scholarly journals Unsteady Pressure Measurements on a Biconvex Airfoil in a Transonic Oscillating Cascade

1986 ◽  
Vol 108 (1) ◽  
pp. 53-59 ◽  
Author(s):  
L. M. Shaw ◽  
D. R. Boldman ◽  
A. E. Buggele ◽  
D. H. Buffum
Keyword(s):  
Mach Number ◽  
Incidence Angle ◽  
Dynamic Pressure ◽  
Leading Edge ◽  
Pressure Transducers ◽  
High Incidence ◽  
Shock Motion ◽  
Plate Analysis ◽  
Phase Angles ◽  
Unsteady Pressure Measurements

Flush-mounted dynamic pressure transducers were installed on the center airfoil of a transonic oscillating cascade to measure the unsteady aerodynamic response as nine airfoils were simultaneously driven to provide 1.2 deg of pitching motion about the midchord. Initial tests were performed at an incidence angle of 0.0 deg and a Mach number of 0.65 in order to obtain results in a shock-free compressible flow field. Subsequent tests were performed at an angle of attack of 7.0 deg and a Mach number of 0.80 in order to observe the surface pressure response with an oscillating shock near the leading edge of the airfoil. Results are presented for interblade phase angles of 90 and −90 deg and at blade oscillatory frequencies of 200 and 500 Hz (semichord reduced frequencies up to about 0.5 at a Mach number of 0.80). Results from the zero-incidence cascade are compared with a classical unsteady flat-plate analysis. Flow visualization results depicting the shock motion on the airfoils in the high-incidence cascade are discussed. The airfoil pressure data are tabulated.

Experimental Investigation on the Annular Sector Cascade of a High Endwall-Angle Turbine

2016 ◽  
Author(s):  
Weiliang Fu ◽  
Jie Gao ◽  
Chen Liang ◽  
Fukai Wang ◽  
Qun Zheng ◽  
...  
Keyword(s):  
Mach Number ◽  
Static Pressure ◽  
Incidence Angle ◽  
Pressure Sensors ◽  
Leading Edge ◽  
Control Program ◽  
Turbine Cascade ◽  
Cascade Flow ◽  
Annular Sector ◽  
The Way

The flow in high endwall-angle turbine is complex, and it is different from the ordinary turbine flow in characteristics. In order to study the flow field characteristics of high endwall-angle turbines, the annular sector cascade experimental study of high endwall-angle turbines is carried out. The blade is studied experimentally in the form of annular sector cascade. The cascade includes 7 blades, and makes up 6 flow passages, in order to simulate full cascade flow. The experimental Mach number is adjusted by the way of changing inlet total pressure, and the Mach number influence (0.7, 0.8 and 0.9) on annular sector cascade flow is studied. Based on it, the inlet incidence angle (−15°, −7.5°, 0°, 7.5° and 15° )is changed with the way of changing sector straight pipes upstream of the cascade, and its influence on turbine flow fields is studied at the Mach number of 0.8. Here, five-hole probes are used to measure aerodynamic parameters distributions downstream of the cascade, and static pressure taps are positioned on the blade surface to measure surface static pressure distribution. The auto-traversing system and pressure sensors were operated by a self-compiled program based control program. The results indicate that there are two passage vortices inside the turbine cascade flow passage under the high Mach number condition, and the passage vortex near the high endwall-angle region is bigger. As Mach number increases, the passage vortices inside turbine cascade passage will become strong, and moves towards the blade mid-span. Besides, it is shown that the way of changing sector straight pipes can achieve the variation of inlet incidence angles. And, the blade profile with big leading-edge radius has good design and off-design performance. Detailed results and analyses are presented in the paper.

scholarly journals Design and Test of a New Axial Compressor for the Nuovo Pignone Heavy Duty Gas Turbines

1996 ◽  
Author(s):  
Erio Benvenuti
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Dynamic Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Leading Edge ◽  
Meridional Flow ◽  
Heavy Duty ◽  
Pressure Transducers ◽  
Easy Integration

This axial compressor design was primarily focused to increase the power rating of the current Nuovo Pignone PGT10 Heavy-Duty gas turbine by 10%. In addition, the new 11-stage design favourably compares with the existing 17-stage compressor in terms of simplicity and cost. By seating the flowpath and blade geometry, the new aerodynamic design can be applied to gas turbines with different power ratings as well. The reduction in the stage number was achieved primarily through the meridional flow-path redesign. The resulting higher blade peripheral speeds achieve larger stage pressure ratios without increasing the aerodynamic loadings. Wide chord blades keep the overall length unchanged thus assuring easy integration with other existing components. The compressor performance map was extensively checked over the speed range required for two-shaft gas turbines. The prototype unit was installed on a special PGT10 gas turbine setup, that permitted the control of pressure ratio independently from the turbine matching requirements. The flowpath instrumentation included strain-gages, dynamic pressure transducers and stator vane leading edge aerodynamic probes to determine individual stage characteristics. The general blading vibratory behavior was proved fully satisfactory. With minor adjustments to the variable stator settings the front stage aerodynamic matching was optimized and the design performance was achieved.

scholarly journals Experimental Methods Applied in a Study of Stall Flutter in an Axial Flow Fan

2004 ◽  
Vol 11 (5-6) ◽  
pp. 597-613 ◽  
Author(s):  
John D. Gill ◽  
Vincent R. Capece ◽  
Ronald B. Fost
Keyword(s):  
Experimental Data ◽  
Dynamic Pressure ◽  
Axial Flow ◽  
Leading Edge ◽  
Axial Flow Fan ◽  
Pressure Transducers ◽  
High Frequency Response ◽  
Stall Flutter ◽  
Response Pressure ◽  
Flutter Testing

Flutter testing is an integral part of aircraft gas turbine engine development. In typical flutter testing blade mounted sensors in the form of strain gages and casing mounted sensors in the form of light probes (NSMS) are used. Casing mounted sensors have the advantage of being non-intrusive and can detect the vibratory response of each rotating blade. Other types of casing mounted sensors can also be used to detect flutter of rotating blades. In this investigation casing mounted high frequency response pressure transducers are used to characterize the part-speed stall flutter response of a single stage unshrouded axial-flow fan. These dynamic pressure transducers are evenly spaced around the circumference at a constant axial location upstream of the fan blade leading edge plane. The pre-recorded experimental data at 70% corrected speed is analyzed for the case where the fan is back-pressured into the stall flutter zone. The experimental data is analyzed using two probe and multi-probe techniques. The analysis techniques for each method are presented. Results from these two analysis methods indicate that flutter occurred at a frequency of 411 Hz with a dominant nodal diameter of 2. The multi-probe analysis technique is a valuable method that can be used to investigate the initiation of flutter in turbomachines.

A Navier-Stokes Analysis of the Stall Flutter Characteristics of the Buffum Cascade

2000 ◽  
Author(s):  
Stefan Weber ◽  
Max F. Platzer
Keyword(s):  
Mach Number ◽  
Turbulence Modeling ◽  
Separation Bubble ◽  
Leading Edge ◽  
Navier Stokes ◽  
Implicit Time ◽  
Fully Implicit ◽  
High Incidence ◽  
Stall Flutter ◽  
Good Agreement

Numerical stall flutter prediction methods are highly needed as modern jet engines require blade designs close to the stability boundaries of the performance map. A Quasi-3D Navier-Stokes code is used to analyze the flow over the oscillating cascade designed and manufactured by Pratt & Whitney, and studied at the NASA Glenn Research Center by Buffum et al. The numerical method solves for the governing equations with a fully implicit time-marching technique in a single passage by making use of a direct-store, periodic boundary condition. For turbulence modeling the Baldwin-Lomax model is used. To account for transition, the criterion to predict the onset location suggested by Baldwin and Lomax is incorporated. Buffum et al. investigated two incidence cases for three different Mach numbers. The low-incidence case at a Mach number of 0.5 exhibited the formation of small separation bubbles at reduced oscillation frequencies of 0.8 and 1.2. For this case the present approach yielded good agreement with the steady and oscillatory measurements. At high-incidence at the same Mach number of 0.5 the measured steady-state pressure distribution and the separation bubble on the upper surface was also found in good agreement with the experiment. But computations for oscillations at high-incidence failed to predict the negative damping contribution caused by the leading edge separation.

The Effects of Inlet Reynolds Number, Exit Mach Number and Incidence Angle on Leading Edge Film Cooling Effectiveness of a Turbine Blade in a Linear Transonic Cascade

2015 ◽  
Author(s):  
Cong Liu ◽  
Hui-ren Zhu ◽  
Zhong-yi Fu ◽  
Run-hong Xu
Keyword(s):  
Reynolds Number ◽  
Mach Number ◽  
Film Cooling ◽  
Incidence Angle ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Exit Mach Number

This paper experimentally investigates the film cooling performance of a leading edge with three rows of film holes on an enlarged turbine blade in a linear cascade. The effects of blowing ratio, inlet Reynolds number, isentropic exit Mach number and off-design incidence angle (i<0°) are considered. Experiments were conducted in a short-duration transonic wind tunnel which can model realistic engine aerodynamic conditions and adjust inlet Reynolds number and exit Mach number independently. The surface film cooling measurements were made at the midspan of the blade using thermocouples based on transient heat transfer measurement method. The changing of blowing ratio from 1.7 to 3.3 leads to film cooling effectiveness increasing on both pressure side and suction side. The Mach number or Reynolds number has no effect on the film cooling effectiveness on pressure side nearly, while increasing these two factors has opposite effect on film cooling performance on suction side. The increasing Mach number decreases the film cooling effectiveness at the rear region mainly, while at higher Reynolds number condition, the whole suction surface has significantly higher film cooling effectiveness because of the increasing cooling air mass flow rate. When changing the incidence angle from −15° to 0°, the film cooling effectiveness of pressure side decreases, and it presents the opposite trend on suction side. At off-design incidence of −15° and −10°, there is a low peak following the leading edge on the pressure side caused by the separation bubble, but it disappears with the incidence and blowing ratio increased.

Unsteady Aerodynamic Analysis of an Oscillating Cascade at Large Incidence

1999 ◽  
Author(s):  
Jong-Shang Liu ◽  
Durbha V. Murthy
Keyword(s):  
Mach Number ◽  
Fluid Dynamic ◽  
Leading Edge ◽  
Two Dimensional ◽  
Edge Region ◽  
Unsteady Aerodynamic ◽  
High Incidence ◽  
Flutter Analysis ◽  
High Mach ◽  
Good Agreement

The flutter analysis capability of the quasi-3D aeroelastic computational fluid dynamic (CFD) code UNSFLO is evaluated by comparing to unsteady pressure results in an oscillating cascade. The geometry is two-dimensional and the oscillation is well-controlled. Time unsteady UNSFLO results are compared with data at inlet Mach number from 0.2 to 0.8 with incidence 0° and 10°. Three reduced frequencies, 0.4, 0.8, and 1.2 with inter-blade phase angle of 180 degrees were tested. The calculated steady state loadings show good agreement with data at zero incidence. The correlations become worse for high incidence angles because of the separation. The calculated aerodynamic work capture the chordwise distribution except in the near leading edge region. The correlations become also worse for high incidence from the leading edge to midchord especially at high Mach number.

Assessment of the Academic CFD Code Galatea-I With the DLR-F11 Model in High Lift Configuration

2016 ◽  
Author(s):  
Sotirios S. Sarakinos ◽  
Georgios N. Lygidakis ◽  
Ioannis K. Nikolos
Keyword(s):  
Mach Number ◽  
Incidence Angle ◽  
Three Dimensional ◽  
Leading Edge ◽  
Navier Stokes ◽  
Test Case ◽  
High Lift ◽  
Low Mach Number ◽  
Flow Phenomena ◽  
Number Flow

In this study the development and assessment of an academic CFD (Computational Fluid Dynamics) code, named Galatea-I, is reported. The proposed solver employs the RANS (Reynolds-Averaged Navier-Stokes) approach, modified by the artificial compressibility method, along with the SST (Shear Stress Transport) turbulence model to predict steady or unsteady turbulent incompressible flow phenomena on three-dimensional unstructured hybrid grids, composed of prismatic, tetrahedral and pyramidal elements. Parallel processing and an agglomeration multigrid method have been included for the acceleration of the solver’s methodologies. Galatea-I is evaluated against a test case of the HiLiftPW-2 (Second High Lift Prediction Workshop). In particular, the low Mach number flow at 7° incidence angle over the DLR-F11 aircraft configuration of Case 1 of the aforementioned workshop was examined; it considers a three-element wing with a leading edge slat and a trailing edge flap attached on a body pod, without including though any of the support brackets used in the wind tunnel experiments. The obtained results are close to the available experimental data, as well as the numerical results of other reference solvers, indicating the proposed methodology’s potential to predict accurately such low Mach number flows over complex geometries.

Design and Test of a New Axial Compressor for the Nuovo Pignone Heavy-Duty Gas Turbines

1997 ◽  
Vol 119 (3) ◽  
pp. 633-639 ◽  
Author(s):  
Erio Benvenuti
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Dynamic Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Leading Edge ◽  
Heavy Duty ◽  
Pressure Transducers ◽  
Heavy Duty Gas Turbine ◽  
Easy Integration

This axial compressor design was primarily focused to increase the power rating of the current Nuovo Pignone PGT10 Heavy-Duty gas turbine by 10 percent. In addition, the new 11-stage design favorably compares with the existing 17-stage compressor in terms of simplicity and cost. By scaling the flowpath and blade geometry, the new aerodynamic design can be applied to gas turbines with different power ratings as well. The reduction in the stage number was achieved primarily through the meridional flowpath redesign. The resulting higher blade peripheral speeds achieve larger stage pressure ratios without increasing the aerodynamic loadings. Wide chord blades keep the overall length unchanged thus assuring easy integration with other existing components. The compressor performance map was extensively checked over the speed range required for two-shaft gas turbines. The prototype unit was installed on a special PGT10 gas turbine setup, that permitted the control of pressure ratio independently from the turbine matching requirements. The flowpath instrumentation included strain gages, dynamic pressure transducers, and stator vane leading edge aerodynamic probes to determine individual stage characteristics. The general blading vibratory behavior was proved fully satisfactory. With minor adjustments to the variable stator settings, the front stage aerodynamic matching was optimized and the design performance was achieved.

A Navier–Stokes Analysis of the Stall Flutter Characteristics of the Buffum Cascade

2000 ◽  
Vol 122 (4) ◽  
pp. 769-776 ◽  
Author(s):  
Stefan Weber ◽  
Max F. Platzer
Keyword(s):  
Mach Number ◽  
Turbulence Modeling ◽  
Separation Bubble ◽  
Leading Edge ◽  
Navier Stokes ◽  
Implicit Time ◽  
Fully Implicit ◽  
High Incidence ◽  
Stall Flutter ◽  
Good Agreement

Numerical stall flutter prediction methods are much needed, as modern jet engines require blade designs close to the stability boundaries of the performance map. A Quasi-3D Navier–Stokes code is used to analyze the flow over the oscillating cascade designed and manufactured by Pratt & Whitney, and studied at the NASA Glenn Research Center by Buffum et al. The numerical method solves for the governing equations with a fully implicit time-marching technique in a single passage by making use of a direct-store, periodic boundary condition. For turbulence modeling, the Baldwin–Lomax model is used. To account for transition, the criterion to predict the onset location suggested by Baldwin and Lomax is incorporated. Buffum et al. investigated two incidence cases for three different Mach numbers. The low-incidence case at a Mach number of 0.5 exhibited the formation of small separation bubbles at reduced oscillation frequencies of 0.8 and 1.2. For this case the present approach yielded good agreement with the steady and oscillatory measurements. At high incidence at the same Mach number of 0.5 the measured steady-state pressure distribution and the separation bubble on the upper surface was also found in good agreement with the experiment. But computations for oscillations at high incidence failed to predict the negative damping contribution caused by the leading edge separation. [S0889-504X(00)01304-0]

scholarly journals Shock Characteristics Measured Upstream of Both a Forward-Swept and an Aft-Swept Fan

2007 ◽  
Author(s):  
Gary G. Podboy ◽  
Martin J. Krupar ◽  
Daniel L. Sutliff ◽  
Csaba Horvath
Keyword(s):  
Flow Visualization ◽  
Leading Edge ◽  
Suction Side ◽  
Surface Flow ◽  
Rotor Blades ◽  
Laser Doppler Velocimeter ◽  
Pressure Measurements ◽  
Pressure Transducers ◽  
Unsteady Pressure ◽  
Unsteady Pressure Measurements

Three different types of diagnostic data — blade surface flow visualization, shroud unsteady pressure, and laser Doppler velocimeter (LDV) — were obtained on two fans, one forward-swept and one aft-swept, in order to learn more about the shocks which propagate upstream of these rotors when they are operated at transonic tip speeds. Flow visualization data are presented for the forward-swept fan operating at 13831 RPMc and for the aft-swept fan operating at 12500 and 13831 RPMc (corresponding to tip rotational Mach numbers of 1.07 and 1.19, respectively). The flow visualization data identify where the shocks occur on the suction side of the rotor blades. These data show that at the takeoff speed, 13831 RPMc, the shocks occurring in the tip region of the forward-swept fan are further downstream in the blade passage than with the aft-swept fan. Shroud unsteady pressure measurements were acquired using a linear array of 15 equally-spaced pressure transducers extending from two tip axial chords upstream to 0.8 tip axial chords downstream of the static position of the tip leading edge of each rotor. Such data are presented for each fan operating at one subsonic and five transonic tip speeds. The unsteady pressure data show relatively strong detached shocks propagating upstream of the aft-swept rotor at the three lowest transonic tip speeds, and weak, oblique pressure disturbances attached to the tip of the aft-swept fan at the two highest transonic tip speeds. The unsteady pressure measurements made with the forward-swept fan do not show strong shocks propagating upstream of that rotor at any of the tested speeds. A comparison of the forward-swept and aft-swept shroud unsteady pressure measurements indicates that at any given transonic speed the pressure disturbance just upstream of the tip of the forward-swept fan is much weaker than that of the aft-swept fan. The LDV data suggest that at 12500 and 13831 RPMc, the forward-swept fan swallowed the passage shocks occurring in the tip region of the blades, whereas the aft-swept fan did not. Due to this difference, the flows just upstream of the two fans were found to be quite different at both of these transonic speeds. Nevertheless, despite distinct differences just upstream of the two rotors, the two fan flows were much more alike about one axial blade chord further upstream. As a result, the LDV data suggest that it is unwise to attempt to determine the effect that the shocks have on far field noise by focusing only on measurements (or CFD predictions) made very near the rotor. Instead, these data suggest that it is important to track the shocks throughout the inlet.

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