Good roll control was maintained to the highest angles tested, but rudder effectiveness diminished with increasing angle of attack and became adverse for angles above 40 deg. Removal of the lower rudder had little effect on the lateral flight characteristics for angles of attack less than about 20 deg but caused the lateral flight behavior to become worse in the high angle-of-attack range. The addition of small fuselage forebody strakes improved the static directional stability and lateral flight behavior of both configurations.

Hassell, James L., Jr.↗