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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 91 records · Page 5
All sourcesNASA NTRSDOE OSTI
NASA NTRS1974

Instantaneous distortion in a Mach 2.5, 40-percent-internal-contraction inlet and its effect on turbojet stall margin

An experimental wind tunnel investigation was conducted to determine the effects of time-variant distortions produced in a supersonic inlet on a J85-GE-13 turbojet engine. Results are presented principally in terms of instantaneous distortion amplitudes and total-pressure contours measured through compressor stall. They indicate that, although a time-averaged distortion may be far from a stall-inducing value, corresponding instantaneous distortion amplitudes can approach or exceed this value. Seven engine stall events were studied. In six of these events instantaneous distortion peaks of sufficient magnitude to cause stall were observed in the time period shortly before stall was detected.

Burstadt, P. L.↗
NASA NTRS1981

A flight investigation of the ultra-deep-stall descent and spin recovery characteristics of a 1/6 scale radiocontrolled model of the Piper PA38 Tomahawk

Ultradeep stall descent and spin recovery characteristics of a 1/6 scale radio controlled model of the Piper PA38 Tomahawk aircraft was investigated. It was shown that the full scale PA38 is a suitable aircraft for conducting ultradeep stall research. Spin recovery was accomplished satisfactorily by entry to the ultradeep stall mode, followed by the exit from the ultradeep stall mode. It is concluded that since the PA38 has excellent spin recovery characteristics using normal recovery techniques (opposite rudder and forward control colum pressure), recovery using ultradeep stall would be beneficial only if the pilot suffered from disorientation.

Blanchard, W. S., Jr.↗
NASA NTRS1983

A theory of rotating stall of multistage axial compressors

A theoretical analysis was made of rotating stall in axial compressors of many stages, finding conditions for a permanent, straight-through traveling disturbance, with the steady compressor characteristic assumed known, and with simple lag processes ascribed to the flows in the inlet, blade passages, and exit regions. For weak disturbances, predicted stall propagation speeds agree well with experimental results. For a locally-parabolic compressor characteristic, an exact nonlinear solution is found and discussed. For deep stall, the stall-zone boundary is most abrupt at the trailing edge, as expected. When a complete characteristic having unstalling and reverse-flow features is adopted, limit cycles governed by a Lienard's equation are found. Analysis of these cycles yields predictions of recovery from rotating stall; a relaxation oscillation is found at some limiting flow coefficient, above which no solution exists. Recovery is apparently independent of lag processes in the blade passages, but instead depends on the lags originating in the inlet and exit flows, and also on the shape of the given characteristic diagram. Small external lags and tall diagrams favor early recovery. Implications for future research are discussed.

Moore, F. K.↗
NASA NTRS1985

Stall recovery control strategy methodology and results

The nonrecoverable stall condition in high performance turbofan engines is characterized by thrust loss, rising turbine temperatures, high pressure compressor rotating stall, and loss of engine control. Stall recovery control is presently investigated by means of an engine system computer model capable of either surge or rotating stall postinstability operation. Several techniques are examined which can yield rapid recovery from stall; a composite strategy which involves reduction of engine speed to idle while simultaneously opening the 10th- and 14th-stage compressor bleed ports allowed recovery to speeds slightly higher than idle while combustor fuel flow continued.

Hopf, W. R.↗
NASA NTRS1985

A theory of post-stall transients in axial compression systems. II - Application

Using the theory developed in Part I, calculations have been carried out to show the evolution of the mass flow, pressure rise, and rotating-stall cell amplitude during compression system post-stall transients. In particular, it is shown that the unsteady growth or decay of the stall cell can have a significant effect on the instantaneous compressor pumping characteristic and hence on the overall system behavior. A limited parametric study is carried out to illustrate the impact of different system features on transient behavior. It is shown, for example, that the ultimate mode of system response, surge or stable rotating stall, depends not only on the B parameter, but also on the compressor length-to-radius ratio. Small values of this latter quantity tend to favor the occurrence of surge, as do large values of B. Based on the analytical and numerical results, several specific topics are suggested for future research on post-stall transients.

Greitzer, E. M.↗
NASA NTRS1988

Stall flutter analysis of propfans

Three semi-empirical aerodynamic stall models are compared with respect to their lift and moment hysteresis loop prediction, limit cycle behavior, easy implementation, and feasibility in developing the parameters required for stall flutter prediction of advanced turbines. For the comparison of aeroelastic response prediction including stall, a typical section model and a plate structural model are considered. The response analysis includes both plunging and pitching motions of the blades. In model A, a correction of the angle of attack is applied when the angle of attack exceeds the static stall angle. In model B, a synthesis procedure is used for angles of attack above static stall angles, and the time history effects are accounted for through the Wagner function.

Reddy, T. S. R.↗
NASA NTRS1989

Analysis of an unswept propfan blade with a semiempirical dynamic stall model

The time history response of a propfan wind tunnel model with dynamic stall is studied analytically. The response obtained from the analysis is compared with available experimental data. The governing equations of motion are formulated in terms of blade normal modes which are calculated using the COSMIC-NASTRAN computer code. The response analysis considered the blade plunging and pitching motions. The lift, drag and moment coefficients for angles of attack below the static stall angle are obtained from a quasi-steady theory. For angles above static stall angles, a semiempirical dynamic stall model based on a correction to angle of attack is used to obtain lift, drag and moment coefficients. Using these coefficients, the aerodynamic forces are calculated at a selected number of strips, and integrated to obtain the total generalized forces. The combined momentum-blade element theory is used to calculate the induced velocity. The semiempirical stall model predicted a limit cycle oscillation near the setting angle at which large vibratory stresses were observed in an experiment. The predicted mode and frequency of oscillation also agreed with those measured in the experiment near the setting angle.

Reddy, T. S. R.↗
NASA NTRS1989

Flow visualization studies of the Mach number effects on the dynamic stall of an oscillating airfoil

Compressibility effects on the dynamic stall of a NACA 0012 airfoil undergoing sinusoidal oscillatory motion were studied using a stroboscopic schlieren system. Schlieren pictures and some quantitative data derived from them are presented and show the influence of free-stream Mach number and reduced frequency on the dynamic-stall vortex. This study shows that a dynamic stall vortex always forms near the leading edge and convects on the airfoil upper surface at approximately 0.3 times the free stream velocity for all cases studied. The results also demonstrate that initiation of the dynamic stall vortex is delayed to higher angles of attack with increased reduced frequency, but that dynamic stall occurs at lower angles of incidence with increasing Mach numbers.

Chandrasekhara, M. S.↗
NASA NTRS1989

Design and development of a compressible dynamic stall facility

A dynamic stall facility offering a unique new capability for studies of compressibility effects on dynamic stall is described. This facility features complete visual access by mounting the test airfoil between optical-quality glass windows which are rotated in unison to produce the oscillating airfoil motion associated with helicopter rotor dynamic stall. By using the density gradients associated with the rapidly changing dynamic stall flow field, this facility permits simultaneous detailed investigation of the flow on the surface as well as in the flow field surrounding airfoils experiencing dynamic stall.

Carr, L. W.↗
NASA NTRS1993

A model for the selective amplification of spatially coherent waves in a centrifugal compressor on the verge of rotating stall

A simple model for the stability zones of a low speed centrifugal compressor is developed, with the goal of understanding the driving mechanism for the changes in stalling behavior predicted for, and observed in, the Purdue Low Speed Centrifugal Research Compressor Facility. To this end, earlier analyses of rotating stall suppression in centrifugal compressors are presented in a reduced form that preserves the essential parameters of the model that affect the stalling behavior of the compressor. The model is then used to illuminate the relationship between compressor geometry, expected mode shape, and regions of amplification for weak waves which are indicative of the susceptibility of the system to rotating stall. The results demonstrate that increasing the stagger angle of the diffuser vanes, and consequently the diffusion path length, results in the compressor moving towards a condition where higher-order spatial modes are excited during stall initiation. Similarly, flow acceleration in the diffuser section caused by an increase in the number of diffuser vanes also results in the excitation of higher modes.

Lawless, Patrick B.↗
NASA NTRS1938

NACA Stall-Warning Indicator

The stall-warning indicator employs a total-head tube located close to the wing surface in a region wherein local stalling occurs before the main portion of the wing stalls. The artificial production of a localized stalled region is accomplished by means of a sharp leading edge extending a few inches along the span. An abrupt drop in the total pressure relative to a static reference taken at some convenient point occurs at the stall in this region. Thin drop in total pressure caused the contraction of a pressure cell to which the total-head tube and the static orifice are connected and closed an electrical circuit that actuates a warning signal.

Thopmson, F L↗
NASA NTRS1948

Wind-tunnel Investigation of High-lift and Stall-control Devices on a 37 Degree Sweptback Wing of Aspect Ratio 6 at High Reynolds Numbers

Results are presented of an investigation in the Langley 19-foot pressure tunnel of the longitudinal characteristics of a semispan model wing having 37 degrees sweepback of the leading edge, an aspect ratio of 6, and NACA 641-212 airfoil section perpendicular to the 27-percent-chord line. Several types of stall-control devices including extensible round-nose leading-edge flaps, a leading-edge slat, and a drooped leading edge were investigated; partial- and full-span trailing-edge split and double slotted flaps were also tested. In addition, various combinations of the aforementioned leading- and trailing-edge flaps were investigated. The tests covered a range of Reynolds numbers between 2.00 x 10(6) and 9.35 x 10(6). The wing with or without trailing-edge splity of double slotted flap was longitudinally unstable near maximum lift due to tip stalling. The addition of an outboard half-span leading-edge flap or a leading-edge slat to the plain wing or wing with inboard half-span split flaps eliminated tip stalling and resulted in stable moment variations at the stall. The drooped leading edge, on the other hand, was only effective when used in conjunction with an upper-surface fence. The combination of an outboard leading-edge device and inboard half-span double slotted flap resulted in an undesirable loop in the pitching-moment curve near maximum lift in spite of an inboard stall. The loop is attributed to the section characteristics of the double slotted flap. Air-flow surveys behind the wing indicated that a suitably placed horizontal tail would eliminate the loop in the moment curve.

Koven, William↗
NASA NTRS1993

Interferometric investigations of compressible dynamic stall over a transiently pitching airfoil

The dynamic stall flow field over NACA 0012 airfoil pitching transiently from 0 - 60 at a constant rate under compressible flow conditions has been studied using the real-time technique of point diffraction interferometry. This investigation using nonintrusive diagnostics provides a quantitative description of the overall flow field, including the finer details of dynamic stall vortex formation, growth and the concomitant changes in the pressure distribution. Analysis of several hundred interferograms obtained for a range of flow conditions shows that the peak leading edge suction pressure coefficient that stall is nearly constant for a given free stream Mach number at all nondimensional pitch rates. Also, this value is below that seen in steady flow at static stall for the same Mach number, indicating that dynamic effects significantly effect the separation behavior. Further, for a given Mach number, the dynamic stall vortex seems to form rapidly at nearly the same angle of attack for all pitch rates studied. As the vortex is shed, it induces an anti-clockwise trailing edge vortex, which grows in a manner similar to that of a starting vortex. The measured peak suction pressure coefficient drops as the free stream Mach number increases. For free stream Mach numbers above 0.4, small multiple shocks appear near the leading edge.

Chandrasekhara, M. S.↗
NASA NTRS1995

Preliminary analysis of dynamic stall effects on a 91-meter wind turbine rotor

Analytical investigation of dynamic stall on HAWT (horizontal-axis wind turbines) rotor loads was conducted. Dynamic stall was modeled using the Gormont approach on the MOD-2 rotor, treating the blade as a rigid body teetering about a fixed axis. Blade flapwise bending moments at station 370 were determined with and without dynamic stall for spatial variations in local wind speed due to wind shear and yaw. The predicted mean flapwise bending moments were found to be in good agreement with test results. Results obtained with and without dynamic stall showed no significant difference for the mean flapwise bending moment. The cyclic bending moments calculated with and without dynamic stall effects were substantially the same. None of the calculated cyclic loads reached the level of the cyclic loads measured on the MOD-2 using the Boeing five-minute-average technique.

Wilson, Robert E.↗
NASA NTRS1995

Control of unsteady separated flow associated with the dynamic stall of airfoils

An effort to understand and control the unsteady separated flow associated with the dynamic stall of airfoils was funded for three years through the NASA cooperative agreement program. As part of this effort a substantial data base was compiled detailing the effects various parameters have on the development of the dynamic stall flow field. Parameters studied include Mach number, pitch rate, and pitch history, as well as Reynolds number (through two different model chord lengths) and the condition of the boundary layer at the leading edge of the airfoil (through application of surface roughness). It was found for free stream Mach numbers as low as 0.4 that a region of supersonic flow forms on the leading edge of the suction surface of the airfoil at moderate angles of attack. The shocks which form in this supersonic region induce boundary-layer separation and advance the dynamic stall process. Under such conditions a supercritical airfoil profile is called for to produce a flow field having a weaker leading-edge pressure gradient and no leading-edge shocks. An airfoil having an adaptive-geometry, or dynamically deformable leading edge (DDLE), is under development as a unique active flow-control device. The DDLE, formed of carbon-fiber composite and fiberglass, can be flexed between a NACA 0012 profile and a supercritical profile in a controllable fashion while the airfoil is executing an angle-of-attack pitch-up maneuver. The dynamic stall data were recorded using point diffraction interferometry (PDI), a noninvasive measurement technique. A new high-speed cinematography system was developed for recording interferometric images. The system is capable of phase-locking with the pitching airfoil motion for real-time documentation of the development of the dynamic stall flow field. Computer-aided image analysis algorithms were developed for fast and accurate reduction of the images, improving interpretation of the results.

Wilder, M. C.↗
NASA NTRS2000

Airfoil Dynamic Stall and Rotorcraft Maneuverability

The loading of an airfoil during dynamic stall is examined in terms of the augmented lift and the associated penalties in pitching moment and drag. It is shown that once stall occurs and a leading-edge vortex is shed from the airfoil there is a unique relationship between the augmented lift, the negative pitching moment, and the increase in drag. This relationship, referred to here as the dynamic stall function, shows limited sensitivity to effects such as the airfoil section profile and Mach number, and appears to be independent of such parameters as Reynolds number, reduced frequency, and blade sweep. For single-element airfoils there is little that can be done to improve rotorcraft maneuverability except to provide good static C(l(max)) characteristics and the chord or blade number that is required to provide the necessary rotor thrust. However, multi-element airfoils or airfoils with variable geometry features can provide augmented lift in some cases that exceeds that available from a single-element airfoil. The dynamic stall function is shown to be a useful tool for the evaluation of both measured and calculated dynamic stall characteristics of single element, multi-element, and variable geometry airfoils.

Bousman, William G.↗
NASA NTRS2002

A Comparative Study of Three Methodologies for Modeling Dynamic Stall

During the past two decades, there has been an increased reliance on the use of computational fluid dynamics methods for modeling rotors in high speed forward flight. Computational methods are being developed for modeling the shock induced loads on the advancing side, first-principles based modeling of the trailing wake evolution, and for retreating blade stall. The retreating blade dynamic stall problem has received particular attention, because the large variations in lift and pitching moments encountered in dynamic stall can lead to blade vibrations and pitch link fatigue. Restricting to aerodynamics, the numerical prediction of dynamic stall is still a complex and challenging CFD problem, that, even in two dimensions at low speed, gathers the major difficulties of aerodynamics, such as the grid resolution requirements for the viscous phenomena at leading-edge bubbles or in mixing-layers, the bias of the numerical viscosity, and the major difficulties of the physical modeling, such as the turbulence models, the transition models, whose both determinant influences, already present in static maximal-lift or stall computations, are emphasized by the dynamic aspect of the phenomena.

Sankar, L.↗
NASA NTRS1946

Effects of Wing Flaps and Wing Duct Inlet on the Lift and Stalling Characteristics of a 1/4-Scale Partial-Span Model of the Republic XF-12 Airplane in the Langley 19-Foot Pressure Tunnel

An investigation was conducted in the Langley 19-foot pressure tunnel to determine the lift, drag, pitching-moment and stalling characteristics fo a 1/4 -scale partial-span model of the left wing of the Republic XF-12 airplane. The effects of a duct inlet, located between the nacelles at the leading edge of the wing, on those characteristics were also investigated. The Reynolds numbers for the investigation covered a range from 4,500,000 to 8,600,000. The results of the investigation indicated that maximum lift coefficients of 1.36, 1.71, and 2.11 were measured on the model with flaps neutral and deflected 20 deg and 55 deg, respectively at a Reynolds number of 8,600,000. When the duct inlet was replaced by a basic airfoil nose the flap-neutral maximum-lift coefficient was increased from 1.36 to 1.41. The results also showed that at maximum lift with flaps neutral or deflected 55 deg. most of the area between the nacelles were stalled while only small areas on other portions of the model were stalled; when the duct inlet was replaced by the basic airfoil nose the stall was delayed to a slightly higher angle of attack but the nature of the stall was relatively unaffected.

Graham, Robert R.↗