Engineering Papers⌕ Search

SEARCH · Engineering Papers

Results for “compressor surge and stall”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

63 records · Page 4
All sourcesNASA NTRSDOE OSTI
NASA NTRS1972

Single stage evaluation of highly loaded high Mach number compressor stages. 6: Data and performance of cantilevered stator

A compressor stage with a rotor tip speed of 1600 ft/sec was tested to evaluate its performance with a cantilevered stator and a rotating inner shroud beneath the stator. Both the rotor blades and the stator vanes were composed of multiple circular arc airfoil sections. Comparison of data taken during this test of the cantilevered stator and previous tests with the same compressor and airfoil geometry in a shroud stator configuration showed only slight differences in stage performance with no significant effect on overall efficiency. However, the severity of the stator wake near the rotating hub was decreased at all flows including the near surge condition. Stall and wise open discharge corrected weight flows were the same as for the shrouded stator configuration.

Merrow, A. S.↗
NASA NTRS1953

Investigation of acceleration characteristics of a single-spool turbojet engine

Operation of a single-spool turbojet engine with constant exhaust-nozzle area was investigated at one flight condition. Data were obtained by subjecting the engine to approximate-step changes in fuel flow, and the information necessary to show the relations of acceleration to the sensed engine variables was obtained. These data show that maximum acceleration occurred prior to stall and surge. In the low end of the engine-speed range the margin was appreciable; in the high-speed end the margin was smaller but had not been completely defined by these data. Data involving acceleration as a function of speed, fuel flow, turbine-discharge temperature, compressor-discharge pressure, and thrust have been presented and an effort has been made to show how a basic control system could be improved by addition of an override in which the acceleration characteristic is used not only to prevent the engine from entering the surge region but also to obtain acceleration along the maximum acceleration line during throttle bursts.

Oppenheimer, Frank L↗
NASA NTRS1982

Characteristics of the flow in the annulus-wall region of an axial-flow compressor rotor blade passage

Three-dimensional characteristics of the mean velocity and turbulence structure in the annulus-wall region of a moderately loaded compressor rotor have been investigated experimentally by employing triaxial hot-wire probes. Results are presented which indicate that the flow within and downstream of the rotor passage is highly complex due to the interaction of the annulus-wall boundary layer, blade boundary layers, tip leakage flow, and passage secondary flow. An understanding of the reported viscous flow interactions is important for establishing improved aerodynamic design criteria and efficiency, for predicting noise levels, and for determining the stall, surge, vibration, and flutter characteristics of turbomachinery.

Davino, R. M.↗
NASA NTRS1984

Unsteady flow in turbomachinery: An overview

The importance of understanding and modeling the unsteady flow phenomena in turbomachinery is discussed. Historical events in the application and development of gas turbines for aircraft propulsion are traced. Technology advancements over the years are highlighted with focus on the compression system components. Trends in compressor research within the National Advisory Committee for Aeronautics (NACA)/National Aeronautics and Space Administration (NASA) are noted. The impact of technology advancements on the increased occurrences of unsteady flow related problems in advanced engine development programs is discussed. The impact of the new and more demanding requirements being imposed on the propulsion system to meet advanced aircraft mission needs are also noted. Brief discussions on the present day understanding and modeling capability of the unsteady flow phenomena are presented to include discussions on rotating stall, surge, flutter, forced response and noise generation.

Source record↗
NASA NTRS1988

Application of bifurcation theory to axial flow compressor instability

When a compression system goes unstable, the mode of response depends on the operating and system parameters, such as throttle setting and B parameter. Previous numerical work on the model developed by Moore and Greitzer (1986) has provided a limited picture of the parametric effects. Applying bifurcation theory to a single-harmonic version of the model has supplied much more complete information, defining the boundaries of each mode of response in the parameter space. Specifically, this is shown in a plot of B vs. throttle setting, which compares well with the corresponding map produced experimentally. The importance of the shape of the rotating stall characteristic is emphasized. The analysis shows a qualitative difference between classic surge and deep surge.

Mccaughan, F. E.↗
NASA NTRS2003

Task II: Three-dimensional Rotating Stall Inception and Effects of Rotating Tip Clearance Asymmetry in Axial Compressors

The effects of two types of flow non-uniformity on stall inception behavior were assessed with linearized stability analyses of two compressor flow models. Response to rotating tip clearance asymmetries induced by a whirling rotor shaft or rotor height variations were investigated with a two-dimensional flow model. A 3-D compressor model was also developed to study the stability of both full-span and part-span rotating stall modes in annular geometries with radial flow variations. The studies focussed on (1) understanding what compressor designs were sensitive to these types of circumferential and spanwise flow non-uniformities, and (2) situations where 2-D stability theories were inadequate because of 3-D flow effects. Rotating tip clearance non-uniformity caused the greatest performance loss for shafts whirling at the rotating stall frequency. A whirling shaft displacement of 1 percent chord caused the stalling mass flow to rise by as much as 10 percent and the peak pressure rise to decrease by 6 percent. These changes were an order of magnitude larger than for equivalent-sized stationary or rotor-locked clearance asymmetries. Spanwise flow non-uniformity always destabilized the compressor, so that 2-D models over-predicted that stall margin compared to 3-D theory. The difference increased for compressors with larger spanwise variations of characteristic slope and reduced characteristic curvature near the peak. Differences between 2-D and 3-D stall point predictions were generally unacceptable (2 - 4 percent of flow coefficient) for single-stage configurations, but were less than 1 percent for multistage compressors. 2-D analyses predicted the wrong stall mode for specific cases of radial inlet flow distortion, mismatching and annulus area contraction, where higher-order radial modes led to stall. The stability behavior of flows with circumferential or radial non-uniformity was unified through a single stability criterion. The stall point for both cases was set by the integral around the annulus of the pressure rise characteristic slope, weighted by the amplitude of the mode shape. For the case of steady circumferential variations, this criterion reduced to the integrated mean slope (IMS) condition associated with steady inlet distortions. The rotating tip clearance asymmetry model was also used to demonstrate the feasibility of actively controlling the shaft position to suppress rotating stall. In axisymmetric mean flow, this method only stabilized the first harmonic mode, increasing the operating range until surge or higher harmonic modes became unstable.

Suder, Kenneth↗
NASA NTRS1947

Performance of the 19XB 10-Stage Axial-Flow Compressor with Altered Blade Angles

Previous performance data of the 19XB axial-flow compressor indicated that the outlet guide vanes and possibly the inlet guide vanes were stalling. Calculations were made to determine if these adverse conditions could be eliminated and if the manufacturer's design specifications could be more nearly approached by altering the blade angles of the first few compression stages as well as the outlet guide vanes. With the blade angles altered, experimental data were taken at compressor speeds of 8500 to 17,000 rpm with inlet-air conditions of 7.4 inches of mercury absolute and 59 0 F. The temperature-rise efficiency increased with speed from 0.70 at 8500 rpm to 0.74 at 13,600 rpm and dropped gradually to 0.70 at 17,000 rpm. At the design speed of 17,000 rpm, the pressure ratio at the peak efficiency point was 3.63. The maximum pressure ratio at design speed was 4.15 at an equivalent weight flow of 29.8 pounds per second. The altered compressor operated very .near the design specifications of pressure ratio and equivalent weight flow. At the high speeds, the peak adiabatic temperature-rise efficiency was increased 0.02 to 0,06 by altering the blade angles. The peak pressure ratio was increased 0.29 at design speed (17,000 rpm) and 0.05 and 0.13 at 11,900 and 13,600 rpm, respectively. The equivalent weight flow through the altered compressor was reduced 2 pounds per second at 15,300 and 17,000 rpm, as was expected from the design calculations. As extreme caution was taken not to surge the compressor violently, the point of minimum air flow may not have been reached in the present investigation and in a previous investigation. A true comparison of the pressure ratios obtained at the high speeds therefore cannot be made.

Downing, Richard M.↗
NASA NTRS1995

Concept Designed and Developed for Distortion- Tolerant, High-Stability Engine Control

Engine Control Future aircraft turbine engines, both commercial and military, must be able to successfully accommodate expected increased levels of steady-state and dynamic engine-face distortion. Advanced tactical aircraft are likely to use thrust vectoring to enhance their maneuverability. As a result, the engines will see more extreme aircraft angles-of-attack and sideslip levels than are currently encountered with present-day aircraft. Also, the mixed-compression inlets needed for the High Speed Civil Transport will likely encounter disturbances similar to those seen by tactical aircraft, in addition to planar pulse, inlet buzz, and high distortion levels at low flight speed and off-design operation. The current approach of incorporating a sufficient component design stall margin to tolerate these increased levels of distortion would significantly reduce performance. The objective of the High Stability Engine Control (HISTEC) program is to design, develop, and flight demonstrate an advanced, high-stability, integrated engine-control system that uses measurement-based, real-time estimates of distortion to enhance engine stability. The resulting distortion-tolerant control reduces the required design stall margin, with a corresponding increase in performance and decrease in fuel burn. The HISTEC concept has been designed and developed, and the software implementing the concept has successfully accommodated time-varying distortion. The NASA Lewis Research Center is currently overseeing the development and validation of the hardware and software necessary to flight test the HISTEC concept. HISTEC is a contracted effort with Pratt & Whitney of West Palm Beach, Florida. The HISTEC approach includes two major systems: A Distortion Estimation System (DES) and Stability Management Control (SMC). DES is an aircraft-mounted, high-speed processor that estimates the amount and type of distortion present and its effect on the engine. It uses high-response pressure measurements at the engine face to calculate indicators of the type and extent of distortion in real time. From these indicators, DES determines the effects of distortion on the propulsion systems and the corresponding engine match point necessary to accommodate it. DES output consists of fan and compressor pressure ratio trim commands that are passed to the SMC. In addition, DES uses maneuver information, consisting of angle-of-attack and sideslip from the flight control, to anticipate high inlet distortion conditions. The SMC, which is contained in the engine-mounted, Improved Digital Electronic Engine Control (IDEEC), includes advanced control laws to directly control the fan and compressor transient operating line (pressure ratio). These advanced control laws, with a multivariable design, have the potential for higher bandwidth and the resulting more precise control of engine match. The ability to measure and assess the distortion effects in real time coupled with a high-response controller improves engine stability at high levels of distortion. The software algorithms implementing DES have been designed, developed, and demonstrated, and integration testing of the DES and SMC software has been completed. The results show that the HISTEC system will be able to sense inlet distortion, determine the effect on engine stability, and accommodate distortion by maintaining an adequate margin for engine surge. The Pratt &Whitney Comprehensive Engine Diagnostic Unit was chosen as the DES processor. An instrumented inlet case for sensing distortion was designed and fabricated. HISTEC is scheduled for flight test on the ACTIVE F-15 aircraft at the NASA Dryden Flight Research Center in Edwards, California, in late 1996.

Source record↗
NASA NTRS1999

Optical Flow-Field Techniques Used for Measurements in High-Speed Centrifugal Compressors

The overall performance of a centrifugal compressor depends on the performance of the impeller and diffuser as well as on the interactions occurring between these components. Accurate measurements of the flow fields in each component are needed to develop computational models that can be used in compressor design codes. These measurements must be made simultaneously over an area that covers both components so that researchers can understand the interactions occurring between the two components. Optical measurement techniques are being used at the NASA Lewis Research Center to measure the velocity fields present in both the impeller and diffuser of a 4:1 pressure ratio centrifugal compressor operating at several conditions ranging from design flow to surge. Laser Doppler Velocimetry (LDV) was used to measure the intrablade flows present in the impeller, and the results were compared with analyses obtained from two three-dimensional viscous codes. The development of a region of low throughflow velocity fluid within this high-speed impeller was examined and compared with a similar region first observed in a large low-speed centrifugal impeller at Lewis. Particle Image Velocimetry (PIV) is a relatively new technique that has been applied to measuring the diffuser flow fields. PIV can collect data rapidly in the diffuser while avoiding the light-reflection problems that are often encountered when LDV is used. The Particle Image Velocimeter employs a sheet of pulsed laser light that is introduced into the diffuser in a quasi-radial direction through an optical probe inserted near the diffuser discharge. The light sheet is positioned such that its centerline is parallel to the hub and shroud surfaces and such that it is parallel to the diffuser vane, thereby avoiding reflections from the solid surfaces. Seed particles small enough to follow the diffuser flow are introduced into the compressor at an upstream location. A high-speed charge-coupled discharge (CCD) camera is synchronized to the laser pulse rate; this allows it to capture images of seed particle position that are separated by a small increment in time. A crosscorrelation of a particle's position in two consecutive images provides an estimate of flow velocity and direction. Multiple image pairs obtained in rapid succession at a particular flow condition provide enough measurements for statistical significance. PIV provides simultaneous velocity measurements over the entire plane that is illuminated by the light sheet instead of at a single point, as is the case when LDV is used. PIV has a further advantage in that the laser light pulse can be triggered by an external source such as a high-response pressure transducer. This feature will allow PIV to synchronize flow imaging to physical phenomena such as rotating stall or stall precursor waves. We hope that this technique can be used to obtain images of the flow field during and just prior to stall.

Skoch, Gary J.↗