Dissertations / Theses on the topic 'Transonic tunnel'

The measurement of wind tunnel flow quality for the transonic flow regime has been plagued by the inability to interpret complex unsteady flow field information obtained in the free stream. Traditionally hot wire anemometry and fluctuating pressure techniques have been used to quantify the unsteady characteristics of a wind tunnel. This research focuses on the application of these devices to the transonic flow regime. Utilizing hot wire anemometry, one can decompose the unsteady flow field with a three sensor technique, to obtain fluctuations associated with the velocity, density, and total temperature. Implementing thermodynamic and kinematic equations, new methods for expanding the measured velocity, density, and total temperature fluctuations to obtain additional fluctuations are investigated. The derived static pressure fluctuations are compared to the static pressure fluctuations obtained with a conventional fluctuating static pressure probe. The results of this comparison are good, which implies that the individual velocity, density, and total temperature components are time accurate. In the process of obtaining a high quality fluctuating flow field information, it was necessary to evaluate the calibration of the hot wire sensors. A direct calibration approach was compared to a conventional non-dimensional technique. These two calibration techniques should have resulted in the same hot wire sensitivities. There were significant differences in the hot wire sensitivities as obtained from the two approaches. The direct approach was determined to have less errors due to the added heat transfer information required of the indirect approach. Both calibration techniques demonstrated that the velocity and density sensitivities were in general not equal. This suggests that the velocity and density information cannot be combined to form a mass flow. A comparison of several hot wire techniques was included to highlight the errors obtained when assuming that these sensitivities are the same. An evaluation of the free stream flow quality associated with a Laminar Flow Control experiment was carried out in the Langley Research Center 8-Foot Transonic Pressure Tunnel (8' TPT). The facility was modified with turbulence manipulators and a liner that provided a flow field around a yawed super-critical airfoil that is conducive to transition research. These devices are evaluated to determine the sources of disturbances associated with the LFC experiment.
Ph. D.

Two instruments mounted in a piggyback arrangement were developed for time-resolved measurements of dynamic flow quality in a transonic cryogenic wind tunnel. The first one is a dual hot-wire aspirating probe for measurement of stagnation pressure and temperature. The second is a miniature high-frequency response angle probe consisting of surface mounted pressure sensors. The aspirating probe was tested in the 0.3-m Transonic Cryogenic Tunnel (TCT) at NASA-Langley Research Center. Stagnation pressure and temperature measurements were taken in the free-stream of the settling chamber and test section. Data were also obtained in the unsteady wake shed from an airfoil oscillating at 5 Hz. The investigation revealed the presence of large stagnation pressure and temperature fluctuations in the settling chamber occurring at the blade passing frequency of the tunnel driving fan. The fluctuations in the test section are of a much more random nature and have amplitudes much lower than those in the test section. The overall results are consistent with previous tunnel disturbance measurements in the 0.3-m TCT. In the unsteady wake shed from the oscillating airfoil, stagnation temperature fluctuations as high as 42 K rms were observed. The high-frequency angle probe is a four sensor, pyramid type probe capable of simultaneously measuring time resolved stagnation and static pressures and two orthogonal flow angles. Using measurements from both probes, all flow parameters of interest can be deduced. Aerodynamic behavior of a full size model of the probe was established in an open air jet of known conditions.
M.S.

The wall interference effects present on three-dimensional models during wind tunnel testing are difficult to correct using post-test model data correction methods. Further, at transonic speeds, with the use of ventilated test sections these corrections become complex to apply and inaccurate. The high quality of wind tunnel testing that is required today means that such methods are no longer satisfactory. The flexible walled wind tunnel has in recent years shown its ability to obtain two-dimensional aerofoil data free from the effects of wall boundary restraint. This work at Southampton was aimed at extending the use of the two-dimensional Transonic Self-Streaming Wind Tunnel to the relief of wall interference effects on three-dimensional models. The compromise of using only two-wall single curvature movement avoids the problems that are inherent with the additional complexity of fully three-dimensional adaptive tunnels. A method of assessing the wall-induced interference velocity components from tunnel boundary pressure data, without reference to the model, has been developed and validated against other wall interference assessment methods. The algorithm, suitable for use in adaptive tunnels, is used with a wall movement influence coefficient method of wall contour prediction resulting in the apparent removal of wall interference effects along a streamlining target line. The residual wall interference velocity components calculated to be present after streamlining on two half-wing models are significantly lower than their straight test section values. Providing the model span is not too large in comparison with the breadth of the test section, the spanwise interference velocity component is negligible. A calibrated force-balance wing-body model has been used to demonstrate the first successful streamlining around a three-dimensional model in the Transonic Self-Streamlining Wind Tunnel. The measured model force data obtained with streamlined walls compares favourably with that derived using a standard post-test model data correction method.

A high-frequency combination probe was used to measure dynamic flow quality in the test section of the NASA Langley 0.3-m Transonic Cryogenic Tunnel. The probe measures fluctuating stagnation (total) temperature and pressure, static pressure, and flow angles in two orthogonal planes. Simultaneous unsteady temperature and pressure measurements were also made in the settling chamber of the tunnel. The data show that the stagnation temperature fluctuations remain constant, and the stagnation pressure fluctuations increase by a factor of two, as the flow accelerates from the settling chamber to the test section. In the test section, the maximum rms value of the normalized fluctuating velocity is 0.7 percent. Correlation coefficients l failed to show vortlcity, entropy, or sound as the dominant mode of turbulence in the tunnel.

At certain tunnel operating conditions, periodic disturbances are seen in the data taken in the test section. A possible cause for the disturbances is found to be acoustic coupling of the test section and plenum chamber via the perforated side walls in the tunnel. The experimental data agree well with the acoustic coupling theory.

Master of Science

Correcting wind tunnel data for wall interference is a critical part of relating the acquired data to a free-air condition. Accurately determining and correcting for the interference caused by the presence of boundaries in wind tunnels can be difficult especially for facilities employing ventilated boundaries. In this work, three varying levels of ventilation at the National Transonic Facility (NTF) were modeled and calibrated with a general slotted wall (GSW) linear boundary condition to validate the computational model used to determine wall interference corrections. Free-air lift, drag, and pitching moment coefficient predictions were compared for a range of lift production and Mach conditions to determine the uncertainty in the corrections process and the expected domain of applicability. Exploiting a previously designed statistical validation method, this effort accomplishes the extension of a calibration and validation for a boundary pressure wall interference corrections method. The foundational calibration and validation work was based on blockage interference only, while this present work extends the assessment of the method to encompass blockage and lift interference production. The validation method involves the establishment of independent cases that are then compared to rigorously determine the degree to which the correction method can converge free-air solutions for differing interference fields. The process involved first establishing an empty-tunnel calibration to gain both a centerline Mach profile of the facility at various ventilation settings, and to gain a baseline wall pressure signature undisturbed by a test article. The wall boundary condition parameters were then calibrated with a blockage and lift interference producing test article, and final corrected performance coefficients were compared for varying test section ventilated configurations to validate the corrections process and assess its domain of applicability. During the validation process discrimination between homogeneous and discrete implementations of the boundary condition was accomplished and final results indicated comparative strength in the discrete implementation's ability to capture experimental flow physics. Final results indicate that a discrete implementation of the General Slotted Wall boundary condition is effective in significantly reducing variations caused by differing interference fields. Corrections performed with the discrete implementation of the boundary condition collapse differing measurements of lift coefficient to within 0.0027, drag coefficient to within 0.0002, and pitching moment coefficient to within 0.0020.
Doctor of Philosophy
The purpose of conducting experimental tests in wind tunnels is often to acquire a quantitative measure of test article aerodynamic characteristics in such a way that those specific characteristics can be accurately translated into performance characteristics of the real vehicle that the test article intends to simulate. The difficulty in accurately simulating the real flow problem may not be readily apparent, but scientists and engineers have been working to improve this desired equivalence for the better part of the last half-century. The primary aspects of experimental aerodynamics simulation that present difficulty in attaining equivalence are: geometric fidelity, accurate scaling, and accounting for the presence of walls. The problem of scaling has been largely addressed by adequately matching conditions of similarity like compressibility (Mach number), and viscous effects (Reynolds number). However, accounting for the presence of walls in the experimental setup has presented ongoing challenges for ventilated boundaries; these challenges include difficulties in the correction process, but also extend into the determination of correction uncertainties. Exploiting a previously designed statistical validation method, this effort accomplishes the extension of a calibration and validation effort for a boundary pressure wall interference corrections method. The foundational calibration and validation work was based on blockage interference only, while this present work extends the assessment of the method to encompass blockage and lift interference production. The validation method involves the establishment of independent cases that are then compared to rigorously determine the degree to with the correction method can converge free-air solutions for differing interference scenarios. The process involved first establishing an empty-tunnel calibration to gain both a centerline Mach profile of the facility at various ventilation settings, and to gain a baseline wall pressure signature undisturbed by a test article. The wall boundary condition parameters were then calibrated with a blockage and lift interference producing test article, and final corrected performance coefficients were compared for varying test section ventilated configurations to validate the corrections process and assess its domain of applicability. During the validation process discrimination between homogeneous and discrete implementations of the boundary condition was accomplished and final results indicated comparative strength in the discrete implementation's ability to capture experimental flow physics. Final results indicate that a discrete implementation of the General Slotted Wall boundary condition is effective in significantly reducing variations caused by differing interference fields. Corrections performed with the discrete implementation of the boundary condition collapse differing measurements of lift coefficient to within 0.0027, drag coefficient to within 0.0002, and pitching moment coefficient to within 0.0020.

An experimental and numerical investigation was conducted to determine the film cooling effectiveness of a normal slot and angled slot under realistic engine Mach number conditions. Freestream Mach numbers of 0.65 and 1.3 were tested. For the normal slot, hot gas ingestion into the slot was observed at low blowing ratios (M < 0.25). At high blowing ratios (M > 0.6) the cooling film was observed to â lift offâ from the surface. For the 30o angled slot, the data was found to collapse using the blowing ratio as a scaling parameter (x/Ms). Results from the current experiment were compared with the subsonic data published to confirm this test procedure. For the angled slot, at the supersonic freestream Mach number, the current experiment shows that at the same x/Ms, the film cooling effectiveness increases by as much as 25% as compared to the subsonic case. The results of the experiment also show that at the same x/Ms, the film cooling effectiveness of the angled slot is considerably higher than that of the normal slot, at both subsonic and supersonic Mach numbers. The flow physics for the slot tests considered here are also described with computational fluid dynamic (CFD) simulations in the subsonic and supersonic regimes.
Master of Science

The aerodynamics of bodies in compressible ground effect flowfields from low-subsonic to supersonic Mach numbers have been investigated numerically and experimentally. A study of existing literature indicated that compressible ground effect has been addressed sporadically in various contexts, without being researched in any comprehensive detail. One of the reasons for this is the difficulty involved in performing experiments which accurately simulate the flows in question with regards to ground boundary conditions. To maximise the relevance of the research to appropriate real-world scenarios, multiple bodies were examined within the confines of their own specific flow regimes. These were: an inverted T026 wing in the low-to-medium subsonic regime, a lifting RAE 2822 aerofoil and ONERA M6 wing in the transonic regime, and a NATO military projectile at supersonic Mach numbers. Two primary aims were pursued. Firstly, experimental issues surrounding compressible ground effect flows were addressed. Potential problems were found in the practice of matching incompressible Computational Fluid Dynamics (CFD) simulations to wind tunnel experiments for the inverted wing at low freestream Mach numbers (<0.3), where the inverted wing was found to experience significant compressible effects even at Mach 0.15. The approach of matching full-scale CFD simulations to scale model testing at an identical Reynolds number but higher Mach number was analysed and found to be prone to significant error. An exploration was also conducted of appropriate ways to conduct experimental tests at transonic and supersonic Mach numbers, resulting in the recommendation of a symmetry (image) method as an effective means of approximating a moving ground boundary in a small-scale blowdown wind tunnel. Issues of scale with regards to Reynolds number persisted in the transonic regime, but with careful use of CFD as a complement to experiments, discrepancies were quantified with confidence. The second primary aim was to use CFD to gain a broader understanding of the ways in which density changes in the flowfield affect the aerodynamic performance of the bodies in question, in particular when a shock wave reflects from the ground plane to interact again with the body or its wake. The numerical approach was extensively verified and validated against existing and new experimental data. The lifting aerofoil and wing were investigated over a range of mid-to-high subsonic Mach numbers (1>M???>0.5), ground clearances and angles of incidence. The presence of the ground was found to affect the critical Mach number, and the aerodynamic characteristics of the bodies across all Mach numbers and clearances proved to be highly sensitive to ground proximity, with a step change in any variable often causing a considerable change to the lift, moment and drag coefficients. At the lowest ground clearances in both two and three dimensional studies, the aerodynamic efficiency was generally found to be less than that of unbounded (no ground) flight for shock-dominated flowfields at freestream Mach numbers greater than 0.7. In the fully-supersonic regime, where shocks tend to be steady and oblique, a supersonic spinning NATO projectile travelling at Mach 2.4 was simulated at several ground clearances. The shocks produced by the body reflected from the ground plane and interacted with the far wake, the near wake, and/or the body itself depending on the ground clearance. The influence of these wave reflections on the three-dimensional flowfield, and their resultant effects on the aerodynamic coefficients, was determined. The normal and drag forces acting on the projectile increased in exponential fashion once the reflections impinged on the projectile body again one or more times (at a height/diameter ground clearance h/d<1). The pitching moment of the projectile changed sign as ground clearance was reduced, adding to the complexity of the trajectory which would ensue.

Le développement de lanceurs spatiaux réutilisables nécessite une connaissance appro-fondie des effets des écoulements transsoniques sur la structure du lanceur, commele tremblement transsonique. En effet, l’intégrité mécanique du lanceur peut êtrecompromise par des interactions onde de choc/couche limite. Ces interactions peuventinduire, par exemple, des forces latérales responsables des moments de roulis et detangage, ou une excitation modale de certains éléments de structure pouvant conduireà leur endommagement, voire leur rupture. Ce travail rapporte des études numériqueset expérimentales sur la caractérisation de l’écoulement transsonique autour d’un aileronà profil losangique, conçu pour les lanceurs dédiés aux nanosatellites, avec un intérêtparticulier pour le tremblement transsonique. Ce phénomène a été longuement étudié.Malheureusement, les mécanismes intimes à l’origine du tremblement et la dynamiquedu phénomène sont encore débattus. De plus, il y a un manque d’études sur lesprofils losangiques. Des visualisations strioscopiques résolues en temps, des mesuresde pressions stationnaire et instationnaire pariétales ainsi que des mesures LDV sontréalisées expérimentalement dans une soufflerie transsonique. Les résultats sont comparésà des prédictions numériques basées sur des approches RANS instationnaire et LES. Lestraits tridimensionnels du tremblement transsonique et son caractère chaotique sur unprofil diamant sans flèche sont mis en évidence expérimentalement
The development of reusable space launchers requires a comprehensive knowledge oftransonic flow effects on the launcher structure, such as buffet. Indeed, the mechanicalintegrity of the launcher can be compromised by shock wave/boundary layer interactions.These interactions can induce, amongst others, lateral forces responsible for rolling andpitching moments, or modal excitation of some structural elements that can lead to theirdamage or even failure. This work reports numerical and experimental investigationson the characterization of the transonic flow past a diamond airfoil, designed fornanosatellite-dedicated launchers, with a particular interest for buffeting. Buffeting hasbeen extensively studied. Unfortunately, the detailed mechanisms that are responsiblefor the buffet inception and its dynamics are still debated. Moreover, there is a lackof studies for diamond airfoils. Here, time-resolved Schlieren visualizations, steadyand unsteady pressure measures and LVD measures are experimentally conducted ina transonic wind tunnel. They are compared with numerical predictions based on un-steady RANS and LES approaches. Three dimensional features of buffet over a diamondairfoil without swept, and the occurrence of a chaotic state, are experimentally highlighted

The current limits of supersonic parachute technology may constrain the ability to safely land future robotic assets on the surface of Mars. This constraint has led to a renewed interest in supersonic inflatable aerodynamic decelerator (IAD) technology, which offers performance advantages over the DGB parachute. Two supersonic IAD designs of interest include the isotensoid and tension cone, named for their respective formative structural theories. Although these concepts have been the subject of various tests and analyses in the 1960s, 1970s, and 2000s, significant work remains to advance supersonic IADs to a technology readiness level that will enable their use on future flight missions. In particular, a review of the literature revealed a deficiency in adequate aerodynamic and aeroelastic data for these two IAD configurations at transonic and subsonic speeds. The first portion of this research amended this deficiency by testing flexible IAD articles at relevant transonic and subsonic conditions. The data obtained from these tests showed that the tension cone has superior drag performance with respect to the isotensoid, but that the isotensoid may demonstrate more favorable aeroelastic qualities than the tension cone. Additionally, despite the best efforts in test article design, there remains ambiguity regarding the accuracy of the observed subscale behavior for flight scale IADs. Due to the expense and complexity of large-scale testing, computational fluid-structure interaction (FSI) analyses will play an increasingly significant role in qualifying flight scale IADs for mission readiness. The second portion of this research involved the verification and validation of finite element analysis (FEA) and computational fluid dynamic (CFD) codes for use within an FSI framework. These verification and validation exercises lend credence to subsequent coupled FSI analyses involving more complex geometries and models. The third portion of this research used this FSI framework to predict the static aeroelastic response of a tension cone IAD in supersonic flow. Computational models were constructed to mimic the wind tunnel test articles and flow conditions. Converged FSI responses computed for the tension cone agreed reasonably well with wind tunnel data when orthotropic material models were used and indicated that current material models may require unrealistic input parameters in order to recover realistic deformations. These FSI analyses are among the first results published that present an extensive comparison between FSI computational models and wind tunnel data for a supersonic IAD.

This thesis describes the design and initial testing of a transonic turbine cascade facility. It is specifically concerned with the best way to obtain flow periodicity and repeatability through the cascade by the use of tailboards at the cascade exit. The problem of how to achieve flow periodicity and repeatability has not been completely resolved. An examination of the literature available on transonic turbine cascade testing indicates some researchers use no tailboards, some use a solid tailboard, and still others use a porous tailboard. In this thesis, the flow through the turbine cascade is tested for three different cascade exit configurations; no tailboard, a solid tailboard, and a porous tailboard. The cascade is also tested with the tailboard at different angles, to see what effect the angle of the tailboard has on the flow through the cascade. The data acquisition and flow visualization systems are discussed and some preliminary results are given.
Master of Science

The path to future aero-engines with more efficient engine architectures requires advanced thermal management technologies to handle the demand of refrigeration and lubrication. Oil systems, holding a double function as lubricant and coolant circuits, require supplemental cooling sources to the conventional fuel based cooling systems as the current oil thermal capacity becomes saturated with future engine developments. The present research focuses on air/oil coolers, which geometrical characteristics and location are designed to minimize aerodynamic effects while maximizing the thermal exchange. The heat exchangers composed of parallel fins are integrated at the inner wall of the secondary duct of a turbofan. The analysis of the interaction between the three-dimensional high velocity bypass flow and the heat exchangers is essential to evaluate and optimize the aero-thermodynamic performances, and to provide data for engine modeling. The objectives of this research are the development of engine testing methods alternative to flight testing, and the characterization of the aerothermal behavior of different finned heat exchanger configurations. A new blow-down wind tunnel test facility was specifically designed to replicate the engine bypass flow in the region of the splitter. The annular sector type test section consists on a complex 3D geometry, as a result of three dimensional numerical flow simulations. The flow evolves over the splitter duplicated at real scale, guided by helicoidally shaped lateral walls. The development of measurement techniques for the present application involved the design of instrumentation, testing procedures and data reduction methods. Detailed studies were focused on multi-hole and fine wire thermocouple probes. Two types of test campaigns were performed dedicated to: flow measurements along the test section for different test configurations, i.e. in the absence of heat exchangers and in the presence of different heat exchanger geometries, and heat transfer measurements on the heat exchanger. As a result contours of flow velocity, angular distributions, total and static pressures, temperatures and turbulence intensities, at different bypass duct axial positions, as well as wall pressures along the test section, were obtained. The analysis of the flow development along the test section allowed the understanding of the different flow behaviors for each test configuration. Comparison of flow variables at each measurement plane permitted quantifying and contrasting the different flow disturbances. Detailed analyses of the flow downstream of the heat exchangers were assessed to characterize the flow in the fins¿ wake region. The aerodynamic performance of each heat exchanger configuration was evaluated in terms of non dimensional pressure losses. Fins convective heat transfer characteristics were derived from the infrared fin surface temperature measurements through a new methodology based on inverse heat transfer methods coupled with conductive heat flux models. The experimental characterization permitted to evaluate the cooling capacity of the investigated type of heat exchangers for the design operational conditions. Finally, the thermal efficiency of the heat exchanger at different points of the flight envelope during a typical commercial mission was estimated by extrapolating the convective properties of the flow to flight conditions.
Villafañe Roca, L. (2013). Experimental Aerothermal Performance of Turbofan Bypass Flow Heat Exchangers [Tesis doctoral no publicada]. Universitat Politècnica de València. https://doi.org/10.4995/Thesis/10251/34774
TESIS

A data reduction method has been developed and tested for a high frequency angle probe. The angle probe is designed for unsteady aerodynamic measurements in transonic cryogenic wind tunnels. The probe measures time-resolved total pressure, static pressure, angle of attack, and yaw angle from readings of four pressure transducers. The unique feature of this probe, as compared to a conventional multi-hole directional probe, is that the four high frequency response silicon pressure transducers are mounted flush on the probe tip. The data reduction method is basically an interpolation routine of calibration curves. The calibration curves consist of experimentally determined non-dimensional flow coefficients.

Two experiments were conducted to test the probe and the data reduction method. The first experiment tested the angle probe in a Karman vortex street shed from a cylinder. In the second experiment, the angle probe was placed in an open air jet with an exit Mach number of 0.42. Plots of the time-resolved measurements and the Fast Fourier Transform analysis were made for each test.
Master of Science

碩士
國立成功大學
航空太空工程學系
84
The objective of this work is to calibrate the transonic pilot tunnel located in ASTRC. The calibration process basically follows that of the large transonic tunnel built by Fluidyne. Various operational parameters were studied which include settling chamber stagnation pressure, wall porosity, Mach flap angle, choke flap angle and the test section wall divergence angle. How do these parameters affect the flow quality were assessed by measuring the centerline Mach number distributions as the control mechanism associated with each parameter was activated. Aside from these tunnel control parameters, the deviation of plenum chamber Mach number to the centerline Mach number was considered important also. All the Mach numbers obtained were calculated using the measured static pressure and the settling chamber stagnation pressure via an isentropic relationship. The calibration results show that the wall effects are significant for high-speed flow and the test section flow speed can be accelerated when either Mach flap or test section wall angle was enlarged. The twice deviation in centerline Mach number was found to be around 0.004∼0.006 for low speed flow(M=0.3∼0.7) and 0.006∼ 0.011 for high speed flow(M=0.7∼1.1), respectively. Generally speaking, a larger Mach number test section flow will cause a larger deviation in the centerline Mach number distribution.

碩士
國立成功大學
航空太空工程學系
82
The purpose of the study is to calibrate the flow quality of the ASTRC/TWT. The calibration included the flow uniformity (M_rms/Mcl), a parameter DM, and the aerodynamic noise level (Δ Cp). In studying the flow uniformity and the DM parameter, the centerline pipes were used to measure the static pressure distribution at the centerline of the wind tunnel. In studying the aerodynamic noise, four pressure transducers were flush- mounted on the 10-degree cone model which was used as the standard model to measure the aerodynamic noise in different flow condition. We used the technique of power spectrum analysis to identify the frequency components of the noise. The results showed that M_rms/Mcl is less than 0.005 with suitable parameter (MF,τ,θw) settings at subsonic. On the other hand, the flow quality at supersonic is inferior to that at subsonic. In the study of aerodynamic noise both the noise level (ΔCp) measured on the wall and 10-degree cone show the same trend. At M=0.8 or 0.9 the power spectrum results wind tunnel structure is about 159Hz. At M=0.8 or 0.9, varing the wall porosity of the test section wall show significant effect on the noise level of the test section.

A transonic wind tunnel was designed, constructed and calibrated in order to provide a valuable tool for the study of transonic flow phenomena. The wind tunnel makes use of flow properties surrounding the propagation of a shock wave along a tube in order to create the transonic flow. As a result, the wind tunnel is a modified shock tube, with its layout being optimised for maximum flow time. The flow times are dependent on the Mach number of the transonic flow being created, with the longest realistic flow time being approximately sixty milliseconds. The majority of the shock tube was built from commercially available steel construction tubing which was then attached to a pressure vessel of similar cross sectional dimension. A test section containing windows was constructed and placed in a position along the length of the tube to maximise the available test flow time. The position optimisation was calculated based on standard shock wave theory. The incident shock wave, as well as any resulting flow features, were visualised using schlieren photography. The test piece was designed to be set at angles of attack of up to ten degrees, both positive and negative. The main purpose of the testing carried out was to validate the functioning of the wind tunnel rather than obtaining more data on the test piece. An RAE2822 aerofoil was used as the test piece due to the large amount of aerodynamic data available on it, especially in the transonic flow region, thus making it an excellent tool for validation. In addition, the Fluent computational fluid dynamics package made use of the same aerofoil to validate their numerical results when the package was under development. This meant that for any numerical result obtained for the RAE2822 aerofoil using the Fluent package, there was a high degree of confidence. This fact provided a great tool for comparing results obtained experimentally in the wind tunnel with results obtained numerically. The short duration testing time was found to be adequate for establishing semi-steady state flow at any transonic flow Mach number. The bursting of the weak diaphragm at the end of the driven section of the shock tube resulted in the upstream propagation of a disturbance with a much lower velocity than would be seen if the incident shock wave reflected off a solid boundary and thus its arrival at the test section was delayed, resulting in a significant increase in testing time. The results obtained experimentally compared well to results obtained numerically. Transonic shock waves that were set up on the test piece had very similar shapes, features and chord-wise positions in both experimental and numerical results, showing that the geometric layout of the test section was correct. Furthermore, it was shown that a short duration flow time wind tunnel could be constructed using a shock tube and that accurate results could be obtained through its use.

碩士
國立成功大學
航空太空工程學系
87
The phenomenon of wind tunnel wall interference is due to the presence of the wall of wind tunnel test section. Consequently the flow field around a testing model situated in the test section of wind tunnel is normally somewhat different from that corresponding to the free-flight condition. With the use of ventilated walls for transonic wind tunnel, the interference problem becomes even more complex. The purpose of this study is to investigate the subsonic wall interference due to the presence of 2-D perforated walls of a transonic wind tunnel in ASTRC, NCKU. The porosity parameter method was employed to study the interference effect. This method was deduced from the linear potential flow theory subject to the concern of perforated walls. Experiments were made for a NACA0012 airfoil, that pressure distributions on the airfoil at Mach numbers from 0.3 to 0.8 and angles of attack from to were obtained. In addition, experiments on the effect of Mach flap position to wall interference were made at M=0.3. Base on the data of pressure distributions obtained, the aerodynamic coefficients were then calculated. Comparing the results obtained with the reference aerodynamic coefficients of NACA0012 lead us to suggest a formula to correct the interference effect of our wind tunnel. The results show that the formula proposed is able to adjust the aerodynamic coefficients measured to the reference data at low Mach number and low-AOA. But at high subsonic or high-AOA, this formula does not give satisfactory result.

碩士
國立成功大學
航空太空工程學系碩博士班
90
The purpose of this study is to develop the measurement techniquesusing pressure sensitive paint (PSP) and liquid crystal for transonic wind tunnel experiment. Acquiring temperature and pressure distribution of modelsurface is very important for studying the flow phenomena associated with theaerodynamic performance of the model. Traditional temperature and pressure measurements are taken with temperature sensors and pressure taps embedded in the model surface. Temperature and pressure taps do not give good spatial resolution due to the need for individual sensors. If the shape of the model is too small or complicated in geometry, it will limit the quantity of the sensors.Also, making a model can be very expensive if installation of a large amount of taps or sensors has to be considered. Using PSP and liquid crystal techniques, one can acquire temperature and pressure distributions,respectively, with high spatial resolution and simple model preparation. This research includes two parts. (1) Develop the PSP technique with the qualitative and quantitative analysis, and apply the technique in transonic wind tunnel experiment, with the aim to investigate the phenomena of flow around an ogive body at angles of attack. (2) Apply the liquid crystal technique in the transonic wind tunnel experiments, for studying the surface patterns concerning a finite cylinder on a flat plate and an ogive body. Attempts are made to reduce the temperature distributions from the obtained liquid crystal images.