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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 271 records · Page 15
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NASA NTRSDate not supplied

Flow and noise from supersonic plug nozzles

Nozzles with external plugs are candidates for propulsion of future supersonic aircraft. A model-scale experimental study is conducted exploring the flow and noise characteristics of various plug nozzles. For a given outer nozzle, the plug geometry is varied from conic with various half-angles (lengths) to plugs designed by the method of characteristics (MoC) as well as truncated and porous plugs. So far, noise characteristics and schlieren flow visualization data have been acquired. Limited numerical simulations have also been done addressing thrust performance. The simulations show that the MoC plugs out-perform the conic cases, especially around high Mach number cruise conditions. The experimental data, however, show that the MoC plugs are noisy relative to the conic cases at low Mach number landing and takeoff (LTO) conditions. The increased noise (with MoC at LTO) is accompanied by unsteady turbulent structures and increased jet spreading, as seen in schlieren pictures. At LTO, the longest conic plug is found to be the least noisy. Furthermore, a porous surface on the long conic plug effectively suppresses broadband shock associated noise (BBSN) at higher Mach numbers. Thus, a dilemma remains. MoC design gives best thrust performance at cruise but it is noisy at LTO. A long conic plug, on the other hand, is quieter at LTO but may have unacceptable thrust at cruise.

jets↗
NASA NTRS1973

Study of high altitude plume impingement

The radiation intensities are determined in the base region of the space shuttle due to solid particle radiation emanating from the solid rocket motors of the shuttle. Results of an analysis of the Titan 3 and simulated solid rocket motor radiation intensities are presented. The gas particle flow fields of the Titan 3 nozzle and plume and a space shuttle solid rocket motor nozzle and plume are described. The gaseous Titan 3 flow fields are discussed utilizing the results of flow fields generated by a gaseous and two phase method-of-characteristics computer programs. A two phase computer flow field analysis program was developed. An outflow correction theory is developed which will be used to modify existing convection heat transfer methods for better heat transfer predictions on bodies immersed in rocket exhaust plumes.

Mcanally, J. V.↗
NASA NTRS1980

Supersonic flow development in slotted wind tunnels

The development of test section slot shapes for achieving smooth supersonic Mach number distribution without overexpansion or waviness has, in the past, been largely an experimentally iterative or 'cutand try' procedure for each wind tunnel. To overcome the obvious disadvantages of time and expense involved in such an experimental approach, a simple analytical method has been developed to predict the supersonic flow development in a two-dimensional slotted tunnel given only the variation of open area ratio with downstream distance and the Mach number corresponding to the plenum static pressure. The well known method of characteristics is used with the constraint that it be compatible with the quadratic cross-flow pressure drop boundary condition at the slotted wall. The predicted results from this method agree remarkably well with the experimental calibration data available for some of the existing facilities. The flow mechanism responsible for causing overexpansion in the centerline Mach number distribution with some slot shapes has been brought to light.

Ramaswamy, M. A.↗
NASA NTRS1985

Linear and nonlinear dynamic analysis of redundant load path bearingless rotor systems

The bearingless rotorcraft offers reduced weight, less complexity and superior flying qualities. Almost all the current industrial structural dynamic programs of conventional rotors which consist of single load path rotor blades employ the transfer matrix method to determine natural vibration characteristics because this method is ideally suited for one dimensional chain like structures. This method is extended to multiple load path rotor blades without resorting to an equivalent single load path approximation. Unlike the conventional blades, it isk necessary to introduce the axial-degree-of-freedom into the solution process to account for the differential axial displacements in the different load paths. With the present extension, the current rotor dynamic programs can be modified with relative ease to account for the multiple load paths without resorting to the equivalent single load path modeling. The results obtained by the transfer matrix method are validated by comparing with the finite element solutions. A differential stiffness matrix due to blade rotation is derived to facilitate the finite element solutions.

Murthy, V. R.↗
NASA NTRS1974

Calculation of aerodynamic characteristics of STOL aircraft

Method predicts lift and pitching moment characteristics of STOL aircraft with externally-blown, jet-augmented wing-flap combinations using potential-flow approach which involves combination of two flow models. Method can accommodate multiple engines per wing panel and part-span flaps.

Dillenius, M. F. E.↗
NASA NTRS2010

Aeroelastic Computations of a Compressor Stage Using the Harmonic Balance Method

The aeroelastic characteristics of a compressor stage were analyzed using a computational fluid dynamic (CFD) solver that uses the harmonic balance method to solve the governing equations. The three dimensional solver models the unsteady flow field due to blade vibration using the Reynolds-Averaged Navier-Stokes equations. The formulation enables the study of the effect of blade row interaction through the inclusion of coupling modes between blade rows. It also enables the study of nonlinear effects of high amplitude blade vibration by the inclusion of higher harmonics of the fundamental blade vibration frequency. In the present work, the solver is applied to study in detail the aeroelastic characteristics of a transonic compressor stage. Various parameters were included in the study: number of coupling modes, blade row axial spacing, and operating speeds. Only the first vibration mode is considered with amplitude of oscillation in the linear range. Both aeroelastic stability (flutter) of rotor blade and unsteady loading on the stator are calculated. The study showed that for the stage considered, the rotor aerodynamic damping is not influenced by the presence of the stator even when the axial spacing is reduced by nearly 25 percent. However, the study showed that blade row interaction effects become important for the unsteady loading on the stator when the axial spacing is reduced by the same amount.

Reddy, T. S. R.↗
NASA NTRS1977

Acoustic waves in the solar atmosphere. I - The hydrodynamic code

This paper studies large-amplitude radiatively damped acoustic waves in the solar atmosphere. A modified method of characteristics is described for the solution of the time-dependent hydrodynamic equations in a gravitational atmosphere. A procedure for the detection of shocks is outlined. Several tests of the accuracy of the method are described. The evolution of the wave and the height of shock formation are computed for several values of the period and the initial acoustic flux in isothermal atmospheres with temperatures of 4000 and 5000 K as well as in a model solar atmosphere.

Ulmschneider, P.↗
NASA NTRS1982

A computer program for the calculation of the flow field including boundary layer effects for mixed-compression inlets at angle of attack

A computer program was developed which is capable of calculating the flow field in the supersonic portion of a mixed compression aircraft inlet operating at angle of attack. The supersonic core flow is computed using a second-order three dimensional method-of-characteristics algorithm. The bow shock and the internal shock train are treated discretely using a three dimensional shock fitting procedure. The boundary layer flows are computed using a second-order implicit finite difference method. The shock wave-boundary layer interaction is computed using an integral formulation. The general structure of the computer program is discussed, and a brief description of each subroutine is given. All program input parameters are defined, and a brief discussion on interpretation of the output is provided. A number of sample cases, complete with data listings, are provided.

Vadyak, J.↗
NASA NTRS1982

Study of highly sweptback wings by the free vortex sheet method

The aerodynamic characteristics of highly sweptback wings with separations induced vortex flows have been numerically investigated using the free vortex sheet method, developed by Boeing Company, under a contract with NASA/Langley Research Center. The models studied included delta and straked wings, and wings with leading edge extensions. Also, PAN-AIR code has been used to design a fixed leading edge extension into a thick delta wing. The theoretical results predicted have been compared with the experimental data wherever available, and the code capabilities and limitations explored. New fuselage effects also have been considered in some cases.

C Subba Reddy↗
NASA NTRS1961

A Method for Obtaining the Nonlinear Aerodynamic Stability Characteristics of Bodies of Revolution from Free-Flight Tests

A method is presented for obtaining the nonlinear aerodynamic stability characteristics of bodies of revolution from free-flight test.s The necessary conditions for the application of this method are: (1) that the roll rate and damping encountered in a single cycle of oscillation be small, and (2) that the resulting motion be reasonably planar. Four approximations to the nonlinear restoring moment are considered and solutions are obtained in closed form: 1. A single-term polynomial in an arbitrary power of the angle of attack. 2. A two-term polynomial having linear and cubic terms. 3. A three-term polynomial having linear, quadratic, and cubic terms. 4. A three-term polynomial having linear, quadratic, and cubic terms. An iteration procedure is formulated to allow the use of each of these approximations for obtaining the aerodynamic coefficients of bodies of revolution from free-flight test data. It is found that although the equations that are solved pertain strictly to planar motion, the solutions are applicable to motions that deviate to a fairly large degree from planar motion.

Kirk, Donn B.↗
NASA NTRS1987

Numerical optimization of conical flow waveriders including detailed viscous effects

A family of optimized hypersonic waveriders is generated and studied wherein detailed viscous effects are included within the optimization process itself. This is in contrast to previous optimized waverider work, wherein purely inviscid flow is used to obtain the waverider shapes. For the present waveriders, the undersurface is a streamsurface of an inviscid conical flowfield, the upper surface is a streamsurface of the inviscid flow over a tapered cylinder (calculated by the axisymmetric method of characteristics), and the viscous effects are treated by integral solutions of the boundary layer equations. Transition from laminar to turbulent flow is included within the viscous calculations. The optimization is carried out using a nonlinear simplex method. The resulting family of viscous hypersonic waveriders yields predicted high values of lift/drag, high enough to break the L/D barrier based on experience with other hypersonic configurations. Moreover, the numerical optimization process for the viscous waveriders results in distinctly different shapes compared to previous work with inviscid-designed waveriders. Also, the fine details of the viscous solution, such as how the shear stress is distributed over the surface, and the location of transition, are crucial to the details of the resulting waverider geometry. Finally, the moment coefficient variations and heat transfer distributions associated with the viscous optimized waveriders are studied.

Bowcutt, Kevin G.↗
NASA NTRS1988

Influence of geometric nonlinearities on skin-stiffener interface stresses

A method for computing skin-stiffener interface stresses in stiffened composite panels is developed. Both geometrically linear and nonlinear analyses are considered. Particular attention is given to the flange termination region where stresses are expected to exhibit unbounded characteristics. The method is based on a finite-element analysis and an elasticity solution. The results indicate that the inclusion of geometric nonlinearities is very important for an accurate determination of the interface stresses. Membrane flattening of the panel tends to reduce the tendency of the stiffener to separate.

Cohen, D.↗
NASA NTRS1987

Calculation of skin-stiffener interface stresses in stiffened composite panels

A method for computing the skin-stiffener interface stresses in stiffened composite panels is developed. Both geometrically linear and nonlinear analyses are considered. Particular attention is given to the flange termination region where stresses are expected to exhibit unbounded characteristics. The method is based on a finite-element analysis and an elasticity solution. The finite-element analysis is standard, while the elasticity solution is based on an eigenvalue expansion of the stress functions. The eigenvalue expansion is assumed to be valid in the local flange termination region and is coupled with the finite-element analysis using collocation of stresses on the local region boundaries. Accuracy and convergence of the local elasticity solution are assessed using a geometrically linear analysis. Using this analysis procedure, the influence of geometric nonlinearities and stiffener parameters on the skin-stiffener interface stresses is evaluated.

Cohen, David↗
NASA NTRS1984

Mach 6 flow-field and boundary-layer surveys beneath the forebody of an airbreathing missile

A wind tunnel investigation has been conducted to determine the flow characteristics of the shock layer beneath the forebody of a hypersonic, airbreathing missile incorporating an aft-mounted inlet. In the inviscid part of the flow field, the measured parameters were in agreement with those predicted by a three-dimensional Method of Characteristics theory. At test conditions matching full scale Mach and Reynolds numbers at an altitude for maximum L/D cruise, boundary layer transition occurred downstream of the inlet face. While this means that a full size vehicle would have a more easily separated boundary layer, it is noted that the Reynolds number discrepancy between wind tunnel and actual flight conditions may result in a movement of the transition point to somewhere upstream of the inlet.

Johnston, P. J.↗
NASA NTRS1973

Mach Reflection, Mach Disc, and the Associated Nozzle Free Jet Flows

The numerical method involving both the method of integral relations and the method of characteristics have been applied to investigate the steady flow phenomena associated with the accurrence of Mach reflection and Mach disc from nozzle flows. The solutions of triple-shock intersection are presented. The regime where Mach configuration appears is defines for the inviscid analysis. The method of integral relations developed for the blunt body problem is modified and extended to the attached shock wave and to internal nozzle flow problems.

Chang, I.↗
NASA NTRS1980

A study of the sonic-boom characteristics of a blunt body at a Mach number of 6

An experimental and analytical study of the sonic boom static pressure signatures generated by a blunt body at mach 6 has shown that finite difference computer programs can be used to give reasonable estimates of the signatures. The calculated near field static pressure signature was extrapolated to the far field by a program using the method of characteristics. A comparison of this extrapolated signature with the signature predicted by far field sonic boom theory (linearized) shows that peak overpressures are about the same, at least up to mach, but the far field theory overestimates the length of the signature.

Ashby, G. C., Jr.↗
NASA NTRS1999

Rotordynamics and Design Methods of an Oil-Free Turbocharger

The feasibility of supporting a turbocharger rotor on air foil bearings is investigated based upon predicted rotordynamic stability, load accommodations, and stress considerations. It is demonstrated that foil bearings offer a plausible replacement for oil-lubricated bearings in diesel truck turbochargers. Also, two different rotor configurations are analyzed and the design is chosen which best optimizes the desired performance characteristics. The method of designing machinery for foil bearing use and the assumptions made are discussed.

Howard, Samuel A.↗
NASA NTRS1984

An analysis of shock coalescence including three-dimensional effects with application to sonic boom extrapolation

A method for analyzing shock coalescence which includes three dimensional effects was developed. The method is based on an extension of the axisymmetric solution, with asymmetric effects introduced through an additional set of governing equations, derived by taking the second circumferential derivative of the standard shock equations in the plane of symmetry. The coalescence method is consistent with and has been combined with a nonlinear sonic boom extrapolation program which is based on the method of characteristics. The extrapolation program, is able to extrapolate pressure signatures which include embedded shocks from an initial data line in the plane of symmetry at approximately one body length from the axis of the aircraft to the ground. The axisymmetric shock coalescence solution, the asymmetric shock coalescence solution, the method of incorporating these solutions into the extrapolation program, and the methods used to determine spatial derivatives needed in the coalescence solution are described. Results of the method are shown for a body of revolution at a small, positive angle of attack.

Darden, C. M.↗