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At least 217 records · Page 12
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

Experimental investigation of stability and stall flutter of a free-floating wing V/STOL model

An experimental investigation was made of the static and dynamic stability characteristics of a one-fourth scale model of a tilt-propeller free-wing V/STOL aircraft. The effects of wing pivot location, wing chord, trailing-edge angle, propeller tilt angle, and thrust were studied, and a limited evaluation was made of high-lift devices. A dynamically similar wing was used to measure frequency and damping ratio from transient response data in the linear aerodynamic regime and the results were compared to quasi-steady and unsteady aerodynamic theory. It was found that at high trim angles of attack, in the nonlinear aerodynamic regime, stall flutter oscillations occurred with typical amplitudes ranging from 15 to 20 deg. Wing control tab deflection was effective in initiating and terminating stall flutter but variations in configuration or operating conditions did not greatly influence the occurrence of characteristics of the oscillations.

Ormiston, R. A.↗
NASA NTRS1972

Analytic prediction of dynamic stall characteristics.

A new approach is presented for solving the problem of predicting dynamic stall characteristics. In this approach, the unsteady aerodynamic characteristics are related theoretically to static aerodynamic characteristics which are readily available for a great number of airfoil shapes. Using these static experimental data as an input, the developed analytical method predicts dynamic stall characteristics that are in good agreement with available experimental data. The analysis is also extended to include frequency and amplitude modulation effects.

Ericsson, L. E.↗
NASA NTRS1973

Experimental study of flow about a stalled two-dimensional airfoil

Wind tunnel tests to determine the flow characteristics of a stalled wing are described. The advantages of the plane wing over the ring wing for wind tunnel investigations are presented. The general flow pattern for stalled ring airfoils is developed. Photographs of the test equipment and installation are provided.

Coles, D. E.↗
NASA NTRS1974

Steady-state inlet temperature distortion effects on the stall limits of a J85-Ge-13 turbojet engine

The effects of circumferential and radial temperature distortions and combined temperature and pressure distortion on the performance of a J85-GE-13 turbojet engine were investigated. In terms of loss of compressor pressure ratio at stall, the most severe pattern of temperature distortion was hub radial while a circumferential pattern of diametrically opposed sectors was least severe. The effect of combined temperature and pressure distortion was dependent on the location of the high temperature and low pressure regions. The most serious stall pressure ratio losses occurred when these regions coincided.

Mehalic, C. M.↗
NASA NTRS1974

Investigation of the stall-induced shock wave (hammershock) at the inlet to the engine

The peak static pressures measured at the inlet to the engine during stall are presented for a turbojet and two turbofan engines. It is shown for one turbofan and the turbojet that the static pressure ratio across the hammershock does not exceed significantly the normal shock pressure ratio necessary to stop the flow. The second turbofan engine did not follow this rule. Possible reasons for the departure are discussed. For the two turbofan engines the influence of the stall method on the hammershock intensity was investigated. Data related to the spatial distribution of pressure in the hammershock are also presented.

Kurkov, A. P.↗
NASA NTRS1974

Analysis of dynamic stall using unsteady boundary-layer theory

The unsteady turbulent boundary layer and potential flow about a pitching airfoil are analyzed using numerical methods to determine the effect of pitch rate on the delay in forward movement of the rear flow reversal point. An explicit finite difference scheme is used to integrate the unsteady boundary layer equations, which are coupled at each instant of time to a fully unsteady and nonlinear potential flow analysis. A substantial delay in forward movement of the reversal point is demonstrated with increasing pitch rate, and it is shown that the delay results partly from the alleviation of the gradients in the potential flow, and partly from the effects of unsteadiness in the boundary layer itself. The predicted delay in flow-reversal onset, and its variation with pitch rate, are shown to be in reasonable agreement with experimental data relating to the delay in dynamic stall. From the comparisons it can be concluded (a) that the effects of time-dependence are sufficient to explain the failure of the boundary layer to separate during the dynamic overshoot, and (b) that there may be some link between forward movement of the reversal point and dynamic stall.

Scruggs, R. M.↗
NASA NTRS1974

Effect of inlet ingestion of a wing tip vortex on turbojet stall margin

A two-dimensional inlet and a turbojet engine were investigated in a Mach 0.4 stream so as to ingest the tip vortex of a forward mounted wing. Results show that ingestion of a wing tip vortex by a turbojet engine can cause a large reduction in engine stall margin. The loss in stall compressor pressure ratio was primarily dependent on vortex location and rotational direction and not on total-pressure variations across the compressor face.

Mitchell, G. A.↗
NASA NTRS1974

Investigation of the stall-induced shock wave /hammershock/ at the inlet to the engine

The peak static pressures measured at the inlet to the engine during stall are presented for a turbojet and two turbofan engines. It is shown for one turbofan and the turbojet that the static pressure ratio across the hammershock does not exceed significantly the normal shock pressure ratio necessary to stop the flow. The second turbofan engine did not follow this rule. Possible reasons for the departure are discussed. For the two turbofan engines, the influence of the stall method on the hammershock intensity has been investigated. Data related to the spatial distribution of pressure in the hammershock are also presented.

Kurkov, A. P.↗
NASA NTRS1975

Effect of inlet ingestion of a wing tip vortex on compressor face flow and turbojet stall margin

A two-dimensional inlet was alternately mated to a coldpipe plug assembly and a J85-GE-13 turbojet engine, and placed in a Mach 0.4 stream so as to ingest the tip vortex of a forward mounted wing. Vortex properties were measured just forward of the inlet and at the compressor face. Results show that ingestion of a wing tip vortex by a turbojet engine can cause a large reduction in engine stall margin. The loss in stall compressor pressure ratio was primarily dependent on vortex location and rotational direction and not on total-pressure distortion.

Mitchell, G. A.↗
NASA NTRS1975

Starting vortex, separation bubbles and stall - A numerical study of laminar unsteady flow around an airfoil

The stalling characteristics of an airfoil in a laminar viscous incompressible fluid are investigated. The governing equations in terms of the vorticity and stream function are solved using an implicit finite-difference scheme and point successive relaxation procedure. The development of the impulsively started flow, the initial generation of circulation, and the behavior of the forces at large times are studied. Following the impulsive start, the lift is at first very large and then it rapidly drops. The subsequent growth of circulation and lift is associated with the starting vortex. After incipient separation, the lift increases owing to enlargement of the separation bubble and intensification of the flow rotation in it. The extension of this bubble beyond the trailing edge causes it to rupture and brings about the stalling characteristics of the airfoil. Subsequently, new bubbles are formed near the upper surface of the airfoil and are swept away. The behavior of the lift acting on the airfoil is explained in terms of the strength and sense of these bubbles.

Mehta, U. B.↗
NASA NTRS1975

Force and pressure measurements on an airfoil oscillating through stall, part 2

Details of force, moment, and pressure distributions on a two dimensional, four foot chord, NACA 0012 airfoil, oscillating in pitch through stall, in a 7 ft. x 10 ft. low speed wind tunnel, are presented. Tests were run with the airfoil in a closed test section and also in a test section having four longitudinal slots in each sidewall set to provide minimum tunnel interference on the wing in steady flow. In unsteady flow, differences between the results for the closed and 2% open case are small. The dynamic stall process is not triggered by the bursting of a laminar separation bubble but rather by the separation of the turbulent boundary layer downstream of the bubble.

Parker, A. G.↗
NASA NTRS1976

Force and pressure measurements on an airfoil oscillating through stall

Details of force, moment, and pressure distributions on a two-dimensional, 4-ft chord, NACA 0012 airfoil, oscillating in pitch through stall, in a 7 ft x 10 ft low-speed wind tunnel are presented. Tests were run with the airfoil in a closed test section and also in a test section having four longitudinal slots in each sidewall set to provide minimum tunnel interference on the wing in steady flow. In steady flow, differences between the results for the closed and 2% open case are small. The dynamic stall process is not triggered by the bursting of a laminar separation bubble but rather by the separation of the turbulent boundary layer downstream of the bubble.

Parker, A. G.↗
NASA NTRS1976

Stall flutter and nonlinear divergence of a two-dimensional flat plate wing

Tests were conducted in a small wind tunnel to study the torsional stall flutter behavior of a two-dimensional flat-plate wing pivoted about the midchord. The nonlinear static divergence equilibrium properties of the wing were well predicted from the measured static moment characteristics. Large amplitude limit cycles ranging from plus or minus 11 degrees to plus or minus 100 degrees were observed. Stall flutter occurred above a critical value of a reduced frequency of about 2. Self-excitation occurred for initial angles of attack between 0 and 8 degrees. Nondimensional harmonic coefficients were extracted from the free transient vibration tests for amplitudes up to 80 degrees.

Dugundji, J.↗
NASA NTRS1976

The effect of varying freestream velocity on dynamic stall characteristics

A low speed wind tunnel equipped with an axial gust generator to simulate the aerodynamic environment of a helicopter rotor was used to study the dynamic stall of a pitching blade. The objective of this investigation was to find out to what extent harmonic velocity perturbations in the freestream affect dynamic stall. The study involved making measurements of the aerodynamic moment on a two-dimensional, pitching blade model in both constant and pulsating airstreams. Using an operational analog computer to perform on-line data reduction, plots of moment versus angle of attack and work done by the moment were obtained. The data taken in the varying freestream were then compared to constant freestream data, and to the results of two analytical methods. These comparisons showed that the velocity perturbations had a significant effect on the pitching moment which could not be consistently predicted by the analytical methods, but had no drastic effect on the blade stability.

Pierce, G. A.↗
NASA NTRS1977

Laser velocimeter systems analysis applied to a flow survey above a stalled wing

A laser velocimeter operating in the backscatter mode was used to survey the flow above a stalled wing. Polarization was used to separate the two orthogonal velocity components of the fringe-type laser velocimeter, and digital counters were used for data processing. The velocities of the kerosene seed particles were measured with less than 2 percent uncertainty. The particle velocity measurements were collected into histograms. The flow field survey was carried out above an aspect-ratio-8 stalled wing with an NACA 0012 section. The angle of attack was 19.5 deg, the Mach number was 0.49, and the Reynolds number was 1,400,000. The flow field was characterized by the periodic shedding of discrete vortices from near the crest of the airfoil.

Young, W. H., Jr.↗
NASA NTRS1978

Simulation study of the oscillatory longitudinal motion of an airplane at the stall

Hybrid simulation of the longitudinal motion of a straight-wing airplane at the stall was studied to investigate the effect of hysteresis in the development of lift and pitching moments on the wing as a function of angle of attack on the occurrence of longitudinal oscillations at the stall. Flight data for the simulated airplane and for various other airplanes are shown for the simulated airplane and for comparison. The results show that oscillations similar to those measured in flight may be obtained by incorporating hysteresis in the lift and pitching-moment curves.

Phillips, W. H.↗
NASA NTRS1978

Analysis of supersonic stall bending flutter in axial-flow compressor by actuator disk theory

An analytical model was developed for predicting the onset of supersonic stall bending flutter in axial-flow compressors. The analysis is based on two-dimensional, compressible, unsteady actuator disk theory. It is applied to a rotor blade row by considering a cascade of airfoils. The effects of shock waves and flow separation are included in the model. Calculations show that the model predicts the onset, in an unshrouded rotor, of a bending flutter mode that exhibits many of the characteristics of supersonic stall bending flutter. The validity of the analysis for predicting this flutter mode is demonstrated.

Adamczyk, J. J.↗
NASA NTRS1979

The prediction of two-dimensional airfoil stall progression

A generalized boundary condition potential flow calculation method was combined with a momentum integral boundary layer method and a base flow theory of separation to predict airfoil viscous-inviscid interference up to and beyond stall. The resultant program considers laminar and turbulent separation and is, therefore, applicable to thin or thick airfoil stall. The calculated flow field includes the airfoil and the separation bubble recombination region behind the airfoil. Calculated pressure distributions and equivalent airfoil shapes, including the displacement thickness of the viscous regions, are compared with flow field measurements for several airfoils. The measured displacement thicknesses and wake centerlines corroborate the calculated shape. The comparison also suggests the use of the analytical solution to evaluate the measurements.

Gross, L. W.↗