scholarly journals Effect of radial inflow distortion on the performance of a highly loaded tandem stage

2021 ◽  
pp. 095765092110316
Author(s):  
Amit Kumar ◽  
AM Pradeep
Keyword(s):  
Pressure Ratio ◽  
Pressure Rise ◽  
Adverse Pressure Gradient ◽  
Significant Loss ◽  
The Novel ◽  
Weight Optimization ◽  
Turbine Engine ◽  
Stall Margin ◽  
Engine Size ◽  
Highly Loaded

Engine size and weight optimization have always been high-priority design objectives for designers. Compressors occupy a relatively large part of the gas turbine engine. Owing to the adverse pressure gradient in the compressor, achieving the required pressure ratio within fewer stages has been a challenging task for compressor designers. Tandem blading is one of the novel concepts, which could be used to increase the pressure ratio by means of higher flow turning through the blade passages. This paper presents the performance characteristics of a tandem stage based on results from experiments and numerical analyses. The investigation is further extended to analyze the effect of a radial hub and tip distortion on the performance of the tandem stage. The experimental results are very well supported with some interesting numerical results, particularly near the hub and tip region. It is observed that the tandem stage demonstrates higher pressure rise and stall margin under clean inflow. The tandem stage is also observed to be more sensitive to radial distortion leading to a significant loss in the total pressure and the stall margin.

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%.

scholarly journals The Development of Supersonic Axial Compressor Boost Stages for Small Gas Turbines

1969 ◽  
Author(s):  
C. H. Muller ◽  
A. Sabatiuk
Keyword(s):  
Gas Turbines ◽  
Static Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Performance Goals ◽  
Ample Evidence ◽  
Rotor Type ◽  
Highly Loaded ◽  
Percent Efficiency

The axial supersonic compressors of the “shock-in-rotor” type are under development for application to small gas turbines. A passage flow approach and passage criteria were used to design and develop the airfoils for the highly loaded rotor and stator blading of these 4 lb/sec machines. The overall stage performance values demonstrated to date are 2.06:1 pressure ratio at 78 percent adiabatic efficiency and 2.56:1 at 74.4 percent efficiency. The loss data and static pressure rise measured for the rotors and exit stators provide ample evidence that the higher design performance goals can be met.

Validation of a Modified Parallel Compressor Model for Prediction of the Effects of Inlet Swirl on Compressor Performance and Operability

2015 ◽  
Author(s):  
Reginald S. Floyd ◽  
Milton Davis
Keyword(s):  
Pressure Ratio ◽  
Pressure Rise ◽  
Engine Performance ◽  
The United States ◽  
Turbine Engine ◽  
Inlet Distortion ◽  
Inlet Swirl ◽  
Compressor Performance ◽  
Inlet Conditions ◽  
Map Data

Engine inlet distortion complications have plagued the turbine engine development community for decades, and engineers have developed countless methods to identify and combat the harmful effects of inlet distortion. One such type of distortion that has gained much attention in recent years is known as inlet swirl, which results in a significant flow angularity at the face of the engine. This flow angularity can affect the pressure rise and flow capacity of the fan or compressor, and subsequently affect compressor and engine performance. Previous modeling and simulation efforts to predict the effect inlet swirl can have on fan and compressor performance have made great strides, yet still leave a lot to be desired. In particular, a one-dimensional parallel compressor model called DYNTECC (Dynamic Turbine Engine Compressor Code) has been used to analyze the effects of inlet swirl on fan and performance operability of the Honeywell F109 turbofan engine. However, when compared to experimental swirl data gathered at the United States Air Force Academy (USAFA), the model predictions were found to be inaccurate. This paper documents work done to compare the initial predictions generated by DYNTECC to the latest set of experimental swirl data, analyze the potential shortcomings of the initial model, and modify the existing model to more accurately reflect test data. Extensive work was completed to create a methodology that can calibrate the model to existing clean inlet fan map data. In addition, an in depth study of fan/compressor stalling criteria was conducted, and the model was modified to use an alternate stalling criteria that more accurately predicted the point of stall for various swirl inlet conditions. The prediction of the fan stall pressure ratio for all inlet swirl conditions tested is within 2% of the ground test stall point at the same referred fan speed and referred mass flow.

Multi Stage Axial Compressor Design and Performance Evaluation

2011 ◽  
Author(s):  
Young Seok Kang ◽  
Tae Choon Park ◽  
Oh Sik Hwang ◽  
Soo Seok Yang
Keyword(s):  
Performance Test ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Multi Stage ◽  
Compressor Design ◽  
Test Result ◽  
Highly Loaded ◽  
Small Aircraft

Recently, needs for Unmanned Air Vehicle (UAV) and small aircraft are increasing and demands for small turbo jet or turbo fan engines are also increasing. Then, size and weight are the two main restrictions in UAV or small aircraft propulsion system applications. One method for resolving such a problem is to increase the pressure rise per stage and to reduce the number of stages. Nowadays, matured compressor aerodynamic design techniques enable us to design highly loaded axial compressors. This paper covers from the design step of a highly loaded transonic axial compressor to the performance test result and its analysis. At the fore part of the paper, aerodynamic process of a multi stage axial compressor is introduced. To satisfy both of the mass flow and pressure rise, the compressor should rotate at a high rotational speed. Therefore the transonic flow field forms in the rotor stages and it is designed with a relatively high pressure rise per stage to satisfy its design target. Basically, one dimensional and quasi three dimensional compressor design were carried with compressor design codes. The compressor stage consists of 3 stages, and the bulk pressure ratio is 2.5. The first stage is burdened with the highest pressure ratio and less pressure rises occur in the following stages. Also it is designed that tip Mach number of the first rotor row does not exceed 1.3. The final design was confirmed by iterating three dimensional CFD calculations to satisfy design target and some design intentions. In the latter part of the paper, its performance test processes are briefly introduced. The performance test result showed that the overall compressor performance targets; pressure ratio and efficiency are well achieved. From the test results, we found some clues for further improvement and optimization of the compressor aerodynamic performance.

Effect of Backward Sweep on Aerodynamic Performance of a 1.5-Stage Highly Loaded Axial Compressor

2020 ◽  
Author(s):  
Song Huang ◽  
Chuangxin Zhou ◽  
Chengwu Yang ◽  
Shengfeng Zhao ◽  
Mingyang Wang ◽  
...  
Keyword(s):  
Total Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Aerodynamic Performance ◽  
Blade Loading ◽  
Stall Margin ◽  
Axial Compressors ◽  
Design Speed ◽  
Total Pressure Ratio ◽  
Highly Loaded

Abstract As a degree of freedom in the three-dimensional blade design of axial compressors, the sweep technique significantly affects the aerodynamic performance of axial compressors. In this paper, the effects of backward sweep rotor configurations on the aerodynamic performance of a 1.5-stage highly loaded axial compressor at different rotational design speeds are studied by numerical simulation. The aim of this work is to improve understanding of the flow mechanism of backward sweep on the aerodynamic performance of a highly loaded axial compressor. A commercial CFD package is employed for flow simulations and analysis. The study found that at the design rotational speed, compared with baseline, backward sweep rotor configurations reduce the blade loading near the leading edge but slightly increases the blade loading near the trailing edge in the hub region. As the degree of backward sweep increases, the stall margin of the 1.5-stage axial compressor increase first and then decrease. Among different backward sweep rotor configurations, the 10% backward sweep rotor configuration has the highest stall margin, which is about 2.5% higher than that of baseline. This is due to the change of downstream stator incidence, which improves flow capacity near the hub region. At 80% rotational design speed, backward sweep rotor configurations improve stall margin and total pressure ratio of the compressor. It’s mainly due to the decreases of the rotor incidence near the middle span, which results in the decreases of separation on the suction surface. At 60% rotational design speed, detached shock disappears. Backward sweep rotor configurations deteriorate stall margin of the compressor, but increase total pressure ratio and adiabatic efficiency when the flow rate is lower than that at peak efficiency condition. Therefore, it’s necessary to consider the flow field structure of axial compressors at whole operating conditions in the design process and use the design freedom of sweep to improve the aerodynamic performance.

Design and Aerodynamic Analysis of a Highly Loaded Helium Compressor

2011 ◽  
Author(s):  
Tingfeng Ke ◽  
Qun Zheng
Keyword(s):  
Rotor Blade ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Flow Coefficient ◽  
Compressor Stage ◽  
Stall Margin ◽  
Air Compressor ◽  
Velocity Triangle ◽  
Large Flow ◽  
Highly Loaded

A design study of the multistage axial helium compressor of a 300MW nuclear gas turbine is presented in this paper. To design the helium compressor, we apply a new velocity triangle, which is different from that of the conventional air compressor. And the new velocity triangle is of large flow coefficient and negative pre-swirl etc. Comparing with the design based on conventional air compressor technique, the new helium compressor stage loading is increased by nearly three times when inlet flow coefficient increases from 0.6 to 2.04, thus stage numbers could decrease from 16 to 6. The higher loadings result in exacerbated corner separations. An increased reaction can relief the three dimensional separation and achieve the expectant stage pressure ratio. The ratio of the maximum thickness to the chord length of the rotor blade from hub to shroud is optimized and three dimensional profiling is applied. In such a way, the separation in the highly loaded helium compressor channel can be controlled and the optimum efficiency achieved. The performance investigation of the designed stage shows that the highly loaded design has acceptable stall margin.

The Effects of Tip Leakage Flow on the Performance of Multistage Compressors Used in Small Core Engine Applications

2015 ◽  
Vol 138 (5) ◽  
Author(s):  
Reid A. Berdanier ◽  
Nicole L. Key
Keyword(s):  
Experimental Data ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Tip Clearance ◽  
Tip Leakage Flow ◽  
Stall Margin ◽  
Tip Leakage ◽  
Small Core ◽  
Thermal Growth

Large rotor tip clearances and the associated tip leakage flows are known to have a significant effect on overall compressor performance. However, detailed experimental data reflecting these effects for a multistage compressor are limited in the open literature. As design trends lead to increased overall compressor pressure ratio for thermal efficiency benefits and increased bypass ratios for propulsive benefits, the rear stages of the high-pressure compressor will become physically small. Because rotor tip clearances cannot scale exactly with blade size due to the margin needed for thermal growth considerations, relatively large tip clearances will be a reality for these rear stages. Experimental data have been collected from a three-stage axial compressor to assess performance with three-tip clearance heights representative of current and future small core machines. Trends of overall pressure rise, stall margin, and efficiency are evaluated using clearance derivatives, and the summarized data presented here begin to narrow the margin of tip clearance sensitivities outlined by previous studies in an effort to inform future compressor designs. Furthermore, interstage measurements show stage matching changes and highlight specific differences in the performance of rotor 1 and stator 2 compared to other blade rows in the machine.

Testing Protruding Studs As a Form of Casing Treatment on a Transonic Turbofan: A Computational Study

2017 ◽  
Author(s):  
M. David Collao ◽  
Robert S. Webster ◽  
Kidambi Sreenivas
Keyword(s):  
Rotor Blade ◽  
Computational Study ◽  
Pressure Ratio ◽  
Pressure Rise ◽  
Slight Reduction ◽  
Stall Margin ◽  
Blade Passage ◽  
Casing Treatment ◽  
The Subject ◽  
Momentum Region

This paper presents the findings of an ongoing CFD study of using protruding studs as a form of casing treatment on a transonic turbofan stage. Simulations have been performed on the subject turbomachine with and without the casing treatment in order to validate computations with available experimental results and to compute any difference in performance. The results of the simulations with the casing treatment suggest that protruding studs have the potential to extend the stall margin of the turbofan while resulting in a slight reduction in pressure rise and efficiency. From the use of an initial configuration of studs, the computed increase in stall margin based on mass flow rate was 5.46%, and the greatest decrease in pressure ratio and adiabatic efficiency were 0.25% and 1.59%, respectively. Flowfield visualizations of simulations at computed near-stall conditions without casing treatment show regions of low momentum flow near the casing in the rotor blade passage, and low momentum regions near the hub in the stator section. Visualization from simulations with casing treatment at computed near-stall conditions show a large blockage imposed by the studs in the rotor blade passage, and a low momentum region near the casing in the stator section. Computed performance maps obtained from using other configurations of studs suggest that further increase in stall margin is possible at other levels of protrusion of the studs.

Design of a Highly Loaded Transonic Two-Stage Fan Using Swept and Bowed Blading

2011 ◽  
Author(s):  
Hailiang Jin ◽  
Donghai Jin ◽  
Fang Zhu ◽  
Ke Wan ◽  
Xingmin Gui
Keyword(s):  
Total Pressure ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Leading Edge ◽  
Navier Stokes ◽  
Two Stage ◽  
Stall Margin ◽  
Stator Vane ◽  
Total Pressure Ratio ◽  
Highly Loaded

This paper presents the design of a highly loaded transonic two-stage fan using several advanced three-dimensional blading techniques including forward sweep and “hub bending” in rotors and several bowed configurations in stators. The effects of these blading techniques on the performance of the highly loaded transonic two-stage fan were investigated on the basis of three-dimensional Navier-Stokes predictions. The results indicate that forward sweep has insignificant impact on the total pressure ratio and adiabatic efficiency of the fan. The throttling range of the fan is found to be improved by forward sweep because the shock in the forward swept rotor is expelled later upstream to the leading edge than that in the unswept one. Hub bending design technique increases the efficiency in the hub region of R1 due to the reduction of the low momentum zone in the hub region near the trailing edge. The stator vane design has a pronounced impact on the performance of the fan. The total pressure ratio, adiabatic efficiency, and stall margin of the schemes with the bowed vanes are increased significantly compared to the scheme with the straight vanes. The large corner stall in the straight S1 vane is reduced effectively by the bowed S1 vanes. Moreover, the strong corner stall in the straight S2 vane is fully eliminated by the bowed S2 vanes. Among the bowed vane schemes, the scheme with positive bowed (P. B.) hub and negative bowed (N. B.) tip vanes has the best efficiency and stall margin performances thanks to the superiority of the performance over the midspan regions of the bowed vanes.

Wave-Rotor Pressure-Gain Combustion Analysis for Power Generation and Gas Turbine Applications

2012 ◽  
Author(s):  
Manikanda Rajagopal ◽  
Abdullah Karimi ◽  
Razi Nalim
Keyword(s):  
Gas Turbine ◽  
Pressure Ratio ◽  
Turbulent Flame ◽  
Pressure Rise ◽  
Gas Turbine Engine ◽  
Operating Conditions ◽  
Specific Power ◽  
Inlet Temperature ◽  
Wave Rotor ◽  
Turbine Engine

A wave-rotor pressure-gain combustor (WRPGC) ideally provides constant-volume combustion and enables a gas turbine engine to operate on the Humphrey-Atkinson cycle. It exploits pressure (both compression and expansion) waves and confined propagating combustion to achieve pressure rise inside the combustor. This study first presents thermodynamic cycle analysis to illustrate the improvements of a gas turbine engine possible with a wave rotor combustor. Thereafter, non-steady reacting simulations are used to examine features and characteristics of a combustor rig that reproduces key features of a WRPGC. In the thermodynamic analysis, performance parameters such as thermal efficiency and specific power are estimated for different operating conditions (compressor pressure ratio and turbine inlet temperature). The performance of the WRPGC is compared with the conventional unrecuperated and recuperated engines that operates on the Brayton cycle. Fuel consumption may be reduced substantially with WRPGC introduction, while concomitantly boosting power. Simulations have been performed of the ignition of propane by a hot gas jet and subsequent turbulent flame propagation and shock-flame interaction.

Sign in / Sign up

Export Citation Format

Share Document