Browsing by Author "Grech, Nicholas"

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    Aerodynamic design of separate-jet exhausts for future civil aero engines, Part I: parametric geometry definition and CFD approach
    (ASME, 2016-03-15) Goulos, Ioannis; Stankowski, Tomasz; Otter, John; MacManus, David G.; Grech, Nicholas; Sheaf, Christopher
    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 axi-symmetric 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 (LHS). 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 sub-domains 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 towards the identification of the fundamental aerodynamic mechanisms that govern the exhaust system performance for a user-specified engine cycle
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    Aerodynamic design of separate-jet exhausts for future civil aero-engines, Part II: design space exploration, surrogate modeling, and optimization
    (American Society of Mechanical Engineers, 2016-03-15) Goulos, Ioannis; Otter, John; Stankowski, Tomaz; MacManus, David; Grech, Nicholas; Sheaf, Christopher
    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).
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    Aerodynamic interference for aero-engine installations
    (AIAA, 2016-01-02) Stankowski, Tomasz P.; MacManus, David G.; Sheaf, Christopher; Grech, Nicholas
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    Aerodynamics of aero-engine installation
    (AIAA, 2016-01-02) Stankowski, Tomasz P.; MacManus, David G.; Sheaf, Christopher; Grech, Nicholas
    Small internal combustion engines, particularly those ranging in power from 1 kW to 10 kW, propel many remotely piloted aircraft (RPA) platforms that play an increasingly significant role in the Department of Defense. Efficiency of these engines is low compared to conventional scale engines and thermal losses are a significant contributor to total energy loss. Existing thermal energy loss models are based on data from much larger engines. Whether these loss models scale to the engine size class of interest, however, has yet to be established. The Small Engine Research Bench (SERB) was used to measure crank angle resolved gas temperature inside the combustion chamber of a small internal combustion engine (ICE). A 55 cc, two stroke, spark-ignition ICE was selected for this study. The engine was modified for optical analysis using sapphire rods 1.6 mm in diameter on opposite sides of the combustion chamber. The engine modification was found to have no measurable impact on indicated mean effective pressure or heat rejection through the cylinder. FTIR absorption thermometry was used to collect mid-infrared absorption spectra. The FTIR was allowed to scan continuously while simultaneously recording the scanning mirror position and crank angle associated with each data point, then data was re-sorted by crank angle. Measured spectra were compared with lines generated using CDSD-4000 and HITEMP line list databases. The line of best fit corresponded to the mean gas temperature through the combustion chamber. In this way temperature was determined as a function of crank angle for three operating conditions: 4,300, 6,000, and 7,500 revolutions per minute, all at wide open throttle. High cycle-to-cycle variation in the regions of combustion and gas exchange degraded temperature measurements at the affected crank angles. Future research will attempt to improve signal to noise in these measurements.
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    Design optimisation of separate-jet exhausts for the next generation of civil aero-engines
    (ISABE, 2017-09-08) Goulos, Ioannis; Otter, John J.; Stankowski, Tomasz; MacManus, David G.; Grech, Nicholas; Sheaf, Christopher
    This paper presents the development and application of a computational framework for the aerodynamic design of separate-jet exhaust systems for Very-High-Bypass-Ratio (VHBR) gas-turbine aero-engines. An analytical approach is synthesised comprising a series of fundamental modelling methods. These address the aspects of engine performance simulation, parametric geometry definition, viscous/compressible flow solution, design space exploration, and genetic optimisation. Parametric design is carried out based on minimal user-input combined with the cycle data established using a zero-dimensional (0D) engine analysis method. A mathematical approach is developed based on Class-Shape Transformation (CST) functions for the parametric geometry definition of gas-turbine exhaust components such as annular ducts, nozzles, after-bodies, and plugs. This proposed geometry formulation is coupled with an automated mesh generation approach and a Reynolds Averaged Navier–Stokes (RANS) flow-field solution method, thus forming an integrated aerodynamic design tool. A cost-e ective Design Space Exploration (DSE) and optimisation strategy has been structured comprising methods for Design of Experiment (DOE), Response Surface Modelling (RSM), as well as genetic optimisation. The integrated framework has been deployed to optimise the aerodynamic performance of a separate-jet exhaust system for a large civil turbofan engine representative of future architectures. The optimisations carried out suggest the potential to increase the engine’s net propulsive force compared to a baseline architecture, through optimum re-design of the exhaust system. Furthermore, the developed approach is shown to be able to identify and alleviate adverse flow-features that may deteriorate the aerodynamic behaviour of the exhaust system.
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    Gas turbine sub-idle performance modelling : groundstart altitude relight, and windmilling
    (Cranfield University, 2013-01) Grech, Nicholas; Pachidis, Vassilios; Zachos, Pavlos K.; Rowe, Arthur; Brown, Steve; Tunstall, Richard
    Engine performance modelling is a major part of the engine design process, in which specialist solvers are employed to predict, understand and analyse the engine’s behaviour at various operating conditions. Sub-idle whole engine performance synthesis solvers are not as reliable and accurate as design point solvers. Lack of knowledge and data result in component characteristics being reverse-engineered or extrapolated from above-idle data. More stringent requirements on groundstart and relight capabilities, has prompted the need to advance the knowledge on low-speed engine performance, thereby requiring more robust sub-idle performance synthesis solvers. The objective of this study, was to improve the accuracy and reliability of a current aero gas turbine sub-idle performance solver by studying each component in isolation through numerical simulations. Areas researched were: low-speed and locked-rotor com- pressor characteristics, low-power combustion efficiency, air blast atomizer and combustor performance at sub-idle, torque-based whole engine sub-idle performance synthesis, and mixer performance at far off-design conditions. The observations and results from the numerical simulations form the contribution to knowledge of this research. Numerical simulations of compressor blades under highly negative incidence angles show the complex nature of the flow, with the results used to determine a suitable flow deviation model, a method to extract blade aerodynamic char- acteristics in highly separated flows, and measure the blockage caused by highly separated flow with operating condition and blade geometry. The study also concluded that the use of Blade Element Theory is not accurate enough to be used at such far off-design con- ditions. The linearised parameter-based whole engine performance solver was converted to used torque-based parameters, which validated against engine test data, shows that it is suitable for low-power simulations with the advantage of having the potential to start engine simulations from static conditions. A study of air-blast atomization at windmilling relight conditions has shown that current established correlations used to predict spray characteristics are not suitable for altitude relight studies, tending to overestimate the atomization quality. Also discovered is the highly influential interaction of compressor wakes with the combustor and atomizer under altitude relight conditions, resulting in more favourable lighting conditions than previous assumptions and models have shown. This is a completely new discovery which will result in a change in the way combustors are designed and sized for relight conditions, and the way combustion rig tests are conducted. The study also has valuable industrial contributions. The locked-rotor numerical data was used within a stage-stacking compressible flow code to estimate the compressor sub- idle map, of which results were used within a whole engine performance solver and results validated against actual engine test data. The atomization studies at relight were used to factor in the insensitivity of current spray correlations, which together with a newly de- veloped sub-idle combustion efficiency sub-routine, are used to determine the combustion efficiency at low-power settings. The interaction of compressor wakes with the atomizer showed that atomizer performance at relight is underestimated, resulting in oversized combustors. By using the knowledge gained within this research, combustor size can be reduced, resulting in lower NOx at take-off and a smaller and lighter core, with a com- bustor requiring less cooling air. The component research has advanced the knowledge and modelling capability of sub-idle performance solvers, increasing their reliability and encouraging their use for future aero gas turbine engines.
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    Nacelle design for ultra-high bypass ratio engines with CFD based optimisation
    (Elsevier, 2020-09-09) Robinson, Matthew; MacManus, David G.; Christie, Robert; Sheaf, Christopher; Grech, Nicholas
    As the size of aero-engines has increased in recent years, the need for slimmer and shorter nacelles has become more pressing. A more aggressive design space must therefore be explored for nacelle designs which are expected to perform worse in the off design conditions such as spillage than current nacelle designs. In this work, a novel design space has been explored through the use of an optimisation method which evaluated nacelle aerodynamic performance based on computational fluid dynamics simulations. A multi-objective optimisation was undertaken where cruise drag, drag rise Mach number, spillage drag and two metrics based on the pressure distribution of the nacelle were optimised. Comparable optimal designs were picked from the Pareto sets of optimisations carried out at different nacelle lengths and radial offsets and some key outcomes established from their aerodynamics and geometries. It was determined that a reduction in the length of the nacelle from 3.8 highlight radii to 3.1 radii resulted in a significantly worse aerodynamic performance which included an increase in peak surface isentropic Mach number at cruise of 0.1 and up to four times as much spillage drag. It was however also established from the optimisation results that as the required drag rise Mach number was decreased the overall performance of short nacelles improved significantly.
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    Parametric design of non-axisymmetric separate-jet aero-engine exhaust systems
    (Elsevier, 2019-07-16) Otter, John J.; Christie, Robert; Goulos, Ioannis; MacManus, David G.; Grech, Nicholas
    Future civil air vehicles are likely to feature propulsion systems which are more closely integrated with the airframe. For a podded underwing configuration, this close coupling is expected to require non-axisymmetric design capabilities for the aero-engine exhaust system. This work presents the development of a novel parametric representation of non-axisymmetric aero-engine exhaust system geometries based on Intuitive Class Shape Transformation (iCST) curves. An exhaust design method was established and aerodynamic analyses of a range of non-axisymmetric configurations was demonstrated. At typical flight conditions, the introduction of non-axisymmetric separate jet nozzles was shown to increase the engine net propulsive force by 0.12% relative to an axisymmetric nozzle.