Developments in Hot-Streak Simulators for Turbine Testing

2009 ◽  
Vol 131 (3) ◽  
Author(s):  
Thomas Povey ◽  
Imran Qureshi
Keyword(s):  
Temperature Profile ◽  
Gas Injection ◽  
Test Facility ◽  
Inlet Temperature ◽  
Temperature Profiles ◽  
Gas Temperature ◽  
High Pressure Turbine ◽  
Injection Temperature ◽  
Hot Streak ◽  
The Impact

The importance of understanding the impact of hot-streaks, and temperature distortion in general, on the high pressure turbine is widely appreciated, although it is still generally the case that turbines are designed for uniform inlet temperature—often the predicted peak gas temperature. This is because there is an insufficiency of reliable experimental data both from operating combustors and from rotating turbine experiments in which a combustor representative inlet temperature profile has accurately been simulated. There is increasing interest, therefore, in experiments that attempt to address this deficiency. Combustor (hot-streak) simulators have been implemented in six rotating turbine test facilities for the study of the effects on turbine life, heat transfer, aerodynamics, blade forcing, and efficiency. Three methods have been used to simulate the temperature profile: (a) the use of foreign gas to simulate the density gradients that arise due to temperature differences, (b) heat exchanger temperature distortion generators, and (c) cold gas injection temperature distortion generators. Since 2004 three significant new temperature distortion generators have been commissioned, and this points to the current interest in the field. The three new distortion generators are very different in design. The generator designs are reviewed, and the temperature profiles that were measured are compared in the context of the available data from combustors, which are also collected. A universally accepted terminology for referring to and quantifying temperature distortion in turbines has so far not developed, and this has led to a certain amount of confusion regarding definitions and terminology, both of which have proliferated. A simple means of comparing profiles is adopted in the paper and is a possible candidate for future use. New whole-field combustor measurements are presented, and the design of an advanced simulator, which has recently been commissioned to simulate both radial and circumferential temperature nonuniformity profiles in the QinetiQ/Oxford Isentropic Light Piston Turbine Test Facility, is presented.

The Effect of Hot-Streaks on HP Vane Surface and Endwall Heat Transfer: An Experimental and Numerical Study

2005 ◽  
Vol 129 (1) ◽  
pp. 32-43 ◽  
Author(s):  
T. Povey ◽  
K. S. Chana ◽  
T. V. Jones ◽  
J. Hurrion
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Rate ◽  
Transfer Rate ◽  
Cooling System ◽  
Numerical Study ◽  
Guide Vane ◽  
Inlet Temperature ◽  
Temperature Profiles ◽  
Gas Temperature ◽  
Hot Streak

Pronounced nonuniformities in combustor exit flow temperature (hot-streaks), which arise because of discrete injection of fuel and dilution air jets within the combustor and because of endwall cooling flows, affect both component life and aerodynamics. Because it is very difficult to quantitatively predict the effects of these temperature nonuniformities on the heat transfer rates, designers are forced to budget for hot-streaks in the cooling system design process. Consequently, components are designed for higher working temperatures than the mass-mean gas temperature, and this imposes a significant overall performance penalty. An inadequate cooling budget can lead to reduced component life. An improved understanding of hot-streak migration physics, or robust correlations based on reliable experimental data, would help designers minimize the overhead on cooling flow that is currently a necessity. A number of recent research projects sponsored by a range of industrial gas turbine and aero-engine manufacturers attest to the growing interest in hot-streak physics. This paper presents measurements of surface and endwall heat transfer rate for a high-pressure (HP) nozzle guide vane (NGV) operating as part of a full HP turbine stage in an annular transonic rotating turbine facility. Measurements were conducted with both uniform stage inlet temperature and with two nonuniform temperature profiles. The temperature profiles were nondimensionally similar to profiles measured in an engine. A difference of one-half of an NGV pitch in the circumferential (clocking) position of the hot-streak with respect to the NGV was used to investigate the affect of clocking on the vane surface and endwall heat transfer rate. The vane surface pressure distributions, and the results of a flow-visualization study, which are also given, are used to aid interpretation of the results. The results are compared to two-dimensional predictions conducted using two different boundary layer methods. Experiments were conducted in the Isentropic Light Piston Facility (ILPF) at QinetiQ Farnborough, a short-duration engine-sized turbine facility. Mach number, Reynolds number, and gas-to-wall temperature ratios were correctly modeled. It is believed that the heat transfer measurements presented in this paper are the first of their kind.

Analysis on the Effect of a Non-Uniform Inlet Profile on Heat Transfer and Fluid Flow in Turbine Stages

2010 ◽  
Author(s):  
Salvadori Simone ◽  
Francesco Montomoli ◽  
Francesco Martelli ◽  
Kam S. Chana ◽  
Imran Qureshi ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Temperature Profile ◽  
Thermal Load ◽  
Thermal Loading ◽  
Radial Profile ◽  
Experimental Studies ◽  
Test Facility ◽  
Inlet Temperature ◽  
Inlet Conditions ◽  
Hot Streak

This paper presents an investigation of the aerothermal performance of a modern unshrouded high pressure (HP) aeroengine turbine subject to non-uniform inlet temperature profile. The turbine used for the study was the MT1 turbine installed in the QinetiQ Turbine Test Facility (TTF) based in Farnborough (UK). The MT1 turbine is a full scale transonic HP turbine, and is operated in the test facility at the correct non-dimensional conditions for aerodynamics and heat transfer. Datum experiments of aero-thermal performance were conducted with uniform inlet conditions. Experiments with nonuniform inlet temperature were conducted with a temperature profile that had a non-uniformity in the radial direction defined by (Tmax−Tmin)/T = 0.355, and a non-uniformity in the circumferential direction defined by (Tmax−Tmin)/T = 0.14. This corresponds to an extreme point in the engine cycle, in an engine where the non-uniformity is dominated by the radial distribution. Accurate experimental area surveys of the turbine inlet and exit flows were conducted, and detailed heat transfer measurements were obtained on the blade surfaces and end-walls. These results are analysed with the unsteady numerical data obtained using the in-house HybFlow code developed at the University of Firenze. Two particular aspects are highlighted in the discussion: prediction confidence for state of the art computational fluid dynamics (CFD) and impact of real conditions on stator-rotor thermal loading. The efficiency value obtained with the numerical analysis is compared with the experimental data and a 0.8% difference is found and discussed. A study of the flow field influence on the blade thermal load has also been detailed. It is shown that the hot streak migration mainly affects the rotor pressure side from 20% to 70% of the span, where the Nusselt number increases by a factor of 60% with respect to the uniform case. Furthermore, in this work it has been found that a nonuniform temperature distribution is beneficial for the rotor tip, contrary to the results found in the open literature. Although the hot streak is affected by the pressure gradient across the tip gap, the radial profile (which dominates the temperature profile being considered) is not fully mixed out in passing through the HP stage, and contributes significantly to cooling the turbine casing. A design approach not taking into account these effects will underestimate to rotor life near the tip and the thermal load at mid-span. The temperature profile that has been used in both the experiments and CFD is the first simulation of an extreme cycle point (more than twice the magnitude of distortion all previous experimental studies): it represents an engine-take-off condition combined with the full combustor cooling. The research was part of the EU funded TATEF2 (Turbine Aero-Thermal External Flows 2) programme.

Aerodynamics and Heat Transfer for a Cooled One and One-Half Stage High-Pressure Turbine—Part I: Vane Inlet Temperature Profile Generation and Migration

2011 ◽  
Vol 134 (1) ◽  
Author(s):  
R. M. Mathison ◽  
C. W. Haldeman ◽  
M. G. Dunn
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Temperature Profile ◽  
Film Cooling ◽  
Leading Edge ◽  
Inlet Temperature ◽  
Temperature Profiles ◽  
Fluid Temperature ◽  
High Pressure Turbine ◽  
Wide Range

As controlled laboratory experiments using full-stage turbines are expanded to replicate more of the complicated flow features associated with real engines, it is important to understand the influence of the vane inlet temperature profile on the high-pressure vane and blade heat transfer as well as its interaction with film cooling. The temperature distribution of the incoming fluid governs not only the input conditions to the boundary layer but also the overall fluid migration. Both of these mechanisms have a strong influence on surface heat flux and therefore component life predictions. To better understand the role of the inlet temperature profile, an electrically heated combustor emulator capable of generating uniform, radial, or hot streak temperature profiles at the high-pressure turbine vane inlet has been designed, constructed, and operated over a wide range of conditions. The device is shown to introduce a negligible pressure distortion while generating the inlet temperature conditions for a stage-and-a-half turbine operating at design-corrected conditions. For the measurements described here, the vane is fully cooled and the rotor purge flow is active, but the blades are uncooled. Detailed temperature measurements are obtained at rake locations upstream and downstream of the turbine stage as well as at the leading edge and platform of the blade in order to characterize the inlet temperature profile and its migration. The use of miniature butt-welded thermocouples at the leading edge and on the platform (protruding into the flow) on a rotating blade is a novel method of mapping a temperature profile. These measurements show that the reduction in fluid temperature due to cooling is similar in magnitude for both uniform and radial vane inlet temperature profiles.

Analysis on the Effect of a Nonuniform Inlet Profile on Heat Transfer and Fluid Flow in Turbine Stages

2011 ◽  
Vol 134 (1) ◽  
Author(s):  
Salvadori Simone ◽  
Francesco Montomoli ◽  
Francesco Martelli ◽  
Kam S. Chana ◽  
Imran Qureshi ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Temperature Profile ◽  
Thermal Load ◽  
Thermal Loading ◽  
Radial Profile ◽  
Experimental Studies ◽  
Test Facility ◽  
Inlet Temperature ◽  
Inlet Conditions ◽  
Hot Streak

This paper presents an investigation of the aerothermal performance of a modern unshrouded high-pressure (HP) aero-engine turbine subject to nonuniform inlet temperature profile. The turbine used for this study was the MT1 turbine installed in the QinetiQ turbine test facility based in Farnborough (UK). The MT1 turbine is a full scale transonic HP turbine, and is operated in the test facility at the correct nondimensional conditions for aerodynamics and heat transfer. Datum experiments of aerothermal performance were conducted with uniform inlet conditions. Experiments with nonuniform inlet temperature were conducted with a temperature profile that had a nonuniformity in the radial direction defined by (Tmax−Tmin)/T¯=0.355, and a nonuniformity in the circumferential direction defined by (Tmax−Tmin)/T¯=0.14. This corresponds to an extreme point in the engine cycle, in an engine where the nonuniformity is dominated by the radial distribution. Accurate experimental area surveys of the turbine inlet and exit flows were conducted, and detailed heat transfer measurements were obtained on the blade surfaces and end-walls. These results are analyzed with the unsteady numerical data obtained using the in-house HybFlow code developed at the University of Firenze. Two particular aspects are highlighted in the discussion: prediction confidence for state of the art computational fluid dynamics (CFD) and impact of real conditions on stator-rotor thermal loading. The efficiency value obtained with the numerical analysis is compared with the experimental data and a 0.8% difference is found and discussed. A study of the flow field influence on the blade thermal load has also been detailed. It is shown that the hot streak migration mainly affects the rotor pressure side from 20% to 70% of the span, where the Nusselt number increases by a factor of 60% with respect to the uniform case. Furthermore, in this work, it has been found that a nonuniform temperature distribution is beneficial for the rotor tip, contrary to the results found in open literature. Although the hot streak is affected by the pressure gradient across the tip gap, the radial profile (which dominates the temperature profile being considered) is not fully mixed out in passing through the HP stage, and contributes significantly to cooling the turbine casing. A design approach not taking into account these effects will underestimate the rotor life near the tip and the thermal load at midspan. The temperature profile that has been used in both experiments and CFD is the first simulation of an extreme cycle point (more than twice the magnitude of distortion of all previous experimental studies): It represents an engine-take-off condition combined with the full combustor cooling. This research was part of the EU funded Turbine AeroThermal External Flows 2 program.

Developing a Combustor Simulator for Investigating High Pressure Turbine Aerodynamics and Heat Transfer

2004 ◽  
Author(s):  
M. D. Barringer ◽  
K. A. Thole ◽  
M. D. Polanka
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Film Cooling ◽  
Pressure Ratio ◽  
Test Facility ◽  
Temperature Profiles ◽  
High Pressure Turbine ◽  
Blow Down ◽  
Turbine Engine ◽  
The Impact

Within a gas turbine engine, the high pressure turbine vanes are subjected to very harsh conditions from the highly turbulent and hot gases exiting the combustor. The temperature and pressure fields exiting the combustor dictate the heat transfer and aero losses that occur in the turbine passages. To better understand these effects, the goal of this work is to develop an adjustable combustor exit profile simulator for the Turbine Research Facility (TRF) at the Air Force Research Laboratory (AFRL). The TRF is a high temperature, high pressure, short duration blow-down test facility that is capable of matching several aerodynamic and thermal non-dimensional engine parameters including Reynolds number, Mach number, pressure ratio, corrected mass flow, gas-to-metal temperature ratio, and corrected speed. The research objective was to design, install, and verify a non-reacting simulator device that provides representative combustor exit total pressure and temperature profiles to the inlet of the TRF turbine test section. This required the upstream section of the facility to be redesigned into multiple concentric annuli that serve the purpose of injecting high momentum dilution jets and low momentum film cooling jets into a central annular chamber, similar to a turbine engine combustor. The design of the simulator allows for variations in injection levels to generate turbulence and pressure profiles. It also can vary the dilution and film cooling temperatures to create a variety of temperature profiles consistent with real combustors. To date, the design and construction of the simulator device has been completed. All of the hardware has been trial fitted and the flow control shutter systems have been successfully installed and tested. Currently, verification testing is being performed to investigate the impact of the generated temperature, pressure, and turbulence profiles on turbine heat transfer and secondary flow development.

The Effect of Hot-Streaks on HP Vane Surface and Endwall Heat Transfer: An Experimental and Numerical Study

2005 ◽  
Author(s):  
T. Povey ◽  
K. S. Chana ◽  
T. V. Jones ◽  
J. Hurrion
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Rate ◽  
Transfer Rate ◽  
Cooling System ◽  
Numerical Study ◽  
Inlet Temperature ◽  
Temperature Profiles ◽  
Wall Heat Transfer ◽  
Hot Streak ◽  
End Wall

Pronounced non-uniformities in combustor exit flow temperature (hot-streaks), which arise because of discrete injection of fuel and dilution air jets within the combustor and because of end-wall cooling flows, affect both component life and aerodynamics. Because it is very difficult to quantitatively predict the affects of these temperature non-uniformities on the heat transfer rates, designers are forced to budget for hot-streaks in the cooling system design process. Consequently, components are designed for higher working temperatures than the mass-mean gas temperature, and this imposes a significant overall performance penalty. An inadequate cooling budget can lead to reduced component life. An improved understanding of hot-streak migration physics, or robust correlations based on reliable experimental data, would help designers minimise the overhead on cooling flow that is currently a necessity. A number of recent research projects sponsored by a range of industrial gas turbine and aero-engine manufacturers attest to the growing interest in hot-streak physics. This paper presents measurements of surface and end-wall heat transfer rate for an HP nozzle guide vane (NGV) operating as part of a full HP turbine stage in an annular transonic rotating turbine facility. Measurements were conducted with both uniform stage inlet temperature and with two non-uniform temperature profiles. The temperature profiles were non-dimensionally similar to profiles measured in an engine. A difference of one half of an NGV pitch in the circumferential (clocking) position of the hot-streak with respect to the NGV was used to investigate the affect of clocking on the vane surface and end-wall heat transfer rate. The vane surface pressure distributions, and the results of a flow-visualisation study, which are also given, are used to aid interpretation of the results. The results are compared to two-dimensional predictions conducted using two different boundary layer methods. Experiments were conducted in the Isentropic Light Piston Facility (ILPF) at QinetiQ Farnborough, a short duration engine-size turbine facility. Mach number, Reynolds number and gas-to-wall temperature ratios were correctly modelled. It is believed that the heat transfer measurements presented in this paper are the first of their kind.

Passive Shroud Cooling Concepts for HP Turbines: Experimental Investigations

2006 ◽  
Author(s):  
Vasudevan Kanjirakkad ◽  
Richard Thomas ◽  
Howard Hodson ◽  
Erik Janke ◽  
Frank Haselbach ◽  
...  
Keyword(s):  
Film Cooling ◽  
Alternative Methods ◽  
Passive Cooling ◽  
Experimental Investigations ◽  
Inlet Temperature ◽  
Temperature Profiles ◽  
High Pressure Turbine ◽  
Turbine Inlet Temperature ◽  
Critical Areas ◽  
Cooling Method

The cooling of rotor shrouds in the first stage of a high-pressure turbine requires special attention as flatter turbine inlet temperature profiles and more highly loaded blades result in increased thermal and mechanical stresses. The use of film cooling and/or internal convective cooling makes the rotor shroud heavier and oversized, restricting the maximum rotational speed. Alternative methods are therefore sought to achieve improved cooling of the shroud. This paper discusses the low speed experimental investigation of two ‘passive’ cooling concepts known as ‘rail cooling’ and ‘platform cooling’. It has been shown experimentally that the modified cooling method, namely the platform cooling, substantially improves the rotor shroud coolant distribution in the critical areas whilst employing significantly lower amounts of coolant.

Numerical Simulation of Temperature Profile in a Compartment Fire with a Roof Opening

2012 ◽  
Vol 538-541 ◽  
pp. 989-992 ◽  
Author(s):  
Jin Mei Li ◽  
Qiang Li ◽  
Yan Lei Dong ◽  
Chang Hai Li
Keyword(s):  
Numerical Simulation ◽  
Numerical Simulations ◽  
Temperature Profile ◽  
Source Size ◽  
Temperature Profiles ◽  
Compartment Fire ◽  
Gas Temperature ◽  
Fire Source ◽  
Smoke Layer ◽  
Cooling Function

Fifteen numerical simulations are presented in this article to investigate the influence of roof opening size and fire source size on gas temperature profiles in a compartment. The fire source size has a significant impact on the temperature hot smoke layer. The temperature of hot smoke layer increases as the increase of fire source size. The roof opening has cooling function to gas temperature in the compartment especially for large roof opening. The temperatures of hot smoke layer decrease with the roof opening size increase in all cases.

scholarly journals Effect of Inlet Temperature Non-Uniformity on High-Pressure Turbine Performance

2010 ◽  
Author(s):  
Craig I. Smith ◽  
Dongil Chang ◽  
Stavros Tavoularis
Keyword(s):  
High Pressure ◽  
Heat Load ◽  
Stokes Equations ◽  
Computation Time ◽  
Radial Pressure ◽  
Navier Stokes ◽  
Inlet Temperature ◽  
High Pressure Turbine ◽  
Navier Stokes Equations ◽  
Hot Streak

The temperature of the flow entering a high-pressure turbine stage is inherently non-uniform, as it is produced by several discrete, azimuthally-distributed combustors. In general, however, industrial simulations assume inlet temperature uniformity to simplify the preparation process and reduce computation time. The effects of a non-uniform inlet field on the performance of a commercial, transonic, single-stage, high-pressure, axial turbine with a curved inlet duct have been investigated numerically by performing URANS (Unsteady Reynolds-Averaged Navier-Stokes equations) simulations with the SST (Shear Stress Transport) turbulence model. By adjusting the alignment of the experimentally-based inlet temperature field with respect to the stator vanes, two clocking configurations were generated: an aligned case, in which each hot streak impinged on a vane and a misaligned case, in which each hot streak passed between two vanes. In the aligned configuration, the hot streaks produced higher time-averaged heat load on the vanes and lower heat load on the blades. As the aligned hot streaks impinged on the stator vanes, they also spread spanwise due to the actions of the casing passage vortices and the radial pressure gradient; this resulted in a stream entering the rotor with relatively low temperature variations. The misaligned hot streaks were convected undisturbed past the relatively cool vane section. Relatively high time-averaged enthalpy values were found to occur on the pressure side of the blades in the misaligned configuration. The non-uniformity of the time-averaged enthalpy on the blade surfaces was lower in the aligned configuration. The flow exiting the rotor section was much less non-uniform in the aligned case, but differences in calculated efficiency were not significant.

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