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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 73 records · Page 4
All sourcesNASA NTRSDOE OSTI
NASA NTRS1984

Optimization and mechanisms of mistuning in cascades

In the present inverse design procedure for the optimum mistuning of a high bypass ratio shroudless fan that is modeled as a cascade of blades (each with a single torsional degree-of-freedom), linearized supersonic aerodynamic theory is used to compute the unsteady aerodynamic forces in the 'influence coefficient' form at a typical blade section. The mistuning pattern is then numerically optimized in order to achieve a specified increase in the aeroelastic stability margin with a minimum amount of mistuning. If the blades are self-damped, an optimized mistuning pattern can be found that achieves a given stability margin for a much lower level of mistuning than required for the alternate mistuning pattern, which requires only two blade frequencies and is relatively insensitive to implementation errors.

Crawley, E. F.↗
NASA NTRS1984

Unsteady transonic flow in cascades

There is a need for methods to predict the unsteady air loads associated with flutter of turbomachinery blading at transonic speeds. The results of such an analysis in which the steady relative flow approaching a cascade of thin airfoils is assumed to be transonic, irrotational, and isentropic is presented. The blades in the cascade are allowed to undergo a small amplitude harmonic oscillation which generates a small unsteady flow superimposed on the existing steady flow. The blades are assumed to oscillate with a prescribed motion of constant amplitude and interblade phase angle. The equations of motion are obtained by linearizing about a uniform flow the inviscid nonheat conducting continuity and momentum equations. The resulting equations are solved by employing the Weiner Hopf technique. The solution yields the unsteady aerodynamic forces acting on the cascade at Mach number equal to 1. Making use of an unsteady transonic similarity law, these results are compared with the results obtained from linear unsteady subsonic and supersonic cascade theories. A parametric study is conducted to find the effects of reduced frequency, solidity, stagger angle, and position of pitching axis on the flutter.

Surampudi, S. P.↗
NASA NTRS1976

A steady and oscillatory kernel function method for interfering surfaces in subsonic, transonic and supersonic flow

The theory, results and user instructions for an aerodynamic computer program are presented. The theory is based on linear lifting surface theory, and the method is the kernel function. The program is applicable to multiple interfering surfaces which may be coplanar or noncoplanar. Local linearization was used to treat nonuniform flow problems without shocks. For cases with imbedded shocks, the appropriate boundary conditions were added to account for the flow discontinuities. The data describing nonuniform flow fields must be input from some other source such as an experiment or a finite difference solution. The results are in the form of small linear perturbations about nonlinear flow fields. The method was applied to a wide variety of problems for which it is demonstrated to be significantly superior to the uniform flow method. Program user instructions are given for easy access.

Cunningham, A. M., Jr.↗
NASA NTRS1999

The Mean Curvature of the Influence Surface of Wave Equation With Sources on a Moving Surface

The mean curvature of the influence surface of the space-time point (x, t) appears in linear supersonic propeller noise theory and in the Kirchhoff formula for a supersonic surface. Both these problems are governed by the linear wave equation with sources on a moving surface. The influence surface is also called the Sigma - surface in the aeroacoustic literature. This surface is the locus, in a frame fixed to the quiescent medium, of all the points of a radiating surface f(x, t) = 0 whose acoustic signals arrive simultaneously to an observer at position x and at the time t. Mathematically, the Sigma- surface is produced by the intersection of the characteristic conoid of the space-time point (x, t) and the moving surface. In this paper, we derive the expression for the local mean curvature of the Sigma - space of the space-time point for a moving rigid or deformable surface f(x, t) = 0. This expression is a complicated function of the geometric and kinematic parameters of the surface f(x, t) = 0. Using the results of this paper, the solution of the governing wave equation of high speed propeller noise radiation as well as the Kirchhoff formula for a supersonic surface can be written as very compact analytic expression.

Farassat, F.↗
NASA NTRS1983

Application of unsteady laminar triple-deck theory to viscous-inviscid interactions from an oscillating flap in supersonic and subsonic flow

Unsteady triple-deck theory is applied to analyze the local viscous-inviscid interaction of an idealized oscillating flap with a laminar boundary layer in either supersonic or subsonic external flow. For small flap amplitudes and small-to-moderate nondimensional frequencies, linearized analytical solutions by means of Fourier transformation are given for the pressure and shear distributions ahead of and behind the flap hinge. In the supersonic case, the predicted unsteady viscous effects reduce the pressure amplitude and spread it out upstream while causing it to lag the quasi-steady inviscid prediction (flap motion); this also results in an unsteady delay of incipient separation at the hinge to a higher flap amplitude. In contrast, the lowered unsteady viscous pressure distribution in purely subsonic flow slightly leads the flap motion, resulting in incipient separation at a smaller flap amplitude than the quasi-steady value. These trends are shown to be in qualitative agreement with available experimental data.

Inger, G. R.↗
NASA NTRS1990

Broadband shock-associated noise of moderately imperfectly expanded supersonic jets

Recently a stochastic model theory of broadband shock-associated noise from supersonic jets was developed. The theory is capable of predicting the near and far field noise spectra of slightly imperfectly expanded jets. In this paper, the noise prediction formulas are extended to the moderately imperfectly expanded Mach number range. It is found that the shock noise intensity depends only on the amplitude of the quasi-periodic shock cell structure and not on the strong first shock. This makes it possible to extend the noise prediction formulas by simply modifying the shock cell strength in the original formulas. Here this step is implemented semiempirically, and a set of shock noise prediction formulas are developed.

Tam, C. K. W.↗
NASA NTRS1956

Supersonic flow past nonlifting bumped and indented bodies of revolution

Linear-theory values of the supersonic pressure field and wave drag of nonlifting bumped and indented bodies of revolution are presented and compared with the corresponding properties of a smooth body. The comparisons indicate that relatively small surface irregularities cause large pressure disturbances and changes in wave drag. The interference drag produced by bumps and indentations is discussed and a correction for the curvature of the characteristics is applied.

FLOW, SUPERSONIC↗
NASA NTRS2019

Effects of Wall Cooling on Supersonic Modes in High-Enthalpy Hypersonic Boundary Layers over a Cone

High-enthalpy hypersonic boundary-layer transition plays an important role in many entry/descent vehicles. Aerothermodynamic performance of these vehicles strongly depends on the transition location on the surface. However, detailed transition flow physics in these chemically reacting boundary layers are poorly understood and transition estimates during the design phase rely heavily on empirically-derived transition criteria such as 𝑹𝒆(𝜽)/𝑴(𝒆). One of the most intriguing characteristics in hypersonic boundary layers is the presence of unstable supersonic modes, first identified in the 1990s. Due to a recent surge in hypersonic applications, there has been renewed interest in studying the flow physics of supersonic modes using either theory or direct numerical simulations. This paper investigates the rise of supersonic modes in a high-enthalpy hypersonic flow over a 𝟓° half-angle cone at various wall temperatures using both quasiparallel linear stability theory and linear PSE. It was found that supersonic modes exist in all wall temperature conditions including the adiabatic wall case. Cooler wall temperature causes the second Mack mode to become an unstable supersonic mode naturally downstream of the upper-branch neutral location when the wall is sufficiently cooled. In terms of the integrated growth, the second mode is still the dominant mode. Nonetheless, supersonic modes can cause additional series of relatively weaker growth than that of the second mode beyond the peak amplitude location. According to the present linear PSE results, contrary to what was speculated in the literature, supersonic-mode pressure disturbance structure radiated into the freestream is nonacoustic in nature and the formation of unstable supersonic modes is mainly associated with the synchronization of phase speed between the instability and acoustic waves in a nonp ayer, not due to nonlinear modal interaction as suggested in the literature.

Chang, Chau-Lyan↗
NASA NTRS2010

Loop Shaping Control Design for a Supersonic Propulsion System Model Using Quantitative Feedback Theory (QFT) Specifications and Bounds

This paper covers the propulsion system component modeling and controls development of an integrated mixed compression inlet and turbojet engine that will be used for an overall vehicle Aero-Propulso-Servo-Elastic (APSE) model. Using previously created nonlinear component-level propulsion system models, a linear integrated propulsion system model and loop shaping control design have been developed. The design includes both inlet normal shock position control and jet engine rotor speed control for a potential supersonic commercial transport. A preliminary investigation of the impacts of the aero-elastic effects on the incoming flow field to the propulsion system are discussed, however, the focus here is on developing a methodology for the propulsion controls design that prevents unstart in the inlet and minimizes the thrust oscillation experienced by the vehicle. Quantitative Feedback Theory (QFT) specifications and bounds, and aspects of classical loop shaping are used in the control design process. Model uncertainty is incorporated in the design to address possible error in the system identification mapping of the nonlinear component models into the integrated linear model.

Connolly, Joseph W.↗
NASA NTRS1992

Theoretical aspects of supersonic jet noise

The topics covered include the following: the three components of supersonic jet noise; shock cell structure of imperfectly expanded jets; large turbulence structures/instability waves; supersonic jet noise theory; generation of turbulent mixing noise; comparisons between predicted peak noise frequency and direction of radiation with measurements; Strouhal number of maximum SPL of hot supersonic jets; near field sound pressure level contours; generation of broadband shock associated noise; calculated and measured far field shock noise spectra; generation of screech tones; and calculated and measured Strouhal number of screech tones.

Tam, Christopher K. W.↗
NASA NTRS1978

Design and analysis of supersonic aircraft

Integrated system of computer programs, developed for design and analysis of supersonic configurations, uses linearized-theory methods to calculate surface pressures and uses supersonic area-rule concepts in combination with linearized theory to calculate aerodynamic force coefficients. Programs include constraints on linear theory methods to provide physical realism. Package proves useful for any type of supersonic configuration.

Coleman, J. L.↗
NASA NTRS1985

An investigation of vortex-induced aerodynamic characteristics of supersonic cruise configurations

The linear lifting surface theory which predicts the life in supersonic flow, even though the drag is usually underpredicted, is described. A method for calculating the nonlinear wave drag was developed to remedy this deficiency. The calculated sectional drag is modified by adding the difference between the exact two dimensional (2-D) and the linear 2-D wave drag at the calculated sectional lift coefficient. Improvement in the supersonic drag prediction is shown. The VORCAM code was modified for the FORTRAN 77 language and its input stream was rearranged. The Boeing code was adapted to the computer system. All CDC special features in the code are replaced with standard FORTRAN algorithms. It is suggested that because of the nonlinearity the solution appears to be nonunique crowding of two vortices, a mechanism of vortex asymmetry, is investigated.

Lan, C. E.↗