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.© 2005 ASME

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

Modern lean burn combustors now employ aggressive swirlers to enhance fuel-air mixing and improve flame stability The flow at combustor exit can therefore have high residual swirl A good deal of research concerning the flow within the combustor is available in open literature The impact of swirl on the aerodynamic and heat transfer characteristics of a HP turbine stage is not well understood, however A combustor swirl simulator has been designed and commissioned in the Oxford Turbine Research Facility (OTRF), previously located at QinetiQ, Farnborough UK The swirl simulator is capable of generating an engine-representative combustor exit swirl pattern At the turbine inlet plane, yaw and pitch angles of over +/-40 degrees have been simulated The turbine research facility used for the study is an engine scale, short duration, rotating transonic turbine, in which the non-dimensional parameters for aerodynamics and heat transfer are matched to engine conditions The research turbine was the unshrouded MT1 design By design, the centre of the vortex from the swirl simulator can be clocked to any circumferential position with respect to HP vane, and the vortex-to-vane count ratio is 1:2 For the current investigation, the clocking position was such that the vortex centre was aligned with the vane leading edge (every second vane) Both the aligned vane and the adjacent vane were characterised This paper presents measurements of HP vane surface and endwall heat transfer for the two vane positions The results are compared with measurements conducted without swirl The vane surface pressure distributions are also presented The experimental measurements are compared with full-stage three-dimensional unsteady numerical predictions obtained using the Rolls Royce in-house code Hydra The aerodynamic and heat transfer characterisation presented in this paper is the first of its kind, and it is hoped to give some insight into the significant changes in the vane flow and heat transfer that occur in the current generation of low NOx combustors The findings not only have implications for the vane aerodynamic design, but also for the cooling system designCopyright © 2011 by Rolls-Royce plc

HP Vane Aerodynamics and Heat Transfer in the Presence of Aggressive Inlet Swirl

Swirling flows are now widely being used in modern gas turbine combustors to improve the combustion characteristics, flame stability and reduce emissions. Residual swirl at combustor exit will affect the performance of the downstream high-pressure (HP) turbine. In order to perform a detailed investigation of the effect of swirl on a full-scale HP turbine stage, a combustor swirl simulator has been designed and commissioned in the Oxford Turbine Research Facility (OTRF), previously located at QinetiQ, Farnborough UK, as the Turbine Test Facility (TTF). The swirl simulator is capable of generating an engine-representative combustor exit swirl distributions at the turbine inlet, with yaw and pitch angles of up to +/-40 degrees. The turbine test facility is an engine scale, short duration, rotating transonic turbine facility, which simulates engine representative M, Re, Tu, non-dimensional speed and gas-to-wall temperature ratio at the turbine inlet. The test turbine is a highly loaded unshrouded design (the MT1 turbine). This paper presents time-averaged experimental heat transfer measurements performed on the rotor casing surface, and on rotor blade surface at 10%, 50% and 90% span. Time-averaged rotor casing static pressure measurements are also presented. Experimental measurements with and without inlet swirl are compared. The measurements are discussed with the aid of three-dimensional steady and unsteady CFD simulations of the turbine stage. Numerical simulations were conducted using the Rolls-Royce in-house code HYDRA, with and without inlet swirl.Copyright © 2011 by Rolls-Royce plc

Effect of Aggressive Inlet Swirl on Heat Transfer and Aerodynamics in an Unshrouded Transonic HP Turbine

This paper presents experimental measurements and computational predictions of surface and end wall heat transfer for a high-pressure (HP) nozzle guide vane operating as part of a full HP turbine stage in an annular rotating turbine facility, with and without inlet temperature distortion (hot streaks). A detailed aerodynamic survey of the vane surface is also presented. The test turbine was the unshrouded MT1 turbine, installed in the Turbine Test Facility (previously called Isentropic Light Piston Facility) at QinetiQ, Farnborough, UK. This is a short-duration facility, which simulates engine-representative M, Re, nondimensional speed, and gas-to-wall temperature ratio at the turbine inlet. The facility has recently been upgraded to incorporate an advanced second-generation combustor simulator, capable of simulating well-defined, aggressive temperature profiles in both the radial and circumferential directions. This work forms part of the pan-European research program, TATEF II. Measurements of HP vane and end wall heat transfer obtained with inlet temperature distortion are compared with results for uniform inlet conditions. Steady and unsteady computational fluid dynamics (CFD) predictions have also been conducted on vane and end wall surfaces using the Rolls-Royce CFD code HYDRA to complement the analysis of experimental results. The heat transfer measurements presented in this paper are the first of their kind in that the temperature distortion is representative of an extreme cycle point, and was simulated with good periodicity and with well-defined boundary conditions in the test turbine.

Effect of Simulated Combustor Temperature Nonuniformity on HP Vane and End Wall Heat Transfer: An Experimental and Computational Investigation

In modern gas turbine engines there exist significant temperature gradients in the combustor exit flow. These gradients arise because both fuel and dilution air are introduced within the combustor as discrete jets. The effects of this non-uniform temperature field on the aerodynamics and heat transfer rate distributions of nozzle guide vanes and turbine blades is difficult to predict, although an increased understanding of the effects of temperature gradients would enhance the accuracy of estimates of turbine component life and efficiency. Low-frequency measurements of heat transfer rate have been conducted on an annular transonic intermediate-pressure (IP) nozzle guide vane operating downstream of a high-pressure (HP) rotating turbine stage. Measurements were conducted with both uniform and non-uniform inlet temperature profiles. The non-uniform temperature profile included both radial and circumferential gradients of temperature. Experiments were conducted in the isentropic light piston facility at QinetiQ Pyestock, a short-duration engine-size turbine facility with 1.5 turbine stages, in which Mach number, Reynolds number and gas-wall temperature ratios are correctly modelled. Experimental heat transfer results are compared with predictions performed using boundary layer methods.

http://journals.sagepub.com/doi/pdf/10.1243/095765003322315487

Heat transfer measurements on an intermediate-pressure nozzle guide vane tested in a rotating annular turbine facility, and the modifying effects of a non-uniform inlet temperature profile

This paper describes the results of systematic investigations undertaken to study the effect of free-stream turbulence on the heat transfer coefficient distribution around gas turbine rotor blades and nozzle guide vanes. The heat transfer coefficient distribution around the blade surface was obtained under a uniform heat flux boundary condition. Experiments were conducted in the Reynolds number range 2.0-8.1 x 10 to the 5th (exit Mach number range 0.182 to 0.600) with the free-stream turbulence level in the range 1.0-21.3 percent. A new type of active turbulence generator was used for generating high turbulence levels. Correlations were obtained for the effect of free-stream turbulence on the local heat transfer coefficient in the laminar, transitional, and turbulent boundary layer regions.

V. Krishnamoorthy

Papers

The Effect of Free-Stream Turbulence on Gas Turbine Blade Heat Transfer