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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.

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At least 217 records · Page 12
All sourcesNASA NTRSDOE OSTI
NASA NTRS1990

Improved tangent-cone method for the Aerodynamic Preliminary Analysis System (APAS) version of the hypersonic arbitrary-body program

The Aerodynamic Preliminary Analysis System (APAS) utilizes a modified version of the Hypersonic Arbitrary-Body Program (HABP) Mark 3 code in its analysis rationale. Four methods are considered for incorporation into the code as the tangent-cone method. The combination of second-order slender body theory and the approximate solution of Hammitt and Murthy shows the best agreement with the exact numerical solutions and is thus included in the APAS production version of the HABP code.

Cruz, Christopher I.↗
NASA NTRS1986

Wall temperature control of low-speed body drag

The use of thermal means to control drag under turbulent boundary layer conditions is examined. Numerical calculations are presented for both skin friction and (unseparated) pressure drag for turbulent boundary-layer flows over a fuselage-like body with wall heat transfer. In addition, thermal control of separation on a bluff body is investigated. It is shown that a total drag reduction of up to 20 percent can be achieved for wall heating with a wall-to-total-freestream temperature ratio of 2. For streamlined slender bodies, partial wall heating of the forebody can produce almost the same order of total drag reduction as the full body heating case. For bluff bodies, the separation delay from partial wall cooling of the afterbody is approximately the same as for the fully cooled body.

Lin, J. C.↗
NASA NTRS1975

Analytical and experimental study of the effects of wing-body aerodynamic interaction on space shuttle subsonic flutter

The effects on flutter of the aerodynamic interaction between the space shuttle bodies and wing, 1/80th-scale semispan models of the orbiter wing, the complete shuttle and intermediate component combinations were tested in the NASA Langley Research Center 26-inch Transonic Blowdown Wind Tunnel. Using the double lattice method combined with slender body theory to calculate unsteady aerodynamic forces, subsonic flutter speeds were computed for comparison. Using calculated complete vehicle modes, flutter speed trends were computed for the full scale vehicle at an altitude of 15,200 meters and a Mach number of 0.6. Consistent with findings of the model studies, analysis shows the shuttle to have the same flutter speed as an isolated cantilevered wing.

Chipman, R. R.↗
NASA NTRS1959

Subsonic Wing Loadings on a 45 deg Sweptback Wing and Body Combination at High Angles of Attack

A study has been made of the subsonic pressure distributions and loadings for a 45 deg sweptback-wing and body combination at angles of attack up to 36 deg. The wing had an aspect ratio of 5.5, a taper ratio of 0.53, and NACA 64A010 sections normal to the quarter-chord line and was mounted on a slender body of fineness ratio 12.5. Test results are presented for Mach numbers of 0.30 and 0.50 with corresponding Reynolds numbers of 1.5 and 2.0 million, respectively. The stall patterns and spanwise loadings at high angles of attack for the present model are correlated with those for other 45 deg sweptback wing and body combinations having aspect ratios between 4.0 and 8.0. A tentative approach is presented for extrapolating the Weissinger span-loading method to higher angles of attack, and for deriving the spanwise-load distributions for 45 deg sweptback wings at angles of attack above 20 deg. The investigation also included tests of the body in combination with only one panel of the swept wing. The problem of estimating the normal-force coefficient for the single panel at high angles of attack is considered.

Axelson, John A↗
NASA NTRS1960

Supersonic Aerodynamic Characteristics of a Low-Drag Aircraft Configuration having an Arrow Wing of Aspect Ratio 1.86 and a Body of Fineness Ratio 20

A free-flight rocket-propelled-model investigation was conducted at Mach numbers of 1.2 to 1.9 to determine the longitudinal and lateral aero-dynamic characteristics of a low-drag aircraft configuration. The model consisted of an aspect-ratio -1.86 arrow wing with 67.5 deg. leading-edge sweep and NACA 65A004 airfoil section and a triangular vertical tail with 60 deg. sweep and NACA 65A003 section in combination with a body of fineness ratio 20. Aerodynamic data in pitch, yaw, and roll were obtained from transient motions induced by small pulse rockets firing at intervals in the pitch and yaw directions. From the results of this brief aerodynamic investigation, it is observed that very slender body shapes can provide increased volumetric capacity with little or no increase in zero-lift drag and that body fineness ratios of the order of 20 should be considered in the design of long-range supersonic aircraft. The zero-lift drag and the drag-due-to-lift parameter of the test configuration varied linearly with Mach number. The maximum lift-drag ratio was 7.0 at a Mach number of 1.25 and decreased slightly to a value of 6.6 at a Mach number of 1.81. The optimum lift coefficient, normal-force-curve slope, lateral-force-curve slope, static stability in pitch and yaw, time to damp to one-half amplitude in pitch and yaw, the sum of the rotary damping derivatives in pitch and also in yaw, and the static rolling derivatives all decreased with an increase in Mach number. Values of certain rolling derivatives were obtained by application of the least-squares method to the differential equation of rolling motion. A comparison of the experimental and calculated total rolling-moment-coefficient variation during transient oscillations of the model indicated good agreement when the damping-in-roll contribution was included with the static rolling-moment terms.

Gillespie, Warren, Jr.↗
NASA NTRS1976

Economy of flight at supersonic speeds

Prandtl's theory is used to determine the airflow over bodies and wings adapted to supersonic flight. By making use of these results, and by incorporating in them an allowance for the probable skin friction, some estimates of expected lift-drag ratios are made for various flight speeds with the best configuration. At each speed a slender body and wings having the best angle of sweepback are considered. For the range of supersonic speeds shown an airplane of normal density and loading would be required to operate at an altitude of the order of 60,000 feet. The limiting value of 1-1/2 times the speed of sound corresponds to a flight speed of 1000 miles per hour. At this speed about 1.5 miles per gallon of fuel are expected. It is interesting to note that this value corresponds to a value of more than 15 miles per gallon when the weight is reduced to correspond to that of an ordinary automobile.

Jones, R. T.↗
NASA NTRS1987

Control of vortical separation on conical bodies

In a variety of aeronautical applications, the flow around conical bodies at incidence is of interest. Such applications include, but are not limited to, highly maneuverable aircraft with delta wings, the aerospace plane and nose portions of spike inlets. The theoretical model used has three parts. First, the single line vortex model is used within the framework of slender body theory, to compute the outer inviscid field for specified separation lines. Next, the three dimensional boundary layer is represented by a momentum equation for the cross flow, analogous to that for a plane boundary layer; a von Karman Pohlhausen approximation is applied to solve this equation. The cross flow separation for both laminar and turbulent layers is determined by matching the pressure at the upper and lower separation points. This iterative procedure yields a unique solution for the separation lines and consequently for the position of the vortices and the vortex lift on the body. Lastly, control of separation is achieved by blowing tangentially from a slot located along a cone generator. It is found that for very small blowing coefficients, the separation can be postponed or suppressedy completely.

Mourtos, Nikos J.↗
NASA NTRS1975

A free flight investigation of transonic sting interference

Transonic sting interference has been studied in a supersonic wind tunnel to obtain free flight and sting support data on identical models. The two principal configurations, representing fuselage bodies, were cigar shaped with tail fins. The others were a sharp 10-deg cone, a sphere, and a blunt entry body. Comparative data indicated that the sting had an appreciable effect on drag for the fuselage-like configurations; drag rise occurred 0.02 Mach number earlier in free flight, and drag level was 15% greater. The spheres and the blunt bodies were insensitive to the presence of stings regardless of their size. The 10-deg cones were in between, experiencing no drag difference with a minimum diameter sting, but a moderate difference with the largest diameter sting tested. All data tend to confirm the notion that for the more slender bodies the sting not only affects flow but the forebody flow as well.

Jaffe, P.↗
NASA NTRS1974

Second-order small disturbance theory for hypersonic flow over power-law bodies

A mathematical method for determining the flow field about power-law bodies in hypersonic flow conditions is developed. The second-order solutions, which reflect the effects of the second-order terms in the equations, are obtained by applying the method of small perturbations in terms of body slenderness parameter to the zeroth-order solutions. The method is applied by writing each flow variable as the sum of a zeroth-order and a perturbation function, each multiplied by the axial variable raised to a power. The similarity solutions are developed for infinite Mach number. All results obtained are for no flow through the body surface (as a boundary condition), but the derivation indicates that small amounts of blowing or suction through the wall can be accommodated.

Townsend, J. C.↗
NASA NTRS1975

Second-order small-disturbance solutions for hypersonic flow over power-law bodies

Similarity solutions were found which give the adiabatic flow of an ideal gas about two-dimensional and axisymmetric power-law bodies at infinite Mach number to second order in the body slenderness parameter. The flow variables were expressed as a sum of zero-order and perturbation similarity functions for which the axial variations in the flow equations separated out. The resulting similarity equations were integrated numerically. The solutions, which are universal functions, are presented in graphic and tabular form. To avoid a singularity in the calculations, the results are limited to body power-law exponents greater than about 0.85 for the two-dimensional case and 0.75 for the axisymmetric case. Because of the entropy layer induced by the nose bluntness (for power-law bodies other than cones and wedges), only the pressure function is valid at the body surface. The similarity results give excellent agreement with the exact solutions for inviscid flow over wedges and cones having half-angles up to about 20 deg. They give good agreement with experimental shock-wave shapes and surface-pressure distributions for 3/4-power axisymmetric bodies, considering that Mach number and boundary-layer displacement effects are not included in the theory.

Townsend, J. C.↗
NASA NTRS1981

Experimental and theoretical supersonic lateral-directional stability characteristics of a simplified wing-body configuration with a series of vertical-tail arrangements

An experimental investigation was conducted to provide a systematic set of lateral-directional stability data for a simplified wing-body model with a series of vertical-tail arrangements. The study was made at Mach numbers from 1.60 to 2.86 at nominal angles of attack from -8 to 12 deg and Reynolds number of 8.2 million per meter. Comparisons at zero angle of attack were made with three existing theoretical methods (MISLIFT - a second-order shock expansion and panel method; APAS - a slender body and first order panel method; and PAN AIR - a higher order panel method) and comparisons at angle of attack were made with PAN AIR. The results show that PAN AIR generally provides accurate estimates of these characteristics at moderate angles of attack for complete configurations with either single or twin vertical tails. APAS provides estimates for complete configurations at zero angle of attack. However, MISLIFT only provides estimates for the simplest body-vertical-tail configurations at zero angle of attack.

Lamb, M.↗
NASA NTRS1986

The equivalent angle-of-attack concept for engineering analysis

An account is given of methods for the estimation of a wing-body-tail missile configuration's aerodynamic performance by means of the 'component buildup' method, in which the overall aerodynamic loads for an airframe are built up from the assumed characteristics of isolated airframe components and then subjected to additional loads generated by component interference. Attention is given to the cases of missile airframes in steady flow at arbitrary angles of attack and bank; the unifying thread of the present treatment is slender body theory, together with its nonlinear extensions through the equivalent angle-of-attack concept. The estimation of the forces and moments acting on each of the fins is emphasized, so that control cross-coupling effects as well as longitudinal and lateral directional effects can be determined.

Hemsch, M. J.↗
NASA NTRS1959

Approximate Analytical Solutions for Hypersonic Flow Over Slender Power Law Bodies

Approximate analytical solutions are presented for two-dimensional and axisymmetric hypersonic flow over slender power law bodies. Both zero order (M approaches infinity) and first order (small but nonvanishing values of 1/(M(Delta)(sup 2) solutions are presented, where M is free-stream Mach number and Delta is a characteristic slope. These solutions are compared with exact numerical integration of the equations of motion and appear to be accurate particularly when the shock is relatively close to the body.

Mirels, Harold↗
NASA NTRS1983

Control of the forebody vortex orientation by asymmetric air injection. Part A: Application to enhance departure/spin recovery of fighter aircraft and Part B: Details of the flow structure

A concept developed to provide powerful directional control effectiveness for a fighter aircraft at high angles of attack is described. The concept utilizes the energy concentrated in the strong forebody vortices (which form on slender bodies of high relative incidence) by controlling the lateral orientation of the vortices with respect to the body. The objective was to utilize the side force associated with asymmetric vortices, in a controlled manner, to enhance the ability of the fighter to recover from a departure from controlled flight. The results from water tunnel and wind tunnel experiments show that a small amount of tangential blowing along the forebody near the apex can effectively alter the forebody vortex system and generate large restoring yawing moments. Six degree of freedom digital simulation results show that this concept can substantially enhance recovery characteristics of fighter aircraft with long, slender forebodies. Also, the results of experiments which were conducted on a cone model are discussed where the principal test objective was to develop an understanding of the fluid mechanics involved in the process of vortex control. Knowledge gained in these more generic tests should allow the concept to be applied to a wider range of configurations.

Skow, A. M.↗
NASA NTRS1987

Subsonic wind-tunnel measurements of a slender wing-body configuration employing a vortex flap

A wind tunnel study at Mach 0.4 was conducted for a slender wing-body configuration with a leading edge vortex flap of curved planform that is deflectable about a 74 degree swept hinge line. The basic data consist of a unique combination of longitudinal aerodynamic, surface pressure, and vortex flap hinge-moment measurements on a common model. The longitudinal aerodynamic, pressure and hinge-moment data are presented without analysis in tabular format. Plots of the tabulated pressure data are also given.

Frink, Neal T.↗
NASA NTRS1957

Three-Dimensional Transonic Flow Theory Applied to Slender Wings and Bodies

The present paper re-examines the derivation of the integral equations for transonic flow around slender wings and bodies of revolution, giving special attention to conditions resulting from the presence of shock waves and to the reduction of the relations to the special forms necessary for the discussion of sonic flow, that is, flow at free-stream Mach number 1.

Heaslet, Max A↗