1 edition of Modifications to the inlet flow field of a transonic compressor rotor found in the catalog.
Modifications to the inlet flow field of a transonic compressor rotor
1985 by Naval Postgraduate School, Available from National Technical Information Service in Monterey, Calif, Springfield, Va .
Written in English
The flow to the rotor of a single stage axial transonic compressor was found to be in disagreement with design requirements. The incidence angle to the rotor blading was measured to be towards stall in the hub and tip regions when close to design near center-span. Hardware modifications were made, guided by flow field calculations. Predictions were compared with measurements. The most significant improvement was obtained by increasing the flow rate at open throttle. No practical method to correct the flow near the hub was found. Keywords: MERIDL computer program; Transonic compressor; Inlet flow field; Distortion screens.
|Contributions||Naval Postgraduate School (U.S.). Dept. of Aeronautics|
|The Physical Object|
|Pagination||vi, 44 p. :|
|Number of Pages||44|
TEACC was validated against experimental data from the transonic NASA rotor, Rotor IB, for a clean inlet and for an inlet distortion produced by a the inlet to produce a highly distorted total pressure flow field to the compressor inlet. High distortion levels may cause the sive modifications to parallel compressor theory through model. A time-accurate three-dimensional Navier-Stokes solver of the unsteady flow field in a transonic fan was carried out using "Fluent-parallel" in a parallel supercomputer. The numerical simulation focused on a transonic fan with inlet square wave total pressure . payne, thomas (lt,usn). title:inlet flow-field measurements of a transonic compressor rotor prior to and during steam-induced rotating stall: bochette, nikolaus (ens,usn). title:computational analysis of flow through a transonic compressor rotor: brunner, matthew (ens,usn).
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Modifications to the Inlet Flow Field of a Transonic Compressor Rotor PERSONAL AUTiiOR(S) Friedrich Neuhoff '3 TPE inalOT 3 TIME COVERED [14 DATE OF REPORT (Year Mvonr.
Day) I~S PAGE COUNT 16 SUPPLEMIENTARY NO0TATION 17 I COSATi CODES Is8 S (I BJECT TiRMIS (Continue on reverse of necessary anda apnEty by blocic numnoer). The flow field of a new, transonic centrifugal compressor rotor with m/s tip speed has been analyzed experimentally and theoretically.
Due to the high speed, the flow was transonic in the rotor inlet region as well as at the diffuser inlet : Hartmut Krain, Bettina Hoffmann, Manfred Beversdorff.
Flow Field Unsteadiness in the Tip Region of a Transonic Compressor Rotor 1 March | Journal of Fluids Engineering, Vol.
No. 1 Unsteady Flow and Shock Motion in a Transonic Compressor RotorCited by: The paper describes the flow conditions found inside an optimized transonic, high pressure ratio centrifugal compressor rotor operated with a rotor tip speed of m/s. The rotor was coupled with a vaned diffuser of rather small exit radius ratio.
Measurements and calculations were carried out on. Effects of Inlet Distortion on the Flow Field in a Transonic Compressor Rotor J. Turbomach (April,) Detailed Flow Study of Mach Number High Transonic Flow With a Shock Wave in a Pressure Ratio 11 Centrifugal Compressor Impeller. Detailed laser anemometer measurements were acquired upstream, within, and downstream of a transonic, axial compressor rotor operating at design and off-design conditions.
The rotor was operated at85, 80, and 60 of design speed which provided inlet relative Mach numbers at the blade tip of,and respectively.
Effects of Inlet Distortion on the Flow Field in a Transonic Compressor Rotor J. Turbomach (April,) The Development of Fast Response Aerodynamic Probes for Flow. This is a transonic flow case for which experimental results are available in the literature to compare .At inflow Mach number of and angle of attack of degrees, the adapted meshes and Mach number contours for levels 2 and 3 are shown in Figure 3, along with the initial mesh increased sharpness of the shock regions is evident from the results.
Takado, J, Sonoda, T, & Nakamura, S. "The Effect of the Inlet Boundary Layer on the Tip Flow Field in a Transonic Fan Rotor." Proceedings of the ASME International Gas Turbine and Aeroengine Congress and Exhibition.
The S2 flow path design method of the transonic compressor is used to design the one stage fan in order to replace the original designed blade cascade which has two-stage transonic fan rotors.
In the modification design, the camber line is parameterized by a quartic polynomial curve and the thickness distribution of the blade profile is controlled by the double-thrice polynomial. A single-stage axial-flow compressor which incorporates rotor inlet counterswirl to improve stage performance is discussed.
Results for two rotor configurations are presented, including design and experimental test data. In this compressor design, inlet guide vanes were used to add counterswirl to the inlet of the rotor.
The effects of circumferential distortions in inlet total pressure on the flow field in a low-aspect-ratio, high-speed, high-pressure-ratio, transonic compressor rotor are investigated in this paper.
The flow field was studied experimentally and numerically with and without inlet total pressure distortion. Mixed flow compressor for small turbofan engine designed for a mass flow rate of kg/sec, total to a total pressure ratio of with inlet conditions of stagnation pressure and stagnation.
By imposing counter-swirl at the inlet of the rotor, the relative Mach number of the inlet can be reduced, the distribution along the span of the reaction can be fine-tuned and the work capacity could be further improved. The above design investigations offer an effective approach to further improve the performance of the LR transonic compressor.
NASA Rotor 37 is one of the most used transonic compressor test cases. Rotor 37 is a low aspect ratio inlet rotor for a core compressor. It was originally tested as a stage by Reid and Moore in and [1 and 2]. The rotor was re-tested at NASA Glenn by Suder et al.
in  and more recently by Shabbir et al. in . As shown, no modifications of the inlet flow field occurred after the optimization, but the flow field structure in the passage is clearly different. The negative curvature of the section upstream of the shock led to the reduction of the upstream relative Mach number from to Experimental study of the flow field within a transonic axial compressor rotor by laser velocimetry and comparison with Through-Flow calculations, J.
of Eng. For Power,Vol. No. The rig test confirmed the numerical assessment of the effects of aerodynamic sweep on the low-aspect-ratio, high-through-flow, transonic compressor rotor.
Detailed analyses of the measured and calculated flow fields indicate that two mechanisms are primarily responsible for the differences in aerodynamic performance among these rotors. This project prepares the structure and instrumentation to investigate the inlet distortion and effects of steam ingestion on a transonic axial compressor.
A compressor test facility, including mechanical equipment, data acquisition system, and remote digital control system, was configured to test a transonic compressor rotor, similar to what.
The flow in the blade tip vicinity of the transonic first stage of a multi-stage axial flow compressor with variable inlet guide vane (IGV) and casing treatment (CT) above the rotor is investigated experimentally and numerically with focus on the effects of the CT on flow structures and compressor performance.
The flow is investigated using classical pneumatic instrumentation and hot wire anemometry. The latter allows one to determine the average flow values as well as the instantaneous blade-to-blade flow field. These measurements were performed upstream and downstream of a low-speed axial compressor stage rotor.
A triangular inlet total pressure. Hotwire measurements of the inlet flow field were taken to determine an inlet turbulence intensity of % during both subsonic as well as transonic compressor operation. A 95% speed line was established from data taken from open throttle to near stall.
The S2 flow path design method of the transonic compressor is used to design the one stage fan in order to replace the original designed blade cascade which has two-stage transonic fan rotors. The shape optimization of blade sweep in a transonic axial compressor rotor of NASA Rotor 37 has been performed using response surface method and the three-dimensional Wavier-Stokes analysis.
Unsteady flow characteristics in a modern transonic axial compressor operating near stall are studied in detail. Measured data from high-response pressure probes show that the tip clearance vortex oscillates substantially near stall.
Instantaneous flow structure varies substantially among different blade passages even with uniform inlet flow. A typical value for the rotor inlet relative flow at the tip could be a Mach number around “” – “”, depending on design. Today’s high efficiency transonic axial-flow compressors can provide a total pressure ratio in the order of around “”, realized by combining high rotor speeds (tip speed in the order of around m.
Sanders, A. and Fleeter, S. Experimental Investigation of Rotor-Inlet Guide Vane Interactions in Transonic Axial-Flow Compressor.
AIAA Journal of Propulsion and. The influence of this inlet distortion on the flow field and the flutter stability of a highly loaded transonic fan rotor (NASA rotor 67) is investigated.
The static deflection of the manufactured blade was considered using an accurate mesh morphing algorithm to update the fan performance characteristic considering the deformed blade structure. The investigated compressor cascade called V– was designed for the hub section of a stator blade in a highly loaded axial compressor (Hoheisel ).The design conditions with an inlet Mach number of Ma 1 = and an inlet flow angle of β 1 = ° lead to a fully subsonic cascade flow.
The Reynolds number based on blade chord length of Re 1 = causes boundary layer. Flow analysis for NASA ro a transonic axial compressor rotor has been discussed in details in AGARD advisory report  and by Reid and Moore . As the test data are available in these references, many researchers have made efforts to validate their computational codes, and also to optimize the rotor.
1 ISABE INFLUENCE OF INLET GUIDE VANE WAKES ON PERFORMANCE AND STABILITY OF A TRANSONIC COMPRESSOR C. Biela, C. Brandstetter, F. Holzinger, er. Moore RD, Reid L. Performance of single-stage axial flow transonic compressor with rotor and stator aspect-ratios of andrespectively, and with a design pressure ratio ofNASA TPThe three dimensional unsteady flow field behind a transonic compressor rotor with a design pressure ratio of at a tip Mach number of has been resolved on the blade-passing time scale, using the M.I.T.
Blowdown Compressor Facility. Quantities determined were total and static pressures, tangential. Greitzer, “ Upstream attenuation and quasi-steady rotor lift fluctuations in asymmetric flows in axial flow compressors,” ASME,PaperGT noticed that the compressor behavior would have a strong impact on the upstream inlet flow distortion.
A 3D, single-passage steady flow Navier-Stokes solver was used to predict complete performance characteristics, including the numerical instability point, for both rotors. The predictions are generally in good agreement with the test data (characteristics, radial profiles and rotor over tip measurements) at all conditions modelled for rotor The flow field ahead, within, and behind the rotor of a transonic axial compressor designed for a total pressure ratio of at a relative tip Mach number of has been studied in detail using.
The purpose of this paper is to investigate the axial spacing between an IGV and a rotor, and effects of interface location in the spacing on the performance prediction of the axial compressor with the 3-dimensional fluid analysis.
A multi-stage transonic axial compressor with an inlet. In order to be representative of these differing flow regimes the results of a range of tests at varying tip-gaps and speeds from subsonic to transonic are presented.
A highly loaded transonic axial splittered rotor is used as the test article in this study. Three experiments with cold tip gaps of [mm], [mm] and [mm] are presented.
Performance and laser measurement results are presented for a transonic centrifugal compressor stage, equipped with a backswept rotor designed for m/s tip speed and a mean relative inlet tip Mach number of The flow field features of the rotor are analysed in detail with the help of laser measurements and the results obtained from a 3D.
Wartzek F., Holzinger F., Brandstetter C., Schiffer HP. () Realistic Inlet Distortion Patterns Interacting with a Transonic Compressor Stage. In: Radespiel R., Niehuis R., Kroll N., Behrends K. (eds) Advances in Simulation of Wing and Nacelle Stall.
FOR Notes on Numerical Fluid Mechanics and Multidisciplinary Design, vol. Table 1 Axial Flow Compressor Characteristics Type of Application Type of Flow Inlet Relative Velocity Mach Number Pressure Ratio per Stage Efﬁ ciency per Stage Industrial Subsonic 88%% Aerospace Transonic 80%% Research Supersonic 75%% The aerospace engines have been the leaders in.Which region is more sensitive to the nonuniform flow and how the mechanism of distorted flow influences the compressor flow field are topics of great importance and should be studied carefully.
In this paper, the exploration of how the inlet total pressure distortion influences the flow field of rotor and stator end walls will base on the. A detailed experimental investigation to understand and quantify the development of loss and blockage in the flow field of a transonic, axial flow compressor rotor has been undertaken.
Detailed laser anemometer measurements were acquired upstream, within, and downstream of a transonic, axial compressor rotor operating at design and off-design Author: Kenneth L.