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At least 361 records · Page 20
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NASA NTRSDate not supplied

Testing of Two Mars Powered Descent Vehicle Concepts in the Langley Unitary Plan Wind Tunnel

Testing was conducted in the NASA Langley Unitary Plan Wind Tunnel in order to investigate the aerodynamic interference of sub-scale versions of two Mars powered descent vehicle concepts at supersonic Mach numbers (2.4 and 3.5): a model based on a blunt hypersonic inflatable aerodynamic decelerator (HIAD) and the second representing a more slender rigid vehicle with body flaps (CobraMRV). Each model was designed to accommodate up to eight nozzles, with the HIAD model having five different nozzle configurations to investigate the effects of nozzle location, cant angle, and area ratio. The models were tested with high pressure air as the nozzle plume gas, and included the following instrumentation: high-speed video, discrete steady state and high-frequency pressure, pressure sensitive paint, and a new flow-through force and moment balance for the HIAD model. The high-speed imagery showed the overall expected growth of the shock layer to increasing thrust levels. The discrete stagnation pressure data on the HIAD model with four and eight blowing nozzles was sensitive to thrust coefficient with the nozzles canted at 0 degrees, whereas having nozzles canted outward 20 degrees and/or being located closer to the heatshield shoulder largely removed that sensitivity. The CobraMRV model stagnation pressure was more sensitive to tunnel Mach number and sideslip angle, due to the nozzle arrangement and plume interference. Pressure sensitive paint data quality was compromised by paint damage from frequent model changes, especially for the HIAD model. However, pressure coefficient data on the CobraMRV model showed the same trends as the discrete pressure measurements, as well as a sensitivity to non-zero sideslip angles. The aerodynamic force coefficients were derived from the pressure sensitive paint data. On the HIAD models, the primary force coefficient decreased with increasing thrust due to the nozzle plumes blocking flow to the heatshield area surrounding the nozzle exist. On the CobraMRV model, the force coefficient was relatively insensitive to thrust coefficient at the lower Mach number when sideslip angle was 0 degrees The force coefficient decreases with increasing thrust coefficient when the sideslip angle is 10 degrees. Balance data quality was negatively impacted by thermal drift issues that were not apparent in pre-test calibration measurements, thus preventing usable test data. Results and lessons learned will be used to take further technology development steps, including more advanced ground test techniques and flight testing.

Karl T Edquist↗
NASA NTRSDate not supplied

Computational Modeling of Two Mars Powered Descent Vehicle Concepts Tested in the Langley Unitary Plan Wind Tunnel

Future human Mars missions will require powered descent starting at supersonic conditions, something which has never been done before at Mars. Computational powered descent flowfield simulations have been completed at full-scale Mars conditions, but the available ground test data are not suitable for calibrating computational uncertainties for aerodynamic interference on proposed Mars descent vehicles. Testing was conducted in the NASA Langley Unitary Plan Wind Tunnel in order to investigate the aerodynamic interference of sub-scale versions of two Mars powered descent vehicle concepts at supersonic Mach numbers (2.4 and 3.5): a model based on a blunt hypersonic inflatable aerodynamic decelerator (HIAD) and the second representing a more slender rigid vehicle with body flaps (CobraMRV). This paper covers computational flowfield predictions completed at wind tunnel conditions and comparisons to the test data. On the blowing HIAD models, the time-averaged pressure inboard of the nozzles was generally well-predicted, especially if the nozzles are canted outward, when the nozzles are located further from the nose. At intermediate CobraMRV thrust coefficients, CFD pressures are more accurately predicted than they are for the HIAD models, largely due to the nozzle locations and pointing directions. Overall, the CFD pressure coefficients were predicted within 0.2 of the steady pressure measurements for all blowing models, with smaller discrepancies at higher HIAD thrust, and larger discrepancies at lower HIAD CobraMRV thrust. All HIAD models were predicted to have a gradually decreasing axial force coefficient as the total thrust increases, in agreement with available pressure sensitive paint data. On models with canted nozzles or with nozzles further from the nose, the force coefficient was slightly higher for a given thrust. On the CobraMRV model, the CFD also shows consistent results between solvers and follows trends revealed in the data; the aerodynamic force coefficient remains near the non-blowing value at a tunnel Mach number of 2.4 regardless of thrust, and increases above that level at a Mach number of 3.5, consistent with the discrete pressure data. CFD analysis at tunnel and flight conditions will continue as flight system designs concepts mature.

Supersonic Retropropulsion↗
NASA NTRSDate not supplied

Testing of Two Mars Powered Descent Vehicle Concepts in the Langley Unitary Plan Wind Tunnel

Testing was conducted in the NASA Langley Unitary Plan Wind Tunnel in order to investigate the aerodynamic interference of sub-scale versions of two Mars powered descent vehicle concepts at supersonic Mach numbers (2.4 and 3.5): a model based on a blunt hypersonic inflatable aerodynamic decelerator (HIAD) and the second representing a more slender rigid vehicle with body flaps (CobraMRV). Each model was designed to accommodate up to eight nozzles, with the HIAD model having five different nozzle configurations to investigate the effects of nozzle location, cant angle, and area ratio. The models were tested with high pressure air as the nozzle plume gas, and included the following instrumentation: high-speed video, discrete steady state and high-frequency pressure, pressure sensitive paint, and a new flow-through force and moment balance for the HIAD model. The high-speed imagery showed the overall expected growth of the shock layer to increasing thrust levels. The discrete stagnation pressure data on the HIAD model with four and eight blowing nozzles was sensitive to thrust coefficient with the nozzles canted at 0 degrees, whereas having nozzles canted outward 20 degrees and/or being located closer to the heatshield shoulder largely removed that sensitivity. The CobraMRV model stagnation pressure was more sensitive to tunnel Mach number and sideslip angle, due to the nozzle arrangement and plume interference. Pressure sensitive paint data quality was compromised by paint damage from frequent model changes, especially for the HIAD model. However, pressure coefficient data on the CobraMRV model showed the same trends as the discrete pressure measurements, as well as a sensitivity to non-zero sideslip angles. The aerodynamic force coefficients were derived from the pressure sensitive paint data. On the HIAD models, the primary force coefficient decreased with increasing thrust due to the nozzle plumes blocking flow to the heatshield area surrounding the nozzle exist. On the CobraMRV model, the force coefficient was relatively insensitive to thrust coefficient at the lower Mach number when sideslip angle was 0 degrees The force coefficient decreases with increasing thrust coefficient when the sideslip angle is 10 degrees. Balance data quality was negatively impacted by thermal drift issues that were not apparent in pre-test calibration measurements, thus preventing usable test data. Results and lessons learned will be used to take further technology development steps, including more advanced ground test techniques and flight testing.

Supersonic Retropropulsion↗
NASA NTRSDate not supplied

Computational Modeling of Two Mars Powered Descent Vehicle Concepts Tested in the Langley Unitary Plan Wind Tunnel

Future human Mars missions will require powered descent starting at supersonic conditions, something which has never been done before at Mars. Computational powered descent flowfield simulations have been completed at full-scale Mars conditions, but the available ground test data are not suitable for calibrating computational uncertainties for aerodynamic interference on proposed Mars descent vehicles. Testing was conducted in the NASA Langley Unitary Plan Wind Tunnel in order to investigate the aerodynamic interference of sub-scale versions of two Mars powered descent vehicle concepts at supersonic Mach numbers (2.4 and 3.5): a model based on a blunt hypersonic inflatable aerodynamic decelerator (HIAD) and the second representing a more slender rigid vehicle with body flaps (CobraMRV). This paper covers computational flowfield predictions completed at wind tunnel conditions and comparisons to the test data. On the blowing HIAD models, the time-averaged pressure inboard of the nozzles was generally well-predicted, especially if the nozzles are canted outward, when the nozzles are located further from the nose. At intermediate CobraMRV thrust coefficients, CFD pressures are more accurately predicted than they are for the HIAD models, largely due to the nozzle locations and pointing directions. Overall, the CFD pressure coefficients were predicted within 0.2 of the steady pressure measurements for all blowing models, with smaller discrepancies at higher HIAD thrust, and larger discrepancies at lower HIAD CobraMRV thrust. All HIAD models were predicted to have a gradually decreasing axial force coefficient as the total thrust increases, in agreement with available pressure sensitive paint data. On models with canted nozzles or with nozzles further from the nose, the force coefficient was slightly higher for a given thrust. On the CobraMRV model, the CFD also shows consistent results between solvers and follows trends revealed in the data; the aerodynamic force coefficient remains near the non-blowing value at a tunnel Mach number of 2.4 regardless of thrust, and increases above that level at a Mach number of 3.5, consistent with the discrete pressure data. CFD analysis at tunnel and flight conditions will continue as flight system designs concepts mature.

Supersonic Retropropulsion↗
NASA NTRS1991

Definition of the unsteady vortex flow over a wing/body configuration

A problem of current interest in computational aerodynamics is the prediction of unsteady vortex flows over aircraft at high angles of attack. A six-month experimental effort was conducted at the John H. Harper Wind Tunnel to acquire qualitative and quantitative information on the unsteady vortex flow over a generic wing-body configuration at high angles of attack. A double-delta flat-plate wing with beveled edges was combined with a slender sharp-nosed body-of-revolution fuselage to form the generic configuration. This configuration produces a strong attached leading edge vortex on the wing, as well as sharply-peaked flow velocity spectra above the wing. While it thus produces flows with several well-defined features of current interest, the model was designed for efficiency of representation in computational codes. A moderate number of surface pressure ports and two unsteady pressure sensors were used to study the pressure distribution over the wing and body surface at high angles of attack; the unsteady pressure sensing did not succeed because of inadequate signal-to-noise ratio. A pulsed copper vapor laser sheet was used to visualize the vortex flow over the model, and vortex trajectories, burst locations, mutual induction of vortex systems from the forebody, strake, and wing, were quantified. Laser Doppler velocimetry was used to quantify all 3 components of the time-average velocity in 3 data planes perpendicular to the freestream direction. Statistics of the instantaneous velocity were used to study intermittency and fluctuation intensity. Hot-film anemometry was used to study the fluctuation energy content in the velocity field, and the spectra of these fluctuations. In addition, a successful attempt was made to measure velocity spectra, component by component, using laser velocimetry, and these were compared with spectra measured by hot-film anemometry at several locations.

Liou, S. G.↗
NASA NTRS1974

An experimental study of the validity of the heat-field concept for sonic-boom alleviation

An experimental program was carried out in the NASA-Langley 4 ft x 4 ft supersonic pressure tunnel to investigate the validity of the heat-field concept for sonic boom alleviation. The concept involves heating the flow about a supersonic aircraft in such a manner as to obtain an increase in effective aircraft length and yield an effective aircraft shape that will result in a shock-free pressure signature on the ground. First, a basic body-of-revolution representing an SST configuration with its lift equivalence in volume was tested to provide a baseline pressure signature. Second, a model having a 5/2-power area distribution which, according to theory, should yield a linear pressure rise with no front shock wave was tested. Third, the concept of providing the 5/2-power area distribution by using an off-axis slender fin below the basic body was investigated. Then a substantial portion (approximately 40 percent) of the solid fin was replaced by a heat field generated by passing heated nitrogen through the rear of the fin.

Swigart, R. J.↗
NASA NTRS1977

A study of the sonic-boom characteristics of a blunt body at a Mach number of 4.14

An experimental and theoretical study has shown that the applicability of far-field sonic-boom theory previously demonstrated for more slender shapes may now be extended to bodies with ratios of diameter to length as great as 2 and to Mach numbers at least as high as 4.14. This finding is of special significance in view of the limitations to the use of existing methods for the extrapolation of close-in experimental data.

Carlson, H. W.↗
NASA NTRS1966

Semiconductor forms biomedical radiation probe

Semiconductor radiation dosimeter in the form of a slender probe is easily inserted into body tissue. The probe has a signal-to-noise ratio that is acceptable to recording equipment and provides realistic measurements of the spatial and energy distributions of radiant electrons and protons.

Burns, F. P.↗
NASA NTRS1962

Tables of Aerodynamic Coefficients Obtained from Developed Newtonian Expressions for Complete and Partial Conic and Spheric Bodies at Combined Angles of Attack and Sideslip with Some Comparisons with Hypersonic Experimental Data

Closed-form expressions and tables composed from these expressions are presented for complete and partial conic and spheric bodies at combined angles of attack and sideslip in Newtonian flow. Aerodynamic coefficients of these bodies are tabulated for various body segments over a range of angles of attack from 1 deg to 85 deg and angles of sideslip from 0 deg to 15 deg. Some comparisons between Newtonian predictions and hypersonic experimental aerodynamic characteristics were made for conic bodies hawing various surface slopes, nose bluntnesses, and body cross sections to indicate the range of validity of the theory. In general, the theory is shown to agree quite well with experimental results for sharp-nose complete cones and for configurations hawing large blunted noses and steep surface slopes. However, agreement between theory and experiment generally is poor for the more slender, slightly blunted complete or half conic bodies and also for sharp-nose half conic bodies where real-flow phenomena such as forebody interference, viscous forces, leeward surface contributions, or leading-edge pressure reductions may have significant effect. The agreement between theory and experiment for the bodies considered can be improved by using the stagnation pressure coefficient behind a normal shock rather than 2 as the Newtonian coefficient, although for the sharp-nose half conic bodies there i s no theoretical justification for this modification.

Armstrong, William O.↗
NASA NTRS1992

Prediction and control of slender-wing rock

The unsteady Euler equations and the Euler equations of rigid-body dynamics, both written in the moving frame of reference, are sequentially solved to simulate the limit-cycle rock motion of slender delta wings. The governing equations of the fluid flow and the dynamics of the present multidisciplinary problem are solved using an implicit, approximately-factored, central-difference-like, finite-volume scheme and a four-stage Runge-Kutta scheme, respectively. For the control of wing-rock motion, leading-edge flaps are forced to oscillate anti-symmetrically at prescribed frequency and amplitude, which are tuned in order to suppress the rock motion. Since the computational grid deforms due to the leading-edge flaps motion, the grid is dynamically deformed using the Navier-displacement equations. Computational applications cover locally-conical and three-dimensional solutions for the wing-rock simulation and its control.

Kandil, Osama A.↗
NASA NTRS1953

The similarity law for hypersonic flow and requirements for dynamic similarity of related bodies in free flight

The similarity law for nonsteady, inviscid, hypersonic flow about slender three-dimensional shapes is derived. Conclusions drawn are shown to be valid for rotational flow. Requirements for dynamic similarity of related shapes in free flight are obtained. The law is examined for steady flow about related three-dimensional shapes. Results of an experimental investigation of the pressures acting on two inclined cones are found to check the law as it applies to bodies of revolution.

Hamaker, Frank M↗
NASA NTRS1973

Comparison of the aerodynamic characteristics of an ablating and nonablating blunted conical body

The influence of ablation on the aerodynamic characteristics of a blunted slender cone was investigated. Plastic models were launched in free flight at ablating conditions. The results were compared with results of similar tests using metal nonablating models. Ablation was found to decrease the dynamic stability and the drag, but had little effect on static stability and lift. The plastic models appeared to experience ablation-induced roll.

Kruse, R. L.↗
NASA NTRS1993

Application of an engineering inviscid-boundary layer method to slender three-dimensional vehicle forebodies

An engineering inviscid-boundary layer method has been modified for application to slender three-dimensional (3-D) forebodies which are characteristic of transatmospheric vehicles. An improved shock description in the nose region has been added to the inviscid technique which allows the calculation of a wider range of body geometries. The modified engineering method is applied to the perfect gas solution over a slender 3-D configuration at angle of attack. The method predicts surface pressures and laminar heating rates on the windward side of the vehicle that compare favorably with numerical solutions of the thin-layer Navier-Stokes equations. These improvements extend the 3-D capabilities of the engineering method and significantly increase its design applications.

Riley, Christopher J.↗
NASA NTRS1993

Application of an Engineering Inviscid-Boundary Layer Method to Slender Three-Dimensional Vehicle Forebodies

An engineering inviscid-boundary layer method has been modified for application to slender three-dimensional (3-D) forebodies which are characteristic of transatmospheric vehicles. An improved shock description in the nose region has been added to the inviscid technique which allows the calculation of a wider range of body geometries. The modified engineering method is applied to the perfect gas solution over a slender 3-D configuration at angle of attack. The method predicts surface pressures and laminar heating rates on the windward side of the vehicle that compare favorably with numerical solutions of the thin-layer Navier-Stokes equations. These improvements extend the 3-D capabilities of the engineering method and significantly increase its design applications.

Riley, Christopher J.↗
NASA NTRS1975

Hypersonic ionizing air viscous shock-layer flows over nonanalytic blunt bodies

The equations which govern the viscous shock-layer flow are presented and the method by which the equations are solved is discussed. The predictions of the present finite-difference method are compared with other numerical predictions as well as with experimental data. The principal emphasis is placed on predictions of the viscous flowfield for the windward plane of symmetry of the space shuttle orbiter and other axisymmetric bodies which approximate the shuttle orbiter geometry. Experimental data on two slender sphere-cones at hypersonic conditions are also considered. The present predictions agreed well with experimental data and with the past predictions. Substantial differences were found between present predictions and more approximate methods.

Miner, E. W.↗
NASA NTRS1979

Symmetrical and Asymmetrical separations about a yawed cone

Three-dimensional flow separations about a 5 degree (semiapex angle, theta sub C), 1.4 m long, circular cone up to moderately high relative incidence, alpha/theta sub C approximately 5, were studied in the Mach number range 0.3 M sub infinity 1.8. The cone was tested in the Ames 1.8 by 1.8 m wind tunnel at Reynolds numbers, R sub L infinity, based on the cone length, L, from 4.5 times 10 to the 6th power to 13.5 times 10 to the 6th power, under nominally zero heat transfer conditions. Overall forces and mean surface pressures were compared with earlier measurements. Supportive three-dimensional laser velocimeter measurements of mean and fluctuating velocity in a slightly asymmetric vortex wake about a slender tangent ogive cylinder at incidence having respective nose and overall body fineness ratios of 3.5 and 12, are included.

Peake, D. J.↗
NASA NTRS1996

Natural Rolling Responses of a Delta Wing in Transonic and Subsonic Flows

The unsteady, three-dimensional, full Navier-Stokes (NS) equations and the Euler equations of rigid-body dynamics are sequentially solved to simulate the natural rolling response of slender delta wings of zero thickness at moderate to high angles of attack, to transonic and subsonic flows. The governing equations of fluid flow and dynamics of the present multi-disciplinary problem are solved using the time-accurate solution of the NS equations with the implicit, upwind, Roe flux-difference splitting, finite-volume scheme and a four-stage Runge-Kutta scheme, respectively. The main focus is to analyze the effect of Mach number and angle of attack on the leading edge vortices and their breakdown, the resultant rolling motion, and overall aerodynamic response of the wing. Three cases demonstrate the natural response of a 65 deg swept, cropped delta wing in a transonic flow with breakdown of the leading edge vortices and an 80 deg swept delta wing in a subsonic flow undergoing either damped or self-excited limit-cycle rolling oscillations as a function of angle of attack. Comparisons with an experimental investigation completes this study, validating the analysis and illustrating the complex details afforded by computational investigations.

Menzies, Margaret A.↗