High Pressure Spool Research Articles

Comprehensive Flow Path Design Method for the Adaptive Cycle Engine Considering the Coupling Relation of Multiple Components

Abstract The adaptive cycle engine (ACE) has multiple coupled components on the same spool and complex bypass system, which makes it have more complex intercomponent coupling relation and hard to coordinate in the flow path design. In this study, the coupling relation of the ACE components and the component reference conditions are analyzed and determined, a multicomponent collaborative optimization design method is proposed to enable the quantitative evaluation of flow path design solutions. In this method, two optimization strategies are presented based on the different priorities of the intercomponent size coupling parameters, the intercomponent aerodynamic coupling parameter and the component performance in the optimization problem. ACE flow path solutions for various feasible design speed combinations are generated automatically considering the component performance and intercomponent coupling relation. According to an ACE flow path design case study, the design physical rotational speeds of low-pressure spool (NL,d) and high-pressure spool (NH,d) should be 7000 to 7600 r/min and 10,000 to 15,000 r/min, respectively. At NH,d = 12,000 r/min and NL,d = 7200 r/min, the high-pressure compression components and the fan components could be designed with the lowest aerodynamic load, respectively. NH,d is the key factor affecting the axial length of ACE. This method can be applied to other gas power plant designs.

Features of design calculation of compressor spool flow path of a twin-shaft gas turbine engine core on the basis of 1D and 2D models of their working process

The features of the stages of parameter design calculation for the formation of the initial design of the flow path of the compressor spool of a twin-shaft engine core of a gas turbine engine are presented and substantiated. The article contains recommendations for choosing values of the pump head coefficient, efficiency and other important parameters for the stages of medium-pressure and high-pressure spools at the stage of thermodynamic calculation. At the stage of design gas-dynamic calculation of the compressor at the middle diameter, typical distributions of axial velocity component and degree of reaction along the flow path of compressor spools should be taken into consideration. At the same time, it is necessary to provide requirements for reducing the flow velocity and static pressure coefficients in the rotor wheels and stator blades, head coefficients and Stepanov coefficients. The design gas-dynamic calculation of the compressor along the radius of the flow path is characterized by a variety of flow swirl laws at the inlet to the rotor wheels, distributions of the pressure increase and efficiency over the height of the blades. In conclusion, an example of a three-dimensional model of a compressor flow path taking into consideration the features of design calculation of the parameters of compressor spools of a twin-shaft engine core of a gas turbine engine on the basis of the appropriate flow path diagram in the meridional plane is presented.

Qualification Approach for High-Pressure Spool Shaft of a Low-Bypass Turbofan Engine for Structural Integrity and Life

Qualification Approach for High-Pressure Spool Shaft of a Low-Bypass Turbofan Engine for Structural Integrity and Life

Stability Improvement of Onboard HVdc Grid and Engine Using an Advanced Power Generation Center for the More-Electric Aircraft

In the high-power settings of engine, such as maximum take-off, more power should be extracted from the high-pressure spool (HPS) than the low-pressure spool (LPS) to avoid the overspeed and potential instability of the HPS. However, as revealed in this paper, extracting more power from the HPS will degrade the onboard HVDC grid stability for the conventional single dc bus power generation center. To address this conflict, in this paper, an advanced power generation center (APGC) incorporating an extra back-to-back (BTB) converter is introduced to improve the stability for both engine and HVDC grid. The BTB converter connects the HP power generation channel and the LP channel, providing an additional power flow path between the HP channel and LP channel. A transfer function-based impedance model and a state-space model of the HVDC grid are proposed to study the HVDC grid stability. The analytical findings of the HVDC grid stability and stability improvement using the APGC architecture have been verified through simulation and experimental results.

Static Stiffness Parametric Studies of Newly Conceptualized Inter Shaft Squeeze Film Damper (ISSFD) Rings

Most of the modern gas turbine engine designs requiring super critical operations will often have squeeze film dampers (SFDs) on the high pressure spool bearing as an inevitable feature. However, Squeeze Film Dampers are not very common for inter-shaft bearing applications and are still in the research and development stage. The current effort concentrates on static stiffness parametric studies based on newly conceptualized intershaft squeeze film damper (ISSFD) ring design suitable for space constrained intershaft bearing application of a typical two spool configured system. The work involves static stiffness evaluation of three and four grooves ISSFD rings by varying the groove angle/groove length/subtended angle of groove & hence the overlapping angle of grooves by conducting static tests on a dedicated static stiffness evaluation test rig designed and fabricated for the purpose. The pattern of circumferential variation of stiffness of each ring is studied and the results revealed a typical trend of decreasing stiffness in ISSFD ring with the increase in groove angle and hence the overlapping angle of grooves. The static stiffness parametric studies has resulted in developing an ideal profile of ISSFD ring having six grooves that exhibited axisymmetric pattern of circumferential variation of stiffness. When used in a space constrained intershaft bearing plane of a typical two spool gas turbine engine, the ideal ISSFD ring profile would exhibit the best damping performance. 
 Keywords: Inter Shaft Squeeze Film Damper (ISSFD), Gas Turbine Engine, Static stiffness, two spools, grooves.

A Model Following Based Real-Time Prediction and Monitoring of TFE-731 Turbofan Engine Compressors

An engine is the heart of an aircraft. It produces thrust, drives the generator, pumps the hydraulic system and provides compressed air for all the systems on the aircraft. Its health plays an essential role in flight safety. In the past, the standard operation procedure to evaluate the health status of engines usually depended on some specific parameters, like inter-stage turbine temperature, low-pressure spool rotating speed or high-pressure spool rotating speed. Once part of the parameters pass over certain safety boundaries that were previously set by the manufacturers or the operators, the engine would be regarded as an unhealthy engine. Nevertheless, in practical applications, such threshold-style mechanism cannot reflect engine fault immediately and therefore could lead to potential flight risk. To solve this issue, a precise forecast model of the engine has to be established. Consequently, this research is dedicated to develop algorithms for engine modeling as well as the identification of optimal parameters. For the TFE-731 engine, there are three section models considered, including low pressure compressor (LPC) model, high pressure compressor (HPC) model and overall turbofan dynamics model. Those models are derived with the consideration of physical isentropic compression equation as well as a data-driven regression technique. Experiments show that a precise modeling fitting can be achieved by using regression analysis and nonlinear optimal parameter estimation. Finally, to compare the prediction stability and accuracy, associated training models using neural network (NN) are also presented. Comparison studies verify that the proposed method is able to achieve stable as well as accurate TFE-731 real-time response prediction and monitoring.

Dual Drive Booster for a Two-Spool Turbofan: High Shaft Power Offtake Capability for More Electric Aircraft and Hybrid Aircraft Concepts

The effects of high shaft power offtake (POT) in a direct drive, a geared drive, and a novel turbofan configuration are investigated. A design and off-design performance analysis shows the configuration specific limitations and advantages. The more electric aircraft (MEA) concept promises to offer advantages with respect to aircraft performance, maintenance, and operating costs. The engines for the MEA concept are based on conventional turbofan architectures. These engines are designed for significantly increased shaft POT that is required by the airframe, and the shaft power is usually taken off the high-pressure (HP) spool. This can impair the off-design performance of the engine and lead to compromises during engine design and to operability limitations. Taking the power off the low-pressure (LP) spool mitigates some of the problems but has other limitations. In this work, an alternative novel turbofan architecture is investigated for its potential to avoid the problems related to high shaft POTs. This architecture is called the dual drive booster because it uses a summation gearbox to drive the booster from both the LP and HP spool. The shaft power, if taken off the booster spool, is effectively provided by both the LP and HP spools, which allows the provision of very high power levels. This new concept is benchmarked against a two-spool direct drive and a geared drive turbofan (GTF). Furthermore, it is described, how the new architecture can incorporate an embedded motor generator. The presented concept mitigates some of the problems, which are encountered during high POT in conventional configurations. In particular, the core compressors are less affected by a change in shaft POT. This allows higher POTs and gives more flexibility during engine design and operation. Additionally, the potential to use the new configuration as a gas turbine-electric hybrid engine is assessed, where electrical power boost is applied during critical flight phases. The ability to convert additional shaft power is compared with conventional configurations. Here, the new configuration also shows superior behavior because the core compressors are significantly less affected by power input than in conventional configurations. The spool speed and its variation are more suitable for electrical machines than in conventional configuration with LP spool power transfer. The dual drive booster concept is particularly suited for applications with high shaft POTs and inputs, and should be considered for propulsion of MEAs.

A nonlinear switching control strategy of regulation and safety protection for aero-engines based on the equilibrium manifold expansion model

This article proposes a nonlinear switching control strategy to solve the regulation control problem with safety constraint for aero-engines based on the equilibrium manifold expansion model. That is a kind of nonlinear model with satisfactory accuracy. The design procedure includes design of sub-controllers such as the nonlinear high-pressure spool speed tracking controller, high pressure turbine outlet temperature safety protection controller and the nonlinear dynamic feedback recovery controller. Also, switching law and dynamic controller state reset law are properly designed. Under the proposed strategy, system trajectory starting from temperature safety boundary will never return to the safety boundary. The designed reset law ensures a smooth switching from protection mode to recovery mode. Switch-off moment of the protection controller can be arbitrarily selected according to the tracking performance requirement. Freedom degree of selection is enlarged. The sub-controllers corresponding to objectives of regulation and safety protection are designed separately with more freedom degree and less conservatism. The effectiveness and the robustness of the proposed strategy are verified through a case study for a two-spool turbofan engine.

Dual Drive Booster for a Two-Spool Turbofan: Performance Effects and Mechanical Feasibility

A novel two-spool turbofan engine configuration is described which uses a booster powered by both the low an high pressure spools. Design and off-design performance analysis shows the operating characteristics of the configuration, and a mechanical feasibility study of the gearbox is presented. The trends toward ever higher engine overall pressure ratio and bypass ratio have resulted in a combination of higher pressure ratio and lower blade speed in the booster compressor of conventional two-spool turbofans. This combination gives rise to many stages in the booster and/or lower booster efficiency and also a higher degree of off-design mismatch between the core compressors. The current paper describes an engine architecture which aims to alleviate both these issues by powering the booster compressor from both low and high pressure spools through an epicyclic gear system. We have called this engine architecture the dual drive booster. The concept gives the engine designer greater flexibility to optimize component performance and work split, resulting in the potential for lower cruise specific fuel consumption and higher hot-day takeoff thrust capability than current engine configurations. The gear system is described along with the mathematical derivation of the booster rotational speed in terms of LP- and HP-spool speeds. Both the design point and off-design performance modeling have been conducted and comparison is made between a conventional turbofan and a turbofan fitted with the dual drive booster. The results show a significant enhancement in takeoff thrust due to the better speed match of the booster. The paper also describes the results of a preliminary study into the design and mechanical feasibility of the engine architecture and gear system. The presented concept is an alternative to the conventional turbofan and should be considered during the conceptual design of future aircraft engines.

Concept description and assessment of the main features of a geared intercooled reversed flow core engine

Intercooled turbofan cycles allow higher overall pressure ratios to be reached which gives rise to improved thermal efficiency. Intercooling also allows core mass flow rate to be reduced which facilitates higher bypass ratios. A new intercooled core concept is proposed in this paper which promises to alleviate limitations identified with previous intercooled turbofan designs. Specifically, these limitations are related to core losses at high overall pressure ratios as well as difficulties with the installation of the intercooler. The main features of the geared intercooled reversed flow core engine are described. These include an intercooled core, a rear-mounted high-pressure spool fitted rearwards of the low-pressure spool as opposed to concentrically as well as a mixed exhaust. In these studies, the geared intercooled reversed flow core engine has been compared with a geared intercooled straight flow core engine with a more conventional core layout. This paper compares the mechanical design of the high-pressure spools and shows how different high-pressure compressor and high-pressure turbine blade heights can affect over-tip leakage losses. In the reversed configuration, the reduction in high-pressure spool mean diameter allows for taller high-pressure compressor and turbine blades to be adopted which reduces over-tip leakage losses. The implication of intercooler sizing and configuration, including the impact of different matrix dimensions, is assessed for the reversed configuration. It was found that a 1-pass intercooler would be more compact although a 2-pass would be less challenging to manufacture. The mixer performance of the reversed configuration was evaluated at different levels of mixing effectiveness. This paper shows that the optimum ratio of total pressure in the mixing plane for the reversed flow core configuration is about 1.02 for a mixing effectiveness of 80%. Lower mixing effectiveness would result in a higher optimum ratio of total pressure in the mixing plane and fan pressure ratio.

Performance characteristics and optimisation of a geared intercooled reversed flow core engine

Intercooled turbofan cycles allow higher overall pressure ratios to be reached, which gives rise to improved thermal efficiency. In addition, intercooling allows for the size, weight and exhaust jet velocity of the core to be reduced. For an optimum jet velocity ratio and fixed thrust, the fan pressure ratio and specific thrust are also reduced, which benefits propulsive efficiency. A new intercooled core concept is proposed in this paper, which promises to alleviate limitations identified in previous intercooled turbofan designs. This concept facilitates the installation of the intercooler and reduces core losses at high overall pressure ratios. This engine concept takes advantage of intercooling and the arrangement of the high pressure spool to reach and exceed overall pressure ratios of 80. In addition, given the reduction in core size, bypass ratios beyond 14 have been considered. In order to identify efficiency gains and performance characteristics which are due to the novel arrangement alone, the geared intercooled reversed flow core engine has been compared with a geared intercooled engine with a more conventional core. Finally an optimisation exercise has been carried out to identify the best configuration for both the geared intercooled reversed flow core concept and the conventional core concept. In this paper, it is demonstrated that the geared intercooled reversed flow core concept allows for a 2.3% reduction in block fuel burn. The reductions are due to the improved core efficiency, higher overall pressure ratio as well as efficiency gains from the use of a mixed exhaust. The sensitivity analysis shows that the improvements are highly dependent on pressure losses in the core and bypass stream and that careful design of the mixer chutes and intercooler headers to achieve low losses is essential if the concept gains are to be realised.

Design of Fault Tolerant Control Laws for Jet Engines

Full Authority Digital Electronic Control (FADEC) system is common in all modern jet engines. In jet engine, the simplest control system is to produce desired engine thrust by changing the fuel flow. Since in flight engine thrust measurement is practically not possible, engine low pressure shaft rotational speed (N1/NL), engine pressure ratio (EPR), or exhaust jet temperature (EJT) has been effectively used as an indicator of the engine thrust. In this paper, the model used is twin spool turbofan engine. If any fault occurs in N1 sensor, the entire engine operation will be affected. Since there exists a unique aerodynamic relationship among the spool speeds, if any fault occurs in N1 sensor, engine thrust can be controlled with a certain amount of degradation using high pressure spool speed (N2/NH).Both soft and hard failures are detected using kalman filter, range, rate and comparison techniques. The effectiveness of the proposed approach is demonstrated by means of simulations.

Assembly Variation Identification of High Pressure Spool by Response Surface Based Model Updating Using IV-optimal Designs

This article, in order to improve the assembly of the high-pressure spool, presents an assembly variation identification method achieved by response surface method (RSM)-based model updating using IV-optimal designs. The method involves screening out non-relevant assembly parameters using IV-optimal designs and the preload of the joints is chosen as the input features and modal frequency is the only response feature. Emphasis is placed on the construction of response surface models including the interactions between the bolted joints by which the non-linear relationship between the assembly variation caused by the changes of preload and the output frequency variation is established. By achieving an optimal process of selected variables in the model, assembly variation can be identified. With a case study of the laboratory bolted disks as an example, the proposed method is verified and it gives enough accuracy in variation identification. It has been observed that the first-order response surface models considering the interactions between the bolted joints based on the IV-optimal criterion are adequate for assembly purposes.

Modeling and Performance Analysis of the Rolls-Royce Fuel Cell Systems Limited: 1 MW Plant

This paper is focused on the performance of the 1 MW plant designed and developed by Rolls-Royce Fuel Cell Systems Limited. The system consists of a two stage turbogenerator coupled with pressure vessels containing the fuel cell stack, internal reformer, cathode ejector, anode ejector, and off-gas burner. While the overall scheme is relatively simple, due to the limited number of components, the interaction between the components is complex and the system behavior is determined by many parameters. In particular, two important subsystems such as the cathode and the anode recycle loops must be carefully analyzed also considering their interaction with and influence on the turbogenerator performance. The system performance model represents the whole, and each physical component is modeled in detail as a subsystem. The component models have been validated or are under verification. The model provides all the operating parameters in each characteristic point of the plant and a complete distribution of thermodynamics and chemical parameters inside the solid oxide fuel cell (SOFC) stack and reformer. In order to characterize the system behavior, its operating envelope has been calculated taking into account the effect of ambient temperature and pressure, as described in the paper. Given the complexity of the system, various constraints have to be considered in order to obtain a safe operating condition not only for the system as a whole but also for each of its parts. In particular each point calculated has to comply with several constraints such as stack temperature distribution, maximum and minimum temperatures, and high and low pressure spool maximum rotational speeds. The model developed and the results presented in the paper provide important information for the definition of an appropriate control strategy and a first step in the development of a robust and optimized control system.

Implications of Turbine Erosion for an Aero-Engine’s High-Pressure-Turbine Blade’s Low-Cycle-Fatigue Life-Consumption

Abstract Some in-service deterioration in any mechanical device, such as a military aero-engine, is inevitable. As a result of experiencing any deterioration, an aero-engine will seek a different steady operating point thereby resulting in a variation in the high-pressure spool speeds in order to provide the same thrust to keep aircraft’s performance invariant. Any increase in the high-pressure spool speed results in greater low-cycle fatigue damage for the hot-end components and thereby higher engine’s life-cycle costs. Possessing better knowledge (of the impacts of high-pressure turbine’s erosion upon the low-cycle fatigue life-consumption of aero-engine’s hot-end components) helps the users to take wiser management decisions. For a military aircraft’s mission profile, using bespoke computer simulations, the impacts of turbine erosion for high-pressure turbine-blade’s low-cycle fatigue life-consumption have been predicted.

Implications of Day Temperature for a High-Pressure-Turbine Blade's Low-Cycle-Fatigue Life Consumption

H IGH-PERFORMANCE aircraft used in modern aviation, especially for military purposes, are complex in design and are required to operate under severe stresses and temperatures [1]. Thus, designer and users of these aircraft continually seek greater reliability, increased availability, enhanced performance, improved safety, and low life-cycle costs. Any extensions of the life expectations of aeroengines directly lower the life-cycle cost and depend upon the types of mission undertaken, operating conditions experienced, and rate of in-service engine deterioration. Significant adverse effects of in-service aeroengine deterioration have been already established by earlier investigations [2–4]. However, as the result of a comprehensive literature review, it was realized that in addition to aeroengine deterioration, the condition of incoming airstream is also expected to have significant effects on an aeroengine’s overall performance. In a country such as Pakistan, variation of day temperature over the year (i.e., from 0 to 50 C) significantly alters the condition of the incoming airstream. Any change in day temperature affects the quality of the incoming airstream for aeroengine operation as well as that of the air surrounding the air vehicles, thereby affecting the performance of both by altering atmospheric and aerodynamic characteristics [5]. As a result, an aeroengine will seek a different steady operating point, thereby resulting in a variation of the high-pressure spool speed (HS) for providing the same thrust to keep the aircraft’s performance invariant. Any rise in the HS results in greater low-cycle-fatigue (LCF) damage for the hot-end components and thereby higher engine life-cycle costs. Possessing a better knowledge of the impacts of day temperature upon the LCF life consumption of an aeroengine’s hot-end components helps the users to take wiser management decisions. Hence, this investigation was undertaken. There are many components in an aeroengine, but its performance is highly sensitive to changes in only a few. Among these, the highpressure turbine (HPT) blades are the most sensitive components, because they are subjected to both the highest rotating speeds and gas temperatures [2], and so they were selected for the present investigation.

HYBRID GLOBAL OPTIMIZATION TECHNIQUE FOR OUTPUT-ERROR IDENTIFICATION OF FRACTIONAL ORDER MODEL - APPLICATION TO AERO GAS TURBINE PLANT

A hybrid global optimization technique for identifying an output-error fractional order model is proposed. The proposed technique combines local and interval global optimization methods. It is capable of estimating the fractional derivative orders and the parameters, with guaranteed convergence. The technique is applied to identify a fractional order model for a twin-spool aero gas turbine plant. The input variable is the fuel rate to the gas turbine and the output variable is the high pressure (HP) spool speed. The operating regime is at 90% HP spool speed demand. It is found that the fractional order model identified with the proposed technique gives a much better fit to the temporal data than higher integer order models.

Measurement and Computation of Heat Transfer in High-Pressure Compressor Drum Geometries With Axial Throughflow

This paper makes comparisons between CFD computations and experimental measurements of heat transfer for the axial throughflow of cooling air in a high-pressure compressor spool rig and a plane cavity rig. The heat transfer measurements are produced using fluxmeters and by the conduction solution method from surface temperature measurements. Numerical predictions are made by solving the Navier–Stokes equations in a full three-dimensional, time-dependent form using the finite-volume method. Convergence is accelerated using a multigrid algorithm and turbulence modeled using a simple mixing length formulation. Notwithstanding systematic differences between the measurements and the computations, the level of agreement can be regarded as promising in view of the acknowledged uncertainties in the experimental data, the limitations of the turbulence model and, perhaps more importantly, the modest grid densities used for the computations.

Eliminating the TF30 P-111 + engine rotor-instability problem

The TF30 P-lll + engine developed a high-pressure spool rotor dynamic instability when it went into production in mid-1986. Vibration rejection rates were as high as 50% until the instability was eliminated with the incorporation of an oil-film damper at the high-pressure turbine (HPT) bearing. This paper focuses on the analytical treatment of the instability and includes a summary of engine testing that was done to help diagnose the problem. Correlation between test and analysis implicates the aerodynamic cross-coupled forces from the HPT system as the destabilizing mechanism. Results from the stability model are presented that show the damper completely suppressing the instability as later confirmed by ground-level tests. This experience points out the subtle characteristic of rotor instability and the need for improved quantification of the destabilizing mechanisms that produce it. The predecessor engine model had been in service for years without the problem, yet it developed a rotor instability with the incorporation of a new HPT that had no obvious impact on engine rotor dynamics.