Scaling of Incidence Variations With Inlet Distortion for a Transonic Axial Compressor

2020 ◽  
Vol 142 (2) ◽  
Author(s):  
D. J. Hill ◽  
J. J. Defoe
Keyword(s):  
Total Pressure ◽  
Incidence Angle ◽  
Computational Cost ◽  
Force Model ◽  
Flow Angle ◽  
Total Enthalpy ◽  
Transonic Compressor ◽  
Inlet Distortion ◽  
Scaling Relationships ◽  
The Impact

Abstract This paper numerically explores the manner in which blade row inlet incidence variation scales with various distortion patterns and intensities. The objectives are to (1) identify the most appropriate parameter whose circumferential variation can be used to assess scaling relationships of a transonic compressor and (2) use this parameter to evaluate two types of non-uniform inflow patterns, vertically stratified total pressure distortions and radially stratified total enthalpy and total pressure distortions. A body force model of the blade rows is employed to reduce computational cost; the approach has been shown to capture distortion transfer and to be able to predict upstream flow redistribution with inlet distortion. Diffusion factor is shown to be an inadequate proxy for streamline loss generation in non-uniform flow, leading to the choice of incidence angle variations as the metric for which we assess scaling relationships. Posteriori scaling of circumferential flow angle variation based on the maximum incidence excursion for varying distortion intensity yields an accurate method for prediction of the impact for other distortion intensities; linear regression of the maximum variation in incidence around the annulus as a function of distortion intensity had R2 > 0.97 for all spanwise locations examined in both the rotor and stator for both vertically and radially stratified distortions. However, changes to far upstream distortion shape yield highly non-linear incidence variation scaling; the results suggest that the pitchwise gradients of far upstream total pressure govern the degree of linearity for incidence variation scaling.

scholarly journals Innovations in Body Force Modeling of Transonic Compressor Blade Rows

2018 ◽  
Vol 2018 ◽  
pp. 1-12 ◽  
Author(s):  
David J. Hill ◽  
Jeffrey J. Defoe
Keyword(s):  
Total Pressure ◽  
Body Force ◽  
Computational Cost ◽  
Pressure Rise ◽  
The Body ◽  
Force Model ◽  
Axial Fan ◽  
Compressor Blades ◽  
Time Resolved ◽  
Transonic Compressor

Aeroengine fans and compressors increasingly operate subject to inlet distortion in the transonic flow regime. In this paper, innovations to low-order numerical modeling of fans and compressors via volumetric source terms (body forces) are presented. The approach builds upon past work to accommodate any axial fan/compressor geometry and ensures accurate work input and efficiency prediction across a range of flow coefficients. In particular, the efficiency drop-off near choke is captured. The model for a particular blade row is calibrated using data from single-passage bladed computations. Compared to full-wheel unsteady computations which include the fan/compressor blades, the source term model approach can reduce computational cost by at least two orders of magnitude through a combination of reducing grid resolution and, critically, eliminating the need for a time-resolved approach. The approach is applied to NASA stage 67. For uniform flow, at 90% corrected speed and peak-efficiency, the body force model is able to predict the total-to-total pressure rise coefficient of the stage to within 1.43% and the isentropic efficiency to within 0.03%. With a 120∘ sector of reduced inlet total pressure, distortion transfer through the machine is well-captured and the associated efficiency penalty predicted with less than 2.7% error.

Assessing the Impact of the Inlet Total Pressure Distortion on the Turbofan Thrust

2018 ◽  
pp. 19-30
Author(s):  
Yu. A. Ezrokhi ◽  
E. A. Khoreva
Keyword(s):  
Total Pressure ◽  
External Compression ◽  
Average Value ◽  
Inlet Distortion ◽  
Air Inlet ◽  
Engine Thrust ◽  
Relative Value ◽  
Main Effect ◽  
Pressure Recovery Factor ◽  
The Impact

The paper considers techniques to develop a mathematical model using a method of «parallel compressors». The model is intended to estimate the impact of the air inlet distortion on the primary parameters of the aero-engine.  The paper presents rated estimation results in the context of twin spool turbofan design for two typical cruiser modes of flight of the supersonic passenger jet. In estimation the base values σbase and the average values of the inlet ram recovery σave remained invariable. Thus, parametrical calculations were performed for each chosen relative value of the area of low-pressure region.The paper shows that an impact degree of the inlet distortion on the engine thrust for two modes under consideration is essentially different. In other words, if in the subsonic mode the impact assessment can be confined only to taking into account the influence of decreasing average values of the inlet total pressure, the use of such an assumption in the supersonic cruiser mode may result in considerable errors.With invariable values of the pressure recovery factor at the engine intake, which correspond to the speed of flight for a typical air inlet of external compression σbase, and average value σave, a parameter Δσuneven  has the main effect on the engine thrust, and degree of this effect essentially depends on a difference between σave and σbase values.

Effects of Inlet Distortion on the Flow Field in a Transonic Compressor Rotor

1996 ◽  
Author(s):  
Chunill Hah ◽  
Douglas C. Rabe ◽  
Thomas J. Sullivan ◽  
Aspi R. Wadia
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Viscous Flow ◽  
Total Pressure ◽  
Stokes Equations ◽  
Three Dimensional ◽  
Aerodynamic Loss ◽  
Transonic Compressor ◽  
Inlet Distortion ◽  
Compressor 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. Total pressure distortion was created by screens mounted upstream from the rotor inlet. Circumferential distortions of 8 periods per revolution were investigated at two different rotor speeds. The unsteady blade surface pressures were measured with miniature pressure transducers mounted in the blade. The flow fields with and without inlet total pressure distortion were analyzed numerically by solving steady and unsteady forms of the Reynolds-averaged Navier-Stokes equations. Steady three-dimensional viscous flow calculations were performed for the flow without inlet distortion while unsteady three-dimensional viscous flow calculations were used for the flow with inlet distortion. For the time-accurate calculation, circumferential and radial variations of the inlet total pressure were used as a time-dependent inflow boundary condition. A second-order implicit scheme was used for the time integration. The experimental measurements and the numerical analysis are highly complementary for this study because of the extreme complexity of the flow field. The current investigation shows that inlet flow distortions travel through the rotor blade passage and are convected into the following stator. At a high rotor speed where the flow is transonic, the passage shock was found to oscillate by as much as 20% of the blade chord, and very strong interactions between the unsteady passage shock and the blade boundary layer were observed. This interaction increases the effective blockage of the passage, resulting in an increased aerodynamic loss and a reduced stall margin. The strong interaction between the passage shock and the blade boundary layer increases the peak aerodynamic loss by about one percent.

scholarly journals Numerical and Experimental Investigations on Optimized 3D Compressor Airfoils

2018 ◽  
Author(s):  
Jan Mihalyovics ◽  
Christian Brück ◽  
Dieter Peitsch ◽  
Ilias Vasilopoulos ◽  
Marcus Meyer
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Experimental Studies ◽  
Total Pressure Loss ◽  
Design Parameters ◽  
Flow Conditions ◽  
Flow Angle ◽  
Flow Phenomena ◽  
The Impact ◽  
Stator Design

The objective of the presented work is to perform numerical and experimental studies on compressor stators. This paper presents the modification of a baseline stator design using numerical optimization resulting in a new 3D stator. The Rolls Royce in-house compressible flow solver HYDRA was employed to predict the 3D flow, solving the steady RANS equations with the Spalart-Allmaras turbulence model, and its corresponding discrete adjoint solver. The performance gradients with respect to the input design parameters were used to optimize the stator blade with respect to the total pressure loss over a prescribed incidence range, while additionally minimizing the flow deviation from the axial direction at the stator exit. Non-uniform profile boundary conditions, being derived from the experimental measurements, have been defined at the inlet of the CFD domain. The presented results show a remarkable decrease in the axial exit flow angle deviation and a minor decrease in the total pressure loss. Experiments were conducted on two compressor blade sets investigating the three-dimensional flow in an annular compressor stator cascade. Comparing the baseline flow of the 42° turning stator shows that the optimized stator design minimizes the secondary flow phenomena. The experimental investigation discusses the impact of steady flow conditions on each stator design while focusing on the comparison of the 3D optimized design to the baseline case. The flow conditions were investigated using five-hole probe pressure measurements in the wake of the blades. Furthermore, oil-flow visualization was applied to characterize flow phenomena. These experimental results are compared with the CFD calculations.

scholarly journals Flow Measurements Behind the Inlet Guide Vane of a Centrifugal Compressor

1998 ◽  
Author(s):  
I. Kassens ◽  
M. Rautenberg
Keyword(s):  
Centrifugal Compressor ◽  
Total Pressure ◽  
Guide Vane ◽  
Incidence Angle ◽  
Pressure Ratio ◽  
Inlet Guide Vane ◽  
Flow Measurements ◽  
Flow Angle ◽  
Partial Load ◽  
Measurement Location

In a centrifugal compressor adjustable inlet guide vanes (IGV) in front of the impeller are used to regulate the pressure ratio and the mass flow. The stationary measurement of the velocity profile in front of the impeller with different angles of the IGV displays shock losses at the inlet edge of blade of the impeller. In the partial-load region (e.g. partial-load efficiency) the radial distribution of the flow influences considerably the performance of the impeller. The tested compressor consists of an adjustable IGV with straight vanes, a shrouded impeller and a vaneless, parallel diffuser. In the first measurement location, behind the IGV, total pressure, static pressure and flow angle were measured with a 5-hole cylinder probe. In the second measurement location, in front of the impeller, the measurement of the total pressure was carried out with a Kiel probe and the flow angle with a Cobra probe accordingly the static wall pressure was measured. Taking into consideration the fundamental thermodynamical equations it was possible to determine the velocity profiles because of the measured distributions of the flow angle in these two measurement locations. For different angles of the IGV and with various mass flows the distributions of the deflection defect behind the IGV are described. Starting with the measured distributions of the flow in front of the impeller the flow angles at the impeller inlet are calculated and the distributions of the incidence angle at the impeller inlet are figured out.

Vaneless Diffuser Advanced Model

2008 ◽  
Vol 130 (1) ◽  
Author(s):  
Oleg Dubitsky ◽  
David Japikse
Keyword(s):  
Total Pressure ◽  
Static Pressure ◽  
Zone Model ◽  
Vaneless Diffuser ◽  
Flow Angle ◽  
Average Flow ◽  
New Model ◽  
Future Design ◽  
The Impact ◽  
Advanced Model

A new vaneless diffuser model is presented. Upon thorough examination of the change in average total pressure, the average static pressure and the average flow angle through a vaneless diffuser, it was discovered that existing models fail to provide useful integrity. Consequently, a new model was built. It was learned that it was necessary to use a two-zone model of the flow entering the vaneless diffuser and to carefully model the two-zone degradation as the flow passes through the vaneless diffuser. The new model is presented with detailed testing. The impact upon future design is outlined, and the expectation is established that various future designs will require the integrity of the new model; old models can be used in limited cases with care.

Numerical Simulation of Low Engine Order Excitation in a Transonic Compressor

2012 ◽  
Vol 546-547 ◽  
pp. 206-211
Author(s):  
Bo Zhang ◽  
Yang Wei Liu ◽  
Bao Jie Liu
Keyword(s):  
Total Pressure ◽  
Static Pressure ◽  
Tip Clearance ◽  
Blade Surface ◽  
Aerodynamic Forces ◽  
Transonic Compressor ◽  
Inlet Distortion ◽  
Unsteady Aerodynamic ◽  
Static Pressure Distribution ◽  
Engine Order

A full-annulus unsteady quasi-3D simulation for a transonic compressor was conducted to investigate the low engine order excitation sources. A special inlet total pressure distortion was used in the simulation to account for the effect of non-axisymmetric tip clearance. The unsteady aerodynamic forces were obtained by integrating the static pressure along the blade surface. Furthermore, the effect of downstream rotor detached shock on upstream stator was analyzed. It was shown that the unsteady aerodynamic excitation of the rotor was subjected to the first two harmonic frequencies of the inlet distortion, and the adjacent stators’ basic frequency and lower harmonics. It was found that the detached shock had a significant impact on the surface static pressure distribution of the upstream stators.

Effects of Inlet Distortion on the Flow Field in a Transonic Compressor Rotor

1998 ◽  
Vol 120 (2) ◽  
pp. 233-246 ◽  
Author(s):  
C. Hah ◽  
D. C. Rabe ◽  
T. J. Sullivan ◽  
A. R. Wadia
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Viscous Flow ◽  
Total Pressure ◽  
Stokes Equations ◽  
Three Dimensional ◽  
Aerodynamic Loss ◽  
Transonic Compressor ◽  
Inlet Distortion ◽  
Compressor 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. Total pressure distortion was created by screens mounted upstream from the rotor inlet. Circumferential distortions of eight periods per revolution were investigated at two different rotor speeds. The unsteady blade surface pressures were measured with miniature pressure transducers mounted in the blade. The flow fields with and without inlet total pressure distortion were analyzed numerically by solving steady and unsteady forms of the Reynolds-averaged Navier–Stokes equations. Steady three-dimensional viscous flow calculations were performed for the flow without inlet distortion while unsteady three-dimensional viscous flow calculations were used for the flow with inlet distortion. For the time-accurate calculation, circumferential and radial variations of the inlet total pressure were used as a time-dependent inflow boundary condition. A second-order implicit scheme was used for the time integration. The experimental measurements and the numerical analysis are highly complementary for this study because of the extreme complexity of the flow field. The current investigation shows that inlet flow distortions travel through the rotor blade passage and are convected into the following stator. At a high rotor speed where the flow is transonic, the passage shock was found to oscillate by as much as 20 percent of the blade chord, and very strong interactions between the unsteady passage shock and the blade boundary layer were observed. This interaction increases the effective blockage of the passage, resulting in an increased aerodynamic loss and a reduced stall margin. The strong interaction between the passage shock and the blade boundary layer increases the peak aerodynamic loss by about one percent.

Performance of a Subsonic Compressor Airfoil With Upstream, End-Wall Injection Flow

2021 ◽  
Author(s):  
Aaron J. Pope ◽  
Andrew Oliva ◽  
Aleksandar Jemcov ◽  
Scott C. Morris ◽  
Mark Stephens ◽  
...  
Keyword(s):  
Total Pressure ◽  
High Speed ◽  
Pressure Field ◽  
Incidence Angle ◽  
Suction Side ◽  
Secondary Flows ◽  
Injection Flow ◽  
Wall Injection ◽  
End Wall ◽  
The Impact

Abstract The performance of a compressor stator airfoil with end-wall injection was studied experimentally and computationally. The geometry was a high-speed, subsonic, linear cascade. The independent variables studied were airfoil incidence angle and mass flow rate of end-wall injection upstream of the stator. The end-wall injection was intended to simulate upstream “leakage” through hardware gaps in the end-walls of gas-turbine engines. The exit of the cascade was interrogated experimentally by a five-hole-probe and a total pressure Kiel probe to provide total pressure measurements, which were used to calculate total pressure loss coefficients at the exit of the test section. Computational studies were completed to examine the end-wall flow physics and entropy generating mechanisms through the stator section. The experimental results showed a distinct decrease in the downstream total pressure field with end-wall injection flow, and the impact of the upstream injection on the stator loss coefficient was not a function of the incidence angle. The computational investigation found that the majority of the end-wall injection’s effect on the downstream total pressure field was observed as an increase in the size of the secondary flows on the suction-side of the stator near the upper end-wall.

scholarly journals Design, Optimization, and Analysis of a High-Turning Transonic Tandem Compressor Cascade

1997 ◽  
Author(s):  
Anton Weber ◽  
Wolfgang Steinert
Keyword(s):  
Mach Number ◽  
Total Pressure ◽  
Side Wall ◽  
Design Flow ◽  
Basic Question ◽  
Compressor Cascade ◽  
Flow Angle ◽  
Transonic Compressor ◽  
Blade Row ◽  
Transonic Cascade

As a feasibility study for a stator guide vane a highly loaded transonic compressor stator blade row was designed, optimized, and tested in a transonic cascade facility. The flow entering the turning device with an inlet Mach number of 1.06 has to be turned by more than 60° and diffused extremely to leave the blade row without swirl. Therefore, the basic question was: Is it feasible to gain such a high amount of flow turning with an acceptable level of total pressure losses? The geometric concept chosen is a tandem cascade consisting of a transonic blade row with a flow turning of 10° followed by a subsequent high-turning subsonic cascade. The blade number ratio of the two blade rows was selected to be 1:2 (transonic: subsonic). Design and optimization have been performed using a modern Navier-Stokes flow solver under 2D assumptions by neglecting side wall boundary-layer effects. In the design process it was found to be necessary to guide the wake of the low turning transonic blade near the suction surface of the subsonic blade. Furthermore, it is advantageous to enlarge the blade spacing of the ‘wake’ passage in relation to the neighbouring one of the high turning part. The optimized design geometry of the tandem cascade was tested in the transonic cascade windtunnel of the DLR in Cologne. At design flow conditions the experiments confirmed the design target in every aspect. A flow turning of more than 60°, a static pressure ratio of 1.75, and a total pressure loss coefficient of 0.15 was measured. The working range at design inlet Mach number of 1.06 is about 3.5° in terms of the inlet flow angle. A viscous analysis of various operating points showed excellent agreement with the experimental results.

Sign in / Sign up

Export Citation Format

Share Document