Low Specific Thrust Research Articles

Fan Stage Design and Performance Optimization for Low Specific Thrust Turbofans

In modern turbofan engines, the bypass section of the fan stage alone provides the majority of the total thrust required in cruise, and the size of the fan has a considerable effect on the overall engine weight and nacelle drag. Thrust requirements in different parts of the flight envelope must also be satisfied together with sufficient margins towards stalling. An accurate description of the interdependencies between the relevant performance and design attributes of the fan stage alone—such as efficiency, surge margin, fan-face Mach number, stage loading, flow coefficient, and aspect ratio—are therefore necessary to estimate system-level objectives such as mission fuel burn and the direct operating cost with enough confidence during the conceptual design phase. The contribution of this study is to apply a parametric optimization approach to the conceptual design of fan stages for low specific thrust turbofans based on the streamline curvature method. Trade-offs between fan stage attributes for Pareto-optimal solutions are modeled by training Kriging surrogate models on the results from the parametric optimization. A case study is provided in the end to demonstrate the potential implications of including a higher level of fan-stage parameter interdependency in an engine systems model. Results implied that being able to predict the rotor solidity required to maintain a given average blade loading—in addition to stage efficiency—is of significant importance when it comes to evaluating the trade-off between engine weight and thrust-specific fuel consumption.

Structural optimization of engine nacelle fan cowl door hinge

To increase propulsive efficiency, future nacelle engines are projected to run at low specific thrust with large bypass ratios. Typically, this entails a larger fan diameter and nacelle, as well as increased drag and weight penalties. As a result, there is a need to construct more compact, nacelles’ components in comparison to present designs in order to minimize drag and weight. These designs are naturally more difficult, necessitating the use of a system to explore and identify the design area that is possible. Designing a nacelle is difficult due to the wide range of operational circumstances. This paper provides a multi-objective optimization technique for fan cowl door component of nacelle design utilizing an evolving genetic algorithm. A collection of geometry definitions utilizing Class Shape Transformations, automated simulation and analysis, a generic algorithm, assessments at various nacelle operating circumstances, and the addition of extra aerodynamic restrictions are all part of the unique framework. This framework was used to look into the nacelle design space for nacelle components. For the new nacelle component design challenge, multi-objective optimization was successfully shown, and the entire system was proved to allow the discovery of the feasible nacelle design space. Moreover, in this paper a series of fan cowl door hinge is designed from the scratch and optimized step by step for reducing the weight of the component without compromising for the stress capacity and load transformations within the allowable limit.

Assessment of Performance Boundaries and Operability of Low Specific Thrust GUHBPR Engines for EIS2025

Abstract This paper aims to develop a robust design process by approaching the performance boundaries and evaluating the operability of the pursued geared turbofan (GTF) engine with low specific thrust for entry-into-service (EIS) 2025. A two-spool direct-drive turbofan (DDTF) engine of EIS 2000 was improved according to aircraft specifications and technology boundaries in 2025. A series of optimized engines with consecutive fan diameters were established to seek the ideal engine by balancing specific fuel consumption (SFC), weight, and mission fuel burn. The fan diameter was proved to be a decisive factor for lowering SFC and energy usage. The cycle design optimization process achieved a thermal efficiency of approximately 52%, and a propulsive efficiency of 79.5%, which is 8.19% increase in propulsive efficiency by enlarging fan diameter from 1.6 m to 1.9 m. Meanwhile, the 1.9 m-fan diameter engine achieved a reduction in SFC and fuel burn of 7.47% and 6.58%, respectively, which offers an overall reduction of 30.82% in block fuel burnt and CO2 emission compared to the DDTF engine. A feasibility check verified the viability of the designed optimum engine in terms of fan tip speed, stage loading, and AN2. Dynamic simulation offered a deep understanding of transient behavior and fundamental mechanism of the GTF engine. An important aspect of this paper is the use of advanced ceramic matrix composite (CMC) materials, which led to an improvement of 4.92% in block fuel burn and 2.93% in engine weight.

Exergo-economic analysis of off-design a target drone engine for reconnaissance mission flight

Nowadays, small-size turbojet engines play a significant role on various military and civil missions. In this study, exergoeconomic analysis of TRS-18 turbojet engine was conducted for nine flight phases. First aim is to find out differences between exergoeconomic behaviour of the components originated from their both irreversibility and purchase costs. The second objective is to reveal effects of flight phases on exergoeconomic performance of TRS-18 and its components by employing several exergoeconomic indicators regarding the TRS18. These indicators are exergy destruction cost, relative cost difference, exergoeconomic factor and specific thrust cost. To estimate cost-ineffective components of TRS-18 engine, component-based exergoeconomic assessment was carried out for each mission. Based on results of exergoeconomic analysis for the whole TRS-18 engine, specific thrust cost of the TRS-18 was found at relatively high values ranging 398.2 and 422.04 $/h.kN at loiter III and IV phases where the engine operates at partial power whereas taxi-out, take-off and landing (TTL) phases have the relatively low specific thrust cost values ranging 159.53 and 166 $/h.kN. Moreover, exergoeconomic factor of the whole TRS-18 was found as relatively low values with approximately 21% at TTL phases. Considering component-based results, the lowest exergoeconomic factor, ranging from 25% to 40%, belongs to combustor, whereas gas turbine has relatively high exergoeconomic factor, varying from 89% to 92%. Finally, for the combustor, exergy destruction cost was determined at the relatively highest rates for whole flight phases. It changes between 35.74 $/h at loiter-II phase and 69.79 $/h at take-off. It is thought this study can help in finding cost-effective flight phases by means of comprehending the nature of the relationship between flight phases and the corresponding exergoeconomic parameters. Also, off-design exergo-economy obtained with this paper may be applied to the other gas turbine engines.

Surrogate models for the prediction of the aerodynamic performance of exhaust systems

The aerodynamic performance of the exhaust system is becoming more important in the design of engines for civil aircraft applications. To increase propulsive efficiency and reduce specific fuel consumption, it is expected that future engines will operate with higher bypass ratios, lower fan pressure ratios and lower specific thrust. At these operating conditions, the net thrust and the specific fuel consumption are more sensitive to losses in the exhaust. Thus the performance of the exhaust needs to be accurately assessed as early as possible during the design process. This research investigates low-order models for the prediction of the performance of separate-jet exhaust systems, as a function of the free-stream Mach number, the fan nozzle pressure ratio and the extraction ratio (fan to core pressure ratio). In the current practice the two nozzles are typically considered in isolation and the performance is modelled as a function of their pressure ratio. It is shown that the additional degrees of freedom have a substantial impact on the metrics describing the performance of the exhaust system. These models can be employed at a preliminary design stage coupled with engine performance models, which require as input the characteristics of the exhaust system. Two engines, which are representative of current and future large turbofan architectures are studied. The low-order models investigated, generalized Kriging and radial basis functions, are constructed based on data obtained with computational fluid dynamics simulations. The data represents the characteristics of the exhaust of each engine, and they are provided for the first time for a wide operational envelope. The influence on accuracy of the type of surragate model and its settings have been quantified. Furthermore, the trade-off between the accuracy of the model and the number of samples has been identified. It is found that the exhaust performance metrics can be modelled using a low-order model with sufficient accuracy. Recommendations on the best settings of the model are also provided.

Multi-objective optimisation of short nacelles for high bypass ratio engines

Future turbo-fan engines are expected to operate at low specific thrust with high bypass ratios to improve propulsive efficiency. Typically, this can result in an increase in fan diameter and nacelle size with the associated drag and weight penalties. Therefore, relative to current designs, there is a need to develop more compact, shorter nacelles to reduce drag and weight. These designs are inherently more challenging and a system is required to explore and define the viable design space. Due to the range of operating conditions, nacelle aerodynamic design poses a significant challenge. This work presents a multi-objective optimisation approach using an evolutionary genetic algorithm for the design of new aero-engine nacelles. The novel framework includes a set of geometry definitions using Class Shape Transformations, automated aerodynamic simulation and analysis, a genetic algorithm, evaluations at various nacelle operating conditions and the inclusion of additional aerodynamic constraints. This framework has been applied to investigate the design space of nacelles for high bypass ratio aero-engines. The multi-objective optimisation was successfully demonstrated for the new nacelle design challenge and the overall system was shown to enable the identification of the viable nacelle design space.

Numerical Study on Aeroelastic Instability for a Low-Speed Fan

Over recent years, engine designs have moved increasingly toward low specific thrust cycles to deliver significant specific fuel consumption (SFC) improvements. Such fan blades may be more prone to aerodynamic and aeroelastic instabilities than conventional fan blades. The aim of this paper is to analyze the flutter stability of a low-speed/low pressure ratio fan blade. By using a validated computational fluid dynamics (CFD) model (AU3D), three-dimensional unsteady simulations are performed for a modern low-speed fan rig for which extensive measured data are available. The computational domain contains a complete fan assembly with an intake duct and the downstream outlet guide vanes (OGVs), which is a whole low-pressure (LP) domain. Flutter simulations are conducted over a range of speeds to understand flutter characteristics of this blade. Only the first flap (1F) mode is considered in this work. Measured rig data obtained by using the same fan set but with two different lengths of the intake showed a significant difference in the flutter boundary for the two intakes. AU3D computations were performed for both intakes and were used to explain this difference between the two intakes, and showed that intake reflections play an important role in flutter of this blade. This observation indicates that the experiment with the long intake used for the performance test may be misleading for flutter. In the next phase of this work, two possible modifications for increasing the flutter margin of the fan blade were explored: changing the mode shape of the blade and using acoustic liners in the casing. The results show that it is possible to increase the flutter margin of the blade by either decreasing the ratio of the twisting to plunging motion in 1F mode or by introducing deep acoustic liners in the intake. The liners have to be deep enough to attenuate the flutter pressure waves and hence influence the stability. The results indicate the importance of reflection in flutter stability of the fan blade and clearly show that intake duct needs to be included in flutter study of any fan blade.

Implications of engine deterioration for creep life

World developments have led the armed forces of many countries to become more aware of how their increasingly-stringent financial budgets are spent. One major expenditure for military authorities is upon aero-engines. Deterioration of an engine generally results in a lower specific thrust and higher specific fuel-consumption for the same spool speeds and turbine-entry temperatures. In order to meet the thrust requirement for the same aircraft’s performance, the engines are run at higher spool speeds and/or turbine-entry temperatures. This results in greater rates of fuel consumption, and a shorter life (due to the greater creep damage and low-cycle fatigue incurred). Possessing a better knowledge, of the impacts of engine deterioration on an aircraft’s operational performance as well as its fuel and residual life-expectancy, helps the users to improve engine utilization.

Use of swirl for flow control in propulsion nozzles

In certain off-design conditions it is desirable to vary the effective area of turbofan propulsion nozzles, e.g., to aid component matching or to reduce fan noise on approach. It has been suggested that the mass flux reductions associated with swirl could be used to achieve such variations. This paper addresses the questions of whether swirl can achieve the required mass flux reductions and whether this can be done with minimal thrust losses. A quasicylindrical theory is used to analyze the swirling flow, and comparisons are made with variablearea nozzles for specific case studies. For a low-specific thrust engine of high bypass ratio (BPR = 12) the required fan nozzle mass flux reduction cannot be achieved using swirl, and in any case, swirl gives a far faster rate of decay of thrust with fan nozzle mass flux than does a variable-area nozzle. For a ducted fan of BPR -4, a 10% reduction in core nozzle mass flux can be achieved with swirl for slightly less thrust loss than with a variable-area nozzle. The swirl profile, however, needs to be carefully optimized.

Energy Conserving Aircraft From The Engine Viewpoint

The increasing cost of hydrocarbon fuel inevitably intensifies the quest for lower engine SFC on the conventional aero gas turbine. This should not obscure the fact that minimising all aircraft direct operating costs is the ultimate yardstick for the engine designer. Higher fuel prices may change priorities in engine design and justify more complex and expensive engines particularly for longer range operations. Optimum engine design for shorter range can be significantly different because of the implication of cyclic life on air cooled turbine blades. Lower specific thrust engines are worthy of close consideration particularly for short haul operation. More effort should be devoted to comprehensive studies of new powerplant concepts.

Glow Discharge as an Advanced Propulsion Device

Abstract : The generation of thrust directly from electric power through a gaseous discharge is investigated with a one dimensional model. The thrust consists of 2 parts, the electric pressure and the electric wind. In terms of the ionization constants of gases and of the mobility constants of ions and electrons, both the electric pressure and the electric wind are obtained in simple dimensionless form. At low gas pressures, the thrust contribution from electric pressure is usually much larger than that from electric wind. The latter may be of comparable magnitudes or even much larger at higher (atmospheric) discharge press res and larger inter-electrode distances. The over-all thrust level is proportional to the gas pressure in the discharge (P) squared. For p = 1 mm Hg, the thrust is less than or about equal to 1 dyne/ sq cm. The specific thrust/unit power is no greater than 10 mg wt/kw. Thrust is derived from electric wind at a much larger specific power consumption. The low specific thrust rules out using such a propulsion system as a booster. For application in outer space, it suffers from the limited specific impulse because propellant gas must be carried aboard. The system has other interesting applications such as a positive control device for hypersonic vehicles at an altitude of 100-200 x 10,000 ft.

Some Aspects of Engine Noise

Although a great deal has been said and written in the past eight years or so about engine noise, previous discussion has been very largely concerned with the physical aspects of jet noise and with the possibility of suppressing it by means of special propelling nozzles and ground mufflers. There is however another, and in general a better, way of dealing with jet noise, namely by using engines of low specific thrust, i.e. of low exhaust velocity and one purpose of the paper is to examine the possibilities of such engines. This leads to a consideration of blade noise from fans and turbines for as jet noise is reduced this tends to become the dominant factor. But first of all it is necessary to re-examine the nature of the noise problem.