Engineering topics

Strazisar, Anthony J.

Publications and source records attributed to Strazisar, Anthony J..

At least 19 records

NASA NTRS2020

Aerothermodynamics Technical Working Group: 2008 Turbomachinery Technology Assessment and Recommendations

This report documents the 2008 Turbomachinery Technology Assessment and Recommendations of the NASA-led Aerothermodynamics Technical Working Group (TWG). It includes technology review, assessment, and recommendation for future research and development. The areas covered are summarized as follows. Inlet Flow Distortion Sensitivity and Stability, Tip Leakage Flows in High-Pressure Cores, Endwall Contouring, Turbine Tip Flows, Combustor and Cooled-Turbine Interaction, Highly Loaded Low-Pressure Turbines.

Heidmann, James D.↗

NASA NTRS2006

Endwall Treatment and Method for Gas Turbine

An endwall treatment for a gas turbine engine having at least one rotor blade extending from a rotatable hub and a casing circumferentially surrounding the rotor and the hub, the endwall treatment including, an inlet formed in an endwall of the gas turbine engine adapted to ingest fluid from a region of a higher-pressure fluid, an outlet formed in the endwall and located in a region of lower pressure than the inlet, wherein the inlet and the outlet are in a fluid communication with each other, the outlet being adapted to inject the fluid from the inlet in the region of lower pressure, and wherein the outlet is at least partially circumferentially offset relative to the inlet.

Hathaway, Michael D.↗

NASA NTRS2005

Adaptive Technologies Developed and Demonstrated for Extending the Stable Operating Range of Compressors

The operational envelope of gas turbine engines is constrained by the stability limit of the compression system. The dangers of exceeding this limit are severe, with the potential for engine failure and loss of the aircraft. To avoid such failures, compressor designers provide an adequate stability (stall) margin in the compressor design to account for inlet distortions, degradation due to wear, throttle transients, and other factors that reduce compressor stability from the original design intent. In some cases, the required stall margin results in the compressor operating line being below the maximum efficiency potential of the compression system. Current approaches to increasing stability tend to decrease the efficiency of the compressor. The focus of this work is to increase the stall margin of compressors without decreasing their efficiency.

Hathaway, Michael D.↗

NASA NTRS2004

Active Closed-Loop Stator Vane Flow Control Demonstrated in a Low-Speed Multistage Compressor

Closed-loop flow control was successfully demonstrated on the surface of stator vanes in NASA Glenn Research Center's Low-Speed Axial Compressor (LSAC) facility. This facility provides a flow field that accurately duplicates the aerodynamics of modern highly loaded compressors. Closed-loop active flow control uses sensors and actuators embedded within engine components to dynamically alter the internal flow path during off-nominal operation in order to optimize engine performance and maintain stable operation.

Bright, Michelle M.↗

NASA NTRS2003

Active Flow Separation Control of a Stator Vane Using Surface Injection in a Multistage Compressor Experiment

Micro-flow control actuation embedded in a stator vane was used to successfully control separation and improve near stall performance in a multistage compressor rig at NASA Glenn. Using specially designed stator vanes configured with internal actuation to deliver pulsating air through slots along the suction surface, a research study was performed to identify performance benefits using this microflow control approach. Pressure profiles and unsteady pressure measurements along the blade surface and at the shroud provided a dynamic look at the compressor during microflow air injection. These pressure measurements lead to a tracking algorithm to identify the onset of separation. The testing included steady air injection at various slot locations along the vane. The research also examined the benefit of pulsed injection and actively controlled air injection along the stator vane. Two types of actuation schemes were studied, including an embedded actuator for on-blade control. Successful application of an online detection and flow control scheme will be discussed. Testing showed dramatic performance benefit for flow reattachment and subsequent improvement in diffusion through the use of pulsed controlled injection. The paper will discuss the experimental setup, the blade configurations, and preliminary CFD results which guided the slot location along the blade. The paper will also show the pressure profiles and unsteady pressure measurements used to track flow control enhancement, and will conclude with the tracking algorithm for adjusting the control.

Culley, Dennis E.↗

NASA NTRS2002

The Effect of Variable Chord Length on Transonic Axial Rotor Performance

Aircraft fan and compressor blade leading edges suffer from atmospheric particulate erosion that reduces aerodynamic performance. Recontouring the blade leading edge region can restore blade performance. This process typically results in blades of varying chord length. The question therefore arises as to whether performance of refurbished fans and compressors could be further improved if blades of varying chord length are installed into the disk in a certain order. To investigate this issue the aerodynamic performance of a transonic compressor rotor operating with blades of varying chord length was measured in back-to-back compressor test rig entries. One half of the rotor blades were the full nominal chord length while the remaining half of the blades were cut back at the leading edge to 95% of chord length and recontoured. The rotor aerodynamic performance was measured at 100, 80, and 60% of design speed for three blade installation configurations: nominal-chord blades in half of the disk and short-chord blades in half of the disk; four alternating quadrants of nominal-chord and short-chord blades; nominal-chord and short-chord blades alternating around the disk. No significant difference in performance was found between configurations, indicating that blade chord variation is not important to aerodynamic performance above the stall chord limit if leading edges have the same shape. The stall chord limit for most civil aviation turbofan engines is between 94-96% of nominal (new) blade chord.

Roberts, William B.↗

NASA NTRS2001

Compressor Stability Enhancement Using Discrete Tip Injection

Mass injection upstream of the tip of a high-speed axial compressor rotor is a stability enhancement approach known to be effective in suppressing small in tip-critical rotors. This process is examined in a transonic axial compressor rotor through experiments and time-averaged Navier-Stokes CFD simulations. Measurements and simulations for discrete injection are presented for a range of injection rates and distributions of injectors around the annulus. The simulations indicate that tip injection increases stability by unloading the rotor tip and that increasing injection velocity improves the effectiveness of tip injection. For the tested rotor, experimental results demonstrate that at 70 percent speed the stalling flow coefficient can be reduced by 30 percent using an injected mass- flow equivalent to 1 percent of the annulus flow. At design speed, the stalling flow coefficient was reduced by 6 percent using an injected mass-fiow equivalent to 2 percent of the annulus flow. The experiments show that stability enhancement is related to the mass-averaged axial velocity at the tip. For a given injected mass-flow, the mass-averaged axial velocity at the tip is increased by injecting flow over discrete portions of the circumference as opposed to full-annular injection. The implications of these results on the design of recirculating casing treatments and other methods to enhance stability will be discussed.

Suder, Kenneth L.↗

NASA NTRS2000

Recommendations for Achieving Accurate Numerical Simulation of Tip Clearance Flows in Transonic Compressor Rotors

The tip clearance flows of transonic compressor rotors are important because they have a significant impact on rotor and stage performance. While numerical simulations of these flows are quite sophisticated. they are seldom verified through rigorous comparisons of numerical and measured data because these kinds of measurements are rare in the detail necessary to be useful in high-speed machines. In this paper we compare measured tip clearance flow details (e.g. trajectory and radial extent) with corresponding data obtained from a numerical simulation. Recommendations for achieving accurate numerical simulation of tip clearance flows are presented based on this comparison. Laser Doppler Velocimeter (LDV) measurements acquired in a transonic compressor rotor, NASA Rotor 35, are used. The tip clearance flow field of this transonic rotor was simulated using a Navier-Stokes turbomachinery solver that incorporates an advanced k-epsilon turbulence model derived for flows that are not in local equilibrium. Comparison between measured and simulated results indicates that simulation accuracy is primarily dependent upon the ability of the numerical code to resolve important details of a wall-bounded shear layer formed by the relative motion between the over-tip leakage flow and the shroud wall. A simple method is presented for determining the strength of this shear layer.

VanZante, Dale E.↗

NASA NTRS2000

Accuracy of Numerical Simulations of Tip Clearance Flow in Transonic Compressor Rotors Improved Dramatically

The tip clearance flows of transonic compressor rotors have a significant impact on rotor and stage performance. Although numerical simulations of these flows are quite sophisticated, they are seldom verified through rigorous comparisons of numerical and measured data because, in high-speed machines, measurements acquired in sufficient detail to be useful are rare. Researchers at the NASA Glenn Research Center at Lewis Field compared measured tip clearance flow details (e.g., trajectory and radial extent) of the NASA Rotor 35 with results obtained from a numerical simulation. Previous investigations had focused on capturing the detailed development of the jetlike flow leaking through the clearance gap between the rotating blade tip and the stationary compressor shroud. However, we discovered that the simulation accuracy depends primarily on capturing the detailed development of a wall-bounded shear layer formed by the relative motion between the leakage jet and the shroud.

VanZante, Dale E.↗

NASA NTRS1998

Jet Injection Used to Control Rotating Stall in a High-Speed Compressor

In a joint effort between the Massachusetts Institute of Technology (MIT) and the NASA Lewis Research Center, a new technology was demonstrated to identify and control rotating stall and surge in a single-stage, high-speed compressor. Through the use of highvelocity, high-frequency jet injectors, the instabilities of surge and stall were controlled in a high-speed compressor rig. Through the use of active stall control, modal instabilities that normally occur in the pressure measurements prior to stall were normalized and the range of the compressor was extended. Normally the events of rotating stall and surge instabilities limit the operation of the aeroengine compressor to a region below the surge line. To enhance the performance of the compressor, the Lewis/MIT team used active stall control methods to extend the normal operation of the compressor beyond the original stall point.

Bright, Michelle M.↗

NASA NTRS1995

The effect of adding roughness and thickness to a transonic axial compressor rotor

The performance deterioration of a high speed axial compressor rotor due to surface roughness and airfoil thickness variations is reported. A 0.025 mm (0.001 in.) thick rough coating with a surface finish of 2.54-3.18 RMS microns (100-125 RMS microinches) is applied to the pressure and suction surface of the rotor blades. Coating both surfaces increases the leading edge thickness by 10% at the hub and 20% at the tip. Application of this coating results in a loss in efficiency of 6 points and a 9% reduction in the pressure ratio across the rotor at an operating condition near the design point. To separate the effect of thickness and roughness, a smooth coating of equal thickness is also applied to the blade. The smooth coating surface finish is 0.254-0.508 RMS microns (10-20 RMS microinches), compared to the bare metal blade surface finish of 0.508 RMS microns (20 RMS microinches). The smooth coating results in approximately half of the performance deterioration found from the rough coating. Both coatings are then applied to different portions of the blade surface to determine which portions of the airfoil are most sensitive to thickness/roughness variations. Aerodynamic performance measurements are presented for a number of coating configurations at 60%, 80%, and 100% of design speed. The results indicate that thickness/roughness over the first 10% of blade chord accounts for virtually all of the observed performance degradation for the smooth coating, compared to about 70% of the observed performance degradation for the rough coating. The performance deterioration is investigated in more detail at design speed using laser anemometer measurements as well as predictions generated by a quasi-3D Navier-Stokes flow solver which includes a surface roughness model. Measurements and analysis are performed on the baseline blade and the full-coverage smooth and rough coatings. The results indicate that coating the blade causes a thickening of the blade boundary layers. The interaction between the rotor passage shock and the thickened suction surface boundary layer then results in an increase in blockage which reduces the diffusion level in the rear half of the blade passage, thus reducing the aerodynamic performance of the rotor.

Suder, Kenneth L.↗

NASA NTRS1995

Laser Anemometer Measurements of the Three-Dimensional Rotor Flow Field in the NASA Low-Speed Centrifugal Compressor

A laser anemometer system was used to provide detailed surveys of the three-dimensional velocity field within the NASA low-speed centrifugal impeller operating with a vaneless diffuser. Both laser anemometer and aerodynamic performance data were acquired at the design flow rate and at a lower flow rate. Floor path coordinates, detailed blade geometry, and pneumatic probe survey results are presented in tabular form. The laser anemometer data are presented in the form of pitchwise distributions of axial, radial, and relative tangential velocity on blade-to-blade stream surfaces at 5-percent-of-span increments, starting at 95-percent-of-span from the hub. The laser anemometer data are also presented as contour and wire-frame plots of throughflow velocity and vector plots of secondary velocities at all measurement stations through the impeller.

Hathaway, Michael D.↗

NASA NTRS1993

Experimental and computational investigation of the NASA low-speed centrifugal compressor flow field

An experimental and computational investigation of the NASA Lewis Research Center's low-speed centrifugal compressor (LSCC) flow field was conducted using laser anemometry and Dawes' three-dimensional viscous code. The experimental configuration consisted of a backswept impeller followed by a vaneless diffuser. Measurements of the three-dimensional velocity field were acquired at several measurement planes through the compressor. The measurements describe both the throughflow and secondary velocity field along each measurement plane. In several cases the measurements provide details of the flow within the blade boundary layers. Insight into the complex flow physics within centrifugal compressors is provided by the computational fluid dynamics analysis (CFD), and assessment of the CFD predictions is provided by comparison with the measurements. Five-hole probe and hot-wire surveys at the inlet and exit to the impeller as well as surface flow visualization along the impeller blade surfaces provided independent confirmation of the laser measurement technique. The results clearly document the development of the throughflow velocity wake that is characteristic of unshrouded centrifugal compressors.

Hathaway, Michael D.↗

NASA NTRS1991

Enhancing aeropropulsion research with high-speed interactive computing

NASA-Lewis has committed to a long range goal of creating a numerical test cell for aeropropulsion research and development. Efforts are underway to develop a first generation Numerical Propulsion System Simulation (NPSS). The NPSS will provide a unique capability to numerically simulate advanced propulsion systems from nose to tail. Two essential ingredients to the NPSS are: (1) experimentally validated Computational Fluid Dynamics (CFD) codes; and (2) high performing computing systems (hardware and software) that will permit those codes to be used efficiently. To this end, NASA-Lewis is using high speed, interactive computing as a means for achieving Integrated CFD and Experiments (ICE). The development is described of a prototype ICE system for multistage compressor flow physics research.

Szuch, John R.↗

NASA NTRS1991

Enhancing aeropropulsion research with high-speed interactive computing

Szuch, John R.↗

NASA NTRS1990

Turbomachinery

Simoneau, Robert J.↗

NASA NTRS1989

Laser anemometer measurements in a transonic axial-flow fan rotor

Laser anemometer surveys were made of the 3-D flow field in NASA rotor 67, a low aspect ratio transonic axial-flow fan rotor. The test rotor has a tip relative Mach number of 1.38. The flowfield was surveyed at design speed at near peak efficiency and near stall operating conditions. Data is presented in the form of relative Mach number and relative flow angle distributions on surfaces of revolution at nine spanwise locations evenly spaced from hub to tip. At each spanwise location, data was acquired upstream, within, and downstream of the rotor. Aerodynamic performance measurements and detailed rotor blade and annulus geometry are also presented so that the experimental results can be used as a test case for 3-D turbomachinery flow analysis codes.

Strazisar, Anthony J.↗

NASA NTRS1987

Turbomachinery

The discipline research in turbomachinery, which is directed toward building the tools needed to understand such a complex flow phenomenon, is based on the fact that flow in turbomachinery is fundamentally unsteady or time dependent. Success in building a reliable inventory of analytic and experimental tools will depend on how the time and time-averages are treated, as well as on who the space and space-averages are treated. The raw tools at disposal (both experimentally and computational) are truly powerful and their numbers are growing at a staggering pace. As a result of this power, a case can be made that a situation exists where information is outstripping understanding. The challenge is to develop a set of computational and experimental tools which genuinely increase understanding of the fluid flow and heat transfer in a turbomachine. Viewgraphs outline a philosophy based on working on a stairstep hierarchy of mathematical and experimental complexity to build a system of tools, which enable one to aggressively design the turbomachinery of the next century. Examples of the types of computational and experimental tools under current development at Lewis, with progress to date, are examined. The examples include work in both the time-resolved and time-averaged domains. Finally, an attempt is made to identify the proper place for Lewis in this continuum of research.

Simoneau, Robert J.↗