This paper describes a new engine-parts facility at the University of Oxford for high technology-readiness-level research, new technology demonstration, and for engine component validation. The Engine Component AeroThermal (ECAT) facility has a modular working section which houses a full annulus of engine components. The facility is currently operated with high-pressure nozzle guide vanes from a large civil jet-engine. A high degree of engine similarity is achieved, with matched conditions of Mach number, Reynolds number, and coolant-to-mainstream pressure ratio. For combustor-turbine interaction studies, a combustor simulator module is used, which is capable of both rich-burn and lean-burn combined temperature, swirl and turbulence profiles. The facility is being used for aerothermal optimisation research (e.g., novel cooling systems, aerodynamic optimisation problems, capacity sensitivity studies), computational fluid dynamics validation (aerodynamic predictions, conjugate predictions), and for component validation to accelerate the engine design process. The three key measurement capabilities are: capacity characteristic evaluation to a precision of 0.02%; overall cooling (metal) effectiveness measurements (using a rainbow set of parts if required); and aerodynamic loss evaluation (with realistic cooling, trailing-edge flow etc.). Each of these three capabilities have been separately developed and optimised in other facilities at the University of Oxford in the last 10 years, to refine aspects of facility design, instrumentation design, experimental technique, and theoretical aspects of scaling and reduction of experimental data. The ECAT facility brings together these three research strands with a modular test vehicle for rapid high technology-readiness-level research, demonstration of new technologies, and for engine component validation. The purpose of this paper is to collect in one place — and put in context — the work that led to the development of the ECAT facility, to describe the facility, and to illustrate the accuracy and utility of the techniques by presenting typical data for each of the key measurements. The ECAT facility is a response to the changing requirements of experimental turbomachinery testing, and it is hoped this paper will be of interest to engine designers, researchers, and those involved in major facility developments in both research institutes and engine companies.

This paper describes a new modular experimental facility that was purpose-built to investigate flow interactions between the combustor and first stage nozzle guide vanes (NGVs) of heavy duty power generation gas turbines with multiple can combustors. The first stage turbine NGV is subjected to the highest thermal loads of all turbine components and therefore consumes a proportionally large amount of cooling air that contributes detrimentally to the stage and cycle efficiency. It has become necessary to devise novel cooling concepts that can substantially reduce the coolant air requirement but still allow the turbine to maintain its aerothermal performance. The present work aims to aid this objective by the design and commissioning of a high-speed linear cascade, which consists of two can combustor transition ducts and four first stage NGVs. This is a modular nonreactive air test platform with engine realistic geometries (gas path and near gas path), cooling system, and boundary conditions (inlet swirl, turbulence level, and boundary layer). The paper presents the various design aspects of the high pressure (HP) blow down type facility, and the initial results from a wide range of aerodynamic and heat transfer measurements under highly engine realistic conditions.

The computational and experimental assessment of a lean-burn low-NOx combustor simulator for an engine component test facility is presented. The Engine Component Aero-Thermal (ECAT) facility is a full-scale engine-parts facility, designed for the study of the aero-thermal performance of fully cooled high-pressure nozzle guide vanes (NGVs). The facility operates with non-dimensionally matched engine conditions in terms of Reynolds number, Mach number and coolant-to-mainstream pressure ratio. The combustor simulator is designed to replicate lean-burn conditions of swirl and temperature distortion upstream of the nozzle guide vanes. The purpose is to allow the study of flow capacity, aerodynamic performance (with film cooling), and thermal performance (overall effectiveness) in the presence of engine-realistic inlet distortions. Detailed experimental measurements with multi-hole probes and thermocouples are presented and compared to results from RANS Simulations. Additional simulations were performed to understand how the elevated back pressure and vane potential field affect the non-dimensional profiles of pressure loss, residual swirl and temperature at the combustor-turbine interface. This is perhaps the most comprehensive study to date of a combustor simulator in an engine-scale research facility, providing unique insight into the known challenges of simulator design, scaling issues when moving from low to high Reynolds number, and limitations of CFD in this flow environment.

This paper presents a novel transient method for calibrating heat transfer gauges for convective wall heat flux measurements in high enthalpy flows. The new method relies on the transient heating of sensor substrates under rapid exposure to a hot flow in order to obtain the necessary reference heat flux. Compared to previous calibration facilities, the new facility is simple, inexpensive, easy to adapt for different flow configurations and sensor geometries, and quick to run across a wide range of conditions. In this paper, the design of the new calibration facility is described and the transient calibration method explained. The method is demonstrated by calibrating a Gardon gauge at convective heat transfer coefficients between 300 and 600 W/m2 K. Typical facility data and calibration results are presented in terms of voltage-heat flux sensitivity and calibration correction ratio. These results are shown to agree with theoretical estimates within the estimated calibration uncertainty.

This paper concerns the development and testing of a novel high frequency pressure probe for high temperature turbomachinery flow measurements. The probe has a measurement bandwidth from DC to 100 kHz and has been demonstrated in conditions of up to 6 bar and 1900 K. The 4 mm diameter probes are uncooled whilst in the flow and employ a fast insertion traverse to limit immersion times to of the order of 0.1 seconds. The probe was calibrated against a hot wire anemometer in a known turbulent flow. Data acquired downstream of the turbine in a turbojet engine is presented. The unsteady pressure data is decomposed into periodic and random components. Power spectra, turbulence intensity and length scale are derived. Short duration gas turbine measurements using fast insertion techniques have been under development for some years at the University of Oxford. The current fast-insertion probe is more compact and robust than previous designs. The present work demonstrates that it can resolve useful flow parameters in hostile gas turbine flows. These can be difficult or impossible to obtain using other methods. The rapid probe insertion technique should add to the armoury of diagnostic tools used by the gas turbine developer.