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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

The paper presents a compilation of experimental data on the effects of wake-induced transition on a highly loaded LP turbine cascade intended to be used for further numerical work. Although the underlying physics is not yet completely understood, the benefits of wake passing are already known and employed in the design process of modern gas turbines. For further optimizations, the next step seems to be now to enable numerical simulations detailed enough to capture the major effects while being as uncomplicated as possible at the same time to be cost-effective. The experimental results constituted in this systematic investigation are available for download and should serve as a basic data set for future calculations with different turbulence and transition models, thereby shedding some light on the complexity and modeling required for a suitable numerical treatment of the wake-induced transition process. The data introduced in this test case was acquired using a turbine cascade called T106D-EIZ with increased blade pitch compared to design point conditions in order to achieve a higher loading. A large separation bubble forms on the suction side and allows to study boundary layer development in great detail. The upstream blade row was simulated by a moving bar type wake generator. The measurements comprise hot wire data of the bar wake characteristics in the cascade inlet plane (velocity deficit and turbulence level), boundary layer surveys with surface-mounted hot films sensors and a hot wire probe at various locations and measurements of the total pressure loss coefficient. Unsteady pressure transducers are embedded into the suction side of a cascade blade and in a wake rake to resolve the local pressure distributions over time. They yield quantitative values easily comparable to the output of numerical simulations. The objective of this paper is to enable and to invite interested researchers to validate their code on the data set. From the extensive test program, a very limited number of operating points have been selected to focus the work. The standardized data files include a “reference” case with an exit Reynolds number of 200.000 and an exit Mach number of 0.4 as well as two points with higher Mach or lower Reynolds number for constant wake passing frequencies and background turbulence levels.

An analysis of laser Doppler velocimeter (LDV) data for the axial velocity flowfield in the rotor of the Mach 2 inlet flow supersonic throughflow fan (SSTF) is presented in this paper. The paper starts with a short description of the SSTF test package to highlight the specifics of the SSTF operation. It is followed by a detailed description of a dedicated LDV system for measurement in a supersonic throughflow fan and the experience gained. Most of the experimental data presented were acquired in a low supersonic throughflow regime (inlet Mach number of 1.4). The results and conclusions presented are based mainly on the experimental data only. A limited amount of computational fluid dynamics (CFD) predictions were used for comparison with the experimental results. The CFD methods, however, are not discussed in this paper. As shown in this paper, a reasonably good agreement between the LDV data and the CFD predictions was found for the low supersonic throughflow regime. The design point data (inlet Mach number of 2.0) exhibited an unexpectedly high noise in the velocity data in comparison with the data for low supersonic throughflow operating conditions. For the off-design supersonic regime (shock in rotor), substantial differences exist between the experimental and computational data.

This paper describes the redesign of the HP turbine of the Rolls-Royce Trent 500 engine, making use of non-axisymmetric end walls. The original, datum turbine used conventional axisymmetric end walls, while the vane and (shrouded) rotor aerofoil profiles were nominally the same for the two designs. Previous research on the large scale, low speed linear cascade at Durham University, see Hartland et al [1], had already demonstrated significant potential for reducing turbine secondary losses using non-axisymmetric end walls - by about one third. This paper shows how a methodology was derived from the results of this research and applied to the design of the single stage Trent 500 HP turbine (model rig). In particular the application of a new linear design system for the parametric definition of these end wall shapes, described in Harvey et al [2], is discussed in detail.Copyright © 2001 by ASME

This paper presents a computational investigation of flows in a single axial flow fan with and without casing treatment. It analyses the effect of the recess casing treatment on stall margin improvement as well as its influence on global performance parameters. The paper seeks to offer a contribution to the understanding of the physical processes occurring when approaching stall and the working mechanism by which casing treatments improve stall margin. A Reynolds-Averaged Navier-Stokes CFD code was used for the analysis and the numerical investigation of the overall performance, efficiency and work-input characteristics of the fan were found to agree very well with previously reported experimental results. The effect of casing treatment was investigated using two types of configurations, vaneless and vaned casing. The vaneless casing treatment produced a sizeable stall margin improvement with negligible loss of efficiency. The recess was fitted later with vanes and was shown to offer both a further stall margin improvement and an increase in the pressure rise coefficient without any significant drop in efficiency at design conditions.

In November 1999 the turbine test facility at the Institute for Thermal Turbomachinery and Machine Dynamics was completed. This test stand is a cold flow transonic test turbine facility, which is a unique combination of a 2.3 MW axial test turbine and a directly-coupled compressor that generates about the half of the compressed air needed to drive the turbine. To cover the losses, additional air is provided by an electrically-driven compressor station. A 600 kW suction blower driven by a helicopter gas turbine is used to decrease the exhaust pressure, thus increasing the expansion ratio up to 4.5 (see Erhard and Gehrer “Design and Construction of a Transonic Test-Turbine Facility”, ASME Paper 2000-GT-480). Since then, flow measurements with two different transonic turbine stages have been carried out. This paper will focus on the description of the measurement system, data analysis and the newest test results that have been obtained on a turbine stage developed by Nuovo Pignone. This stage has been configured to be representative of advanced designs for medium/small land based gas turbines and is characterized by high pressure ratios and strongly transonic effects. The characteristic feature of this stage is the convergent- divergent flow channel. Basic data of this turbine stage to obtain flow similarity are: 2.2 MW shaft power at 11500 rpm, pressure ratio 4.5, mass flow 22 kg/s. The paper will contain appropriately plotted results of dimensionless stage characteristic curves over a broad range of operation points.

It is well-known that phase-shifting controllers used for active combustion control must be manually adjusted in order to maintain control over a broad range of operating combustor operating conditions. If one assumes that the thermoacoustic instabilities are linearly stabilizable, then what is needed is a method to determine, and ultimately predict, the frequency response of the plant for any range of operating conditions, so the controller design can be automatically updated to track the changing plant gain/phase relationships that are observed with changing heat release. A unique test-based, design process has been proposed to predict the gain/phase characteristics required of a proportional, phase-shifting controller that can stabilize the thermoacoustic instabilities. In this paper, that process is used to automate the design of a fixed-gain feedback controller that limits the amplitudes of any feedback induced instabilities (to some pre-specified level) while providing the best control of the targeted limit cycling pressure oscillations. The paper describes how a neural network was trained, using the suggested design process, to predict the frequency response of the thermoacoustics in a tube combustor at frequencies adjacent to the limit cycle frequency using certain operating conditions that included a sparsely-sampled temperature profile, total air/fuel flow rate, and equivalence ratio. The neural net training was performed using complex valued, open-loop frequency response function data as the desired signal with the previously mentioned operating conditions as the input signals. (The open loop data was collected for a narrow frequency range surrounding the limit cycle instability by performing a sine dwell at discrete frequencies). Once the neural network was trained, it was used to predict the approximate phase and gain margins as a function of temperature and flow conditions. The margins were then used to automatically update and design a fixed shape feedback controller having the proper phase and magnitude to ensure stability and control in the face of changing operating conditions. A companion paper describes the methodology that underlies the automated design of the feedback controller gain and phase delay.

This paper describes the implementation and results of an experiment to measure the isentropic efficiency of a fully scaled transonic turbine stage in a short-duration blowdown type facility. Short duration test facilities offer the potential for low cost, high accuracy testing of turbomachinery components under fully scaled conditions. The method developed employs the ‘brake’ approach for turbine efficiency measurement. This paper describes newly developed techniques to measure the turbine torque and mass flow in this type of facility. The isentropic efficiency was measured over a test window of 400 to 500 milliseconds. The estimated measurement uncertainty (U95) is 0.45% for absolute measurements, and 0.25% for relative measurements. Corrected mass flow and isentropic efficiency measurements are presented for a range of corrected speed (97–112%) and compared with CFD predictions. Close agreement was observed both in overall magnitudes and trends between the measured data and predictions.

Unsteady stator-rotor interaction in gas turbines has been investigated both experimentally and numerically for some years now. Even though the numerical methods are still in development, today they have reached a certain degree of maturity allowing industry to focus on the results of the computations and their impact on turbine design, rather than on a further improvement of the methods themselves. The key to increase efficiency in modern gas turbines is a better understanding and subsequent optimization of the loss-generation mechanisms. A major part of these are the secondary losses. To this end, this paper presents the time-resolved secondary flow field for the two test cases computed, viz the first and the last turbine stage of a modern heavy duty gas turbine. A companion paper referring to the same computations focuses on the unsteady pressure fluctuations on vanes and blades. The investigations have been performed with the flow solver ITSM3D which allows for efficient calculations that simulate the real blade count ratio. This is a prerequisite to simulate the unsteady phenomena in frequency and amplitude properly.Copyright © 2000 by ASME

This paper is part II of a comprehensive study on the blade leading edge sweep/bend of supersonic and transonic axial compressors. The paper explores and analyzes the kinematic characteristic variables of three-dimensional (3-D) swept shock surfaces. In the research field studying the sweep aerodynamics of axial flow compressors and fans, many types of high loading swept blades are under intensive study. So, in both direct and inverse design methods and experimental validations, an accurate grasp of the sweep characteristic of the blade’s 3-D swept shock surface becomes of more concern than before. Associated with relevant blading variables, this paper studies the forward and zero and backward sweeps of shock surfaces, defines and resolves every kind of useful sweep angle, obtains dimensionless sweep similarity factors, suggests a kind of method for the quantitative classification of 3-D shock structures, and proposes the principle of 3-D shock structure measurements. Two rotor blade leading edge shock surfaces from two high loading single stage fans are analyzed and contrasted. This study is the foundation of the kinematic design of swept shock surfaces.

Understanding the behaviour of turbomachinery components has rapidly improved during the last years. The focus has been set mainly on unsteady component characteristics such as compressor stall and combustion chamber flameout. However, modelling the whole engine during transient operation is still a major workpackage during the development process of modern jet engines. The aim of transient modelling of jet engines is the prediction of the normal engine behaviour during changes of the powersetting, but also to prove engine safeness in abnormal operation, e.g. caused by failures of the control system. Various approaches are used depending on the application purpose. For the most accurate prediction, complex synthesis programs model each component separately and iterate to match the components against each other. For customer use as well as for evaluation whether the real control software in the Electronic Engine Controller operates satisfactorily, more simple programs are derived from this, which include the engine characteristics in a more general way. This gives a very robust tool for simulation of simple transient manoeuvres as well as e.g. flameout or starting predictions. The paper will present the variety of approaches for modelling transient engine behaviour and will discuss the limitations of these. It will also include a description, how the control system of modern jet engines, the FADEC, is set up in these environments. This modelling is needed, if a realistic approach towards reliability of in service operation shall be performed. Synthesis results are given from the development process. The use of advanced modelling of the control system with dedicated tools to improve the capabilities of the whole system is highlighted. An example for this is included with regard to the scheduling of the handling bleed valve of the booster compressor. Finally, the paper will give an outlook on the future planning of the synthesis of transient engine operation within Rolls-Royce Deutschland GmbH (RRD).

The paper is a sequel to an earlier work by Sorokes et al. 1998, “Investigation of the Circumferential Static Pressure Non-Uniformity Caused by a Centrifugal Compressor Discharge Volute.” The earlier work described experimental and computational fluid dynamics analyses of the non-uniform static pressure distortion caused by the discharge volute in a high pressure, centrifugal compressor with vaneless diffusers. This paper describes additional testing and analytical work done using low solidity vaned diffusers (LSD’s) in place of select vaneless diffusers to determine the alternate diffuser’s effectiveness in eliminating or reducing the magnitude of the non-uniform pressure field. As in the earlier studies, the experiments described in this paper were done using a heavily instrumented gas re-injection compressor operating at over 6000 psia discharge pressure. Instrumentation was installed to measure static, total, and dynamic pressure as well as impeller strain and mechanical vibrations. A brief description of the compressor and instrumentation are provided. Concurrent with the experimental work, CFD runs were completed to study the effect of the alternate vaned diffusers. The CFD pressure profile trends agreed well with the experimental results and provided analytical corroboration for the conclusions drawn from the test data. Conclusions are drawn regarding: a) the effectiveness of the LSD’s on the pressure non-uniformity; b) the associated effects on the measured dynamic strains in the impellers; and c) the usefulness of computational fluid dynamics (CFD) in assessing the aerodynamic forces associated with the non-uniformity.

Magnetic bearings are widely used as active suspension devices in rotating machinery, mainly for active vibration control purposes. The concept of active tip clearance control suggests a new application of magnetic bearings as servo-actuators to stabilize rotating stall in axial compressors. This paper presents a first-of-a-kind feasibility study of an active stall control experiment with a magnetic bearing servo-actuator in the NASA Glenn high-speed single-stage compressor test facility. Together with CFD and experimental data a two-dimensional, incompressible compressor stability model was used in a stochastic estimation and control analysis to determine the required magnetic bearing performance for compressor stall control. The resulting requirements introduced new challenges to the magnetic bearing actuator design. A magnetic bearing servo-actuator was designed which fulfilled the performance specifications. Control laws were then developed to stabilize the compressor shaft. In a second control loop, a constant gain controller was implemented to stabilize rotating stall. A detailed closed loop simulation at 100% corrected design speed resulted in a 2.3% reduction of stalling mass flow which is comparable to results obtained in the same compressor by Weigl et al. (1998) using unsteady air injection. The design and simulation results presented here establish the viability of magnetic bearings for stall control in aero-engine high-speed compressors. Furthermore the paper outlines a general design procedure to develop magnetic bearing servo-actuators for high-speed turbomachinery.

This paper provides a high level overview of why the Royal Navy is seeking to procure a 1–2MW Gas Turbine Alternator as part of a full electric propulsion system. It describes the Royal Navy requirements of marine gas turbines and highlights the reasons for the adoption of Integrated Full Electric Propulsion. The paper also describes the major technical and commercial requirements of the 1–2MW gas turbine alternator and the procurement process that was used to select the prime contractor.Copyright © 2000 by ASME

This paper analyses the flow in a transonic turbine stage, using time-dependent numerical results. Unsteady blade-row interaction has repercussions on the time-averaged flow, which are represented by the so-called “deterministic correlations”. These correlations appear in the system of equations governing the time-averaged flow; they can be divided into four types with different physical meanings. Time-dependent results enable direct computation of these correlations in both rotor and stator frames of reference. The computed deterministic correlations are analysed in the paper, in order to bind them to physical phenomena and to evaluate their influence on the time-averaged flow field. This analysis is also intended to help assess the shortcomings of simple mixing-plane methods and more complex approaches using deterministic correlation models. While the first part focuses on one particular type of deterministic correlation, the so-called “spatial correlation”, the second part attempts a more detailed analysis of time-dependent results and gives some clues to the orders of magnitude of the four types of deterministic correlation. The conclusions should be taken with caution; they may partly depend on the present turbine configuration with a specified structure of unsteadiness and on the present turbulence model.

This paper describes the design concepts and object-oriented architecture of Onyx, an extensible domain framework for computational simulation of gas turbine engines. Onyx provides a flexible environment for defining, modifying and simulating the component-based gas turbine models described in Part 1 of this paper. Using advanced object-oriented technologies such as design patterns and frameworks, Onyx enables users to customize and extend the framework to add new functionality or adapt simulation behavior as required. A customizable visual interface provides high-level symbolic control of propulsion system construction and execution. For computationally-intensive analysis, components may be distributed across heterogeneous computing architectures and operating systems. A distributed gas turbine engine model is developed and simulated to illustrate the use of the framework.Copyright © 1999 by ASME

Two primary topics are addressed in this paper: 1. A brief record of the evolution of U. S. Navy (USN) gas turbine qualification and testing requirements that helped produce the successful engines now in service and 2. Examination of these requirements in light of current and projected defense budgets resulting from changes in the world over the past twenty-five years. The paper proposes changes in the current methods the USN employs in qualifying its gas turbines for surface ship applications. The proposed alternate methods consider commercial testing and service experience as well as testing based on MIL-E-17341 requirements. Finally, the marine gas turbine community is invited to begin a dialogue on the feasibility of these alternate methods.

The paper highlights on-going studies and design efforts by a team of several specialists in aspects related to a type of power generation installation, which consists of a directly coupled closed-cycle gas turbines (the Light Water Reactors have a double loupe energy transfer system coupled by a heat exchanger) to an inherently safe nuclear heat source. In particular the NEREUS installation will be discussed. NEREUS is the God for shipping in Greek mythology, but in this case it stands for: a Naturally safe, Efficient, Reactor, Easy to operate, Ultimately simple and Small. This describes very well the aims of the design team. This part will explain the inherently safe character of the nuclear heat source, the way the energy production in the heat source is controlled, the type of nuclear fuel and the treatment of the nuclear waste, the refuelling, the weight, the dimensions and the construction of the nuclear part of the NEREUS installation. Part two will describe the non-nuclear part of the installation, the possible applications, its potential markets and the technical and financial indications from the on-going study. This paper should be read in conjunction with the paper “The NEREUS installation — the Non-Nuclear part (GT)”.

On March 13th, 1998, Dr. Hans Joachim Pabst von Ohain co-inventor of the turbojet passed away at the age of 86. As a young doctoral student, von Ohain conceived of and built a demonstrator turbojet engine. He was hired by the Heinkel Aircraft Company in 1936 and under intense time pressure imposed by Ernst Heinkel, designed the world’s first flight turbojet engine. This paper traces the technical antecedents leading to historic jet-powered flight made on August 27, 1939 by a Heinkel He 178 aircraft powered by von Ohain’s HeS 3B turbojet. During his tenure at Heinkel and thereafter at the Heinkel-Hirth Company, he was responsible for a series of turbojet engines culminating in the advanced second generation HeS 011 with a thrust of 2,860 lbs. This paper is a tribute to an outstanding scientist who made possible the turbojet revolution and who will forever be remembered as the inventor of the world’s first flight turbojet.Copyright © 1999 by ASME

This paper is Part I of a comprehensive study on the blade leading edge sweep/bend of supersonic and transonic axial compressors. The objective of the paper is to develop an analytical shock surface model which links a blade leading edge with the 3-D shock structure, and can be eventually explored as a means to estimate the direct and indirect shock losses. The model surface is composed of helixes on a set of cylinders with aerodynamically determined ascendent for each helix. One kind of the relation between the ascendent and 2-D aerodynamics on S1 stream surface has been established for a leading edge attached shock surface. The succeeding research has revealed that the concepts of a forward/backward swept shock surface, and their qualitative and quantitative natures, can be analysed by the model in uniformity. The model is constructed as such that it can also be used in the inverse method of design process. Two leading edge shock surfaces from two high loading transonic fans are simulated by the model.Copyright © 1999 by ASME

Throughflow theory, as a design problem, is extensively used in the design of transonic axial flow compressors and fans. The losses caused by a simplified passage shock named shock loss model were taken into account in the design of transonic stages. However, the influences of shock waves on flow fields are still under development for transonic design. This problem is studied in the present paper. The basic equations of the throughflow method are re-derived under the concept of discontinuous passage averaging, and then, the distribution shock forces acied on the averaged throughflow fields are obtained. To evaluate these effects on averaged throughflow fields, a passage shock structure model is developed based on the previous models. Finally, the shock influences are assessed and compared in the present paper. It is demonstrated that the distribution shock forces have dramatically effects on the results of design.

In this paper the mixing of stator wakes in turbomachinery is considered. An extension is made to the existing model of Denton (1993), which addresses the effects of acceleration before mixing. Denton showed that if a total pressure wake was accelerated mixing loss diminished and vice versa. Here a total temperature wake is shown to exhibit a reverse trend. An attempt is also made to better understand the work transfer process between a stator wake and a rotor. The paper concentrates on axial turbines, but a brief look at compressors is included. It is argued that the freestream work is not the same as the wake work and the concept of ‘Differential Work’ is introduced. A simple steady velocity triangle based model is proposed to give an estimate of the ratio of wake work to freestream work (μ see later). The model is compared to an unsteady CFD result to offer some verification of the assumptions. It is concluded that the rotodynamic work process tends to reduce total pressure wake depths in turbines and compressors and therefore mixing losses. The mixing loss due to total temperature wakes is less strongly affected by the differential work process.

A study was conducted for NASA Dryden Flight Research Center to investigate the viability of developing an externally powered zero stage booster fan to boost the altitude performance of a high altitude turbojet from 80 kft. to over 90 kft., in order to realize a very high altitude subsonic remotely piloted aircraft (RPA) capable of atmospheric science missions. A modified non-airbreathing torpedo engine powers the booster fan. The powered fan is grafted onto an airbreathing turbojet (the General Electric YJ97 jet engine) and is used to supercharge the engine (boost inlet pressure and airflow) at altitudes above its original design limits. The inlet boost keeps the engine operating and delivering thrust, at subsonic speeds, to mission altitudes over 95,000 ft. If this modified engine were grafted back into the air vehicle it was originally designed for (the Viet Nam era AQM-91A Compass Arrow) the modifications would enable the vehicle to increase its maximum altitude from 80,000 ft to over 88,000 ft for periods up to 45 min. With some additional wing area, the aircraft, which is in the 5000 lbm TOGW class, could carry a 500 lbm payload to altitudes over 92,000 kft with an operational radius over 1500 NMi. This study examined a number of candidate component arrangements to realize this fan-boosted propulsion system. Performance was estimated for the subsystem components, the resulting propulsion systems, and the associated RPA. In addition, the study estimated the time and cost to demonstrate/validate this concept in a full-scale propulsion system wind tunnel test program. This paper describes the concept, summarizes the concept selection efforts and performance analysis conducted during the study, and explains the rationale behind this unusual approach. A description of the proposed development and test program, and the cost estimate, lies beyond the scope of this paper.

In this paper, the two-stage shrouded HPT engine configuration rated at 22000 lbs thrust is used as the baseline from which a single stage HPT unshrouded design is systematically derived to evaluate the potential weight and cost advantage. The baseline thermodynamic cycle at the rated thrust level was modified in order to optimize the turbine inlet temperature, overall pressure ratio, and core flow with a single stage HPT and deliver competitive performance. The comparative study, although preliminary in depth, has led to the advantages and disadvantages associated with an unshrouded single versus a two-stage shrouded HPT design. The results compare design configuration, secondary air system, weight, safety, life, specific fuel consumption (SFC), and future thrust growth capability. The main advantages of the single stage application are reductions in cost and complexity of design, lower turbine gas temperature, and ease of maintenance. The main disadvantages are in reduced turbine polytropic/isentropic efficiency for HPC pressure ratio greater than 9, increased SFC, higher rim speed, higher HPT exit Mach number, higher bypass ratio to achieve the desired thrust level, and possibly higher weight. A quantitative statement on the reduction of engine cost/weight is premature until a detailed design and the associated cost-benefit is performed. The paper concludes by recommending that the design philosophy of the modern unmixed turbofan engine (single or two-stage HPT) leads to a balance between the selected turbine gas temperature versus the by-pass ratio in order to minimize cost and maximize the thrust-to-weight ratio and the cycle efficiency. In either ease, the expected high reliability and reduced engine cost/weight in the context of future thrust-growth capability need to be demonstrated by proven technology which seem to favor the two-stage HPT configuration.Copyright © 1999 by ASME

Development of the Northrop Grumman / Rolls-Royce WR21 Intercooled Recuperated (ICR) Gas Turbine, begun in 1992, is now well advanced and system testing has been completed on eight engine builds at the Royal Navy’s Admiralty Test House located at the Defence Evaluation and Research Agency, Pyestock in the United Kingdom. Test activity is shortly to move to the US Navy’s Test Site at the Naval Surface Warfare Center, Carderock Division – Ship Systems Engineering Station in Philadelphia, PA, where a new test facility has been built to carry out some final development testing and an endurance test. A previous paper on this subject (94-GT-186) defined a test program leading to a design review and the beginning of Qualification Testing. The development program has since evolved and it is the aim of this paper to summarize engine testing to date and set out the plan for conclusion of development testing. The paper will describe the development of the Philadelphia Test Site, as a combined site for the US Navy’s Integrated Power System (IPS) and ICR testing. This will include a description of the advanced, high-accuracy Data Acquisition System (DAS). Finally, the test program and the development and endurance test objectives will be outlined.

This paper describes the development of a computational model for the moderate to long lengthscale flowfield behaviour of high-speed multi-stage compressors. The simulation is a time accurate solution of the fully compressible non-linear equations for fluid flow through an entire compressor. Flow in the non-bladed regions is calculated using a two-dimensional Euler solver whilst the individual blade rows are modelled using multiple one-dimensional flowfields with body forces. The body forces are chosen to give the required steady state blade row performance and a physically based circulation model is used to give the correct non-steady flow dynamics. The capabilities of the simulation are demonstrated in this paper by calculating how the flowfield breakdown of a four stage compressor is affected by operating speed and a range of different inlet flow distortions.Copyright © 1997 by ASME

This paper presents a study of stall inception mechanisms a in low-speed axial compressor. Previous work has identified two common flow breakdown sequences, the first associated with a short lengthscale disturbance known as a ‘spike’, and the second with a longer lengthscale disturbance known as a ‘modal oscillation’. In this paper the physical differences between these two mechanisms are illustrated with detailed measurements. Experimental results are also presented which relate the occurrence of the two stalling mechanisms to the operating conditions of the compressor. It is shown that the stability criteria for the two disturbances are different: long lengthscale disturbances are related to a two-dimensional instability of the whole compression system, while short lengthscale disturbances indicate a three-dimensional breakdown of the flow-field associated with high rotor incidence angles. Based on the experimental measurements, a simple model is proposed which explains the type of stall inception pattern observed in a particular compressor. Measurements from a single stage low-speed compressor and from a multistage high-speed compressor are presented in support of the model.Copyright © 1997 by ASME

The Allison 501-K17 Ship Service Gas Turbine Generator Set (SSGTGS) is used is provide ship board electrical power on several U.S. Navy Class ships, including the DD-963 Spruance Destroyer, the DDG-993 Kidd Guided Missile Destroyer and CG-47 Ticonderoga Guided Missile Cruiser Classes. The first of these units were placed in service during the mid 1970s. The Naval Sea Systems Command (NAVSEA) in conjunction with the Naval Surface Warfare Center, Carderock Division, Ship Systems Engineering Station (NSWCCD-SSES) have undertaken a major upgrade effort to improve the reliability, operation, serviceability and maintainability of the unit. This paper examines the process of this program and details the specific improvements made to the unit as a result of this effort. In addition, this paper outlines the experience gained as a result of installing these upgrades in the Fleet.Copyright © 1997 by ASME

The present study is part of an on-going project concerned with the effects of off-design incidence on losses. Previous work has examined the profile losses (Benner et al., 1995). This study focusses on the secondary flow, the endwall flow in the absence of tip clearance. This paper describes an experimental study in a large-scale, low-speed, linear turbine cascade for which the incidence was varied. Three different values of incidence were investigated: 0, +10 and +20 degrees. Detailed flow field measurements were made upstream and downstream of the cascade and static pressure distributions were measured on the blade surfaces. The data were supplemented by extensive surface oil flow visualization. This paper concentrates on the physical insights into the relationship between the behaviour of secondary flows and incidence which have been gained from the measurements and the flow visualization. A preliminary discussion of the effect of incidence on the losses is also included.Copyright © 1997 by ASME

This paper deals with the numerical simulation of flow through a 1.5 stage axial flow turbine. The 3-row configuration has been experimentally investigated at the University of Aachen where measurements behind the first vane, the first stage and the full configuration were taken. These measurements allow single blade row computations, to the measured boundary conditions taken from complete engine experiments, or full multistage simulations. The results are openly available inside the framework of ERCOFTAC 1996. There are two separate but interrelated parts to the paper. Firstly, two significantly different Navier-Stokes codes are used to predict the flow around the first vane and the first rotor, both running in isolation. This is used to engender confidence in the code that is subsequently used to model the multiple bladerow tests, the other code is currently only suitable for a single blade row. Secondly, the 1.5 stage results are compared to the experimental data and promote discussion of surrounding blade row effects on multistage solutions.

As the operational life of aircraft hardware increases, repetitions of phenomena previously explained as “isolated incidents” have been seen to occur. These problems are not necessarily application specific, nor material’s anomaly, nor design defect, nor operational abuse, nor any single cause, but, usually, an unfortunate combination of degradation in several of these factors, which becomes, statistically, more probable as the operational hours increase. Because they are rare, it can take years to collect enough clues to identify these links and synthesize the event. When the problem involves an aging commercial aircraft turbine engine, that has passed from the designer to the manufacturer, and then to the airframer and the operator, through the hands of the maintainer, the repairer and the spare parts breakout supplier, the combinations of variables becomes too large to attack individually. This paper characterizes a common sense approach to solving these problems. This approach favors a small team with a strong leader dedicated to that task. It is based on a theory that clues may be found in the history of the product and the industry, as well as in the exhibits and demographics of the incidents, and that a broad approach to fact finding, coupled with disciplined treatment of data, will allow the investigators to develop some intuitive sense of their problem, and facilitate linking of the pertinent facts. This paper includes a hypothetical example, based on information in the public domain, of such a problem relating to an aging gas turbine engine. It is not meant to represent an actual case history, but is rather a combination of recognizable events, not uncommon, in either type or quantity, in the gas turbine industry, combined to illustrate this theory.

In this paper, for the first time, a set of guidelines is presented for the systematic design of mixed flow and centrifugal compressors and pumps with suppressed secondary flows and a uniform exit flow field. the paper describes the shape of the optimum pressure distribution for the suppression of secondary flows in the impeller with reference to classical secondary flow theory. The feasibility of achieving this pressure distribution is then demonstrated by deriving guidelines for the design specifications of a three-dimensional inverse design method, in which the blades are designed subject to a specified circulation distribution or 2 pi r (V) over bar(theta). The guidelines will define the optimum choice of the blade loading or partial derivative r (V) over bar(theta)/partial derivative m and the stacking condition for the blades. These guidelines are then used in the design of three different low specific speed centrifugal pump impellers and a high specific speed industrial centrifugal compressor impeller. The flows through all the designed impellers are computed numerically by a three-dimensional viscous code and the resulting flow field is compared to that obtained in the corresponding conventional impeller. The results show consistent suppression of secondary flows in all cases. the design guidelines are validated experimentally by comparing the performance of the inverse designed centrifugal compressor impeller with the corresponding conventional impeller. The overall performance of the stage with the inverse designed impeller with suppressed secondary flows was found to be 5 percent higher than the conventional impeller at the peak efficiency point. Exit flow traverse results at the impeller exit indicate a more uniform exit flow than that measured at the exit from the conventional impeller.

By far the greater part of our understanding about stall and surge in axial compressors comes from work on low-speed laboratory machines. As a general rule, these machines do not model the compressibility effects present in high-speed compressors and therefore doubt has always existed about the application of low-speed results to high-speed machines. In recent years interest in active control has led to a number of studies of compressor stability in engine type compressors. The instrumentation used in these experiments has been sufficiently detailed that, for the first time, adequate data is available to make direct comparisons between high-speed and low-speed compressors. This paper presents new data from an eight-stage fixed geometry engine compressor and compares this with low-speed laboratory data. The results show remarkable similarities in both the stalling and surging behaviour of the two machines, particularly when the engine compressor is run at intermediate speeds. The engine results also show that, as in the laboratory tests, surge is precipitated by the onset of rotating stall. This is true even at very high speeds where it had previously been thought that surge might be the result of a blast wave moving through the compressor. This paper therefore contains new information about high-speed compressors and confirms that low speed testing is an effective means of obtaining insight into the behaviour of high-speed machines.Copyright © 1993 by ASME

The Machinery and Engine Technology (MET) Program of the National Research Council of Canada (NRCC) has established a program for the evaluation of sensors to measure gas turbine engine performance accurately. The precise measurement of fuel flow is an essential part of steady-state gas turbine performance assessment. Prompted by an international engine testing and information exchange program, and a mandate to improve all aspects of gas turbine performance evaluation, the MET Laboratory has critically examined two types of fuel flowmeters, Coriolis and turbine. The two flowmeter types are different in that the Coriolis flowmeter measures mass flow directly, while the turbine flowmeter measures volumetric flow, which must be converted to mass flow for conventional performance analysis. The direct measurement of mass flow, using a Coriolis flowmeter, has many advantages in field testing of gas turbines, because it reduces the risk of errors resulting from the conversion process. Turbine flowmeters, on the other hand, have been regarded as an industry standard because they are compact, rugged, reliable, and relatively inexpensive. This paper describes the project objectives, the experimental installation, and the results of the comparison of the Coriolis and turbine type flowmeters in steady-state performance testing. Discussed are variations between the two types of flowmeters due to fuel characteristics, fuel handling equipment, acoustic and vibration interference and installation effects. Also included in this paper are estimations of measurement uncertainties for both types of flowmeters. Results indicate that the agreement between Coriolis and turbine type flowmeters is good over the entire steady-state operating range of a typical gas turbine engine. In some cases the repeatability of the Coriolis flowmeter is better than the manufacturers specification. Even a significant variation in fuel density (10%), and viscosity (300%), did not appear to compromise the ability of the Coriolis flowmeter to match the performance of the turbine flowmeter. Copyright © 1993 by ASME., ASME 1993 International Gas Turbine and Aeroengine Congress and Exposition, GT 1993, 24 May 1993 through 27 May 1993

This paper deals with a theoretical study of the active control of purely axial 1D instabilities in multi-stage axial-flow compressors. The paper briefly considers the development of a suitable surge model and continues with the derivation of a controller using linear optimal control theory. The controller is specifically designed to suppress the instabilities predicted by the linearised form of the surge model. The control technique involves a bleed being dynamically varied in response to fluctuations of variables. It is found that a stabilising optimal controller can always be obtained. The second part of the paper presents a non-linear simulation of the surge prediction model with the linear controller, using a 4th order Runge-Kutta scheme. This demonstrates how some non-linearities in the flow model can affect the controller operation. It is also shown that, the amplitudes and the frequency response required of the actuator to retain stability can exceed the practical limitations.Copyright © 1993 by ASME

This paper presents results of a study, supported by NASA Contract NAS3-20646 under the Energy Efficient Engine Program, conducted to investigate the development of boundary layers under the influence of velocity distributions simulating the suction sides of two state-of-the-art turbine airfoils. One velocity distribution represented a forward loaded airfoil (“squared-off” design) while the other represented an aft loaded airfoil (“aft loaded” design). These velocity distributions were simulated in a low-speed high aspect ratio wind tunnel specifically designed for boundary layer investigations. It is intended that the detailed data presented in this paper be used to develop improved turbulence model suitable for application to turbine airfoil design.Copyright © 1981 by ASME

The paper describes a general program which has been developed for the prediction of the transient performance of gas turbines. The program is based on the method of continuity of mass flow. It has been applied successfully to a wide range of aero gas turbines, ranging from single to three-spool and from simple jet to bypass types with or without mixed exhausts. The results for three of these engine types are illustrated. Computing times are reasonable, increasing with the complexity of the engine. A parallel paper describes the inclusion of thermal effects in the prediction program.

This paper aims to discuss the factors which led to the selection of the Spey as a marine gas turbine, and the market forces which have made an uprating of the simple cycle unit necessary, and which are now leading to the conception of an advanced cycle version of the gas turbine at even higher power and efficiency. The operational characteristics of warships using these Spey units are compared with those of ships using other propulsion systems. While this paper stands on its own, it can also be regarded as the first part of a series of two; the second part, to be published in Beijing during the September 1985 ASME Gas Turbine Exposition, will deal with that further possible uprating of the Spey to about 22 MW using intercooling and heat exchange to improve its thermal efficiency to about 43%.Copyright © 1985 by ASME

Airborne Early Warning for the Navy fleets has been provided for the past 20 years by the Grumman/Allison E-2 Airframe/T56 engine combination. Although avionic capability has been continually updated to meet the increased threat, the airframe and powerplant have seen only minor changes. Projected mission requirements and future avionic system enhancements require payload increases being limited by the power capability of the present T56 powerplant. Of paramount importance in the E-2 carrier deck operation is the single engine rate of climb capability of the aircraft. This paper discusses the logical evolution of a replacement engine for the E-2C — a derivative T56 engine contracted and designated by the Navy as the T56-A-427 — to meet the projected single engine takeoff and other mission requirements. The T56-A-427 provides 24% power and 13% fuel consumption improvements with identical installation interfaces, and substantially improves E-2C performance characteristics across the flight envelope. Furthermore, the paper shows that meeting these stringent performance requirements with a derivative engine results in a low risk development program and an engine with improved maintainability and reliability, which can capitalize on the in-place logistics support base of the T56 — the world’s longest production run gas turbine engine.Copyright © 1985 by ASME

Several methods are available for the optimization of basic design parameters and the preliminary efficiency prediction of axial flow turbine stages. However, their application is often questionable for stages having low specific speed and/or small volume flow rates. In particular, the question may arise whether a better performance is achieved by a partial admission, impulse stage or by a full admission reaction stage having lower blade height. The paper firstly reviews the available loss correlation methods applicable to partial admission turbines, then a comparison is performed between the efficiency achievable by partial and full admission stages designed for the same operating conditions. The turbine design procedure for both options is fully automatized by an efficiency optimization method similar to the one described in previous authors’ papers. The results of calculations are presented in the paper as a function of similarity parameters (specific speed, size parameter, expansion ratio). It is found that the results obtained with different correlations are relatively similar for “conventional” turbine stages (low expansion ratio, moderate size parameters), while important differences take place for very small sizes and/or in presence of important compressibility effects. The presented results can be useful: 1) to decide whether selecting full or partial admission solutions; 2) to optimize the degree of admission and the other basic design parameters, and 3) to predict with reasonable accuracy the stage efficiency.

Military engine manufacturers have been given a new set of ground rules by the Department of Defense and Congress when they vie for engine contracts. The manufacturers must establish that their engines are Life Cycle Cost (LCC) effective. This directive has resulted in engine manufacturers developing life cycle cost models, including airframes, to account for an overall weapons system. Once the model has been developed, it can be expanded allowing parametric studies of the engine/airframe system thereby developing a cost effective system. The Aircraft Engine Group of General Electric has developed such an LCC computerized model. This paper discusses various applications of the subject model as applied to an engine program before the design is finalized, during the detail design phase, and after the engine has been placed in production. The paper also includes examples of these applications of the model.Copyright © 1981 by ASME