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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 19 records
All sourcesNASA NTRSDOE OSTI
DOE OSTI2012

Air Conditioning Stall Phenomenon Testing, Model Development, and Simulation

Electric distribution systems are experiencing power quality issues of extended reduced voltage due to fault-induced delayed voltage recovery (FIDVR). FIDVR occurs in part because modern air conditioner (A/C) and heat pump compressor motors are much more susceptible to stalling during a voltage sag or dip such as a sub-transmission fault. They are more susceptible than older A/C compressor motors due to the low inertia of these newer and more energy efficient motors. There is a concern that these local reduced voltage events on the distribution system will become more frequent and prevalent and will combine over larger areas and challenge transmission system voltage and ultimately power grid reliability. The Distributed Energy Communications and Controls (DECC) Laboratory at Oak Ridge National Laboratory (ORNL) has been employed to (1) test, (2) characterize and (3) model the A/C stall phenomenon.

air conditioning stall↗
NASA NTRS1957

Surge-Inception Study in a Two-Spool Turbojet Engine

A two-spool turbojet engine was operated in the Lewis altitude wind tunnel to study the inception of compressor surge. In addition to the usual steady-state pressure and temperature measurements, the compressors were extensively instrumented with fast-response interstage pressure transducers. Thus it was possible to obtain maps for both compressors, pressure oscillations during rotating stall, effects of stall on efficiency, and stage-loading curves. In addition, with the transient measurements, it was possible to record interstage pressures and then compute stage performance during accelerations to the stall limit. Rotating stall was found to exist at low speeds in the outer spool. Although the stall arose from poor flow conditions at the inlet-stage blade tips, the low-energy air moved through the machine from the tip at the inlet to the outer spool to the hub at the inlet to the inner spool. This tip stall ultimately resulted in compressor surge in the mid-speed region, and necessitated inter-compressor air bleed. Interstage pressure measurements during acceleration to the compressor stall limit indicated that rotating stall was not a necessary condition for compressor surge and that, at the critical stall point, the circumferential interstage pressure distribution was uniform. The exit-stage group of the inner spool was first t o stall; then, the stages upstream stalled in succession until the inlet stage of the outer spool was stalled. With a sufficiently high fuel rate, the process repeated with a cycle time of about 0.1 second. It was possible to construct reproducible stage stall lines as a function of compressor speed from the stage stall points of several such compressor surges. This transient stall line was checked by computing the stall line from a steady-state stage-loading curve. Good agreement between the stage stall lines was obtained by these two methods.

Wallner, Lewis E.↗
NASA NTRS1999

Rotating Pip Detection and Stall Warning in High-Speed Compressors Using Structure Function

A statistic for both rotating pip and incipient stall detection, called Structure Function is introduced for use in high speed research compressor environments. Experimental studies on stall inception processes have long observed two types of pre-stall compressor activity. Presently there exist methods for indicating modal stall precursive events in the compressor. This is a first application of a new method to detect rotating pip activity prior to stall in research compressors. The algorithm requires a very short sample of data to distinguish pip activity prior to stall, and thus may be used in a real time application. Additionally, this Structure Function algorithm is also used as a single sensor stall warning method under a variety of operating conditions, including clean inlet conditions, radially and circumferentially distorted inlet conditions, and in examples of steady air injection along the casing, and controlled air injection conditions. Structure Function provides a potential advantage over linear spectral techniques and wavelet algorithms for stall detection due to the simplicity of the algorithm and because it does not rely on a priori knowledge of frequency content.

Bright, Michelle M.↗
DOE OSTI2024

FIDVR Capability of Hybrid Grid-Forming PV Power Plants During Feeder Restoration

Grid-forming inverters must optimally transfer power from dc-coupled photovoltaic arrays and batteries into an ac grid. Further, they must be able to restore single-phase induction motors (SPIMs) and withstand fault-induced delayed voltage-recovery (FIDVR) events. These resilience and reliability challenges are addressed here by: (i) engineering a controller to optimally operate dc-coupled hybrid resources; (ii) modeling residential air-conditioning compressors for restoration/FIDVR studies; and (iii) analyzing SPIM thermal-relay performance under limited inverter currents and designing an electronic protection for stalled SPIMs. Furthermore, these contributions are demonstrated via electromagnetic-transient simulations and can be helpful to understand recommendations by the North American Electric Reliability Corporation.

14 SOLAR ENERGY↗
NASA NTRS2001

Active Control of Rotating Stall Demonstrated for a Multistage Compressor With Inlet Distortion

Aircraft compressors can suffer debilitating consequences as a result of rotating stall and surge events caused by inlet distortions. This is particularly true of aircraft during takeoff, when the compressor is operating at peak performance close to the surge line. Significant research has been conducted by the NASA Glenn Research Center in the area of compressor stability enhancement through active and passive control methods. Most recently, an experiment was conducted at the Wright Patterson Air Force Base Research Laboratory on a two-stage fan with inlet guide vanes and inlet distortion. In this joint Small Business Innovation Research effort between Scientific Systems and Glenn, control of rotating stall was demonstrated in a multistage transonic fan. This twostage fan with inlet guide vanes was tested under clean and distorted inlet conditions. The compressor was also configured with a circumferential distortion screen capable of 180 of distortion and with 14 high-velocity injectors upstream of the first rotor. Twelve of these injectors could oscillate up to frequencies of 450 Hz. The additional two injectors were located next to each other and were used in concert with each other as a single, on/off, high-authority actuator. In a first test of injection in this multistage environment, 12 of the valves were opened 50 percent of their full stroke to assess steady injection through the compressor. This baseline injection is shown in the compressor characteristic of the following figure, and stall margin improvements are tracked from this baseline condition. The compressor was then tested with clean inlet conditions using 12 injectors and active control. Pressure disturbances were tracked before rotating stall, and a constant gain control scheme reduced the stalling mass flow by 10.8 percent over the baseline. With the distortion screen present in the inlet, a pole-zero cancellation control scheme was used to achieve a 6.4-percent decrease in stalling mass flow. These improvements also are shown in the figure. In a final experiment, actively controlled, high-frequency injection from the 12 valves was used in conjunction with the high-authority actuators. In this test, the stalling mass flow of the compressor was reduced by 27 percent as indicated in the graph. These results were obtained by injecting less than 2 percent of the total compressor throughflow into the rotor tip region via 14 injection ports. These results mark the first successful demonstration of actively controlled air injection as a stall-control strategy for multistage compressors operating at speeds typical of an actual gas turbine engine. A goal of continuing research is to determine the combination of air-injection parameters and control strategies that are most effective in providing stall control for both clean and distorted inlet flow conditions for multistage environments. Other goals include the demonstration of stall control at many locations along the core compressor and development and application of active stall control strategies that will be integral flightworthy components of onboard engine hardware.

VanSchalkwyk, Christian↗
NASA NTRS2008

Devices and methods of operation thereof for providing stable flow for centrifugal compressors

Centrifugal compressor flow stabilizing devices and methods of operation thereof are disclosed that act upon the flow field discharging from the impeller of a centrifugal compressor and modify the flow field ahead of the diffuser vanes such that flow conditions contributing to rotating stall and surge are reduced or even eliminated. In some embodiments, shaped rods and methods of operation thereof are disclosed, whereas in other embodiments reverse-tangent air injection devices and methods are disclosed.

Skoch, Gary J.↗
NASA NTRS2004

Flutter Stability of the Efficient Low Noise Fan Calculated

The TURBO-AE aeroelastic code has been used to verify the flutter stability of the Efficient Low Noise Fan (ELNF), which is also referred to as the trailing-edge blowing fan. The ELNF is a unique technology demonstrator being designed and fabricated at the NASA Glenn Research Center for testing in Glenn's 9-by-15-Foot Low-Speed Wind Tunnel. In the ELNF, air can be blown out of slots near the trailing edges of the fan blades to fill in the wakes downstream of the rotating blades. This filling of the wakes leads to a reduction of the rotor-stator interaction (tone) noise that results from the interaction of wakes with the downstream stators. The ELNF will demonstrate a 1.6-EPNdB1 reduction in tone noise through wake filling, without increasing the broadband noise. Furthermore, the reduced blade row interaction will decrease the possibility of forced response and enable closer spacing of blade rows, thus reducing engine length and weight. During the design of the ELNF, the rotor blades were checked for flutter stability using the detailed aeroelastic analysis capability of the three-dimensional Navier-Stokes TURBOAE code. The aeroelastic calculations were preceded by steady calculations in which the blades were not allowed to vibrate. For each rotational speed, as the back-pressure was increased, the mass flow rate decreased, and the operating point moved along the constant speed characteristic (speed-line) from choke to stall as shown on the fan map. The TURBO-AE aeroelastic analyses were performed separately for the first two vibration modes (bending and torsion) and covered the complete range of interblade phase angles or nodal diameters at which flutter can occur. The results indicated that the ELNF blades would not encounter flutter at takeoff conditions. The calculations were then repeated for a part-speed condition (70-percent rotational speed), and the results again showed no flutter in the operating region. On the fan map (shown), the predicted flutter point at part speed condition was located beyond the stall line, which means that the ELNF will not encounter flutter since it will never operate beyond the stall line. All the calculations done so far have been for the nonblowing case, and selected calculations will be repeated with air blowing from the trailing edge of the fan.

Bakhle, Milind A.↗
NASA NTRS2011

Correlating CFD Simulation with Wind Tunnel Test for the Full-Scale UH-60A Airloads Rotor

Data from the recent UH-60A Airloads Test in the National Full-Scale Aerodynamics Complex 40- by 80- Foot Wind Tunnel at NASA Ames Research Center are presented and compared to predictions computed by a loosely coupled Computational Fluid Dynamics (CFD)/Comprehensive analysis. Primary calculations model the rotor in free-air, but initial calculations are presented including a model of the tunnel test section. The conditions studied include a speed sweep at constant lift up to an advance ratio of 0.4 and a thrust sweep at constant speed into deep stall. Predictions show reasonable agreement with measurement for integrated performance indicators such as power and propulsive but occasionally deviate significantly. Detailed analysis of sectional airloads reveals good correlation in overall trends for normal force and pitching moment but pitching moment mean often differs. Chord force is frequently plagued by mean shifts and an overprediction of drag on the advancing side. Locations of significant aerodynamic phenomena are predicted accurately although the magnitude of individual events is often missed.

Romandr, Ethan↗
NASA NTRS1993

X-31 flight test update

The goals of the tactical utility flight test are to verify the basic tactical lessons learned in simulation and determine the situational awareness implications of post-stall maneuvering. Tactics and lessons learned from post-stall engagements will supplement current fighter tactics and complement future fighter tactics development. Although the present envelope of the X-31 for carefree post stall (PST) use is limited to 225 KIAS versus 325 KIAS in simulations, the same initial conditions used in 'Pinball II' are being utilized in flight test. Starting conditions up to 325 KIAS are being flown during the build-up phase leading to the full close in air combat (CIC) evaluations. The flight control program has been modified to limit PST entry to below 225 KIAS. Full aft stick application above 225 KIAS cannot command angle of attack (AOA) in excess of 30 deg.

Groves, AL↗
NASA NTRS1957

Measurement of Aerodynamic Forces for Various Mean Angles of Attack on an Airfoil Oscillating in Pitch and on Two Finite-span Wings Oscillating in Bending with Emphasis on Damping in the Stall

The oscillating air forces on a two-dimensional wing oscillating in pitch about the midchord have been measured at various mean angles of attack and at Mach numbers of 0.35 and 0.7. The magnitudes of normal-force and pitching-moment coefficients were much higher at high angles of attack than at low angles of attack for some conditions. Large regions of negative damping in pitch were found, and it was shown that the effect of increasing the Mach number 0.35 to 0.7 was to decrease the initial angle of attack at which negative damping occurred. Measurements of the aerodynamic damping of a 10-percent-thick and of a 3-percent-thick finite-span wing oscillating in the first bending mode indicate no regions of negative damping for this type of motion over the range of variables covered. The damping measured at high angles of attack was generally larger than that at low angles of attack. (author)

Rainey, A Gerald↗
NASA NTRS2013

Modeling Commercial Turbofan Engine Icing Risk With Ice Crystal Ingestion

The occurrence of ice accretion within commercial high bypass aircraft turbine engines has been reported under certain atmospheric conditions. Engine anomalies have taken place at high altitudes that have been attributed to ice crystal ingestion, partially melting, and ice accretion on the compression system components. The result was degraded engine performance, and one or more of the following: loss of thrust control (roll back), compressor surge or stall, and flameout of the combustor. As ice crystals are ingested into the fan and low pressure compression system, the increase in air temperature causes a portion of the ice crystals to melt. It is hypothesized that this allows the ice-water mixture to cover the metal surfaces of the compressor stationary components which leads to ice accretion through evaporative cooling. Ice accretion causes a blockage which subsequently results in the deterioration in performance of the compressor and engine. The focus of this research is to apply an engine icing computational tool to simulate the flow through a turbofan engine and assess the risk of ice accretion. The tool is comprised of an engine system thermodynamic cycle code, a compressor flow analysis code, and an ice particle melt code that has the capability of determining the rate of sublimation, melting, and evaporation through the compressor flow path, without modeling the actual ice accretion. A commercial turbofan engine which has previously experienced icing events during operation in a high altitude ice crystal environment has been tested in the Propulsion Systems Laboratory (PSL) altitude test facility at NASA Glenn Research Center. The PSL has the capability to produce a continuous ice cloud which are ingested by the engine during operation over a range of altitude conditions. The PSL test results confirmed that there was ice accretion in the engine due to ice crystal ingestion, at the same simulated altitude operating conditions as experienced previously in flight. The computational tool was utilized to help guide a portion of the PSL testing, and was used to predict ice accretion could also occur at significantly lower altitudes. The predictions were qualitatively verified by subsequent testing of the engine in the PSL. The PSL test has helped to calibrate the engine icing computational tool to assess the risk of ice accretion. The results from the computer simulation identified prevalent trends in wet bulb temperature, ice particle melt ratio, and engine inlet temperature as a function of altitude for predicting engine icing risk due to ice crystal ingestion.

Jorgenson, Philip C. E.↗
NASA NTRS2014

Modeling of Commercial Turbofan Engine With Ice Crystal Ingestion: Follow-On

The occurrence of ice accretion within commercial high bypass aircraft turbine engines has been reported under certain atmospheric conditions. Engine anomalies have taken place at high altitudes that have been attributed to ice crystal ingestion, partially melting, and ice accretion on the compression system components. The result was degraded engine performance, and one or more of the following: loss of thrust control (roll back), compressor surge or stall, and flameout of the combustor. As ice crystals are ingested into the fan and low pressure compression system, the increase in air temperature causes a portion of the ice crystals to melt. It is hypothesized that this allows the ice-water mixture to cover the metal surfaces of the compressor stationary components which leads to ice accretion through evaporative cooling. Ice accretion causes a blockage which subsequently results in the deterioration in performance of the compressor and engine. The focus of this research is to apply an engine icing computational tool to simulate the flow through a turbofan engine and assess the risk of ice accretion. The tool is comprised of an engine system thermodynamic cycle code, a compressor flow analysis code, and an ice particle melt code that has the capability of determining the rate of sublimation, melting, and evaporation through the compressor flow path, without modeling the actual ice accretion. A commercial turbofan engine which has previously experienced icing events during operation in a high altitude ice crystal environment has been tested in the Propulsion Systems Laboratory (PSL) altitude test facility at NASA Glenn Research Center. The PSL has the capability to produce a continuous ice cloud which is ingested by the engine during operation over a range of altitude conditions. The PSL test results confirmed that there was ice accretion in the engine due to ice crystal ingestion, at the same simulated altitude operating conditions as experienced previously in flight. The computational tool was utilized to help guide a portion of the PSL testing, and was used to predict ice accretion could also occur at significantly lower altitudes. The predictions were qualitatively verified by subsequent testing of the engine in the PSL. In a previous study, analysis of select PSL test data points helped to calibrate the engine icing computational tool to assess the risk of ice accretion. This current study is a continuation of that data analysis effort. The study focused on tracking the variations in wet bulb temperature and ice particle melt ratio through the engine core flow path. The results from this study have identified trends, while also identifying gaps in understanding as to how the local wet bulb temperature and melt ratio affects the risk of ice accretion and subsequent engine behavior.

Jorgenson, Philip C. E.↗
NASA NTRS2014

Modeling of Commercial Turbofan Engine with Ice Crystal Ingestion; Follow-On

The occurrence of ice accretion within commercial high bypass aircraft turbine engines has been reported under certain atmospheric conditions. Engine anomalies have taken place at high altitudes that have been attributed to ice crystal ingestion, partially melting, and ice accretion on the compression system components. The result was degraded engine performance, and one or more of the following: loss of thrust control (roll back), compressor surge or stall, and flameout of the combustor. As ice crystals are ingested into the fan and low pressure compression system, the increase in air temperature causes a portion of the ice crystals to melt. It is hypothesized that this allows the ice-water mixture to cover the metal surfaces of the compressor stationary components which leads to ice accretion through evaporative cooling. Ice accretion causes a blockage which subsequently results in the deterioration in performance of the compressor and engine. The focus of this research is to apply an engine icing computational tool to simulate the flow through a turbofan engine and assess the risk of ice accretion. The tool is comprised of an engine system thermodynamic cycle code, a compressor flow analysis code, and an ice particle melt code that has the capability of determining the rate of sublimation, melting, and evaporation through the compressor flow path, without modeling the actual ice accretion. A commercial turbofan engine which has previously experienced icing events during operation in a high altitude ice crystal environment has been tested in the Propulsion Systems Laboratory (PSL) altitude test facility at NASA Glenn Research Center. The PSL has the capability to produce a continuous ice cloud which is ingested by the engine during operation over a range of altitude conditions. The PSL test results confirmed that there was ice accretion in the engine due to ice crystal ingestion, at the same simulated altitude operating conditions as experienced previously in flight. The computational tool was utilized to help guide a portion of the PSL testing, and was used to predict ice accretion could also occur at significantly lower altitudes. The predictions were qualitatively verified by subsequent testing of the engine in the PSL. In a previous study, analysis of select PSL test data points helped to calibrate the engine icing computational tool to assess the risk of ice accretion. This current study is a continuation of that data analysis effort. The study focused on tracking the variations in wet bulb temperature and ice particle melt ratio through the engine core flow path. The results from this study have identified trends, while also identifying gaps in understanding as to how the local wet bulb temperature and melt ratio affects the risk of ice accretion and subsequent engine behavior.

Jorgenson, Philip C. E.↗
NASA NTRS1941

Flight Investigation of Wing-gun Fairings on a Fighter Type Airplane

Description is given of flight tests conducted on gun fairings, designed to correct the detrimental effects of the projecting and submerged wing guns on an F4F-3 fighter. It was found that the installation of unfaired guns on a clean wing resulted in a premature stall that increased the stalling speed in the carrier-approach and landing conditions of flight by suitably fairing the guns, it was possible to reduce the stalling speeds to values approaching very nearly the clean-wing values.

LOOPING - GROUND↗
NASA NTRS1974

Investigation of nonlinear inviscid and viscous flow effects in the analysis of dynamic stall

A method for analyzing unsteady airfoil stall was refined by including nonlinear effects in the representation of the inviscid flow. Certain other aspects of the potential-flow model were reexamined and the effects of varying Reynolds number on stall characteristics were investigated. Refinement of the formulation improved the representation of the flow and chordwise pressure distribution below stall, but substantial quantitative differences between computed and measured results are still evident for sinusoidal pitching through stall. Agreement is substantially improved by assuming the growth rate of the dead-air region at the onset of leading-edge stall is of the order of the component of the free stream normal to the airfoil chordline. The method predicts the expected increase in the resistance to stalling with increasing Reynolds number. Results indicate that a given airfoil can undergo both trailing-edge and leading-edge stall under unsteady conditions.

Crimi, P.↗
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 NTRS1958

Flight Investigation of a Normal-Acceleration Automatic Longitudinal Control System in a Fighter Airplane

A flight investigation was made to obtain experimental information on the handling qualities of a normal-acceleration type of automatic longitudinal control system. The control system was installed in a subsonic fighter-type airplane. In hands-off (stick-free) flight the normal-acceleration control system attempted to regulate the normal acceleration to a constant value which is dependent on the automatic-control-system trim setting. In maneuvering flight a given pilot's stick deflection produced a proportional change in normal acceleration, the change in acceleration being independent of flight condition. A small side-located controller stick was used by the pilot to introduce signals into the automatic control system. In the flight program emphasis was placed on the acceleration-limiting capabilities of the control system. The handling qualities were investigated in maneuvers such as slow and rapid pull-ups and turns and also in flight operations such as cruising, stalls, landings, aerobatics, and air-to-air tracking. Good acceleration limiting was obtained with the normal-acceleration control system by limiting the magnitude of the input signal that the pilot could introduce into the control system. The same values of control-system gain settings could be used from an acceleration-limiting stand-point at both 10,000 and 30,000 feet for the complete speed range of the airplane. The response characteristics of the airplane-control system combination were also satisfactory at both high and low altitude with these same values of control-system gain setting. In the pilot's opinion, the normal-acceleration control system provided good stability and control characteristics in flight operations such as cruising, stalls, landings, aerobatics, and air-to-air tracking.

Sjoberg, S. A.↗
NASA NTRS2002

Active Control Analysis for Aeroelastic Instabilities in Turbomachines

Turbomachines onboard aircraft operate in a highly complex and harsh environment. The unsteady flowfield inherent to turbomachines leads to several problems associated with safety, stability, performance and noise. In-flight surge or flutter incidents could be catastrophic and impact the safety and reliability of the aircraft. High-Cycle-Fatigue (HCF), on the other hand, can significantly impact safety, readiness and maintenance costs. To avoid or minimize these problems generally a more conservative design method must be initiated which results in thicker blades and a loss of performance. Actively controlled turbomachines have the potential to reduce or even eliminate the instabilities by impacting the unsteady aerodynamic characteristics. By modifying the unsteady aerodynamics, active control may significantly improve the safety and performance especially at off-design conditions, reduce noise, and increase the range of operation of the turbomachine. Active control can also help improve reliability for mission critical applications such as the Mars Flyer. In recent years, HCF has become one of the major issues concerning the cost of operation for current turbomachines. HCF alone accounts for roughly 30% of maintenance cost for the United States Air-Force. Other instabilities (flutter, surge, rotating-stall, etc.) are generally identified during the design and testing phase. Usually a redesign overcomes these problems, often reducing performance and range of operation, and resulting in an increase in the development cost and time. Despite a redesign, the engines do not have the capabilities or means to cope with in-flight unforeseen vibration, stall, flutter or surge related instabilities. This could require the entire fleet worldwide to be stood down for expensive modifications. These problems can be largely overcome by incorporating active control within the turbomachine and its design. Active control can help in maintaining the integrity of the system in unforeseen events and provide for more aggressive designs to reduce the weight and improve efficiency of the turbomachine. Another area where active control can be useful is in controlling and suppressing rotating stall and surge in compressors, thereby increasing its operating range. Although some of these benefits will be offset by the added cost and weight penalty of the control system, the potential benefits in safety, reliability, performance, and noise characteristics are significant enough to warrant research in the area of active control of turbomachines. There is renewed interest within industry to understand unsteady aerodynamic behavior. This improved understanding not only leads to better design of turbomachines, which avoid instabilities but also which helps in understanding the controllability of the instabilities. The proliferation of micro-electro-mechanical-system (MEMS) devices has made available new tools to designers for employing feedback controls at reasonable costs. MEMS have also made the control devices small and unobtrusive enough to be implemented within the turbomachine without significant obstruction to the flow path. This has made active-control very attractive especially for systems requiring extreme confidence.

Keith, Theo G., Jr.↗