Isentropic Light Piston Facility Research Articles

Effect of Temperature Nonuniformity on Heat Transfer in an Unshrouded Transonic HP Turbine: An Experimental and Computational Investigation

Detailed experimental measurements have been performed to understand the effects of turbine inlet temperature distortion (hot-streaks) on the heat transfer and aerodynamic characteristics of a full-scale unshrouded high pressure turbine stage at flow conditions that are representative of those found in a modern gas turbine engine. To investigate hot-streak migration, the experimental measurements are complemented by three-dimensional steady and unsteady CFD simulations of the turbine stage. This paper presents the time-averaged measurements and computational predictions of rotor blade surface and rotor casing heat transfer. Experimental measurements obtained with and without inlet temperature distortion are compared. Time-mean experimental measurements of rotor casing static pressure are also presented. CFD simulations have been conducted using the Rolls-Royce code HYDRA and are compared with the experimental results. 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 transonic facility, which simulates engine-representative M, Re, Tu, N/T, and Tg/Tw to the turbine inlet. The facility has recently been upgraded to incorporate an advanced second-generation temperature distortion generator, capable of simulating well-defined, aggressive temperature distortion both in the radial and circumferential directions, at the turbine inlet.

Direct shaft torque measurements in a transient turbine facility

This paper describes the development and implementation of a shaft torque measurement system for the Oxford Turbine Research Facility (formerly the Turbine Test Facility (TTF) at QinetiQ, Farnborough), or OTRF. As part of the recent EU TATEF II programme, the facility was upgraded to allow turbine efficiency measurements to be performed. A shaft torque measurement system was developed as part of this upgrade. The system is unique in that, to the authors’ knowledge, it provided the first direct measurement of shaft torque in a transient turbine facility although the system has wider applicability to rotating test facilities in which power measurement is a requirement. The adopted approach removes the requirement to quantify bearing friction, which can be difficult to accurately calibrate under representative operating conditions. The OTRF is a short duration (approximately 0.4 s run time) isentropic light-piston facility capable of matching all of the non-dimensional parameters important for aerodynamic and heat studies, namely Mach number, Reynolds number, non-dimensional speed, stage pressure ratio and gas-to-wall temperature ratio. The single-stage MT1 turbine used for this study is a highly loaded unshrouded design, and as such is relevant to modern military, or future civil aero-engine design. Shaft torque was measured directly using a custom-built strain gauge-based torque measurement system in the rotating frame of reference. This paper describes the development of this measurement system. The system was calibrated, including the effects of temperature, to a traceable primary standard using a purpose-built facility. The bias and precision uncertainties of the measured torque were ±0.117% and ±0.183%, respectively. To accurately determine the shaft torque developed by a turbine in the OTRF, small corrections due to inertial torque (associated with changes in the rotational speed) and aerodynamic drag (windage) are required. The methods for performing these corrections are described.

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

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.

A hot-streak (combustor) simulator suited to aerodynamic performance measurements

A new hot-streak (combustor) simulator has been designed and implemented in a turbine test facility at QinetiQ Farnborough (the QinetiQ Isentropic Light Piston Facility, ILPF) to study the impact of temperature distortion on high pressure (HP) turbine efficiency, aerodynamics, and heat transfer. The ILPF is an engine scale, short duration, rotating transonic turbine test facility, in which M, Re, Tg/ Tw, and The hot-streak simulator is a second-generation design, in which cold gas is introduced into a hot mainstream though radial and circumferential slots upstream of the turbine stage. The simulator is rotatable, so the effect of clocking (relative circumferential position of hot streak and nozzle guide vane leading edge) can easily be investigated. An emphasis was placed on accurate measurement of turbine inlet enthalpy flux so that the impact of hot streaks on turbine efficiency could be investigated. The hot-streak simulator differs from all previous systems in that a pronounced radial and circumferential temperature profile has been generated, with a hot-streak to vane count of 1:1. The profile is very well matched (non-dimensionally) to the target profile, which is a combustor temperature profile measured in a modern operating engine at the most extreme point in the cycle. The most accurate area survey of a simulated temperature profile has been conducted to date, and this demonstrates that the simulator offers an exceptionally high degree of circumferential symmetry and run-to-run repeatability. The design and commissioning of the simulator is described, and the measured temperature profiles are compared with the target profile.

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

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.

On a Novel Annular Sector Cascade Technique

An advanced technique for establishing pressure boundary conditions in annular sector cascade experiments has been developed. This novel technique represents an improvement over previous methods and provides the first means by which annular sector boundary conditions that are representative of those which develop in an annular cascade can be established with a high degree of satisfaction. The technique will enable cascade designers to exploit the obvious advantages of annular sector cascade testing: the reduced cost of both facility manufacture and facility operation and the use of engine parts in place of two-dimensional counterparts. By employing an annular sector of deswirl vanes downstream of the annular sector of test vanes, the radial pressure gradient established in the swirling flow downstream of the test vanes is not disturbed. The deswirl vane exit flow—which has zero swirl velocity—can be exhausted without unsteadiness, and without the risk of separation, into a plenum at constant pressure. The pressure ratio across the annular sector of test vanes can be tuned by adjusting the throat area at the deswirl vane exit plane. Flow conditioning systems which utilize the Oxford deswirl vane technology have previously been used to set pressure boundary conditions downstream of fully annular cascades in both model and engine scale (the Isentropic Light Piston Facility at Farnborough) experimental research facilities (Povey, T., Chana, K. S., Oldfield, M. L. G., Jones, T. V., and Owen, A. K., 2001, Proceedings of the ImechE Advances in Fluid Machinery Design Seminar, London, June 13; Povey, T., Chana, K. S., Jones, T. V., and Oldfield, M. L. G., 2003, Advances of CFD in Fluid Machinery Design, ImechE Professional Engineering, London, pp. 65–94). The deswirl vane is particularly suited to the control of highly whirling transonic flows. It has been demonstrated by direct comparison of aerodynamic measurements from fully annular and annular sector experiments that the use of a deswirl vane sector for flow conditioning at the exit of an annular sector cascade represents an attractive novel solution to the boundary condition problem. The annular sector technique is now described.

An Investigation on Turbine Tip and Shroud Heat Transfer

Detailed experimental investigations have been performed to measure the heat transfer and static pressure distributions on the rotor tip and rotor casing of a gas turbine stage with a shroudless rotor blade. The turbine stage was a modern high pressure Rolls-Royce aero-engine design with stage pressure ratio of 3.2 and nozzle guide vane (ngv) Reynolds number of 2.54E6. Measurements have been taken with and without inlet temperature distortion to the stage. The measurements were taken in the QinetiQ Isentropic Light Piston Facility and aerodynamic and heat transfer measurements are presented from the rotor tip and casing region. A simple two-dimensional model is presented to estimate the heat transfer rate to the rotor tip and casing region as a function of Reynolds number along the gap.

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

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.

Measurement and Calculation of Nozzle Guide Vane End Wall Heat Transfer

A three-dimensional steady viscous finite volume pressure correction method for the solution of the Reynolds-averaged Navier–Stokes equations has been used to calculate the heat transfer rates on the end walls of a modern High Pressure Turbine first-stage stator. Surface heat transfer rates have been calculated at three conditions and compared with measurements made on a model of the vane tested in annular cascade in the Isentropic Light Piston Facility at DERA, Pyestock. The NGV Mach numbers, Reynolds numbers, and geometry are fully representative of engine conditions. Design condition data have previously been presented by Harvey and Jones (1990). Off-design data are presented here for the first time. In the areas of highest heat transfer, the calculated heat transfer rates are shown to be within 20 percent of the measured values at all three conditions. Particular emphasis is placed on the use of wall functions in the calculations with which relatively coarse grids (of around 140,000 nodes) can be used to keep computational run times sufficiently low for engine design purposes.

Heat Transfer Predictions for Two Turbine Nozzle Geometries at High Reynolds and Mach Numbers

Predictions of turbine vane and endwall heat transfer and pressure distributions are compared with experimental measurements for two vane geometries. The differences in geometries were due to differences in the hub profile, and both geometries were derived from the design of a high rim speed turbine (HRST). The experiments were conducted in the Isentropic Light Piston Facility (ILPF) at Pyestock at a Reynolds number of 5.3 x 106, a Mach number of 1.2, and a wall-to-gas temperature ratio of 0.66. Predictions are given for two different steady-state three-dimensional Navier–Stokes computational analyses. C-type meshes were used, and algebraic models were employed to calculate the turbulent eddy viscosity. The effects of different turbulence modeling assumptions on the predicted results are examined. Comparisons are also given between predicted and measured total pressure distributions behind the vane. The combination of realistic engine geometries and flow conditions proved to be quite demanding in terms of the convergence of the CFD solutions. An appropriate method of grid generation, which resulted in consistently converged CFD solutions, was identified.