Visualization of Supersonic Flow Using O2-LIPF.

A video image analysis system for concentration measurements and flow visualization in building wakes

A quantitative imaging method using 2D single-pulse laser- induced predissociative fluorescence (LIPF) of OH concentrations is developed to study the flame structure in a swirling methane jet flame. A narrowband tunable KrF excimer laser is used to excite the P2(8) rotational line of the A2(Sigma) implied by (Chi) 2(Pi) (3,0) transition at (lambda) equals 248.46 nm. Though this transition produces a relatively weak signal, LIPF is much less sensitive to collisional quenching in atmospheric flames and suitable for generating quantitative data. OH concentration data are obtained by careful calibration against the flat flame burner data of known fuel-air equivalence ratio and temperature using identical optical setup. In this experiment, the OH fluorescence signal is imaged onto an intensified CCD camera. Since the distribution of OH concentration has good correspondence with the flame, the measured 2D imaging of OH indicates the instantaneous shape of the reaction zone. In the present experiment, the flame structure is first identified from the flame visualization. In the upstream of the swirling flame, combustion is found to take place in two regions; the recirculation vortex inside the recirculation bubble and the thin layer between and the thin layer between the recirculation zone and the ambient air. Higher intensity of OH is found early in recirculation zone. It is superequilibrium OH, because its intensity is higher than that from the recirculated burnt gas and the measured temperature is low. The measured OH concentrations have a direct influence on flame stabilization and NOx pollutant formation.© (1999) COPYRIGHT SPIE--The International Society for Optical Engineering. Downloading of the abstract is permitted for personal use only.

The drive towards lightweight, more efficient engine architectures leads to the development and improvement of key aircraft engine components. The Turbine Vane Frame (TVF) can be seen as an evolution of the Turbine Center Frame (TCF), combining its structural and aerodynamic purposes with the function otherwise provided by the first stage turbine inlet guide vanes. Therefore, the TVF needs to smoothly guide the flow over the radial offset between high-pressure and low-pressure turbine stages, provide shrouded paths for oil lines, and impart to the flow the turning required by the following low pressure turbine stage. The expected beneficial effects on the engine include reduced weight, cost, and fuel burn. The key relevance of this component makes a broad, low-tier approach to its study highly desirable. The HighSpeed Wind Tunnel (HSWT) at the Technical University of Graz delivers excellent opportunities to explore the design space for Turbine Vane Frames. This paper deals with the impact of several important design parameters on the TVF. An engine relevant TVF configuration is replicated in an annular sector cascade test rig. The test rig is operated at relevant Mach number and turbulence level. Representative inlet flow is achieved through flow conditioning. Aerodynamic data is collected with five-hole probe and hot wire anemometry techniques. The insights offered by flow visualization and numerical computations are furthermore leveraged. Particular attention is given to the flow features at the TVF exit.

Based on the hot-wire method for studying the fluctuations of compressible flows, the issues of determining the acoustic characteristics of the flow in the test sections of wind tunnels at transonic and supersonic speeds are considered. It is shown that at supersonic flows, in addition to the Mach waves described by Kovasznay, generated by stationary sources of disturbances on the walls of the test sections, and Mach waves generating the most intense fluctuations, distributed and moving in a supersonic turbulent boundary, described by Laufer, there may be Mach waves, the sources of which is a turbulent boundary layer. Using the hot-wire approach, it is possible to determine the characteristics of each type of these waves and their sources. It is also established that simple sound waves can be both produced by the turbulent boundary layer and those penetrating into the test section from sources located in the prechamber of the wind tunnel through the critical section of the Laval nozzle. In wind tunnels of high subsonic speeds, acoustic disturbances are produced from sound waves identified by intensity, direction and spectral composition using developed methods of thermal anemometry. The characteristics of acoustic disturbances (intensity, direction, location of sources) determined using the hot-wire allow them to be purposefully preserved or reduced, or their influence on phenomena under investigation can be taken into account. The article was prepared based on the materials of the report at the 10th Russian conference “Computational experiment in aeroacoustics and aerodynamics”, September 16–21, 2024, Svetlogorsk, Kaliningrad region, http://ceaa.imamod.ru/.

This paper describes the development of a bendable-wing Micro Air Vehicle (MAV) which is capable of being packed in a minimal volume. The wing is rigid under normal flight loads but compliant when the force is reversed, allowing it to be bent into a small diameter. The bendable-wing technology was developed to accommodate packing requirements on several larger vehicles, but was explored further in the creation of an endurance aircraft for use in the 2006 International Micro Air Vehicle Competition. The design procedure entailed airframe and wing design, component selection and placement, and composite stress analysis on the bendable wing. Stress analysis led to the choice of ±45° fiber orientation on the bendable-wing. Manufacturing was aided by custom CAD software and automated machining. Wind tunnel testing and flow visualization were used to show that flow separation is significant at low angles of attack of the thin under-cambered MAV wing. Visual Image Correlation indicated that a maximum permanent deformation of less than 0.5 mm was present after the wing was removed from the storage container. Qualitative flight testing confirmed that the observed residual wing deformation did not hinder flight and verified that the aircraft could consistently complete the objectives of the endurance mission.

The effect of aerodynamic detuning on the supersonic steady and unsteady blade passage flow field is experimentally investigated on a free surface water table by means of color Schlieren and shadowgraph flow visualization techniques. Two aerodynamic detuning mechanisms are considered: (1) alternate circumferential spacing of adjacent airfoils; (2) the replacement of alternate airfoils with splitters. The steady flow visualization demonstrates the significant effect of aerodynamic detuning on the passage flow field and, in particular, the shock wave-airfoil surface intersection locations. The unsteady flow visualization studies show the importance of the interblade phase angle. A mathematical model is also described and utilized to demonstrate the enhanced aeroelastic stability associated with the altered cascade passage shock wave structure due to these aerodynamic detuning mechanisms.

이차원 초음속 코안다 유동의 특성에 관한 실험적 연구가 진행되었다. 다양한 슬롯높이대 코안다 벽의 곡률반경 비 및 표면거칠기, 그리고 제트의 전압 변화에 대하여 충격파 구조 및 이력현상과 같은 코안다 유동의 특성이 유동가시화를 통하여 관찰되었다. 그 결과 초음속 코안다 제트의 이력현상은 코안다 표면거칠기의 영향을 받는 것으로 나타났다. 또한 강한 압축성 주제트의 추력편향제어를 위하여 이러한 동축류 코안다 유동이 적용되었고, 이 경우 유체역학적 추력제어의 성능은 코안다 플랩의 곡률반경에 대한 슬롯높이의 비 뿐만 아니라 주제트의 전압에 큰 영향을 받음이 확인되었다. The characteristics of two-dimensional supersonic coanda flow was experimentally investigated. For various ratios of slot height to coanda wall's radius of curvature, surface roughnesses, and jet stagnation pressures, the characteristics of the supersonic coanda flow such as shock structures and hysteresis were observed by flow visualization. It was found that the characteristics of hysteresis of the coanda jet was related to the surface roughness of the coanda wall. The study was further extended for application of the tangentially injected coanda jet to control co-flowing highly compressible main jet direction. It was noticed that the stagnation pressure of the main jet as well as the ratio of the slot height to coanda wall's radius of curvature wall was an influencing factor in the performance of the fluidic thrust vectoring method.

A novel optical diagnostic approach, namely, a dual hologram shearing interferometry with enhanced sensitivity, is proposed for visualization, and for measuring the density gradients of the flow in wind tunnels. The technique is especially useful for strong turbulent and/or unsteady regions flows. The features of the technique make it tolerant to vibrational disturbances typical to many wind tunnel facilities. The method was demonstrated by its application to a supersonic flow over a spherically blunted nose cone/cylinder model.

The present study proposes to design, fabricate and investigate a miniature supersonic wind tunnel at an affordable cost, safe to operate, and reliable for education purposes. The configuration consists of a high pressure reservoir, and mini tunnel which will be used for providing the supersonic flow over aerodynamic models. To design the mini supersonic wind tunnel, the method of characteristics is used to ensure that any shocks that arise in contact with the surface do not impact the uniformity and velocity of the flow through the duct. To achieve the supersonic flow, the nozzle contour of the supersonic portion of the wind tunnel is built in such a way that the shocks nullify each other and have zero effect on the velocity of the flow. For compressible flow visualization, an affordable Schlieren setup consisting of a torch, mirrors, and a mobile camera is used for flow visualization. Finally, the visualization results of the present study is in good relation with the theoretical calculations. The present facility is also calibrated by measuring the oblique shock angle in the test section, in the range of Mach number 1.6 to 1.8.

Experimental investigations of the flow around the hypersonic research configuration ELAC 1 are presented. Results of flow visualization in supersonic flow and measurements of the velocity field in incompressible flow are shown. The visualization was carried out by a combination of the vapour-screen method and oil flow-patterns. The application of these techniques showed the complete vortex structure at the leeward side. A vortex topology for angles of attack a ≤ 10° is presented. Furthermore the difference between turbulent and laminar flow was investigated with a large scale model of the first 40% of ELAC 1. Velocity measurements were carried out with Particle-Image velocimetry. A three-dimensional reconstruction of the flow field is presented. With this the velocity field was established at the leeward side of ELAC 1.

A temporally developing turbulent round mixing layer at Mach number M(j) = 1.92 has been simulated, and results compared to a similar nearly incompressible simulation at M(j) = 0.4. The supersonic mixing layer growth rate, defined with a delta99 thickness parameter, is found to be suppressed relative to its low Mach number counterpart. The Reynolds stresses in the supersonic flow are also suppressed in a manner similar to experimental results. Flow visualization, 1D energy spectra, and two point correlations indicate that a realistic turbulent flow is being simulated. At a given thickness in the development of the two mixing layers, the mean velocity, mean temperature, Reynolds stresses, Reynolds shear stress anisotropy, pressure fluctuations, budgets of Reynolds stresses, and turbulent kinetic energy are compared. It is found that even though maximum turbulent Mach number reaches 0.43 in the compressible case, dilatational effects are minimal, and the total viscous dissipation is nearly the same fraction of the production of turbulent kinetic energy for both flows. (Author)

Wave propagation in inhomogeneous media has been studied for such diverse applications as propagation of radiowaves in atmosphere, light propagation through thin films and in inhomogeneous waveguides, flow visualization, and others. In recent years an increased interest has been developed in wave propagation through shocks in supersonic flows. Results of experiments conducted in the past few years has shown such interesting phenomena as a laser beam splitting and spreading. The paper describes a model constructed to propagate a laser beam through shock-like inhomogeneous media. Numerical techniques are presented to compute the beam through such media. The results of computation are presented, discussed, and compared with experimental data.

Comparative analysis of the prediction ability of several modern detailed kinetic models describing the ignition and combustion of hydrogen/air mixture in a supersonic flow is conducted. The possible uncertainties in the rate constants of principle reactions responsible for the chain mechanism development in the H2/O2(air) mixture are inspected. Special attention is paid to matching the lengths of ignition and energy release zones and combustion completeness predicted by considered kinetic models in both one-dimensional homogeneous premixed flow and two-dimensional non-premixed flow in a scramjet model combustor. It is demonstrated that, while the difference in the induction zone length predicted by considered kinetic models does not exceed 30%, the discrepancy in the calculated values of energy release zone length may be as large as a factor of 4. As a consequence, different reaction mechanisms give distinguished magnitudes of combustion completeness at the combustor exit.

Set up of an experimental protocol for the investigation of graphite combustion in supersonic flow

This paper presents the results of an experimental study of initiation of hydrogen combustion in a supersonic flow with fuel supply from the duct wall. Data were obtained on the dynamics of disturbance development from gas-dynamic pulses and the transition to pseudoshock combustion with fuel supply from the combustion chamber walls rather than along the flow axis. Specific features of initiation of pre-detonation combustion for this scheme of hydrogen supply into a supersonic flow have been identified. It has been shown that steady-state pre-detonation combustion regimes differ slightly in dependence on the fuel supply method (along the axis or from the flow periphery), while the dynamics of propagation of wave structures from gas-dynamic pulses is virtually identical in both cases.