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To present the detailed implementation processes of the IDEAL algorithm for two-dimensional compressible flows based on Delaunay triangular mesh, and compare the performance of the SIMPLE and IDEAL algorithms for solving compressible problems. What’s more, the implementation processes of Delaunay mesh generation and derivation of the pressure correction equation are also introduced.

Programming completely in C++.

Five compressible examples are used to test the SIMPLE and IDEAL algorithms, and the comparison with measurement data shows good agreement. The IDEAL algorithm has much better performance in both convergence rate and stability over the SIMPLE algorithm.

The detail solution procedure of implementing the IDEAL algorithm for compressible flows based on Delaunay triangular mesh is presented in this work, seemingly first in the literature.

The mixing of fuel and air plays a pivotal role in enhancing combustion in supersonic regime. Proper mixing stabilizes the flame and prevents blow-off. Blow-off is due to the shorter residence time of fuel and air in the combustor, as the flow is in supersonic regime. The flame is initiated in the local subsonic region created using a flameholder within the supersonic combustor. This study aims to design an effective flameholder which increases the residence time of fuel in the combustor allowing proper combustion preventing blow-off and other instabilities.

The geometry of the strut-based flameholder is altered in the present study to induce a streamwise motion of the fluid downstream of the strut. The streamwise motion of the fluid is initiated by the ramps and grooves of the strut geometry. The numerical simulations were carried out using ANSYS Fluent and are validated against the available experimental and numerical results of cold flow with hydrogen injection using plain strut as the flameholder. In the present study, numerical investigations are performed to analyse the effect on hydrogen injection in strut-based flameholders with ramps and converging grooves using Reynolds-averaged Navier–Stokes equation coupled with Menter’s shear stress transport k-ω turbulence model. The analysis is done to determine the effect of geometrical parameters and flow parameter on the flow structures near the base of the strut where thorough mixing takes place. The geometrical parameters under consideration include the ramp length, groove convergence angle, depth of the groove, groove compression angle and the Mach number. Two different strut configurations, namely, symmetric and asymmetric struts were also studied.

Higher turbulence and complex flow structures are visible in asymmetric strut configuration which develops better mixing of hydrogen and air compared to symmetric strut configuration. The variation in the geometric parameters develop changes in the fluid motion downstream of the strut. The fluid passing through the converging grooves gets decelerated thereby reducing the Mach number by 20% near the base of the strut compared to the straight grooved strut. The shorter ramps are found to be more effective, as the pressure variation in lateral direction is carried along the strut walls downstream of the strut increasing the streamwise motion of the fluid. The decrease in the depth of the groove increases the recirculation zone downstream of the strut. Moreover, the increase in the groove compression angle also increases the turbulence near the base of the strut where the fuel is injected. Variation in the injection port location increases the mixing performance of the combustor by 25%. The turbulence of the fuel jet stream is considerably changed by the increase in the injection velocity. However, the change in the flow field properties within the flow domain is marginal. The increase in fuel mass flow rate brings about considerable change in the flow field inducing stronger shock structures.

The present study identifies the optimum geometry of the strut-based flameholder with ramps and converging grooves. The reaction flow modelling may be performed on the strut geometry incorporating the design features obtained in the present study.

This paper aims to present the results of experimental and numerical research on heat transfer distribution under the impinging jets at various distances from the wall and high jet velocity. This work is a part of the INNOLOT Program financed by National Centre for Research and Development.

The air jets flow out from the common pipe and impinge on a surface which is cooled by them, and in this way, all together create a model of external cooling system of low-pressure gas turbine casing. Measurements were carried out for the arrangement of 26 in-line jets with orifice diameter of 0.9 mm. Heat transfer distribution was investigated for various Reynolds and Mach numbers. The cooled wall, made of transparent PMMA, was covered with a heater foil on which a layer of self-adhesive liquid crystal foil was placed. The jet-to-wall distance was set to h = from 4.5 to 6 d.

The influence of various Reynolds and Mach numbers on cooled flat plate and jet-to-wall distance in terms of heat transfer effectiveness is presented. Experimental results used for the computational fluid dynamics (CFD) model development, validation and comparison with numerical results are presented.

Impinging air jets is a commonly used technique to cool advanced turbines elements, as it produces large convection enhancing the local heat transfer, which is a critical issue in the development of aircraft engines.

The achieved results present experimental investigations carried out to study the heat transfer distribution between the orthogonally impinging jets from long round pipe and flat plate. Reynolds number based on the jet orifice exit conditions was varied between 2,500 and 4,000; meanwhile, for such Re, the flow velocity in jets was particularly very high, changing from M = 0.56 to M = 0.77. Such flow conditions combination, i.e. the low Reynolds number and very high flow velocity cannot be found in the existing literature.

This study is focused on the local entropy generation of steady two‐dimensional symmetric flow past a parabolic cylinder in a uniform stream parallel to its axis. The effect of both Reynolds number (Re) and temperature difference between the cylinder wall and the freestream (ΔT) on thermal, viscous, and total entropy generation is investigated for different values of Prandtl number (Pr) and Mach number (Ma). It was found that the thermal entropy generation increased as the temperature difference increased, while the viscous entropy generation decreased as Reynolds number increased. The results also showed that increasing Prandtl number or Mach number increased both the viscous and thermal contributions to the entropy generation.

THE Press was recently given the opportunity of seeing some of the more modern aerodynamic research facilities of the Ministry of Supply at R.A.E. Farnborough and N.A.E. Bedford. The Bedford establishment is administratively part of the R.A.E., and is concerned at present primarily with aerodynamic research on high‐speed aircraft. Later, work on engines will also be done there. It lies on the borders of three wartime R.A.F. airfields, the original intention being to build a runway which would enable jet aircraft to take off and land again within its length. The establishment has grown up from nothing since the war, and this is reflected in the attractive and orderly disposition of the buildings, their pleasing contemporary architecture, and the general impression of clean design. The site is still in the hands of the contractors, but it can be seen that when it is complete it will be a fine example of what such an establishment should look like. Credit for this must be shared between the Ministry of Supply, the Ministry of Works, and the contractors, who have allowed imagination to play its part in design, without it leading to extravagance. Particularly attractive are the colour schemes in the main administrative block, and the use of colour on the engineering plant itself.

THE problem of providing engines suitable for high Mach number aircraft is a fascinating study which at the moment has only been taken to the stage where many solutions look feasible, thus the choice of engines for the different roles for which high Mach number aircraft may be used is still fairly wide open.

The aim of the research is to conduct a study into a configuration of an aircraft system with a focus on aerodynamics. In addition, trim condition and static stability constraints were included. The main application of this system is suborbital space flights. The presented concept of a modular airplane system (MAS) consists of two vehicles: a Rocket Plane and a Carrier. Both are designed in tailless configurations but coupled formed a classic tail aircraft configuration, where the Rocket Plane works as the empennage. The most important challenge is to define the mutual position of those two tailless vehicles under the assumption that each vehicle will be operating alone in different flight conditions while joined in one object create a conventional aircraft. Each vehicle configuration (separated and coupled) must fulfil static stability and trim requirements.

Aircrafts’ aerodynamic characteristics were obtained using the MGAERO software which is a commercial computing fluid dynamics tool created by AMI Aero. This software uses the Euler flow model. Results from this software were used in the static stability and trim condition analysis.

The main outcome of this investigation is a mutual position of the Rocket Plane and the Carrier that fulfils project requirements. Also, the final configuration of both separated vehicles (Rocket Plane and Carrier) and the complete MAS were defined. In addition, it was observed that in the case of classic aircraft configuration which is created by connecting two tailless vehicles increasing horizontal tail arm reduces static stability. This is related to a significantly higher mass ratio of the horizontal tail (the Rocket Plane) with respect to the whole system. Moving backward, the Rocket Plane has a notable effect on a position of a centre of gravity of the whole system static stability. Moreover, the impact of the mutual vehicles’ position (horizontal tail arm) and inclination angle on the coupled vehicle lift to drag ratio was analysed.

In terms of aerodynamic computation, MGAERO software using an inviscid flow model, therefore, both a friction drag and breakdown of vortex are not considered. But the presented research is for the computation stage of the design, and the MGAERO software guarantees satisfactory accuracy with respect to the relatively low time of computations. The second limitation is that the presented results are for the conceptual stage of the design and dynamic stability constraints were not taken into account.

The ultimate goal of the coupled aircraft project is to conduct flying tests and the presented result is one of the milestones to achieve this goal.

A design process for a conventional aircraft configuration is well known however, there are not many examples of vehicles that consist of two coupled aircrafts where both vehicles have similar mass. The unique part of this paper includes results of the investigation of the mutual position of the vehicles that can fly alone, as well as in coupled form. The impact of the position of the centre of gravity on trim conditions and static stability of the coupled configuration was investigated.

The purpose of this paper is to investigate the computational features of the Flowfield Dependent Variation (FDV) method, a numerical scheme built to simulate flows characterized by multiple speeds, multiple physical phenomena, and by large variations in flow variables.

Fundamentally, the FDV method may be regarded as a variant of the Lax-Wendroff Scheme (LWS) that is obtained by replacing the explicit time derivatives in LWS by a weighted combination of explicit and implicit time derivatives. The weighting factors – referred to as FDV parameters – may be broadly classified as convective and diffusive parameters which, for example, are determined using flow quantities such as the Mach number and Reynolds number, respectively. Hence, the reference to these parameters and the method as “flow field dependent.” A von Neumann Fourier analysis demonstrates that the increased implicitness makes FDV both more stable and less dispersive compared to LWS, a feature crucial to capturing shocks and other phenomena characterized by high gradients in variables. In the current study, the FDV scheme is implemented in a Taylor-Galerkin-based finite element method framework that supports arbitrarily high order, unstructured isoparametric elements in one-, two- and three-dimensional geometries.

At first, the spatial accuracy of the implemented FDV scheme is established using the Method of Manufactured Solutions, wherein the results show that the order of accuracy of the scheme is nearly equal to the order of the shape function polynomial plus one. The dispersion and dissipation errors of FDV, when applied to the compressible Navier-Stokes and energy equations, are investigated using a 2-D, small-amplitude acoustic pulse propagating in a quiescent medium. It is shown that FDV with third-order shape functions accurately captures both the amplitude and phase of the acoustic pulse. The method is then applied to cases ranging from low-Mach number subsonic flows (Mach number M=0.05) to high-Mach number supersonic flows (M=4) with shock-boundary layer interactions. For all cases, fair to good agreement is observed between the current results and those in the literature.

The spatial order of accuracy of the FDV method, its stability and dispersive properties, as well as its applicability to low- and high-Mach number flows are established.

WHEN talking about airbreathing engines it is now generally understood that they are either turbine engines when the maximum flight Mach number is subsonic or moderately supersonic, or ramjets when the Mach number is definitely high. When trying to meet the propulsion requirements from take‐off to a high enough speed the joint use of both engine types has to be considered. In such case most people would think of the ramjet as taking over the propulsion task from the turbine engine when reaching a certain value of the flight Mach number, or more precisely of the air stagnation temperature, above which the turbine engine is no longer able to operate. The most elementary view is that of presenting it as a limitation in the engine structure, with improvements calling for the use of better materials. Bringing thermo‐dynamics into the picture shows that increased air stagnation temperature results in a deterioration in the cycle efficiency of the turbine engine proper and this may result in the specific fuel consumption of the turbojet becoming higher than that of the ramjet. Such a performance limitation can be shifted to higher Mach numbers while using increased turbine intake temperatures. The consideration of the aerodynamics of the internal flow brings out another type of limitation due to the difficulty of keeping the operating line of the turbojet over the flight profile far enough from the surge limit though within the range of good compresser efficiency. Variable geometry in the compressor and turbine stators may produce some improvement in this respect.

Ideal and practical performance of ram‐jet units in steady flight in the stratosphere at Mach numbers from 1·5 to 4 is examined. The effects of combustion, temperature, altitude, intake and exhaust nozzle design are considered.