scholarly journals A Shock Loss Model for Supersonic Compressor Cascades

1997 ◽  
Author(s):  
Gregory S. Bloch ◽  
William W. Copenhaver ◽  
Walter F. O’Brien
Keyword(s):  
Mach Number ◽  
Boundary Layer Separation ◽  
Navier Stokes ◽  
Flow Angle ◽  
Loss Model ◽  
Viscous Profile ◽  
Loss Models ◽  
Profile Loss ◽  
Shock Loss ◽  
Good Agreement

Loss models used in compression system performance prediction codes are often developed from the study of two-dimensional cascades. In this paper, compressible fluid mechanics has been applied to the changes in shock geometry that are known to occur with back pressure for unstarted operation of supersonic compressor cascades. This physics-based engineering shock loss model is applicable to cascades with arbitrary airfoil shapes. Predictions from the present method have been compared to measurements and Navier-Stokes analyses of the L030-4 and L030-6 cascades, and very good agreement was demonstrated for unstarted operation. A clear improvement has been demonstrated over previously published shock loss models for unstarted operation, both in the accuracy of the predictions and in the range of applicability. The dramatic increase in overall loss with increasing inlet flow angle is shown to be primarily the result of increased shock loss, and much of this increase is caused by the detached bow shock. For a given Mach number, the viscous profile loss is nearly constant over the entire unstarted operating range of the cascade, unless a shock-induced boundary layer separation occurs near stall. Shock loss is much more sensitive to inlet Mach number than is viscous profile loss.

A Shock Loss Model for Supersonic Compressor Cascades

1999 ◽  
Vol 121 (1) ◽  
pp. 28-35 ◽  
Author(s):  
G. S. Bloch ◽  
W. W. Copenhaver ◽  
W. F. O’Brien
Keyword(s):  
Mach Number ◽  
Boundary Layer Separation ◽  
Navier Stokes ◽  
Flow Angle ◽  
Loss Model ◽  
Viscous Profile ◽  
Loss Models ◽  
Profile Loss ◽  
Shock Loss ◽  
Good Agreement

Loss models used in compression system performance prediction codes are often developed from the study of two-dimensional cascades. In this paper, compressible fluid mechanics has been applied to the changes in shock geometry that are known to occur with back pressure for unstarted operation of supersonic compressor cascades. This physics-based engineering shock loss model is applicable to cascades with arbitrary airfoil shapes. Predictions from the present method have been compared to measurements and Navier–Stokes analyses of the LO30-4 and L030-6 cascades, and very good agreement was demonstrated for unstarted operation. A clear improvement has been demonstrated over previously published shock loss models for unstarted operation, both in the accuracy of the predictions and in the range of applicability. The dramatic increase in overall loss with increasing inlet flow angle is shown to be primarily the result of increased shock loss, and much of this increase is caused by the detached bow shock. For a given Mach number, the viscous profile loss is nearly constant over the entire unstarted operating range of the cascade, unless a shock-induced boundary layer separation occurs near stall. Shock loss is much more sensitive to inlet Mach number than is viscous profile loss.

Loss Generation in Transonic Turbine Blading

2017 ◽  
Author(s):  
Penghao Duan ◽  
Choon S. Tan ◽  
Andrew Scribner ◽  
Anthony Malandra
Keyword(s):  
Experimental Data ◽  
Boundary Layer ◽  
Mach Number ◽  
High Speed ◽  
Turbine Blades ◽  
Surface Boundary ◽  
Loss Model ◽  
Exit Mach Number ◽  
Loss Characteristic ◽  
Shock Loss

The measured loss characteristic in a high-speed cascade tunnel of two turbine blades of different designs showed distinctly different trend with exit Mach number ranging from 0.8 to 1.4. Assessments using steady RANS computation of the flow in the two turbine blades, complemented with control volume analyses and loss modelling, elucidate why the measured loss characteristic looks the way it is. The loss model categorizes the total loss in terms of boundary layer loss, trailing edge loss and shock loss; it yields results in good agreement with the experimental data as well as steady RANS computed results. Thus RANS is an adequate tool for determining the loss variations with exit isentropic Mach number and the loss model serves as an effective tool to interpret both the computational and experimental data. The measured loss plateau in Blade 1 for exit Mach number of 1 to 1.4 is due to a balance between a decrease of blade surface boundary layer loss and an increase in the attendant shock loss with Mach number; this plateau is absent in Blade 2 due to a greater rate in shock loss increase than the corresponding decrease in boundary layer loss. For exit Mach number from 0.85 to 1, the higher loss associated with shock system in Blade 1 is due to the larger divergent angle downstream of the throat than that in Blade 2. However when exit Mach number is between 1.00 and 1.30, Blade 2 has higher shock loss. For exit Mach number above around 1.4, the shock loss for the two blades is similar as the flow downstream of the throat is completely supersonic. In the transonic to supersonic flow regime, the turbine design can be tailored to yield a shock pattern the loss of which can be mitigated in near equal amount of that from the boundary layer with increasing exit Mach number, hence yielding a loss plateau in transonic-supersonic regime.

Transonic Compressor Rotor Cascade With Boundary-Layer Separation: Experimental and Theoretical Results

1993 ◽  
Author(s):  
R. Fuchs ◽  
W. Steinert ◽  
H. Starken
Keyword(s):  
Boundary Layer ◽  
Mach Number ◽  
Computational Design ◽  
Boundary Layer Separation ◽  
Ratio Test ◽  
Integral Boundary ◽  
Flow Angle ◽  
Laminar Separation ◽  
Transonic Compressor ◽  
Compressor Rotor

A transonic compressor rotor cascade designed for an inlet Mach number of 1.09 and 14 degrees of flow turning has been redesigned for higher loading by an increased pitch-to-chord ratio. Test results, showing the influence of inlet Mach number and flow angle on cascade performance are presented and compared to data of the basic design. Loss-levels of both, the original and the redesigned higher loaded blade were identical at design condition, but the new design achieved even lower losses at lower inlet Mach numbers. The computational design and analysis has been performed by a fast inviscid time-dependent code coupled to a viscous direct/inverse integral boundary-layer code. Good agreement was achieved between measured and predicted surface Mach number distributions as well as exit-flow angles. A boundary-layer visualization method has been used to detect laminar separation bubbles and turbulent separation regions. Quantitative results of measured bubble positions are presented and compared to calculated results.

New Supersonic Loss Model for the Preliminary Design of Transonic Turbine Blades and the Influence of Pitch

2020 ◽  
Vol 142 (4) ◽  
Author(s):  
Luis Teia
Keyword(s):  
Mach Number ◽  
Turbine Blades ◽  
Turbine Rotor ◽  
Published Data ◽  
Efficient Design ◽  
Operational Parameters ◽  
Loss Model ◽  
Shock System ◽  
Transonic Turbine ◽  
Shock Loss

Abstract In order to produce a more efficient design of a compact turbine driving a cryogenic engine turbo-pump for a satellite delivering rocket, a new supersonic loss model is proposed. The new model was constructed based on high-quality published data, composed of Schlieren photographs and experimental measurements, that combined provided a unique insight into the mechanisms driving supersonic losses. Using this as a cornerstone, model equations were formulated that predict the critical Mach number and shock loss and shock-induced mixing loss as functions of geometrical (i.e., blade outlet and uncovered turning angle and trailing edge thickness) and operational parameters (i.e., exit Mach number). A series of highly resolved CFD numerical simulations were conducted on an in-house designed state-of-the-art transonic turbine rotor row (around unity aspect ratio (AR)) to better understand changes in the shock system for varying parameters. The main outcome showed that pitch to chord ratio has a powerful impact on the shock system, and thus on the manner by which shock loss and shock-induced mixing loss is distributed to compose the overall supersonic losses. The numerical loss estimates for two pitch to chord ratios—t⁄c = 0.70 and t⁄c = 0.98—were compared with absolute loss data of a previously published similar blade with satisfactory agreement. Calibrated equations are provided to allow hands-on integration into existing overall turbine loss models, where supersonic losses play a key role, for further enhancement of preliminary turbine design.

Evaluation of Unsteady CFD Methods by Their Application to a Transonic Propfan Stage

2001 ◽  
Author(s):  
S. Schmitt ◽  
F. Eulitz ◽  
L. Wallscheid ◽  
A. Arnone ◽  
M. Marconcini
Keyword(s):  
Numerical Methods ◽  
Navier Stokes ◽  
Test Case ◽  
Flow Angle ◽  
Laser Measurements ◽  
General Flow ◽  
Blade Row ◽  
Flow Features ◽  
Blade Row Interaction ◽  
Good Agreement

The accuracy in predicting the unsteady aerodynamic blade-row-interaction of two state-of-the-art Navier-Stokes codes is evaluated within the current paper. The general flow features of the test case — a transonic research propfan stage — are described in brief as far as necessary to understand the detailed comparisons. The calculated unsteady velocity and flow angle distributions at various axial planes of the stage are compared to data from unsteady laser measurements. The general flow features of the propfan are very well reproduced by the numerical methods and a good agreement is also obtained in comparison to the measured data. One important outcome of the comparison is the good agreement of both numerical methods with the unsteady fluctuations measured in the experiment.

An Iterative Method for Blade Profile Loss Model Adaptation Using Streamline Curvature

2007 ◽  
Author(s):  
Vassilios Pachidis ◽  
Pericles Pilidis ◽  
Ioannis Templalexis ◽  
Luca Marinai
Keyword(s):  
Experimental Data ◽  
Axial Flow ◽  
Public Domain ◽  
Performance Comparison ◽  
Model Adaptation ◽  
Blade Profile ◽  
Loss Model ◽  
Streamline Curvature ◽  
Loss Models ◽  
Profile Loss

The various incidence, deviation and loss models used in through-flow analysis methods, such as Streamline Curvature, are nothing more than statistical curve fits. A closer look at public domain data reveals that these statistical correlations and curve fits are usually based on experimental cascade data that actually display a fairly large scatter, resulting in a relatively high degree of uncertainty. This usually leads to substantial differences between the calculated and actual performances of a given gas turbine engine component. Typically, matching calculated results from a throughflow analysis against experimental data requires the combination of various correlations available in the public domain, through a very tedious, complex and time consuming ‘trial and error’ process. This particular study supports the view that it might actually be much more time-effective to “adopt” a given loss model against experimental data through an iterative, physics-based approach, rather than try to identify the best combination of available correlations. For example, the well-established “Swan’s model” for calculating the blade profile loss factor in subsonic and transonic axial flow compressors depends strongly on approximate correlations for calculating the blade wake momentum thickness, and therefore represents such a case. This study demonstrates this by looking into an iterative approach to blade profile loss model adaptation that can provide a relatively simple and quick, but also physics-based way of ‘calibrating’ profile loss models against available experimental data for subsonic applications. This paper presents in detail all the analysis necessary to support the above concept and discusses Swan’s model in particular as an example. Finally, the paper discusses the performance comparison of a two-dimensional, Streamline Curvature compressor model against experimental data before and after the adaptation of that particular loss model.

New Profile Loss Model for Improved Prediction of Transonic Axial Flow Compressor Performance in Choking Region

2015 ◽  
Author(s):  
Sangjo Kim ◽  
Donghyun Kim ◽  
Kuisoon Kim ◽  
Changmin Son ◽  
Myungho Kim ◽  
...  
Keyword(s):  
Performance Prediction ◽  
Incidence Angle ◽  
Axial Flow ◽  
Operating Conditions ◽  
Loss Model ◽  
Axial Flow Compressor ◽  
Compressor Performance ◽  
Loss Models ◽  
Flow Compressor ◽  
Profile Loss

New off-design profile loss models have been developed by performing thorough investigations on compressor performance prediction using one-dimensional stage-stacking approach and three-dimensional computational flow dynamics (CFD) results. Generally, a loss model incorporating various compressor geometry and operating conditions is required to predict the performance of various types of compressors. In this study, three sets of selected loss models were applied to predict axial flow compressor performance using stage-stacking approach. The results were compared with experimental data as well as CFD results. The comparison shows an interesting observation in choking region where the existing loss models cannot capture the rapid decrease in pressure and efficiency while CFD predicted the characteristics. Therefore, an improved off-design profile loss model is proposed for better compressor performance prediction in choking region. The improved model was derived from the correlation between the normalized total loss and the incidence angle. The choking incidence angle, which is a major factor in determining the off-design profile loss, was derived from correlations between the inlet Mach number, throat width-to-inlet spacing ratio, and minimum loss incidence angle. The revised stage-stacking program employing new profile loss model together with a set of loss models was applied to predict a single and multistage compressors for comparison. The results confirmed that the new profile loss model can be widely used for predicting the performance of single and multistage compressor.

An Iterative Method for Blade Profile Loss Model Adaptation Using Streamline Curvature

2007 ◽  
Vol 130 (1) ◽  
Author(s):  
Vassilios Pachidis ◽  
Pericles Pilidis ◽  
Ioannis Templalexis ◽  
Luca Marinai
Keyword(s):  
Experimental Data ◽  
Flow Analysis ◽  
Public Domain ◽  
Model Adaptation ◽  
Blade Profile ◽  
Loss Model ◽  
Streamline Curvature ◽  
Through Flow ◽  
Loss Models ◽  
Profile Loss

The various incidence, deviation, and loss models used in through-flow analysis methods, such as streamline curvature, are nothing more than statistical curve fits. A closer look at public domain data reveals that these statistical correlations and curve fits are usually based on experimental cascade data that actually display a fairly large scatter, resulting in a relatively high degree of uncertainty. This usually leads to substantial differences between the calculated and actual performances of a given gas turbine engine component. Typically, matching calculated results from a through-flow analysis against experimental data requires the combination of various correlations available in the public domain, through a very tedious, complex, and time consuming “trial and error” process. This particular study supports the view that it might actually be much more time effective to “adopt” a given loss model against experimental data through an iterative, physics-based approach, rather than try to identify the best combination of available correlations. For example, the well-established “Swan’s model” for calculating the blade profile loss factor in subsonic and transonic axial flow compressors depends strongly on approximate correlations for calculating the blade wake momentum thickness, and therefore represents such a case. This study demonstrates this by looking into an iterative approach to blade profile loss model adaptation that can provide a relatively simple and quick, but also physics-based way of “calibrating” profile loss models against available experimental data for subsonic applications. This paper presents in detail all the analysis necessary to support the above concept and discusses Swan’s model in particular as an example. Finally, the paper discusses the performance comparison of a two-dimensional, streamline curvature compressor model against experimental data before and after the adaptation of that particular loss model. This analysis proves the potential of the simulation strategy presented in this paper to “adopt” a given loss model against experimental data through an iterative, physics-based approach.

scholarly journals A Study on Axial Skewing Effect of a Turbine Stator by 3-D Navier-Stokes Analysis

1996 ◽  
Author(s):  
Naixing Chen ◽  
Qian Zhou ◽  
Weiguang Huang
Keyword(s):  
Leading Edge ◽  
Boundary Layer Separation ◽  
Engineering Thermophysics ◽  
Aerodynamic Characteristics ◽  
Radial Force ◽  
Rotor Blades ◽  
Navier Stokes ◽  
Flow Angle ◽  
Turbine Stator ◽  
Stator Vane

The objective of the present paper is to investigate the axial skewing effect of a turbine stator blading on its aerodynamic characteristics systematically, to find the possibility for reducing the secondary losses by using axially skewed bladings and to understand better its flow physics in turbomachinery. In the present paper a typical turbine stator blading is applied as an example to generate a set of axially skewed bladings for systematical study and illustration of their differences of aerodynamics characteristics. The paper gives the procedure for generating different forward-skewed and backward-skewed blades with the same profile sections at the same radius and the numerical method used is also described briefly. The method is based on the 3-D time-marching finite volume Navier-Stokes solution and was developed by the Institute of Engineering Thermophysics, Chinese Academy of Sciences. The turbulence model proposed by Baldwin and Lomax is used here for predicting the effective viscosity. The calculated results and their comparisons are also given in the present paper. On the basis of the analysis it is shown that the appropriate use of skewed blades gives designers another possibility to control the flow in the blade channel. By adopting forward-skewed blades to replace the straight blades can be reduced the blade loading near the leading edge and in the central part of the span. It is also found that the pressure gradient at the endwall of forward-skewed blading yields the radial force that enables to avoid the boundary layer separation from the endwall. The axial skew of blade enables to restrain the strength of the secondary flow. Therefore, the total pressure loss can be reduced. An attention should be paid: if the skewed blade for stator is chosen to be used, the radial distribution of the outlet flow angle from stator vane is required to meet the optimal incidence satisfaction to the rotor blades. Otherwise, it results in the reduction of efficiency.

Investigation of Scramjet Inlet Flow Using a Flux-Splitting Solver

2010 ◽  
Author(s):  
D. M. Holian ◽  
R. R. Mankbadi
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Boundary Layer Separation ◽  
Navier Stokes ◽  
Shock Interactions ◽  
Flux Splitting ◽  
Operation Conditions ◽  
Inlet Geometry ◽  
Proper Operation ◽  
Good Agreement

A detailed analysis is carried out on a rectangular scramjet inlet to analyze the flow field. The focus is on examining boundary layer separation and shock interactions to ensure proper operation of the inlet. We developed herein a flux-splitting Navier Stokes solver to be used for optimizing the inlet geometry and operation conditions. The results seem to be in good agreement with that of FLUENT CFD software and explain the experimental results of Haberle (2008).

Sign in / Sign up

Export Citation Format

Share Document