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A practical method to evaluate quantitatively the uniformity of fuel/air mixing is essential for research and development of advanced low-emission combustion systems. Typically, this is characterised by measuring an unmixedness parameter or a uniformity index. An alternative approach, based on the fuel/air equivalence ratio distribution, is proposed and demonstrated in a simple methane/air venturi mixer. This approach has two main advantages: it is correlated with the fuel/air mixture combustion temperature, and the maximum temperature variation caused by fuel/air non-uniformity can be estimated. Because of these, it can be used as a criterion to check fuel/air mixing quality, or as a target for fuel/air mixer design with acceptable maximum temperature variation. For the situations where the fuel/air distribution non-uniqueness issue becomes important for fuel/air mixing check or mixer design, an additional statistical supplementary criterion should also be used.
This paper proposes an alternating elliptical impingement chamber in the leading edge of a gas turbine to restrain the cross flow and enhance the heat transfer, and investigates the detailed flow and heat transfer characteristics. The chamber consists of straight sections and transition sections. Numerical simulations are performed by solving the three-dimensional (3D) steady Reynolds-Averaged Navier–Stokes (RANS) equations with the Shear Stress Transport (SST) k–
turbulence model. The influences of alternating the cross section on the impingement flow and heat transfer of the chamber are studied by comparison with a smooth semi-elliptical impingement chamber at a cross-flow Velocity Ratio (VR) of 0.2 and Temperature Ratio (TR) of 1.00 in the primary study. Then, the effects of the cross-flow VR and TR are further investigated. The results reveal that, in the semi-elliptical impingement chamber, the impingement jet is deflected by the cross flow and the heat transfer performance is degraded. However, in the alternating elliptical chamber, the cross flow is transformed to a pair of longitudinal vortices, and the flow direction at the centre of the cross section is parallel to the impingement jet, thus improving the jet penetration ability and enhancing the impingement heat transfer. In addition, the heat transfer in the semi-elliptical chamber degrades rapidly away from the stagnation region, while the longitudinal vortices enhance the heat transfer further, making the heat transfer coefficient distribution more uniform. The Nusselt number decreases with increase of VR and TR for both the semi-elliptical chamber and the alternating elliptical chamber. The alternating elliptical chamber enhances the heat transfer and moves the stagnation point up for all VR and TR, and the heat transfer enhancement is more obvious at high cross-flow velocity ratio.
In order to know the characteristics of reducing the exhaust gas infrared signal of the lobed mixer according to the external air mixing ratio, an infrared signal and temperature distribution measurement using a micro-turbojet engine is performed. A certain amount of compressed air is supplied through an external duct mounted on the micro-turbojet engine exhaust to simulate bypass flow, which is mixed with high-temperature core air and ejected to the atmosphere. The exhaust nozzle used in the experiment is a lobed mixer with a lobe of sinusoidal shape and is designed to have a penetration of 0.2. Exhaust gas temperature and infrared signal are measured according to distance from nozzle outlet under conditions of bypass ratio of 0.5, 1.0 and 1.4. Infrared reduction rates are compared to data without compressed air supply. As a result of the experiment, as the bypass ratio increased, the infrared signal of the exhaust gas and the temperature decrease with bypass ratio increase, and in the case of a bypass ratio of 1.4, the effect of reducing the temperature is observed even at a long distance. In addition, we compared the results of previous studies of a simple cone shape without mixer with infrared reduction effect. The results show that the lobed mixer has a greater effect on reducing the temperature of the exhaust gas and reducing the infrared signal than the cone nozzle. The structure of the mixed jet flow is also studied through Schlieren visualisation and 3D temperature distribution.
The commercial Computational Fluid Dynamics (CFD) software STAR-CCM+ was used to simulate the flow and breakup characteristics of a Liquid Jet Injected into the gaseous Crossflow (LJIC) under real engine operating conditions. The reasonable calculation domain geometry and flow boundary conditions were obtained based on a civil aviation engine performance model similar to the Leap-1B engine which was developed using the GasTurb software and the preliminary design results of its low-emission combustor. The Volume of Fluid (VOF) model was applied to simulate the breakup feature of the near field of LJIC. The numerical method was validated and calibrated through comparison with the public test data at atmospheric conditions. The results showed that the numerical method can capture most of the jet breakup structure and predict the jet trajectory with an error not exceeding ±5%. The verified numerical method was applied to simulate the breakup of LJIC at the real engine operating condition. The breakup mode of LJIC was shown to be surface shear breakup at elevated condition. The trajectory of the liquid jet showed good agreement with Ragucci’s empirical correlation.
New alternative jet fuels have provided many advantages in the aviation industry, especially in terms of economics and environment. However, fuel–seal compatibility is one of the major issues that restricts alternative fuel advancement into the market. Thus, to help understand and solve the problem, this study examines the swelling effect of prepared and non-prepared O-rings in different fuels and aromatic species. Stress relaxation experiments were carried out to evaluate seal compatibility under compression, which mimics engine operation conditions. Seals were compressed and immersed in a variety of fuels and their blends for about 90h while maintaining a constant temperature 30°C and constant compression force of 25% seal thickness. The two types of elastomers investigated were fluorosilicone and nitrile O-rings, which are predominantly used in the aviation industry. Meanwhile, three different fuels and aromatic species were utilised as the variables in the experiments. The fuels used were Jet-A1, SPK and SHJFCS, while the aromatic species added were propyl benzene, tetralin and p-xylene. The swelling effects were determined from the P/Po value. Results indicate that Jet-A1 has the highest swelling effect, followed by SHJFCS and SPK. It was observed that the higher the percentage of aromatics in fuel, the higher the rate of swelling. Furthermore, prepared seals had a lower swelling rate than did non-prepared seals. Meanwhile, the intensity of the swelling effect in the Jet-A1-SHJFCS blends was in the order of 60/40, 85/15 and 50/50 blend. The work done in this study will aid in the selection of suitable aromatic species in future fuels. The novelty of this research lies in the determination of the appropriate amount of aromatic content as well as the selection of type of aromatic and its mixture fuel. Moreover, the various proportions of fuel blends with aromatic are investigated. The primary aim of this study is to understand the behaviour of prepared and non-prepared seals, and their compatibility with alternative fuels.
Particulate deposits in aero-engine turbines change the profile of blades, increase the blade surface roughness and block internal cooling channels and film cooling holes, which generally leads to the degradation of aerodynamic and cooling performance. To reveal particle deposition effects in the turbine, unsteady simulations were performed by investigating the migration patterns and deposition characteristics of the particle contaminant in a one-stage, high-pressure turbine of an aero-engine. Two typical operating conditions of the aero-engine, i.e. high-temperature take-off and economic cruise, were discussed, and the effects of particle size on the migration and deposition of fly-ash particles were demonstrated. A critical velocity model was applied to predict particle deposition. Comparisons between the stator and rotor were made by presenting the concentration and trajectory of the particles and the resulting deposition patterns on the aerofoil surfaces. Results show that the migration and deposition of the particles in the stator passage is dominated by the flow characteristics of fluid and the property of particles. In the subsequential rotor passage, in addition to these factors, particles are also affected by the stator–rotor interaction and the interference between rotors. With higher inlet temperature and larger diameter of the particle, the quantity of deposits increases and the deposition is distributed mainly on the Pressure Side (PS) and the Leading Edge (LE) of the aerofoil.
Losses induced by tip clearance limit decisive improvements in the system efficiency and aerodynamic operational stability of aero-engine axial compressors. The tendency towards even lower blade heights to compensate for higher fluid densities aggravates their influence. Generally, it is emphasised that the tip clearance should be minimised but remain large enough to prevent collisions between the blade tip and the casing throughout the entire mission. The present work concentrates on the development of a preliminary aero-engine axial compressor casing design methodology involving meta-modelling techniques. Previous research work at the Institute for Turbomachinery and Flight Propulsion resulted in a Two-Dimensional (2D) axisymmetric finite element model for a generic multi-stage high-pressure axial compressor casing. Subsequent sensitivity studies led to the identification of significant parameters that are important for fine-tuning the tip clearance via specific flange design. This work is devoted to an exploration of the potential of surrogate modelling in preliminary compressor casing design with respect to rapid tip clearance assessments and its corresponding precision in comparison with finite element results. Reputed as data-driven mathematical approximation models and conceived for inexpensive numerical simulation result reproduction, surrogate models show even greater capacity when linked with extensive design space exploration and optimisation algorithms.
Compared with high-fidelity finite element simulations, the reductions obtained in computational time when using surrogate models amount to 99.9%. Validated via statistical methods and dependent on the size of the training database, the precision of surrogate models can reach down to the range of manufacturing tolerances. Subsequent inclusion of such surrogate models in a parametric optimisation process for tip clearance minimisation rapidly returned adaptions of the geometric design variables.
Increasing demand for commercial air travel is projected to have additional environmental impact through increased emissions from fuel burn. This has necessitated the improvement of aircraft propulsion technologies and proposal of new concepts to mitigate this impact. The hybrid-electric aircraft propulsion system has been identified as a potential method to achieve this improvement. However, there are many challenges to overcome. One such challenges is the combination of electrical power sources and the best strategy to manage the power available in the propulsion system. Earlier methods reviewed did not quantify the mass and efficiency penalties incurred by each method, especially at system level. This work compares three power management approaches on the basis of feasibility, mass and efficiency. The focus is on voltage synchronisation and adaptation to the load rating. The three methods are the regulated rectification, the generator field flux variation and the buck-boost. This comparison was made using the propulsion system of the propulsive fuselage aircraft concept as the reference electrical configuration. Based on the findings, the generator field flux variation approach appeared to be the most promising, based on a balance of feasibility, mass and efficiency, for a 2.6MW system.
This paper presents a performance analysis on a novel engine concept, currently under development, in order to achieve hybrid air-breathing rocket technology. A component-level approach has been developed to simulate the performance of the engine at Mach 5, and the thermodynamic interaction of the different working fluids has been analysed. The bypass ramjet duct has also been included in the model. This facilitates the improved evaluation of performance parameters. The impact of ram drag induced by the intake of the engine has also been demonstrated. The whole model is introduced into a multi-platform application for aeroengine simulation to make it accessible to the interested reader. Results show that the bypass duct modelling increases the overall efficiency by approximately 7%. The model calculates the specific impulse at approximately 1800 seconds, which is 4 times higher than any chemical rocket.
An unusual philosophical approach is proposed here to decarbonise larger civil aircraft that fly long ranges and consume a large fraction of civil aviation fuel. These inject an important amount of carbon emissions into the atmosphere, and holistic decarbonising solutions must consider this sector. A philosophical–analytical investigation is reported here on the feasibility of an airliner family to fly over long ranges and assist in the elimination of carbon dioxide emissions from civil aviation.
Backed by state-of-the-art correlations and engine performance integration analytical tools, a family of large airliners is proposed based on the development and integration of the body of a very large two-deck four-engine airliner with the engines, wings and flight control surfaces of a very long-range twin widebody jet. The proposal is for a derivative design and not a retrofit. This derivative design may enable a swifter entry to service.
The main contribution of this study is a philosophical one: a carefully evaluated aircraft family that appears to have very good potential for first-generation hydrogen-fuelled airliners using gas turbine engines for propulsion. This family offers three variants: a 380-passenger aircraft with a range of 3,300nm, a 330-passenger aircraft with a range of 4,800nm and a 230-passenger aircraft with a range of 5,500nm. The latter range is crucially important because it permits travel from anywhere in the globe to anywhere else with only one stop. The jet engine of choice is a 450kN high-bypass turbofan.
The potential to recover waste heat from the exhaust gases of a turboprop engine and produce useful work through an Organic Rankine Cycle (ORC) is investigated. A thermodynamic analysis of the engine’s Brayton cycle is derived to determine the heat source available for exploitation. The aim is to use the aircraft engine fuel as the working fluid of the organic Rankine cycle in order to reduce the extra weight of the waste heat recovery system and keep the thrust-to-weight ratio as high as possible. A surrogate fuel with thermophysical properties similar to aviation gas turbine fuel is used for the ORC simulation. The evaporator design as well as the weight minimisation and safety of the suggested application are the most crucial aspects determining the feasibility of the proposed concept. The results show that there is potential in the exhaust gases to produce up to 50kW of power, corresponding to a 10.1% improvement of the overall thermal efficiency of the engine.