Transport of Unsteadiness Across the Rotor of a Transonic Turbine Stage

2006 ◽  
Author(s):  
Remy De´nos ◽  
V. Jerez Fidalgo ◽  
P. Adami
Keyword(s):  
Total Pressure ◽  
Static Pressure ◽  
Turbulent Viscosity ◽  
Leading Edge ◽  
Pressure Variation ◽  
Frame Of Reference ◽  
Turbine Stage ◽  
Blade Section ◽  
Transonic Turbine ◽  
Absolute Frame

The main sources of excitation for the rotor of a transonic turbine stage are the vane shocks and wakes. Numerical and experimental results are analyzed to understand how these non-uniformities are transported across the rotor. When transposing the pitchwise variations of static and total pressure from the absolute frame of reference into the rotor relative frame of reference in order to derive the rotor relative inlet total pressure, the distortion due to the static pressure dominates. The rotor traverses tangentially the vane trailing edge shock. In particular, the shock sweeps the blade surface from the crown towards the leading edge creating large unsteady variation of static pressure around the blade section. Reflected shocks cause additional fluctuations in the passage. The wake traverses zones of high and low pressure established by the shock. It is convected, distorted and stretched inside the passage. The relative total pressure inside the passage is influenced by both vane shocks and wakes. Downstream of the stage, the signature of the vane non-uniformities is still present. The minimum of the time-averaged pitchwise total pressure variation does not coincide with the vane wake avenue identified thanks to turbulent viscosity plots. It is demonstrated that the vane shock is able to impose total pressure variation downstream of the stage that are larger than that caused by the vane wakes.

Influence of the Hub Endwall Cavity Flow on the Time-Averaged and Time-Resolved Aero-Thermodynamics of an Axial HP Turbine Stage

2002 ◽  
Author(s):  
R. De´nos ◽  
G. Paniagua
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Significant Influence ◽  
Total Pressure ◽  
Rotor Blade ◽  
Static Pressure ◽  
Cavity Flow ◽  
Turbine Stage ◽  
Time Resolved ◽  
Transonic Turbine

This experimental research investigates the influence of the hub endwall cavity flow on the aerodynamics and heat transfer of a high-pressure transonic turbine stage tested under engine representative conditions. The measurements include the hub and tip endwall static pressure downstream of the vane, the static pressure and heat transfer on the rotor blade at 15% span and on the hub platform as well as the stage downstream total pressure and temperature. Both steady and unsteady aspects are addressed. The hub endwall cavity flow has a significant influence on both the time-averaged and time-resolved components of the measured quantities. The effects are shown to be mainly due to an increase of the pitchwise averaged static pressure at hub downstream of the vane when cavity flow ejection is activated.

Design of a Transonic High-Pressure Turbine Stage 2D Section With Reduced Rotor/Stator Interaction

2004 ◽  
Author(s):  
Michele Vascellari ◽  
Re´my De´nos ◽  
Rene´ Van den Braembussche
Keyword(s):  
Rotor Blade ◽  
Static Pressure ◽  
Aerodynamic Force ◽  
Pressure Fluctuations ◽  
Uniform Field ◽  
Turbine Stage ◽  
Rotor Stator Interaction ◽  
Transonic Turbine ◽  
Blade Failure ◽  
Rotor Profiles

In transonic turbine stages, the exit static pressure field of the vane is highly non-uniform in the pitchwise direction. The rotor traverses periodically this non-uniform field and large static pressure fluctuations are observed around the rotor section. As a consequence the rotor blade is submitted to significant variations of its aerodynamic force. This contributes to the high cycle fatigue and may result in unexpected blade failure. In this paper an existing transonic turbine stage section is redesigned in the view of reducing the rotor stator interaction, and in particular the unsteady rotor blade forcing. The first step is the redesign of the stator blade profile to reduce the stator exit pitchwise static pressure gradient. For this purpose, a procedure using a genetic algorithm and an artificial neural network is used. Next, two new rotor profiles are designed and analysed with a quasi 3D Euler unsteady solver in order to investigate their receptivity to the shock interaction. One of the new profiles allows reducing the blade force variation by 50%.

Analysis of the Mixing Plane Interface Between Stator and Rotor of a Transonic Axial Turbine Stage

2000 ◽  
Author(s):  
P. Giangiacomo ◽  
V. Michelassi ◽  
F. Martelli
Keyword(s):  
Mass Flow ◽  
Static Pressure ◽  
Three Dimensional ◽  
Turbine Stage ◽  
Flow Rates ◽  
Tip Leakage ◽  
Mass Flow Rates ◽  
Stator Blade ◽  
Transonic Turbine ◽  
Rotor Mass

A three-dimensional transonic turbine stage is computed by means of a numerical simulation tool. The simulation accounts for the coolant ejection from the stator blade and for the tip leakage of the rotor blade. The stator and rotor rows interact via a mixing plane, which allows the stage to be computed in a steady manner. The analysis is focused on the matching of the stator and rotor mass flow rates. The computations proved that the mixing plane approach allows the stator and rotor mass flow rates to be balanced with a careful choice of the stator-rotor static pressure interface. At the same time, the pitch averaged distribution of the transported quantities at the interface for the stator and rotor may differ slightly, together with the value of the static pressure at the hub.

Influence of bypass ratio on subsonic and correctly expanded sonic co-flowing jets with finite lip thickness

2018 ◽  
Vol 233 (7) ◽  
pp. 2536-2548 ◽  
Author(s):  
S Thanigaiarasu ◽  
R Naren Shankar ◽  
E Rathakrishnan
Keyword(s):  
Total Pressure ◽  
Strong Influence ◽  
Static Pressure ◽  
Pressure Variation ◽  
Jet Mixing ◽  
Free Jet ◽  
Pressure Decay ◽  
Sonic Jet ◽  
Mach Numbers ◽  
Bypass Ratio

The effects of bypass ratio on co-flowing subsonic and correctly expanded sonic jet decay have been studied experimentally. Co-flowing jets with lip thickness 1.0 Dp (where Dp is the diameter of primary nozzle and is equal to 10 mm) with bypass ratios of around 0.7, 1.4, and 6.4 at primary jet exit Mach numbers 0.6, 0.8, and 1.0 have been analyzed. A single free jet equivalent to primary nozzle of the co-flowing nozzle was considered for comparison. Primary jet centerline total pressure decay, spread, and static pressure variation were investigated. The results show that the mixing of the high bypass ratio co-flowing jet with lip thickness 1.0 Dp is superior to low bypass ratio co-flowing jet. Both lip thickness and bypass ratio have a strong influence on the co-flowing jet mixing. Bypass ratio 6.3 experiences a significantly higher mixing than bypass ratio 0.7 and 1.4. Selected jets were also investigated computationally. The computations capture the salient flow physics and reproduce well with the experiments.

scholarly journals Reconstructing Compressor Non-Uniform Circumferential Flow Field From Spatially Undersampled Data: Part 2 — Practical Application for Experiments

2020 ◽  
Author(s):  
Fangyuan Lou ◽  
Douglas R. Matthews ◽  
Nicholas J. Kormanik ◽  
Nicole L. Key
Keyword(s):  
Flow Field ◽  
Total Pressure ◽  
High Speed ◽  
Static Pressure ◽  
Pressure Field ◽  
Leading Edge ◽  
Pressure Profile ◽  
Sampled Data ◽  
Mean Values ◽  
Novel Method

Abstract In the previous part of the paper, a novel method to reconstruct the compressor non-uniform circumferential flow field using spatially under-sampled data points is developed. In this part of the paper, the method is applied to two compressor research articles to further demonstrate the potential of the novel method in resolving the important flow features associated with these circumferential non-uniformities. In the first experiment, the static pressure field at the leading edge of a vaned diffuser in a high-speed centrifugal compressor is reconstructed using pressure readings from nine static pressure taps placed on the hub of the diffuser. The magnitude and phase information for the first three dominant wavelets are characterized. Additionally, the method shows significant advantages over the traditional averaging methods for calculating repeatable mean values of the static pressure. While using the multi-wavelet approximation method, the errors in the mean static pressure with one dropout measurement are 70% less than the pitchwise-averaging method. In the second experiment, the full-annulus total pressure field downstream of Stator 2 in a three-stage axial compressor is reconstructed from a small segment of data representing 20% coverage of the annulus. Results show very good agreement between the reconstructed total pressure profile and the experiment at a variety of spanwise locations from near hub to near shroud. The features associated with blade-row interactions accounting for passage-to-passage variations are resolved in the reconstructed total pressure profile.

scholarly journals Novel Characteristics of Subsonic Coflowing Jets With Varying Lip Thickness

2020 ◽  
Author(s):  
Naren Shankar Radha Krishnan ◽  
Dilip Raja Narayana
Keyword(s):  
Mach Number ◽  
Atmospheric Pressure ◽  
Total Pressure ◽  
Radial Direction ◽  
Static Pressure ◽  
Axial Direction ◽  
Pressure Variation ◽  
Geometrical Constraints ◽  
The Mean ◽  
Mach Numbers

Effect of Mach number on coflowing jet at lip thickness of 0.2 Dp, 1.0 Dp and 1.5 Dp (where Dp is primary nozzle exit diameter, 10 mm) at Mach numbers 1.0, 0.8 and 0.6 were studied experimentally. It was found that an increase in Mach number does not have any profound effect on axial total and static pressure variation for 0.2 Dp. Decreasing the mean diameter is due to the geometrical constraints. In this study, the primary nozzle dimension and secondary duct is maintained constant for comparison. For the case of 0.2 Dp, static pressure is almost equal to atmospheric pressure for all Mach numbers. Whereas for other two lip thickness, increase in Mach number marginally influences axial total pressure and profoundly varies static pressure. It is noted that it varies considerably up to 11.1% in the axial direction and up to 17% in the radial direction for Mach number 1.0. For lower Mach numbers, such variation is not observed. Increase in Mach number increases static pressure variation in the coflowing jet flow field with lip thickness 1.0 Dp and 1.5 Dp.

Co-Flowing Jet Control Using Lip Thickness Variation

2018 ◽  
Vol 0 (0) ◽  
Author(s):  
R. Naren Shankar ◽  
S. Thanigaiarasu ◽  
S. Elangovan ◽  
E. Rathakrishnan
Keyword(s):  
Atmospheric Pressure ◽  
Total Pressure ◽  
Static Pressure ◽  
Pressure Variation ◽  
Thickness Variation ◽  
Nozzle Wall ◽  
Jet Mixing ◽  
Jet Control ◽  
Mach Numbers ◽  
Mixing Behaviour

Abstract The control of co-flowing jets by varying lip thickness has been studied experimentally. Lip thickness is defined as the thickness of primary nozzle wall separating primary jet and secondary jet at the co-flowing nozzle exit. Co-flowing jets from a primary nozzle of diameter 10 mm (1.0 Dp) and a secondary duct with lip thickness (LT) 0.2 Dp, 1.0 Dp and 1.5 Dp at Mach numbers 0.6, 0.8 and 1.0 have been studied. Jet centreline total pressure decay, static pressure variation and jet mixing behaviour were analysed. The results show that the mixing of the co-flowing jet with substantial values of lip thickness is superior to the co-flowing jets with comparatively lower values of lip thickness. Co-flowing jets with lip thickness 1.0 Dp and 1.5 Dp experience a significantly higher mixing than the lip thickness 0.2 Dp jet, for all Mach numbers analyzed in the present study. Moreover, in the case of correctly expanded jets, the local static pressure is assumed to be equal to atmospheric pressure. This assumption becomes invalid for co-flowing jets with substantial lip thickness. The centerline static pressure varies sinusoidally above and below atmospheric pressure by a maximum of 11 %, which is due to wake dominance.

Development of High-Efficiency Half-Ducted Propeller Fan for Air-Conditioners by Blade Tip Shape Modification

2019 ◽  
Author(s):  
Masashi Yoshikawa ◽  
Hiroyuki Toyoda ◽  
Hisashi Daisaka
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Static Pressure ◽  
Leading Edge ◽  
Loss Coefficient ◽  
Total Pressure Loss ◽  
Trailing Edge ◽  
Pressure Loss Coefficient ◽  
Blade Tip ◽  
Total Pressure Loss Coefficient

Abstract We developed a high-efficiency half-ducted propeller fan to reduce the electric power consumption of the outdoor unit of air conditioner by using computational fluid dynamics (CFD). Total pressure loss coefficient on the cylindrical surface of blade tip started increasing at the middle of the blade, and the region of high total pressure loss coefficient was formed after trailing edge. Therefore, we assumed that decreasing this region helped increasing static pressure efficiency. Limiting stream lines on the pressure surface showed that the flow from leading edge leaked at the middle of the blade tip, so it was assumed that the region of the high total pressure loss coefficient arose from the leakage at the middle of the blade tip. We confirmed that static pressure at the middle of blade tip, which was the leakage point, was low. We assumed that low inward force to the flow caused the leakage. On the other hand, static pressure at trailing edge of the blade tip was high. Therefore, it was found that the inward force could be increased by making the static pressure higher at the meddle of the blade tip. In order to make the static pressure higher at the middle of the blade tip, we attempted to move the maximum camber position of the blade tip from trailing edge side to leading edge side. Calculation results showed leakage at the blade tip decreased and the static pressure efficiency increased by 0.5%. Experimental results showed that the static pressure efficiency increased by 1.7 % and sound pressure level was almost the same. For the above reasons, we found leakage of flow from leading edge could be decreased by adjusting the maximum camber position of the blade tip. Decreasing leakage contributed to increasing static pressure efficiency and decreasing electric power consumption.

Performance of Two Transonic Axial Compressor Rotors Incorporating Inlet Counterswirl

1987 ◽  
Vol 109 (1) ◽  
pp. 142-148 ◽  
Author(s):  
C. H. Law ◽  
A. J. Wennerstrom
Keyword(s):  
Static Pressure ◽  
Axial Compressor ◽  
Axial Flow ◽  
Leading Edge ◽  
Axial Compressors ◽  
Streamline Curvature ◽  
Static Pressure Distribution ◽  
Pressure Distributions ◽  
Blade Section ◽  
Flow Compressor

A single-stage axial-flow compressor which incorporates rotor inlet counterswirl to improve stage performance is discussed. Results for two rotor configurations are presented, including design and experimental test data. In this compressor design, inlet guide vanes were used to add counterswirl to the inlet of the rotor. The magnitude of the counterswirl was radially distributed to maximize the overall stage efficiency by minimizing the rotor combined losses (diffusion losses and shock losses). The shock losses were minimized by simultaneously optimizing the rotor blade section geometry, through-blade static pressure distribution, and leading edge aerodynamic/geometric shock sweep angles. Results from both the design and experimental performance analyses are presented and comparisons are made between the experimental data and the analyses and between the performance of both rotor designs. The computation of the flow field for both rotor designs and for the analysis of both tests was performed in an identical fashion using an axisymmetric, streamline-curvature-type code. Results presented include tip section blade-to-blade static pressure distributions and rotor through-blade and exit distributions of various aerodynamic parameters. The performance of this compressor stage represents a significant improvement in axial compressor performance compared to previous attempts to use rotor inlet counterswirl and to current, more conventional, state-of-the-art axial compressors operating under similar conditions.

scholarly journals Influence of Vane-Blade Spacing on Transonic Turbine Stage Aerodynamics: Part II — Time-Resolved Data and Analysis

1998 ◽  
Author(s):  
Judy A. Busby ◽  
Roger L. Davis ◽  
Daniel J. Dorney ◽  
Michael G. Dunn ◽  
Charles W. Haldeman ◽  
...  
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Numerical Solutions ◽  
Unsteady Aerodynamics ◽  
Total Pressure Loss ◽  
Turbine Stage ◽  
Airfoil Surface ◽  
Time Resolved ◽  
Transonic Turbine ◽  
Inlet Conditions

This paper presents results of a combined experimental/computational investigation into the effects of vane-blade spacing on the unsteady aerodynamics of a transonic turbine stage. Time-resolved data were taken in a shock-tunnel facility in which the flow was generated with a short-duration source of heated and pressurized air. This data is compared with the results obtained from four unsteady Navier-Stokes solvers. The time-resolved flow for three axial spacings is examined. For each vane-blade spacing, the inlet conditions were nearly identical and the vane exit flow was transonic. Surface-mounted high-response pressure transducers at midspan were used to obtain the pressure measurements. The computed two-dimensional unsteady airfoil surface pressure predictions are compared with the Kulite pressure transducer measurements. The unsteady and axial spacing effects on loading and performance are examined. In general the numerical solutions compared very favorably with each other and with the experimental data. The overall predicted stage losses and efficiencies did not vary much with vane/blade axial spacing. The computations indicated that any increases in the blade relative total pressure loss were offset by a decrease in vane loss as the axial spacing was decreased. The decrease in predicted vane total pressure loss with decreased axial spacing was primarily due to a reduction in the wake mixing losses. The increase in predicted blade relative total pressure loss with a decrease in axial spacing was found to be mainly due to increased vane wake/blade interaction.

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