Experimental Characterization of the Vane Heat Flux Under Pulsating Trailing-Edge Blowing

2016 ◽  
Author(s):  
J. Saavedra ◽  
G. Paniagua ◽  
B. H. Saracoglu
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Shock Wave ◽  
Heat Flux ◽  
Wall Temperature ◽  
Separation Bubble ◽  
Convective Heat Transfer Coefficient ◽  
Trailing Edge ◽  
Power Extraction ◽  
The Impact

The steady improvement of aircraft engine performance has led towards more compact engine cores with increased structural loads. Compact single-stage high-pressure turbines allow high power extraction, operating in the low supersonic range. The shock waves formed at the airfoil trailing edge contribute substantially to turbine losses, mainly due to the shock-boundary layer interactions as well as high-frequency forces on the rotor. We propose to control the vane trailing edge shock interaction with the downstream rotor, using a pulsating vane-trailing-edge-coolant at the rotor passing frequency. A linear cascade of transonic vanes was investigated at different Mach numbers, ranging from subsonic to supersonic regimes (0.8, 1.1) at two engine representative Reynolds numbers (4 and 6 million). The steady and unsteady heat flux was retrieved using thin-film 2-layered gauges. The complexity of the tests required the development of an original heat transfer post-processing approach. In a single test, monitoring the heat flux data and the wall temperature we obtained the adiabatic wall temperature and the convective heat transfer coefficient. The right-running trailing edge shock wave impacts on the neighboring vane suction side. The impact of the shock wave on the boundary layer creates a separation bubble, which is very sensitive to the intensity and angle of the shock wave. Increasing the coolant blowing rate induces the shock to be less oblique, moving the separation bubble upstream. A similar effect is caused by the pulsations of the coolant.

Experimental Characterization of the Vane Heat Flux Under Pulsating Trailing-Edge Blowing

2017 ◽  
Vol 139 (6) ◽  
Author(s):  
J. Saavedra ◽  
G. Paniagua ◽  
B. H. Saracoglu
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Shock Wave ◽  
Heat Flux ◽  
Wall Temperature ◽  
Separation Bubble ◽  
Convective Heat Transfer Coefficient ◽  
Trailing Edge ◽  
Power Extraction ◽  
The Impact

The steady improvement of aircraft engine performance has led toward more compact engine cores with increased structural loads. Compact single-stage high-pressure turbines allow high power extraction, operating in the low supersonic range. The shock waves formed at the airfoil trailing edge contribute substantially to turbine losses, mainly due to the shock-boundary layer interactions as well as high-frequency forces on the rotor. We propose to control the vane trailing edge shock interaction with the downstream rotor, using a pulsating vane-trailing-edge-coolant at the rotor passing frequency. A linear cascade of transonic vanes was investigated at different Mach numbers, ranging from subsonic to supersonic regimes (0.8, 1.1) at two engine representative Reynolds numbers (4 × 106 and 6 × 106). The steady and unsteady heat flux was retrieved using thin-film two-layered gauges. The complexity of the tests required the development of an original heat transfer postprocessing approach. In a single test, monitoring the heat flux data and the wall temperature we obtained the adiabatic wall temperature and the convective heat transfer coefficient. The right-running trailing edge shock wave impacts on the neighboring vane suction side. The impact of the shock wave on the boundary layer creates a separation bubble, which is very sensitive to the intensity and angle of the shock wave. Increasing the coolant blowing rate induces the shock to be less oblique, moving the separation bubble upstream. A similar effect is caused by the pulsations of the coolant

Analysis of the Unsteady Overtip Casing Heat Transfer in a High Speed Turbine

2013 ◽  
Vol 135 (3) ◽  
Author(s):  
S. Lavagnoli ◽  
G. Paniagua ◽  
C. De Maesschalck ◽  
T. Yasa
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Wall Temperature ◽  
Convective Heat Transfer Coefficient ◽  
Tip Clearance ◽  
Heat Flux Sensor ◽  
Flow Temperature ◽  
Local Flow ◽  
Essential Information ◽  
Transfer Pressure

In modern gas turbine engines, the rotor casing is vulnerable to thermal failures due to large unsteady heat fluxes. The rotor tip flow unsteadiness is induced by the periodic passage of the rotor blades, with an intensity dependent on the tip gap geometry. Hence, the understanding of the physics is of paramount importance to develop appropriate predictive tools and improve the cooling schemes. The present research aims at providing essential information on the flow conditions, which should serve to assess the relative impact of the overtip flow, tip gap magnitude, and work extraction processes on the casing thermal load. This paper presents simultaneous measurements of steady and unsteady heat transfer, pressure and rotor tip clearance in the casing of a transonic turbine stage. The research article was tested in a compression tube facility operating at engine representative conditions (vane Mach number 1.07, vane outlet Reynolds number 1.3 × 106, pressure ratio is 2.92, at 6790 rpm). The rotor blade geometry has a flat tip with a nominal tip clearance of about 0.4% of blade height. The heat transfer, pressure, and tip clearance data were obtained at three circumferential positions around the turbine casing. The heat flux was monitored using a single-layered thin film gauge able to resolve with high fidelity the wall temperature fluctuations. The heat flux sensor was mounted on a probe equipped with a heating device that allows varying the wall temperature. A series of experiments was performed at different heating rates to derive the local adiabatic wall temperature and the adiabatic convective heat transfer coefficient. A high bandwidth capacitive sensor provided the instantaneous value of the single blade tip clearance. A simple zero-dimensional model has been proved effective to predict the local flow temperature while the rotor spins up prior to the test, and estimate the overtip flow temperature during a test.

Condensation of Flowing Vapor on a Horizontal Tube—Numerical Analysis as a Conjugate Heat Transfer Problem

1984 ◽  
Vol 106 (4) ◽  
pp. 841-848 ◽  
Author(s):  
H. Honda ◽  
T. Fujii
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Heat Flux ◽  
Nusselt Number ◽  
Flux Density ◽  
Wall Temperature ◽  
Heat Flux Density ◽  
Tube Wall ◽  
Horizontal Tube ◽  
Uniform Wall

Condensation of flowing vapor on a horizontal tube is numerically analyzed under given conditions of vapor and coolant. Besides the usual boundary layer concept, some approximations are introduced for the determination of shear stress at the vapor-liquid interface. The conjugation of the two-phase boundary layer equations and the heat conduction equation within the tube wall is achieved by using an iterative scheme at the outer surface of the tube wall. The solution thus obtained reveals the effects of vapor velocity, tube material, heat transfer of coolant side, etc., upon circumferential distributions of temperature, heat flux density, and Nusselt number at the outer tube surface. Also the solution compared well with available experimental results for the wall temperature distribution and average Nusselt number. The heat transfer characteristics of steam and refrigerant vapors resemble those of the tubes with uniform wall heat flux density and uniform wall temperature, respectively.

Closed-Form Analytical Solutions for Laminar Natural Convection on Horizontal Plates

2013 ◽  
Vol 135 (10) ◽  
Author(s):  
Abhijit Guha ◽  
Subho Samanta
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Heat Flux ◽  
Natural Convection ◽  
Prandtl Number ◽  
Boundary Layers ◽  
Wall Temperature ◽  
Temperature Profiles ◽  
Integral Analysis ◽  
Laminar Natural Convection

A boundary layer based integral analysis has been performed to investigate laminar natural convection heat transfer characteristics for fluids with arbitrary Prandtl number over a semi-infinite horizontal plate subjected either to a variable wall temperature or variable heat flux. The wall temperature is assumed to vary in the form T¯w(x¯)-T¯∞=ax¯n whereas the heat flux is assumed to vary according to qw(x¯)=bx¯m. Analytical closed-form solutions for local and average Nusselt number valid for arbitrary values of Prandtl number and nonuniform heating conditions are mathematically derived here. The effects of various values of Prandtl number and the index n or m on the heat transfer coefficients are presented. The results of the integral analysis compare well with that of previously published similarity theory, numerical computations and experiments. A study is presented on how the choice for velocity and temperature profiles affects the results of the integral theory. The theory has been generalized for arbitrary orders of the polynomials representing the velocity and temperature profiles. The subtle role of Prandtl number in determining the relative thicknesses of the velocity and temperature boundary layers for natural convection is elucidated and contrasted with that in forced convection. It is found that, in natural convection, the two boundary layers are of comparable thickness if Pr ≤ 1 or Pr ≈ 1. It is only when the Prandtl number is large (Pr > 1) that the velocity boundary layer is thicker than the thermal boundary layer.

On Selection of Reference Temperature of Heat Transfer Coefficient for Complicated Flows

2007 ◽  
Author(s):  
X. C. Li ◽  
J. Zhou ◽  
K. Aung
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Heat Transfer Coefficient ◽  
Surface Temperature ◽  
Transfer Coefficient ◽  
Wall Temperature ◽  
Convective Heat Transfer Coefficient ◽  
Internal Heat ◽  
Reference Temperature ◽  
The Heat Transfer Coefficient

One of the most fundamental concepts in heat transfer is the convective heat transfer coefficient, which is closely related with the flow Reynolds number, flow geometry and the thermal conditions on the heat transfer surface. To define the heat transfer coefficient, a reference temperature is needed besides the surface temperature and heat flux. The reference temperature can be chosen differently, such as the fluid bulk mean temperature (for internal flows) and the temperature at the far field (for external flows). For complicated flows, the adiabatic wall temperature, defined as the wall temperature when the surface heat flux is zero, is commonly adopted as the reference temperature. Other options can also be applied to complicated flows. This paper analyzed some of the potential selections of the reference temperature for different flow settings, including film cooling, jet impingement with cross flows and a mixing flow in a straight duct with or without internal heat source. Both laminar and turbulent flows are considered with different boundary conditions. Dramatic changes of heat transfer coefficient are observed with different reference temperatures. In some special conditions the heat transfer coefficient becomes negative, which means the heat flux has a different direction with the driving temperature difference defined. An innovative method is proposed to calculate the heat transfer coefficient of complicated flows with constant surface temperature.

Impact of Wall Temperature on Turbine Blade Tip Aerothermal Performance

2014 ◽  
Vol 136 (5) ◽  
Author(s):  
Q. Zhang ◽  
L. He
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
Flow Field ◽  
Temperature Dependence ◽  
Turbine Blade ◽  
Wall Temperature ◽  
High Speed ◽  
Convective Heat Transfer Coefficient ◽  
Secondary Flows ◽  
Blade Tip

Currently the aerodynamics and heat transfer over a turbine blade tip tend to be analyzed separately with the assumption that the wall thermal boundary conditions do not affect the over-tip-leakage (OTL) flow field. There are some existing correlations for correcting the wall temperature effect on heat transfer when scaled to engine realistic conditions. But they were either developed to account for the temperature dependence of fluid properties largely empirically, or based on a boundary-layer model. It would be difficult (if not impossible) to define a boundary layer in many parts of a realistic blade passage with marked three-dimensional (3D) end wall and secondary flows (including those within a blade tip and around it). The questions to be asked here are: is the OTL aerodynamics significantly affected by the wall thermal condition? And if it is, how can we count this effect consistently in turbine blade tip design and analysis using modern CFD methods? In the present study the problem has been examined for typical high-pressure turbine blade tip configurations. An extensively developed RANS code (HYDRA) is employed and validated against the experimental data from a high speed linear cascade testing rig. The numerical analysis reveals that the wall–gas temperature ratio could greatly affect the transonic OTL flow field and there is a strong two-way coupling between aerodynamics and heat transfer. The feedbacks of the thermal boundary condition to aerodynamics behave differently at different flow regimes over the tip, clearly indicating a highly localized dependence of the convective heat transfer coefficient (HTC) upon wall temperatures. This implies that to use HTC for blade metal temperature predictions without resorting a fully conjugate solution, the temperature dependence needs to be corrected locally. A nonlinear correction approach has been adopted in the present work, and the results demonstrate its effectiveness for the transonic turbine tip configurations studied.

Analysis of the Heat Transfer Driving Parameters in Tight Rotor Blade Tip Clearances

2014 ◽  
Author(s):  
S. Lavagnoli ◽  
C. De Maesschalck ◽  
G. Paniagua
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Convective Heat Transfer ◽  
Convective Heat ◽  
Wall Temperature ◽  
Convective Heat Transfer Coefficient ◽  
Tip Clearance ◽  
Adiabatic Wall ◽  
Engine Operation ◽  
Flow Parameters

Turbine rotor tips and casings are vulnerable to mechanical failures due to the extreme thermal loads they undergo during engine operation. In addition to the heat flux variations during the transient phase, high-frequency unsteadiness occurs at every rotor passage, with amplitude dependent on the tip gap. The development of appropriate predictive tools and cooling schemes requires the precise understanding of the heat transfer mechanisms. The present paper analyzes the nature of the overtip flow in transonic turbine rotors running at tight clearances, and explores a methodology to determine the relevant flow parameters that model the heat transfer. Steady-state three-dimensional Reynolds-Averaged Navier-Stokes calculations were performed to simulate engine-like conditions considering two rotor tip gaps, 0.1% and 1% of the blade span. At tight tip clearance, the adiabatic wall temperature is not anymore independent of the solid thermal boundary conditions. The adiabatic wall temperature predicted with the linear Newton’s cooling law was observed to rise to non-physical levels in certain regions within the rotor tip gap, resulting in unreliable convective heat transfer coefficients. This paper investigates different approaches to estimate the relevant flow parameters that drive the heat transfer. The present study allows experimentalists to retrieve information on the gap flow temperature and convective heat transfer coefficient based on the use of wall heat flux measurements. Such approach is required to improve the accuracy in the evaluation of the heat transfer data while enhancing the understanding of tight-clearance overtip flows.

Analysis of the Unsteady Overtip Casing Heat Transfer in a High Speed Turbine

2012 ◽  
Author(s):  
S. Lavagnoli ◽  
G. Paniagua ◽  
C. De Maesschalck ◽  
T. Yasa
Keyword(s):  
Heat Transfer ◽  
Heat Flux ◽  
Wall Temperature ◽  
Convective Heat Transfer Coefficient ◽  
Tip Clearance ◽  
Heat Flux Sensor ◽  
Flow Temperature ◽  
Local Flow ◽  
Essential Information ◽  
Transfer Pressure

In modern gas turbine engines, the rotor casing is vulnerable to thermal failures due to large unsteady heat fluxes. The rotor tip flow unsteadiness is induced by the periodic passage of the rotor blades, with an intensity dependent on the tip gap geometry. Hence, the understanding of the physics is of paramount importance to develop appropriate predictive tools and improve the cooling schemes. The present research aims at providing essential information on the flow conditions, which should serve to assess the relative impact of the overtip flow, tip gap magnitude and work extraction processes on the casing thermal load. This paper presents simultaneous measurements of steady and unsteady heat transfer, pressure and rotor tip clearance in the casing of a transonic turbine stage. The research article was tested in a compression tube facility operating at engine representative conditions (vane Mach number 1.07, vane outlet Reynolds number 1.3×106, pressure ratio is 2.92, at 6790 RPM). The rotor blade geometry has a flat tip with a nominal tip clearance of about 0.4% of blade height. The heat transfer, pressure, and tip clearance data were obtained at three circumferential positions around the turbine casing. The heat flux was monitored using a single-layered thin film gauge able to resolve with high-fidelity the wall temperature fluctuations. The heat flux sensor was mounted on a probe equipped with a heating device that allows varying the wall temperature. A series of experiments was performed at different heating rates to derive the local adiabatic wall temperature and the adiabatic convective heat transfer coefficient. A high bandwidth capacitive sensor provided the instantaneous value of the single blade tip clearance. A simple zero-dimensional model has been proved effective to predict the local flow temperature while the rotor spins up prior to the test, and estimate the overtip flow temperature during a test.

scholarly journals Effect of wall temperature on separation bubble size in laminar hypersonic shock/boundary layer interaction flows

2019 ◽  
Vol 11 (11) ◽  
pp. 168781401988555 ◽  
Author(s):  
Amjad A Pasha ◽  
Khalid A Juhany
Keyword(s):  
Boundary Layer ◽  
Heat Flux ◽  
Mach Number ◽  
Wall Temperature ◽  
Separation Bubble ◽  
Layer Interaction ◽  
Shock Boundary Layer Interaction ◽  
Boundary Layer Interaction ◽  
Hypersonic Shock ◽  
Bubble Length

At hypersonic speeds, the external wall temperatures of an aerospace vehicle vary significantly. As a result, there is a considerable heat transfer variation between the boundary layer and the wall of the hypersonic vehicle. In this article, numerical computations are performed to investigate the effect of wall temperature on the separation bubble length in laminar hypersonic shock-wave/boundary-layer interaction flows over double-cone configuration at the Mach number of 12.2. The flow field is described in detail in terms of different shocks, expansion fans, shear layer and separation bubble. The variation of the Prandtl number has a negligible effect on the flow field and wall data. A specific heat ratio of less than 1.4 results in the better prediction of wall pressure and heat flux in the shock/boundary-layer interaction region. It is observed that as the wall temperature is increased, the separation bubble size and hence the separation shock length increases. The high firmness of the laminar boundary-layer at a high Mach number shows that the wall temperature in the shock/boundary-layer interaction region has little effect. The peak wall pressure and heat flux decrease with an increase in wall temperature. An estimation is developed between separation bubble length and wall temperature based on the computed results.

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