This paper presents the development of an integrated approach which targets the aerodynamic design of separate-jet exhaust systems for future gas-turbine aero-engines. The proposed framework comprises a series of fundamental modeling theories which are applicable to engine performance simulation, parametric geometry definition, viscous/compressible flow solution, and design space exploration (DSE). A mathematical method has been developed based on class-shape transformation (CST) functions for the geometric design of axisymmetric engines with separate-jet exhausts. Design is carried out based on a set of standard nozzle design parameters along with the flow capacities established from zero-dimensional (0D) cycle analysis. The developed approach has been coupled with an automatic mesh generation and a Reynolds averaged Navier–Stokes (RANS) flow-field solution method, thus forming a complete aerodynamic design tool for separate-jet exhaust systems. The employed aerodynamic method has initially been validated against experimental measurements conducted on a small-scale turbine powered simulator (TPS) nacelle. The developed tool has been subsequently coupled with a comprehensive DSE method based on Latin-hypercube sampling. The overall framework has been deployed to investigate the design space of two civil aero-engines with separate-jet exhausts, representative of current and future architectures, respectively. The inter-relationship between the exhaust systems' thrust and discharge coefficients has been thoroughly quantified. The dominant design variables that affect the aerodynamic performance of both investigated exhaust systems have been determined. A comparative evaluation has been carried out between the optimum exhaust design subdomains established for each engine. The proposed method enables the aerodynamic design of separate-jet exhaust systems for a designated engine cycle, using only a limited set of intuitive design variables. Furthermore, it enables the quantification and correlation of the aerodynamic behavior of separate-jet exhaust systems for designated civil aero-engine architectures. Therefore, it constitutes an enabling technology toward the identification of the fundamental aerodynamic mechanisms that govern the exhaust system performance for a user-specified engine cycle.

No abstract

The aerodynamic performance of the bypass exhaust system is key to the success of future civil turbofan engines. This is due to current design trends in civil aviation dictating continuous improvement in propulsive efficiency by reducing specific thrust and increasing bypass ratio (BPR). This paper aims to develop an integrated framework targeting the automatic design optimization of separate-jet exhaust systems for future aero-engine architectures. The core method of the proposed approach is based on a standalone exhaust design tool comprising modules for cycle analysis, geometry parameterization, mesh generation, and Reynolds-averaged Navier–Stokes (RANS) flow solution. A comprehensive optimization strategy has been structured comprising design space exploration (DSE), response surface modeling (RSM) algorithms, as well as state-of-the-art global/genetic optimization methods. The overall framework has been deployed to optimize the aerodynamic design of two civil aero-engines with separate-jet exhausts, representative of current and future engine architectures, respectively. A set of optimum exhaust designs have been obtained for each investigated engine and subsequently compared against their reciprocal baselines established using the current industry practice in terms of exhaust design. The obtained results indicate that the optimization could lead to designs with significant increase in net propulsive force, compared to their respective notional baselines. It is shown that the developed approach is implicitly able to identify and mitigate undesirable flow-features that may compromise the aerodynamic performance of the exhaust system. The proposed method enables the aerodynamic design of optimum separate-jet exhaust systems for a user-specified engine cycle, using only a limited set of standard nozzle design variables. Furthermore, it enables to quantify, correlate, and understand the aerodynamic behavior of any separate-jet exhaust system for any specified engine architecture. Hence, the overall framework constitutes an enabling technology toward the design of optimally configured exhaust systems, consequently leading to increased overall engine thrust and reduced specific fuel consumption (SFC).

This paper presents an integrated approach, targeting the comprehensive assessment of combined helicopter–engine designs, within designated operations. The developed methodology comprises a series of individual modeling theories, each applicable to a different aspect of helicopter flight dynamics and performance. These relate to rotor blade modal analysis, three-dimensional flight path definition, flight dynamics trim solution, aeroelasticity and engine performance. The individual mathematical models are elaborately integrated within a numerical procedure, solving for the total mission fuel consumption. The overall simulation framework is applied to the performance analysis of the Aérospatiale SA330 helicopter within two generic, twin-engine medium helicopter missions. An extensive comparison with flight test data on main rotor trim controls, power requirements and unsteady blade structural loads is presented. It is shown that, for the typical range of operating conditions encountered by modern twin-engine medium civil helicopters, the effect of operational altitude on fuel consumption is predominantly influenced by the corresponding effects induced on the engine, rather than on airframe–rotor performance. The implications associated with the implicit coupling between aircraft and engine performance, are discussed in the context of mission analysis. The potential to comprehensively evaluate integrated helicopter–engine systems within complete three-dimensional operations, using modeling fidelity designated for main rotor design applications, is demonstrated. The proposed method essentially constitutes an enabler in terms of focusing the rotorcraft design process on designated operation types, rather than on specific sets of flight conditions.

This paper presents a numerical formulation targeting the rapid estimation of natural vibration characteristics of helicopter rotor blades. The proposed method is based on application of Lagrange's equation of motion to the kinematics of blade flap/lag bending and torsion. Modal properties obtained from Bernoulli-Euler beam and classical torsional vibration theory, are utilised as assumed deformation functions in order to estimate the time variations of strain and kinetic energy for each degree of freedom. Integral expressions are derived, describing the generalised centrifugal force and torsional moment acting on the blade in terms of normal coordinates, for flap/ lag transverse displacement and torsional deformation. Closed form expressions are provided for the direct analysis of hingeless, freely-hinged and spring-hinged articulated rotor blades. Results are presented in terms of natural frequencies and mode shapes for two small-scale rotor blade models. Extensive comparisons are carried out with experimental measurements and nonlinear finite element analysis. Predictions of resonant frequencies are also presented for two full-scale rotor blade models and the results are compared with established multi-body dynamics analysis methods. It is shown that, the proposed approach exhibits excellent numerical behaviour with low computational cost and definitive convergence characteristics. The comparisons suggest very good and in some cases excellent accuracy levels, especially considering the method's simplicity, computational efficiency, and ease of implementation.

The increasing demands for air-taxi operations together with the ambitious targets for reduced environmental impact have driven significant interest in alternative rotorcraft architectures and propulsion systems. The design of Hybrid-Electric Propulsion Systems (HEPSs) for rotorcraft is seen as being able to contribute to those goals. This work aims to conduct a comprehensive design and trade-off analysis of hybrid powerplants for rotorcraft, targeting enhanced payload-range capability and fuel economy. An integrated methodology for the design, performance assessment and optimal implementation of HEPSs for conceptual rotorcraft has been developed. A multi-disciplinary approach is devised comprising models for rotor aerodynamics, flight dynamics, HEPS performance and weight estimation. All models are validated using experimental or flight test data. The methodology is deployed for the assessment of a hybrid-electric tilt-rotor, modelled after the NASA XV-15. This work targets to provide new insight in the preliminary design and sizing of optimally designed HEPSs for novel tilt-rotor aircraft. The paper demonstrates that at present, current battery energy densities (250Wh/kg) severely limit the degree of hybridization if a fixed useful payload and range are to be achieved. However, it is also shown that if advancements in battery energy density to 500Wh/kg are realized, a significant increase in the level of hybridization and hence reduction of fuel burned and carbon output relative to the conventional configuration can be attained. The methodology presented is flexible enough to be applied to alternative rotorcraft configurations and propulsion systems.

This paper presents a mathematical model for the simulation of rotor blade flexibility in real-time helicopter flight dynamics applications that also employs sufficient modeling fidelity for prediction of structural blade loads. A matrix/vector-based formulation is developed for the treatment of elastic blade kinematics in the time domain. A novel, second-order-accurate, finite-difference scheme is employed for the approximation of the blade motion derivatives. The proposed method is coupled with a finite-state induced-flow model, a dynamic wake distortion model, and an unsteady blade element aerodynamics model. The integrated approach is deployed to investigate trim controls, stability and control derivatives, nonlinear control response characteristics, and structural blade loads for a hingeless rotor helicopter. It is shown that the developed methodology exhibits modeling accuracy comparable to that of non-real-time comprehensive rotorcraft codes. The proposed method is suitable for real-time flight simulation, with sufficient fidelity for simultaneous prediction of oscillatory blade loads.Nomenclature A flap/lag/torsion Lagrange

This work focuses on the development and application of a generic methodology targeting the design of optimum rotorcraft operations in terms of fuel burn, gaseous emissions and ground noise impact. An integrated tool capable of estimating the performance and emitted noise of any defined rotorcraft configuration within any designated mission has been deployed. A comprehensive and cost-effective optimization strategy has been structured. The methodology has been applied to two generic–baseline missions representative of current rotorcraft operations. Optimally designed operations for fuel burn, gaseous emissions and ground noise impact have been obtained. A comparative evaluation has been waged between the acquired optimum designs. The respective trade-off arising from the incorporation of flight paths optimized for different objectives has been quantified. Pareto front derived models for fuel burn and emitted noise have been structured for each mission. The Pareto models have been subsequently deployed for the design of operations optimized in a multidisciplinary manner. The results have shown that the proposed methodology is promising with regards to achieving simultaneous reduction in fuel burn, gaseous emissions and ground noise impact for any defined mission. The obtainable reductions are found to be dependent on the designated mission. Finally, the potential to design optimum operations in a multidisciplinary fashion using only a single design criterion is demonstrated.