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Lepicovsky, J.

Publications and source records attributed to Lepicovsky, J..

At least 19 records
All sourcesNASA NTRSDOE OSTI
NASA NTRS2003

Resonance Effects in the NASA Transonic Flutter Cascade Facility

Investigations of unsteady pressure loadings on the blades of fans operating near the stall flutter boundary are carried out under simulated conditions in the NASA Transonic Flutter Cascade facility (TFC). It has been observed that for inlet Mach numbers of about 0.8, the cascade flowfield exhibits intense low-frequency pressure oscillations. The origins of these oscillations were not clear. It was speculated that this behavior was either caused by instabilities in the blade separated flow zone or that it was a tunnel resonance phenomenon. It has now been determined that the strong low-frequency oscillations, observed in the TFC facility, are not a cascade phenomenon contributing to blade flutter, but that they are solely caused by the tunnel resonance characteristics. Most likely, the self-induced oscillations originate in the system of exit duct resonators. For sure, the self-induced oscillations can be significantly suppressed for a narrow range of inlet Mach numbers by tuning one of the resonators. A considerable amount of flutter simulation data has been acquired in this facility to date, and therefore it is of interest to know how much this tunnel self-induced flow oscillation influences the experimental data at high subsonic Mach numbers since this facility is being used to simulate flutter in transonic fans. In short, can this body of experimental data still be used reliably to verify computer codes for blade flutter and blade life predictions? To answer this question a study on resonance effects in the NASA TFC facility was carried out. The results, based on spectral and ensemble averaging analysis of the cascade data, showed that the interaction between self-induced oscillations and forced blade motion oscillations is very weak and can generally be neglected. The forced motion data acquired with the mistuned tunnel, when strong self-induced oscillations were present, can be used as reliable forced pressure fluctuations provided that they are extracted from raw data sets by an ensemble averaging procedure.

Lepicovsky, J.↗
NASA NTRS2002

Methodology of Blade Unsteady Pressure Measurement in the NASA Transonic Flutter Cascade

In this report the methodology adopted to measure unsteady pressures on blade surfaces in the NASA Transonic Flutter Cascade under conditions of simulated blade flutter is described. The previous work done in this cascade reported that the oscillating cascade produced waves, which for some interblade phase angles reflected off the wind tunnel walls back into the cascade, interfered with the cascade unsteady aerodynamics, and contaminated the acquired data. To alleviate the problems with data contamination due to the back wall interference, a method of influence coefficients was selected for the future unsteady work in this cascade. In this approach only one blade in the cascade is oscillated at a time. The majority of the report is concerned with the experimental technique used and the experimental data generated in the facility. The report presents a list of all test conditions for the small amplitude of blade oscillations, and shows examples of some of the results achieved. The report does not discuss data analysis procedures like ensemble averaging, frequency analysis, and unsteady blade loading diagrams reconstructed using the influence coefficient method. Finally, the report presents the lessons learned from this phase of the experimental effort, and suggests the improvements and directions of the experimental work for tests to be carried out for large oscillation amplitudes.

Lepicovsky, J.↗
NASA NTRS2002

Unsteady Pressures in a Transonic Fan Cascade Due to a Single Oscillating Airfoil

An extensive set of unsteady pressure data was acquired along the midspan of a modern transonic fan blade for simulated flutter conditions. The data set was acquired in a nine-blade linear cascade with an oscillating middle blade to provide a database for the influence coefficient method to calculate instantaneous blade loadings. The cascade was set for an incidence of 10 dg. The data were acquired on three stationary blades on each side of the middle blade that was oscillated at an amplitude of 0.6 dg. The matrix of test conditions covered inlet Mach numbers of 0.5, 0.8, and 1.1 and the oscillation frequencies of 200, 300, 400, and 500 Hz. A simple quasiunsteady two-dimensional computer simulation was developed to aid in the running of the experimental program. For high Mach number subsonic inlet flows the blade pressures exhibit very strong, low-frequency, self-induced oscillations even without forced blade oscillations, while for low subsonic and supersonic inlet Mach numbers the blade pressure unsteadiness is quite low. The amplitude of forced pressure fluctuations on neighboring stationary blades strongly depends on the inlet Mach number and forcing frequency. The flowfield behavior is believed to be governed by strong nonlinear effects due to a combination of viscosity, compressibility, and unsteadiness. Therefore, the validity of the quasi-unsteady simplified computer simulation is limited to conditions when the flowfield is behaving in a linear, steady manner. Finally, an extensive set of unsteady pressure data was acquired to help development and verification of computer codes for blade flutter effects.

Lepicovsky, J.↗
NASA NTRS2002

Intermittent Flow Regimes in a Transonic Fan Airfoil Cascade

A study was conducted in the NASA Glenn Research Center linear cascade on the intermittent flow on the suction surface of an airfoil section from the tip region of a modern low aspect ratio fan blade. Experimental results revealed that, at a large incidence angle, a range of transonic inlet Mach numbers exist where the leading-edge shock-wave pattern was unstable. Flush mounted high frequency response pressure transducers indicated large local jumps in the pressure in the leading edge area, which generates large intermittent loading on the blade leading edge. These measurements suggest that for an inlet Mach number between 0.9 and 1.0 the flow is bi-stable, randomly switching between subsonic and supersonic flows. Hence, it appears that the change in overall flow conditions in the transonic region is based on the frequency of switching between two stable flow states rather than on the continuous increase of the flow velocity. To date, this flow behavior has only been observed in a linear transonic cascade. Further research is necessary to confirm this phenomenon occurs in actual transonic fans and is not the byproduct of an endwall restricted linear cascade.

Lepicovsky, J.↗
NASA NTRS2000

On Flowfield Periodicity in the NASA Transonic Flutter Cascade: Experimental Study - Part 1

An extensive study to improve flow uniformity and periodicity in the NASA Transonic Flutter Cascade is presented here. The results are reported in two independent parts dealing with the experimental approach and the analytical approach. The first part, the Experimental Study, focuses first on the data sets acquired in this facility in the past and explains several discrepancies, particularly the questions of actual flow incidence and cascade back pressure levels. Next, available means for control and modifications of the cascade flowfield, boundary layer bleed and tailboard settings are presented in detail. This is followed by experimental data sets acquired in modified test facility configurations that were based on analytical predictions of the cascade flowfield. Finally, several important conclusions about improving the cascade flowfield uniformity and blade load periodicity are summarized. The important conclusions are: (1) boundary layer bleed does not improve the cascade flow periodicity; (2) tunnel wall contours must be carefully matched to the expected shape of cascade streamlines; (3) actual flow incidence for each cascade configuration rather must be measured instead of relying on the tunnel geometry; and (4) the current cascade configuration exhibits a very high blade load uniformity over six blades from blade #2 to blade #7, and the facility is now ready for unsteady pressure data acquisition.

Lepicovsky, J.↗
NASA NTRS1998

Application of Pressure Sensitive Paint to Confined Flow at Mach Number 2.5

Pressure sensitive paint (PSP) is a novel technology that is being used frequently in external aerodynamics. For internal flows in narrow channels, and applications at elevated nonuniform temperatures, however, there are still unresolved problems that complicate the procedures for calibrating PSP signals. To address some of these problems, investigations were carried out in a narrow channel with supersonic flows of Mach 2.5. The first set of tests focused on the distribution of the wall pressure in the diverging section of the test channel downstream of the nozzle throat. The second set dealt with the distribution of wall static pressure due to the shock/wall interaction caused by a 25 deg. wedge in the constant Mach number part of the test section. In addition, the total temperature of the flow was varied to assess the effects of temperature on the PSP signal. Finally, contamination of the pressure field data, caused by internal reflection of the PSP signal in a narrow channel, was demonstrated. The local wall pressures were measured with static taps, and the wall pressure distributions were acquired by using PSP. The PSP results gave excellent qualitative impressions of the pressure field investigated. However, the quantitative results, specifically the accuracy of the PSP data in narrow channels, show that improvements need to be made in the calibration procedures, particularly for heated flows. In the cases investigated, the experimental error had a standard deviation of +/- 8.0% for the unheated flow, and +/- 16.0% for the heated flow, at an average pressure of 11 kpa.

Lepicovsky, J.↗
NASA NTRS1996

Modern Experimental Techniques in Turbine Engine Testing

The paper describes application of two modern experimental techniques, thin-film thermocouples and pressure sensitive paint, to measurement in turbine engine components. A growing trend of using computational codes in turbomachinery design and development requires experimental techniques to refocus from overall performance testing to acquisition of detailed data on flow and heat transfer physics to validate these codes for design applications. The discussed experimental techniques satisfy this shift in focus. Both techniques are nonintrusive in practical terms. The thin-film thermocouple technique improves accuracy of surface temperature and heat transfer measurements. The pressure sensitive paint technique supplies areal surface pressure data rather than discrete point values only. The paper summarizes our experience with these techniques and suggests improvements to ease the application of these techniques for future turbomachinery research and code verifications.

Lepicovsky, J.↗
NASA NTRS1996

Seeding for laser velocimetry in confined supersonic flows with shocks

There is a lack of firm conclusions or recommendations in the open literature to guide laser velocimeter (LV) users in minimizing the uncertainty of LV data acquired in confined supersonic flows with steep velocity gradients. This fact led the NASA Lewis Research Center (LeRC) in Cleveland (Ohio, USA), and the Institute of Propulsion Technology of DLR in Cologne (Germany) to a joint research effort to improve reliability of LV measurements in supersonic flows. Over the years, NASA and DLR have developed different expertise in laser velocimetry, using different LV systems: Doppler and two-spot (L2F). The goal of the joint program is to improve the reliability of LV measurements by comparing results from experiments in confined supersonic flows performed under identical test conditions but using two different LV systems and several seed particle generators. Initial experiments conducted at the NASA LERC are reported in this paper. The experiments were performed in a narrow channel with Mach number 2.5 flow containing an oblique shock wave generated by an immersed 25-dg wedge.

Lepicovsky, J.↗
NASA NTRS1995

Application of Thin-Film Thermocouples to Localized Heat Transfer Measurements

The paper describes a proof-of-concept experiment on thin-film thermocouples used for localized heat transfer measurements applicable to experiments on hot parts of turbine engines. The paper has three main parts. The first part describes the thin-film sensors and manufacturing procedures. Attention is paid to connections between thin-film thermocouples and lead wires, which has been a source of problems in the past. The second part addresses the test arrangement and facility used for the heat transfer measurements modeling the conditions for upcoming warm turbine tests at NASA LeRC. The paper stresses the advantages of a modular approach to the test rig design. Finally, we present the results of bulk and local heat flow rate measurements, as well as overall heat transfer coefficients obtained from measurements in a narrow passage with an aspect ratio of 11.8. The comparison of bulk and local heat flow rates confirms applicability of thin-film thermocouples to upcoming warm turbine tests.

Lepicovsky, J.↗
NASA NTRS1992

LV software support for supersonic flow analysis

The software for configuring an LV counter processor system has been developed using structured design. The LV system includes up to three counter processors and a rotary encoder. The software for configuring and testing the LV system has been developed, tested, and included in an overall software package for data acquisition, analysis, and reduction. Error handling routines respond to both operator and instrument errors which often arise in the course of measuring complex, high-speed flows. The use of networking capabilities greatly facilitates the software development process by allowing software development and testing from a remote site. In addition, high-speed transfers allow graphics files or commands to provide viewing of the data from a remote site. Further advances in data analysis require corresponding advances in procedures for statistical and time series analysis of nonuniformly sampled data.

Bell, W. A.↗
NASA NTRS1990

An experimental investigation of nozzle-exit boundary layers of highly heated free jets

An experimental investigation of the effects of nozzle operating conditions on the development of nozzle-exit boundary layers of highly heated air free jets is reported in this paper. The total pressure measurements in the nozzle-exit boundary layer were obtained at a range of jet Mach numbers from 0.1 to 0.97 and jet total temperatures up to 900 K. The analysis of results shows that the nozzle-exit laminar boundary-layer development depends only on the nozzle-exit Reynolds number. For the nozzle-exit turbulent boundary layer, however, it appears that the effects of the jet total temperature on the boundary-layer integral characteristics are independent from the effects of the nozzle-exit Reynolds number. This surprizing finding has not yet been reported. Further, laminar boundary-layer profiles were compared with the Pohlhausen solution for a flat-wall converging channel and an acceptable agreement was found only for low Reynolds numbers. For turbulent boundary layers, the dependence of the shape factor on relative Mach numbers at a distance of one momentum thickness from the nozzle wall resembles Spence's prediction. Finally, the calculated total pressure loss coefficient was found to depend on the nozzle-exit Reynolds number for the laminar nozzle-exit boundary layer, while for the turbulent exit boundary layer this coefficient appears to be constant.

Lepicovsky, J.↗
NASA NTRS1990

Total temperature effects on centerline Mach number characteristics of freejets

This paper describes a detailed experimental study of Mach number centerline characteristics of unheated and heated freejets. The jet characteristics were obtained at a range of jet Mach numbers from 0.1 to 0.9 and jet total temperatures up to 900 K. Previously published results showed that the jet total temperature significantly affects the nozzle-exit, boundary-layer characteristics. The results of this study indicate that a strong correlation exists between nozzle-exit, boundary-layer conditions and freejet development. It is clear from this investigation that experimental data on freejet development cannot be meaningfully compared from one facility to another without specific knowledge of nozzle-exit, boundary-layer conditions. It was concluded that direct effect of the jet operating conditions (elevated flow temperature) on freejet development is much less important than the indirect effect due to changed nozzle-exit, boundary-layer characteristics.

Lepicovsky, J.↗
NASA NTRS1988

Acoustically excited heated jets. 1: Internal excitation

The effects of relatively strong upstream acoustic excitation on the mixing of heated jets with the surrounding air are investigated. To determine the extent of the available information on experiments and theories dealing with acoustically excited heated jets, an extensive literature survey was carried out. The experimental program consisted of flow visualization and flowfield velocity and temperature measurements for a broad range of jet operating and flow excitation conditions. A 50.8-mm-diam nozzle was used for this purpose. Parallel to the experimental study, an existing theoretical model of excited jets was refined to include the region downstream of the jet potential core. Excellent agreement was found between theory and experiment in moderately heated jets. However, the theory has not yet been confirmed for highly heated jets. It was found that the sensitivity of heated jets to upstream acoustic excitation varies strongly with the jet operating conditions and that the threshold excitation level increases with increasing jet temperature. Furthermore, the preferential Strouhal number is found not to change significantly with a change of the jet operating conditions. Finally, the effects of the nozzle exit boundary layer thickness appear to be similar for both heated and unheated jets at low Mach numbers.

Lepicovsky, J.↗
NASA NTRS1988

Acoustically excited heated jets. 2: In search of a better understanding

The second part of a three-part report on the effects of acoustic excitation on jet mixing includes the results of an experimental investigation directed at resolving the question of poor excitability of some of the heated jets. The theoretical predictions discussed in Part 1 are examined to find explanations for the observed discrepancies between the measured and the predicted results. Additional testing was performed by studying the self excitation of the shock containing hot jets and also by exciting the jet by sound radiated through source tubes located externally around the periphery of the jet. The effects of nozzle-exit boundary layer conditions on jet excitability was also investigated. It is concluded that high-speed, heated jet mixing rates and consequently also the jet excitability strongly depends on nozzle exit boundary layer conditions.

Lepicovsky, J.↗
NASA NTRS1988

Acoustically excited heated jets. 3: Mean flow data

This is Part 3 of a report on the excitability of heated jets under the influence of acoustic excitation. The effects of upstream internal acoustic excitation on jet mixing were described in Part 1. Part 2 described the effects of external excitation on flow mixing. Part 3 contains quantitative results from the measurements of mean Mach number and temperature and consists of radial profiles and centerline distributions measured at selected jet operating conditions for internally excited and unexcited jets. The mean flow data are presented in both graphical and tabulated forms. For the sake of completeness, this part contains temperature probe calibration curves also.

Lepicovsky, J.↗
NASA NTRS1987

Effects of nozzle-exit boundary-layer conditions on excitability of heated free jets

This paper reports an experimental study on the effects of nozzle-exit boundary-layer conditions on the excitability of heated free jets. The results were obtained at a Mach number 0.8 and total temperatures of 300 K and 670 K. External acoustic excitation was used to excite the jet. The excitation frequencies ranged up to 6 KHZ and the excitation levels were up to 150 dB. A level of 147 dB was used for most of the test points. Nozzle-exit boundary-layer characteristics were controlled by boundary layer tripping upstream of the nozzle exit plane. It has been shown that the free jet mixing rate strongly depends on the boundary-layer characteristics at the nozzle exit. Further, it appears that jets with a thin laminar nozzle-exit boundary layer are more selective about the optimum excitation frequency than those with a thick turbulent nozzle-exit boundary layer. Finally, for the jet of Mach number 0.8, it appears that free jet mixing and development may be controlled by flow excitation as well as by nozzle-exit boundary-layer modifications.

Lepicovsky, J.↗