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The latest gas turbine development being undertaken by the Royal Navy is the ACL technology demonstrator programme. The aim of the project is to design and demonstrate a GTA of around 2MW that will have reduced whole life costs when compared to a conventional diesel generator. The project is being managed for the MoD by the Marine Propulsion Systems IPT and being procured through the prime contractor - Turbomeca. The prime contract was awarded in December 2000 and is scheduled to run for five years. This paper describes the main aims and objectives of the project. The paper also gives a detailed overview of the GTA system, describes the key technologies and discusses the outline development programme.

An efficient numerical method is applied to investigate the influence of different tip gap shapes on the performance of an axial-flow compressor stator under the condition of rotating hub or stationary hub. Tip gap shapes considered in this paper include zero gap, uniform gap, and linearly varying gap. The analysis of the results shows, at least for these test cases, tip leakage vortex can be found to generate earlier for stationary hub cases than for rotating hub cases. On the same working conditions, the blade row with zero gap has not the highest efficiency, while the blade row with smaller tip gap can attain higher efficiency due to the role of tip leakage flow the direction of which is opposite to the secondary flow in the passage near the hub. For the same circumferential leakage area in the tip region, the blade row with linearly expanding gap can obtain higher efficiency. The work of this paper can provide a useful reference for optimum design of compressor stator.Copyright © 2001 by ASME

This paper is Part III of a comprehensive study on the blade leading edge sweep/bend of supersonic and transonic axial compressors. For the curved leading edge shock surfaces generated by high loading swept axial flow rotors, using an analytical structure model of 3-D shock surface, this paper investigated the characteristics of the shock surface structure, then calculated the distributions of the flow variables in front of and behind the shock. The research focused on the influences of the relatively forward- or backward-swept rotors on the flow field behind but near the leading edge shock surface. By the analytical model, this paper answered the one hand of the question of the swept rotors’ functions. For the first time it quantitatively showed the existence of the phenomenon of 3-D shock induced span-wise and circumferential mixing in the primary flow.Copyright © 2001 by ASME

This paper describes the design and construction of a new two stage axial turbine test facility, christened “Lisa”. The research objective of the rig is to study the impact (relevance) of unsteady flow phenomena upon the aerodynamic performance, this being achieved through the use of systematic studies of parametric changes in the stage geometry and operating point. Noteworthy in the design of the rig is the use of a twin shaft arrangement to decouple the stages. The inner shaft carries the load from the first stage whilst the outer is used with an integral torque-meter to measure the loading upon the second stage alone. This gives an accurate measurement of the loading upon the aerodynamically representative second stage, which possesses the correct stage inlet conditions in comparison to the full two stage machine which has an unrealistic axial inlet flow at the first stator. A calibrated Venturi nozzle measures the mass flow at an accuracy of below 1%, from which stage efficiencies can be derived. The rig is arranged in a closed loop system. The turbine has a vertical arrangement and is connected through a gear box to a generator system that works as a brake to maintain the desired operating speed. The turbine exit is open to ambient pressure. The rig runs at a low pressure ratio of 1.5. The maximum Mach number at stator exit is 0.3 at an inlet pressure of 1.5 bar. The maximum mass flow is 14 kg/sec. Nominal rotor design speed is 3000 RPM. The tip to hub blade ratio is 1.29, and the nominal axial chord is 50 mm. The rig is designed to accommodate a broad range of measurement techniques, but with a strong emphasis upon unsteady flow methods, for example fast response aerodynamic pressure probes for time-resolved flow measurements. The first section of this paper describes the overall test facility hardware. This is followed by a detailed focus on the torque measurement device including stage efficiency measurements at operating conditions in Lisa. Discussion of measurement techniques completes the paper.

Nowadays there is a big effort toward improving the low pressure turbine efficiency even to the extent of penalising other relevant design parameters. LP turbine efficiency influences SFC more than other modules in the engine. Most of the research has been oriented to reduce profile losses, modifying the suction surface, the pressure surface or the three-dimensional regions of the flow. To date, the pressure surface has received very little attention. The dependence of the profile losses on the behaviour of both pressure and suction surfaces has been investigated for the case of a high lift design that is representative of a modern civil engine LP turbine. The experimental work described in this paper consists on two different sets of experiments: the first one concluded an improved pressure surface definition and the second set was oriented to achieve further improvement in losses modifying the profile suction surface. Three profiles were designed and tested over a range of conditions. The first profile is a thin-solid design. This profile has a large pressure side separation bubble extending from near the leading edge to mid-chord. The second profile is a hollow design with the same suction surface as the first one but avoiding pressure surface separation. The third one is also a hollow design with the same pressure surface as the second profile but more aft loaded suction surface. The study is part of a wider on-going research programme covering the effects of the different design parameters on losses. The paper describes the experiments conducted in a lowspeed linear cascade facility. It gathers together steady and unsteady loss measurements by wake traverse and surface pressure distributions for all the profiles. It is shown that thick profiles generates only around 90% of the losses of a thinsolid profile with the same suction surface. The results support the idea of an optimum position axial position for the peak Mach number. Caution is recommended as profile aft loading would not be a completely secure method for reducing losses. NOMENCLATURE α1 Inlet flow angle. α2 Outlet flow angle. Cax Axial chord. I Incidence.

The Royal Navy (RN) has operated marinised industrial and aero-derivative propulsion gas turbines since the late 1940’s. In order to safeguard the gas turbines (principally from foreign object damage (FOD) and salt ingestion) the RN places a high degree of importance on the gas turbine intake system, a principal element within which is the intake filtration equipment. Since the introduction of marine gas turbines the RN have witnessed and participated in the development of intake filtration systems from knitmesh filters to highly efficient 3 stage (vane- coalescer-vane) separators (spray eliminators). The requirements for RN gas turbine intake systems are described in Naval Engineering Standard (NES 312), and have been brought about by long standing operational experience of these systems and commercial best practice. This paper will briefly outline the RN design criteria for gas turbine intake systems and how this has been modified by field experience. It will then go on to look at intake separators, giving their design criteria and describe from inception, a study into the possible adoption of next generation high velocity spray eliminators describing initial specification, the product design process and development testing. The potential benefits afforded by such a high velocity system will also be discussed. In addition the paper will describe the initial fit of a trial unit onto HMS Coventry, for back to back testing with a unit of conventional design. It is intended to present Part 2 of this paper at ASME 2002 Amsterdam, which will contain a précis of the design, presentation of the data from the sea trial and the results and conclusions from the engine inspections, including an assessment by Rolls Royce the engine Design Authority. At this time the sutability of the equipment for the intakes of the WR21 will be considered.

In this second part of a two-part paper on the application of the NASA multistage aerodynamic simulation CFD tool, the APNASA code, work is presented on how the code was used successfully in the design of a brand new, four-stage axial compressor aimed for a modern turbofan engine. In particular, the code was used to guide the blade geometry changes such that the right stage matching throughout the compressor was achieved. The overall performance is shown to be in good agreement with the map generated prior to the detailed design from a one-dimensional model, including the surge line. This success is attributed to the fact that the compressor has been successfully matched at its design intention. More detailed comparisons of the measurements with the prediction demonstrate capability of APNASA in capturing details of flow, such as the rotor exit pressure and temperature profiles and the stator exit flow angle. Largely as a result of the application of APNASA, the compressor described in this paper has been designed “right the first time”, resulting in significant cost savings in a new engine development program for Honeywell Engines and Systems.Copyright © 2001 by ASME

Aerodynamic investigations in a cold flow annular sector high-pressure turbine cascade test facility and an annular sector cascade facility operating at near-engine conditions are presented. The test section of both facilities is a 36° sector cascade of a modern turbine stator consisting of 6 vanes. The two facilities have been designed in order to gain detailed information concerning film cooled gas turbine vanes. Due to the operation conditions of the hot annular sector cascade it takes over the part of detailed investigations of the influence of film cooling on the heat transfer. In the cold annular sector cascade facility investigations on the aerodynamic behavior of the cascade are performed. Both facilities together will lead to a better understanding of the complicate three-dimensional flow in modern gas turbines. A detailed description of both facilities is given in this paper. Aerodynamic investigations in both facilities were performed. The in- and outlet Mach number and profile Mach number distribution is in good agreement in both of them and shows a periodic flow filed. Aerodynamic performance measurements in the cold flow facility have been conducted by means of a five-hole pneumatic pressure probe traverses 106% of cax downstream of the cascade to gain information about the quality of the flow field across flow passages “+1” and “–1” in terms of yaw angle, pitch angle and primary loss distribution. Comparison with a three dimensional Navier Stokes solvers show a very good agreement with the measurements. In order to deduce the external heat transfer coefficient on the vane a transient test procedure was adopted in the high-pressure hot facility. The dependency of the heat transfer coefficients on the Reynolds number is presented in the paper. The experimental results show reasonable agreement with calculations using a two dimensional boundary layer code.Copyright © 2001 by ASME

Exploitation of stator–stator interaction phenomena can increase overall efficiency in axial turbomachines. The first part of this paper sets out to present results of steady and unsteady flow experiments obtained in a new three stage cold flow low pressure turbine. Observation and understanding of boundary layer development and transition phenomena on the vanes dependent on relative stator position will be the focal point. In addition to this, stator position influence on profile loss, turbine efficiency and the development of secondary flow are examined. The experiments were carried out in the closed circuit test bed of the Institute of Aeronautical Propulsion at Stuttgart University. Annulus geometry and blading design of the research turbine were taken from the low pressure turbine of a modern commercial jet engine. The three stage test rig had identical blade counts for all stators or rotors respectively, whilst the circumferential position of each stator row could be individually adjusted. The second and third stators were optimised with respect to the radial alignment of the vanes. Surface mounted hot film gauges on the vanes and hot film probes were employed to assess the unsteady interaction phenomena. For steady measurements, pneumatic five hole probes and multiple kielhead pressure and temperature probes were used. For both the second and third stator, circumferential position was varied in eight steps over one pitch. Whereas the design point forms the basis of this detailed investigation, some attention was also paid to variations in Reynolds number and wheel speed. The results, such as quasi wall shear stress, stochastic and periodic fluctuations, total pressure etc., are presented in the form of chordwise ensemble-averaged distributions and contour plots and should be compared with the corresponding numerical studies presented in the second part of this paper.

Part I of this paper described how the HP turbine model rig of the Rolls-Royce Trent 500 was redesigned by applying non-axisymmetric end walls to both the vane and blade passages, whilst leaving the turbine operating point and overall flow conditions unaltered. This paper describes the results obtained from testing of the model rig and compares them with those obtained for the datum design (with conventional axisymmetric end walls). Measured improvements in the turbine efficiency are shown to be in line with those expected from the previous linear cascade research at Durham University, see Harvey et al. [1] and Hartland et al. [2]. These improvements are observed at both design and off-design conditions. Hot wire traverses taken at the exit of the rotor show, unexpectedly, that the end wall profiling has caused changes across the whole of the turbine flow field. This result is discussed making reference to a preliminary 3-D CFD analysis. It is concluded that the design methodology described in part I of this paper has been validated, and that non-axisymmetric end wall profiling is now a major new tool for the reduction of secondary loss in turbines (and potentially all axial flow turbomachinery). Further work, though, is needed to fully understand the stage (and multistage) effects of end wall profiling.Copyright © 2001 by ASME