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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 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS1987

Highlights of unsteady pressure tests on a 14 percent supercritical airfoil at high Reynolds number, transonic condition

Steady and unsteady pressures were measured on a two-dimensional supercritical airfoil in the Langley 0.3-m Transonic Cryogenic Tunnel at Reynolds numbers (6-35) x 10 to the 6th. The airfoil was oscillated in pitch at amplitudes from + or - .25 degrees to + or - 1.0 degrees at frequencies from 5 Hz to 60 Hz. The special requirements of testing an unsteady pressure model in a pressurized cryogenic tunnel are discussed. Selected steady measured data are presented and are compared with GRUMFOIL calculations at Reynolds number 6 x 10 to the 6th and 30 x 10 to the 6th. Experimental unsteady results at the same Reynolds numbers are examined for Reynolds-number effects. Measured unsteady results at two mean angles of attack at Reynolds number of 30 x 10 to the 6th are also examined.

Hess, Robert W.↗
NASA NTRS1987

Unsteady aerodynamics of blade rows

The requirements placed on an unsteady aerodynamic theory intended for turbomachinery aeroelastic applications are discussed along with a brief description of the various theoretical models that are available to address these requirements. The main emphasis is placed on the description of a linearized inviscid theory which fully accounts for the effects of a nonuniform mean or steady flow on unsteady aerodynamic response. Although this theory has been developed primarily for blade flutter prediction, more general equations are presented which account for unsteady excitations due to incident external aerodynamic disturbances as well as those due to blade motions. The resulting equations consist of a system of three field equations along with conditions imposed at blade, wake and shock surfaces and in the far field. These equations can be solved to determine the fluctuations in all fluid dynamic properties throughout the required solution domain. Example solutions are presented to demonstrate several effects associated with nonuniform steady flows on the linearized unsteady aerodynamic response to prescribed blade motions.

Verdon, J. M.↗
NASA NTRS1986

Unsteady separation characteristics of airfoils operating under dynamic stall conditions

Unsteady viscous/inviscid interaction phenomena are investigated for airfoils operating under dynamic stall conditions, using coupling procedures between a time-dependent inviscid panel method and two-dimensional unsteady boundary layer codes. Two strategies are pursued: a coupling of the inviscid panel method with the boundary layer code, and a strong coupling of the inviscid panel method with the boundary layer code. Attention is given to the main features of the unsteady time-marching panel method and boundary layer codes, as well as to numerical stability and the phenomenon of unsteady separation.

Geissler, Wolfgang↗
NASA NTRS1987

Advanced turboprop wing installation effects measured by unsteady blade pressure and noise

A single rotation model propeller (SR-7A) was tested at simulated takeoff/approach conditions (Mach 0.2), in the NASA Lewis 9- by 15-Ft Anechoic Wind Tunnel. Both unsteady blade surface pressures and noise measurements were made for a tractor configuration with propeller/straight wing and propeller alone configurations. The angle between the wing chord and propeller axis (droop angle) was varied along with the wing angle of attack to determine the effects on noise and unsteady loading. A method was developed that uses unsteady blade pressure measurements to provide a quantitative indication of propeller inflow conditions, at least for a uniform (across the propeller disk) inflow angle. The wing installation caused a nearly uniform upwash at the propeller inlet as evidenced by the domination of the pressure spectra by the first shaft order. This inflow angle increased at a rate of almost 150 percent of that of the wing angle-of-attack for a propeller-wing spacing of 0.54 wing chords at a constant droop angle. The flyover noise, as measured by the maximum blade passing frequency level, correlates closely with the propeller inflow angle (approx. 0.6 dB per degree of inflow angle) for all droop angles and wing angles of attack tested, including the propeller alone data. Large changes in the unsteady pressure responses on the suction surface of the blade were observed as the advance ratio was varied. The presence of a leading edge vortex may explain this behavior since changes in the location of this vortex would change with loading (advance ratio).

Heidelberg, Laurence J.↗
NASA NTRS1987

Unsteady aerodynamics of an oscillating cascade in a compressible flow field

Fundamental experiments were performed in the NASA Lewis Transonic Oscillating Cascade Facility to investigate and quantify the unsteady aerodynamics of a cascade of biconvex airfoils executing torsion-mode oscillations at realistic reduced frequencies. Flush-mounted, high-response miniature pressure transducers were used to measure the unsteady airfoil surface pressures. The pressures were measured for three interblade phase angles at two inlet Mach numbers, 0.65 and 0.80, and two incidence angles, 0 and 7 deg. The time-variant pressures were analyzed by means of discrete Fourier transform techniques, and these unique data were then compared with predictions from a linearized unsteady cascade model. The experimental results indicate that the interblade phase angle had a major effect on the chordwise distributions of the airfoil surface unsteady pressure, and that reduced frequency, incidence angle, and Mach number had a somewhat less significant effect.

Buffum, Daniel H.↗
NASA NTRS1988

Unsteady three-dimensional flow separation

A concise mathematical framework is constructed to study the topology of steady 3-D separated flows of an incompressible, or a compressible viscous fluid. Flow separation is defined by the existence of a stream surface which intersects with the body surface. The line of separation is itself a skin-friction line. Flow separation is classified as being either regular or singular, depending respectively on whether the line of separation contains only a finite number of singular points or is a singular line of the skin-friction field. The special cases of 2-D and axisymmetric flow separation are shown to be of singular type. In regular separation it is shown that a line of separation originates from a saddle point of separation of the skin-friction field and ends at nodal points of separation. Unsteady flow separation is defined relative to a coordinate system fixed to the body surface. It is shown that separation of an unsteady 3-D incompressible viscous flow at time t, when viewed from such a frame of reference, is topologically the same as that of the fictitious steady flow obtained by freezing the unsteady flow at the instant t. Examples are given showing effects of various forms of flow unsteadiness on flow separation.

Hui, W. H.↗
NASA NTRS1988

An unconditionally stable Runge-Kutta method for unsteady flows

A quasi-three dimensional analysis was developed for unsteady rotor-stator interaction in turbomachinery. The analysis solves the unsteady Euler or thin-layer Navier-Stokes equations in a body fitted coordinate system. It accounts for the effects of rotation, radius change, and stream surface thickness. The Baldwin-Lomax eddy viscosity model is used for turbulent flows. The equations are integrated in time using a four stage Runge-Kutta scheme with a constant time step. Implicit residual smoothing was employed to accelerate the solution of the time accurate computations. The scheme is described and accuracy analyses are given. Results are shown for a supersonic through-flow fan designed for NASA Lewis. The rotor:stator blade ratio was taken as 1:1. Results are also shown for the first stage of the Space Shuttle Main Engine high pressure fuel turbopump. Here the blade ratio is 2:3. Implicit residual smoothing was used to increase the time step limit of the unsmoothed scheme by a factor of six with negligible differences in the unsteady results. It is felt that the implicitly smoothed Runge-Kutta scheme is easily competitive with implicit schemes for unsteady flows while retaining the simplicity of an explicit scheme.

Jorgenson, Philip C. E.↗
NASA NTRS1987

Advanced turboprop wing installation effects measured by unsteady blade pressure and noise

A single rotation model propeller (SR-7A) was tested at simulated takeoff/approach conditions (Mach 0.2), in the NASA Lewis 9- by 15-Ft Anechoic Wind Tunnel. Both unsteady blade surface pressures and noise measurements were made for a tractor configuration with propeller/straight wing and propeller alone configurations. The angle between the wing chord and propeller axis (droop angle) was varied along with the wing angle of attack to determine the effects on noise and unsteady loading. A method was developed that uses unsteady blade pressure measurements to provide a quantitative indication of propeller inflow conditions, at least for a uniform (across the propeller disk) inflow angle. The wing installation caused a nearly uniform upwash at the propeller inlet as evidenced by the domination of the pressure spectra by the first shaft order. This inflow angle increased at a rate of almost 150 percent of that of the wing angle-of-attack for a propeller-wing spacing of 0.54 wing chords at a constant droop angle. The flyover noise, as measured by the maximum blade passing frequency level, correlates closely with the propeller inflow angle (approx. 0.6 dB per degree of inflow angle) for all droop angles and wing angles of attack tested, including the propeller alone data. Large changes in the unsteady pressure responses on the suction surface of the blade were observed as the advance ratio was varied. The presence of a leading edge vortex may explain this behavior since changes in the location of this vortex would change with loading (advance ratio).

Heidelberg, Laurence J.↗
NASA NTRS1988

Unsteady vortex-dominated flows around maneuvering wings over a wide range of Mach numbers

The problem of unsteady flow around maneuvering wings is solved using the unsteady Euler equations. The unsteady conservative Euler equations are derived for the flow relative motion with respect to a moving (translating and rotating) frame of reference. The resulting equations can handle the most general case for unsteady three-dimensional flow around maneuvering wings or wing-body configurations undergoing six degrees of freedom motion; three translations and three rotations. The equations are solved using two computational schemes; an explicit multistage finite-volume scheme and an implicit approximately factored finite-volume scheme. The computational applications cover two cases. The first case is for a locally conical supersonic flow of rolling oscillation of a sharp-edged delta wing at zero angle of attack. The second case is for a pitching oscillation around a mean angle of attack of a NACA 0012 airfoil in transonic flow.

Kandil, Osama A.↗
NASA NTRS1988

Unsteady aerodynamics of rotor-stator interaction in a turbine stage

The two-dimensional, compressible, unsteady, thin-layer Navier-Stokes equations have been used to study the turbulent flow field of a turbine rotor/stator configuration in the high pressure oxidizer turbopump of the Space Shuttle main engine. The calculations were performed on a system of patched and overlaid grids. The grid system consists of 'O-' and 'H-' grids, with the inner 'O-'grid enclosing the blade surface for an accurate resolution of the leading and trailing edges, while the outer 'H-'grid makes the treatment of boundary conditions easier. The integration scheme used is an iterative, factored, implicit method with numerical fluxes evaluated by the third-order accurate upwind-biased Osher scheme. Computed results in the form of pressure contours, Mach number contours, time-averaged surface pressure, unsteady pressure amplitude and unsteady velocity vectors are presented. Many flow phenomena, such as wake cutting by rotor airfoils and evolution of unsteady passage vortices, are discussed.

Yang, R.-J.↗
NASA NTRS1987

Unsteady diffuser vane pressure and impeller wake measurements in a centrifugal pump

Unsteady surface pressure measurements on a vaned diffuser of a centrifugal pump, and wake measurements of the flow exiting a centrifugal impeller into a vaneless diffuser are presented. Frequency spectra and ensemble averages are given for the unsteady measurements. Two different impellers were used, the pump impeller of the HPOTP (High Pressure Oxygen Turbopump) of the SSME (Space Shuttle Main Engine) and a two-dimensional impeller. The magnitude of the unsteady total pressure measured in the stationary frame at the impeller exit was found to be of the same order of magnitude as the total pressure rise across the pump. The magnitude of the unsteady diffuser vane pressures was observed to be significantly different on suction and pressure side of the vane, attaining its largest value on the suction side the leading edge while decreasing along the vane.

Arndt, N.↗
NASA NTRS1988

Three-dimensional unsteady transonic viscous-inviscid interaction using the Euler and boundary-layer equations

The objective of this study is the development of a numerical technique which can provide three-dimensional, time-accurate, compressible, turbulent flow solutions in a practical and relatively economical way. The approach taken is that of the method of viscous-inviscid interaction. The Euler equations are assumed to govern the outer inviscid portion of the flow, and the viscous layer close to the solid wall is described by a set of integral boundary-layer equations. The viscous solutions are obtained in a direct fashion with a weighted-average phase error scheme. The method of equivalent sources is used for viscous-inviscid coupling. Steady-state and unsteady computations for an AGARD airfoil and a wing show that satisfactory engineering solutions can be obtained for attached, high Reynolds number flows using this method. Quasi-unsteady interactions are shown to produce similar results to those provided by true-unsteady interactions. Considerable computer resources can be saved for some cases by using quasi-unsteady interactions.

Whitfield, David L.↗
NASA NTRS1989

Role of computational fluid dynamics in unsteady aerodynamics for aeroelasticity

In the last two decades there have been extensive developments in computational unsteady transonic aerodynamics. Such developments are essential since the transonic regime plays an important role in the design of modern aircraft. Therefore, there has been a large effort to develop computational tools with which to accurately perform flutter analysis at transonic speeds. In the area of Computational Fluid Dynamics (CFD), unsteady transonic aerodynamics are characterized by the feature of modeling the motion of shock waves over aerodynamic bodies, such as wings. This modeling requires the solution of nonlinear partial differential equations. Most advanced codes such as XTRAN3S use the transonic small perturbation equation. Currently, XTRAN3S is being used for generic research in unsteady aerodynamics and aeroelasticity of almost full aircraft configurations. Use of Euler/Navier Stokes equations for simple typical sections has just begun. A brief history of the development of CFD for aeroelastic applications is summarized. The development of unsteady transonic aerodynamics and aeroelasticity are also summarized.

Guruswamy, Guru P.↗
NASA NTRS1989

Unsteady stagnation-point heat transfer during passage of a concentrated vortex

The unsteady boundary layer due to a single rectilinear vortex filament approaching a 2-D stagnation point is investigated. Assuming the vortex remains far from the surface, incompressible potential flow theory is used to determine the time dependent inviscid flow field. The unsteady boundary layer equations are solved by an alternating-direction-implicit finite-difference method. Two mechanisms which cause fluctuations in heat transfer are the unsteady velocity field in the boundary layer and secondly, the unsteady total temperature at the edge of the boundary layer. The relative importance of these mechanisms is dependent upon the total temperature fluctuations relative to the imposed temperature difference. As a vortex approaches a stagnation point it may be forced to one side of the stagnation line or the other, depending on its initial position. Results are presented for both of these cases.

Rigby, David L.↗
NASA NTRS1989

An unconditionally stable Runge-Kutta method for unsteady flows

A quasi-three-dimensional analysis was developed for unsteady rotor-stator interaction in turbomachinery. The analysis solves the unsteady Euler or thin-layer Navier-Stokes equations in a body-fitted coordinate system. It accounts for the effects of rotation, radius change, and stream surface thickness. The Baldwin-Lomax eddy viscosity model is used for turbulent flows. The equations are integrated in time using a four-stage Runge-Kutta scheme with a constant time step. Implicit residual smoothing was employed to accelerate the solution of the time accurate computations. The scheme is described and accuracy analyses are given. Results are shown for a supersonic through-flow fan designed for NASA Lewis. The rotor:stator blade ratio was taken as 1:1. Results are also shown for the first stage of the Space Shuttle Main Engine high pressure fuel turbopump. Here the blade ratio is 2:3. Implicit residual smoothing was used to increase the time step limit of the unsmoothed scheme by a factor of six with negligible differences in the unsteady results. It is felt that the implicitly smoothed Runge-Kutta scheme is easily competitive with implicit schemes for unsteady flows while retaining the simplicity of an explicit scheme.

Jorgenson, Philip C. E.↗
NASA NTRS1989

Unsteady stagnation-point heat transfer during passage of a concentrated vortex

The unsteady boundary layer due to a single rectilinear vortex filament approaching a two-dimensional stagnation point is investigated. The unsteady boundary-layer equations are solved by an alternating-direction-implicit finite-difference method. Two mechanisms which cause fluctuations in heat transfer are the unsteady velocity field in the boundary layer and secondly, the unsteady total temperature at the edge of the boundary layer. The relative importance of these mechanisms is dependent upon the total temperature fluctuations relative to the imposed temperature difference.

Rigby, David L.↗
NASA NTRS1990

Numerical solutions of the linearized Euler equations for unsteady vortical flows around lifting airfoils

A linearized unsteady aerodynamic analysis is presented for unsteady, subsonic vortical flows around lifting airfoils. The analysis fully accounts for the distortion effects of the nonuniform mean flow on the imposed vortical disturbances. A frequency domain numerical scheme which implements this linearized approach is described, and numerical results are presented for a large variety of flow configurations. The results demonstrate the effects of airfoil thickness, angle of attack, camber, and Mach number on the unsteady lift and moment of airfoils subjected to periodic vortical gusts. The results show that mean flow distortion can have a very strong effect on the airfoil unsteady response, and that the effect depends strongly upon the reduced frequency, Mach number, and gust wave numbers.

Scott, James R.↗
NASA NTRS1988

Calculation of unsteady Euler flows in turbomachinery using the linearized Euler equations

The model used by the present unsteady cascade flow calculation method is based on the linearized unsteady Euler equations and accounts for blade loading and geometry, shock motion, and wake motion. The steady flow that must be determined prior to the unsteady flow is ascertained by means of a Newton iteration procedure. A noteworthy feature of the procedure is the use of shock-fitting to determine steady and unsteady shock positions; the use of the Euler equations in conjunction with the Rankine-Hugoniot shock-jump conditions is found to directly model the generation of entropy and vorticity at shocks. Results are presented for both channel and cascade flows.

Hall, Kenneth C.↗