scholarly journals Effects of Transient Heat Transfer on Compressor Stability

2018 ◽  
Vol 140 (12) ◽  
Author(s):  
A. Kiss ◽  
Z. Spakovszky
Keyword(s):  
Heat Transfer ◽  
Pressure Ratio ◽  
Strong Dependence ◽  
Transient Heat Transfer ◽  
Compressor Blades ◽  
Flow Angle ◽  
Stall Margin ◽  
Point Reduction ◽  
Angle Deviation ◽  
Blade Row

The effects of heat transfer between the compressor structure and the primary gas path flow on compressor stability are investigated during hot engine re-acceleration transients. A mean line analysis of an advanced, high-pressure ratio compressor is extended to include the effects of heat transfer on both stage matching and blade row flow angle deviation. A lumped capacitance model is used to compute the heat transfer of the compressor blades, hub, and casing to the primary gas path. The inputs to the compressor model with heat transfer are based on a combination of full engine data, compressor test rig measurements, and detailed heat transfer computations. Nonadiabatic transient calculations show a 8.0 point reduction in stall margin from the adiabatic case, with heat transfer predominantly altering the transient stall line. 3.4 points of the total stall margin reduction are attributed to the effect of heat transfer on blade row deviation, with the remainder attributed to stage rematching. Heat transfer increases loading in the front stages and destabilizes the front block. Sensitivity studies show a strong dependence of stall margin to heat transfer magnitude and flow angle deviation at low speed, due to the effects of compressibility. Computations for the same transient using current cycle models with bulk heat transfer effects only capture 1.2 points of the 8.0 point stall margin reduction. Based on this new capability, opportunities exist early in the design process to address potential stability issues due to transient heat transfer.

scholarly journals Effects of Transient Heat Transfer on Compressor Stability

2018 ◽  
Author(s):  
A. Kiss ◽  
Z. Spakovszky
Keyword(s):  
Heat Transfer ◽  
Pressure Ratio ◽  
Strong Dependence ◽  
Transient Heat Transfer ◽  
Compressor Blades ◽  
Flow Angle ◽  
Stall Inception ◽  
Stall Margin ◽  
Angle Deviation ◽  
Blade Row

The effects of heat transfer between the compressor structure and primary gas path flow on compressor stability are investigated during hot engine re-acceleration transients, or so called “Bodie” transients. A mean line analysis of an advanced, high-pressure ratio compressor is extended to include the effects of heat transfer on both stage matching and blade row flow angle deviation. A lumped capacitance model is used to compute the heat transfer of the compressor blades, hub, and casing to the primary gas path. The inputs to the compressor model with heat transfer are based on a combination of full engine data, compressor test rig measurements, and detailed heat transfer computations. Transient calculations with heat transfer show a 8.0 point reduction in stall margin from the adiabatic case, with heat transfer predominantly altering the transient stall line. 3.4 points of the total stall margin reduction are attributed to the effect of heat transfer on blade row deviation and the remainder is attributed to stage re-matching. It is found that heat transfer increases loading in the front stages and destabilizes the front block. Furthermore, the stage re-matching due to heat transfer alters the slope of the compressor characteristic and promotes modal-type stall inception. Sensitivity studies show a strong dependence of stall margin to heat transfer magnitude and flow angle deviation at low speed, due to the effects of compressibility. Computations for the same transient using current cycle models with bulk heat transfer effects, such as NPSS, only capture 1.2 points of the 8.0 point stall margin reduction. This implies that, using this new capability, opportunities exist early in the design process to address potential stability issues due to transient heat transfer.

Design and Test of a Novel Highly-Loaded Compressor

2019 ◽  
Author(s):  
Chengwu Yang ◽  
Ge Han ◽  
Shengfeng Zhao ◽  
Xingen Lu ◽  
Yanfeng Zhang ◽  
...  
Keyword(s):  
Pressure Ratio ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Flow Angle ◽  
Stall Margin ◽  
High Pressure Ratio ◽  
Numerical Predictions ◽  
Highly Loaded

Abstract The blades of rear stages in small size core compressors are reduced to shorter than 20 mm or even less due to overall high pressure ratio. The growing of tip clearance-to-blade height ratio of the rear stages enhance the leakage flow and increase the possibility of a strong clearance sensitivity, thus limiting the compressor efficiency and stability. A new concept of compressor, namely diffuser passage compressor (DP), for small size core compressors was introduced. The design aims at making the compressors robust to tip clearance leakage flow by reducing pressure difference between pressure and suction surfaces. To validate the concept, the second stage of a two-stage highly loaded axial compressor was designed with DP rotor according to a diffuser map. The diffuser passage stage has the same inlet condition and loading as the conventional compressor (CNV) stage, of which the work coefficient is around 0.37. The predicted performance and flow field of the DP were compared with the conventional axial compressor in detail. The rig testing was supplemented with the numerical predictions. Results reveal that the throttle characteristic of DP indicates higher pressure rise and the loss reduction in tip clearance is mainly responsible for the performance improvement. For the compressor with DP, the pressure and flow angle are more uniform on exit plane. What’s more, the rotor with diffused passage reveals more robust than the conventional rotor at double clearance gap. Furthermore, the experimental data indicate that DP presents higher pressure rise at design and part speeds. At design speed, the stall margin was extended by 7.25%. Moreover, peak adiabatic efficiency of DP is also higher than that of CNV by about 0.7%.

Impact of Surface Roughness on Compressor Cascade Performance

2010 ◽  
Vol 132 (6) ◽  
Author(s):  
Seung Chul Back ◽  
June Hyuk Sohn ◽  
Seung Jin Song
Keyword(s):  
Gas Turbines ◽  
Velocity Ratio ◽  
Compressor Blade ◽  
Compressor Cascade ◽  
Roughness Effects ◽  
Flow Angle ◽  
Angle Deviation ◽  
Blade Row ◽  
Rough Region ◽  
Linear Compressor Cascade

This paper presents an experimental investigation of roughness effects on aerodynamic performance in a low-speed linear compressor cascade. Equivalent sandgrain roughnesses of 12 μm, 180 μm, 300 μm, 425 μm, and 850 μm have been tested. In nondimensional terms, these roughnesses represent compressor blade roughnesses found in actual gas turbines. Downstream pressure, velocity, and angle have been measured with a five-hole probe at 0.3 chord downstream of the blade trailing edge. For the tested roughnesses of 180 μm, 300 μm, 425 μm, and 850 μm, the axial velocity ratio across the blade row decreases by 0.1%, 2.1%, 2.5%, and 5.4%, respectively. For the same cases, the exit flow angle deviation increases by 24%, 38%, 51%, and 70%, respectively. Finally, the mass-averaged total pressure loss increases by 12%, 44%, 132%, and 217%, respectively. Also, the loss increases more rapidly in the transitionally rough region. Thus, among the three parameters, the loss responds most sensitively to changes in compressor blade roughness.

scholarly journals CENTRIFUGAL COMPRESSOR EFFICIENCY CALCULATION WITH HEAT TRANSFER

2017 ◽  
Vol 18 (2) ◽  
pp. 225-237 ◽  
Author(s):  
Valeriu Dragan
Keyword(s):  
Heat Transfer ◽  
Centrifugal Compressor ◽  
Pressure Ratio ◽  
Adiabatic Wall ◽  
Flow Angle ◽  
Blade Cooling ◽  
Performance Variation ◽  
Measuring Efficiency ◽  
Compressor Power

In this paper we present a case study of apparent performance variation ofan optimized centrifugal compressor design when its metal parts are cold - before the conjugated heat transfer between the fluid and parts reaches an energetic equilibrium. The methods used are numerical, using full viscous 3D computational fluid dynamics with heat transfer. Three cases were considered, an adiabatic wall baseline, an all-blade cooling at 293 K and a more realistic stator row cooling at 293 K. Results indicate an apparent yet erroneous isentropic efficiency reading increase beyond 100% - which was to be expected due to the fluid cooling. However the isentropic and polytropic efficiencies could be estimated and were used to more accurately assess the performance of the compressor. Power consumption decreased to approximately 97% of the original load while the pressure ratio was marginally increased. This alone does not, however, explain the non-physical efficiency readings, which are mainly due to the assumptionsand manner under which the efficiency itself is calculated. The paper  presents a more robust approach to measuring efficiency, regardless of the heat transfer within the turbomachinery itself. Possible applications of the study may range from cold-start regime simulation to the optimization of inter-cooling setup or even flow angle control without mechanically actuated OGV

Transient heat-transfer analysis for sub-scale noz...

2005 ◽  
Author(s):  
Dr. Jae-Seok Yoo ◽  
Mr. Byung-Hun Kim ◽  
Dr. Young-Soon Jang ◽  
Dr. Yeong-Moo Yi
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Analysis ◽  
Transient Heat Transfer
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