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.

Highly loaded single stage gas turbines are being developed to minimize the turbine size and weight. Such highly loaded turbines often result in transonic flows, which imply a reduction in the efficiency due to the shock losses. The efficiency of a turbine is defined as the ratio between the real work extracted by the turbine rotor from the fluid and the maximum available enthalpy for a given pressure ratio. The relationship between turbine performance and design parameters is not yet fully comprehended due to the complexity of the flow field and unsteady flow field interactions. Hence, experimental and numerical studies remain necessary to understand the flow behavior at different conditions to advance the state of the art of the prediction tools. The purpose of the current research is to develop a methodology to determine the efficiency with an accuracy better than 1 % in a cooled and uncooled high pressure (HP) turbine tested in a short duration facility with a running time of about 0.4s. Such low level of uncertainty requires the accurate evaluation of a large number of quantities simultaneously, namely the mass flow of the mainstream, the coolant, and leakage flows properties, the inlet total pressure and total temperature, the stage exit total pressure, the shaft power, the mechanical losses and the heat transfer. The experimental work is carried out in a compression tube facility that allows testing the turbine at the temperature ratios, Re and Mach numbers encountered in real engines. The stage mass flow is controlled by a variable sonic throat located downstream of the stage exit. Due to the absence of any brake, the turbine power is converted into rotor acceleration. The accurate measurement of this acceleration as well as those of the inertia and the rotational speed provides the shaft power. The inertia of the whole rotating assembly was accurately determined by accelerating and decelerating the shaft with a known energy. The mass-flow is derived from the measured turbine inlet total pressure and the vane sonic throat. The turbine sonic throat was evaluated based on a zero-dimensional model of the turbine. The efficiencies of two transonic turbines are measured at design and off-design conditions. The turbine design efficiency is obtained as 91.8 %. The repeatability of the measurements for 95% confidence level varies between 0.3 % and 1.1 % of the efficiency depending on the test case. The theoretical uncertainty level of 1.2 % is mainly affected by the uncertainty of exit total pressure measurements. Additionally, the effect of vane trailing edge shock formations and their interactions with the rotor blade are analyzed based on the experimental data, the numerical tools and the loss correlations. The changes of blade and vane performances are measured at mid-span for three different pressure ratios which influence the vane and rotor shock mechanisms. Moreover, the unsteady forces on the rotor blades and the rotor disk were calculated by integration of the unsteady static pressure field on the rotor surface.

Approved for public release, distribution is unlimited / For over half a century the NPS Aerolabʼ Low Speed Wind Tunnel located in Halligan Hall of the Naval Postgraduate school has served to provide students and faculty with meaningful aerodynamic data for research and problem analysis. New data acquisition hardware was installed three years ago but never fully verified, and contained no integrated software program to collect data from the strain-gauge balance pedestal. Existing National Instruments based hardware for the NPS low-speed wind tunnel was reconfigured to obtain data from the strain-gauge pedestal. Additionally, a data acquisition software program was written in LabVIEW⠭ to accommodate the hardware. The Virtual Instruments (VI) program collects and plots accurate data from all four strain gauges in real-time, producing non-dimensional force and moment coefficients. A research study on the performance of an OH-6A helicopter fuselage was conducted. NPS Aerolabʼ wind tunnel tests consisted of drag, lift, and pitching moment measurements of the OH-6A along yaw and angle-of-attack sweeps. The results of the NPS wind tunnel data were compared against testing conducted on a full-scale OH-6A helicopter in NASA Ames' 40 ft. x 80ft. wind tunnel, along with the U.S. Army's Light Observation Helicopter (LOH) wind tunnel tests. Results of current testing substantiate the LabVIEW⠭ code. / Ensign, United States Navy

Laminar turbulent transition is known for a long time as a critical phenomenon influencing the thermal load encountered by hypersonic vehicle during their planetary re-entry trajectory. Despite the efforts made by several research laboratories all over the world, the prediction of transition remains inaccurate, leading to oversized thermal protection system and dramatic limitations of hypersonic vehicles performances. One of the reasons explaining the difficulties encountered in predicting transition is the wide variety of parameters playing a role in the phenomenon. Among these parameters, surface roughness is known to play a major role and has been investigated in the present thesis.<p><p>A wide bibliographic review describing the main parameters affecting transition and their coupling is proposed. The most popular roughness-induced transition predictions correlations are presented, insisting on the lack of physics included in these methods and the difficulties encountered in performing ground hypersonic transition experiments representative of real flight characteristics. This bibliographic review shows the importance of a better understanding of the physical phenomenon and of a wider experimental database, including real flight data, for the development of accurate prediction methods.<p><p>Based on the above conclusions, a hypersonic experimental test campaign is realized for the characterization of the flow field structure in the vicinity and in the wake of 3D roughness elements. This fundamental flat plate study is associated with numerical simulations for supporting the interpretation of experimental results and thus a better understanding of transition physics. Finally, a model is proposed in agreement with the wind tunnel observations and the bibliographic survey.<p><p>The second principal axis of the present study is the development of a hypersonic in-flight roughness-induced transition experiment in the frame of the European EXPERT program. These flight data, together with various wind tunnel measurements are very important for the development of a wide experimental database supporting the elaboration of future transition prediction methods. / Doctorat en Sciences de l'ingénieur / info:eu-repo/semantics/nonPublished

No description available.

In this thesis, techniques for testing aircraft models in ground effect in a low speed wind tunnel are investigated. Although these types of tests have been done before, the current study is unique in that forces are measured with an overhead balance instead of an internal balance. This has the advantage that the types of models that are difficult to mount on a sting with an internal balance, can often be mounted with a strut protruding from the top of the model. Positioning a sting-mounted model close to the ground at a high angle-of-attack is also usually difficult if not impossible. Finally, drag measurements are often more accurate when measured with an overhead balance rather than with an internal sting-type balance. The disadvantages associated with this method of testing are identified and solutions suggested. These include accurate moment transfers and correcting for support tares and interference. The thesis also investigates general procedures associated with ground effect testing such as proper boundary corrections and the necessity of a rolling floor. A simplified preliminary test series was performed in order to identify shortcomings in the existing equipment and procedures. This series is explained in Chapter 2. Chapter 4 and 5 describe changes made to the existing equipment following this test series. These include a novel telescopic fairing to shroud the mounting strut and an internal pitching mechanism. The correction techniques and general theory are summarised in Chapter 3. The author concludes in Chapter 6 that with the application of the techniques described in this thesis, the test engineer should be able to obtain accurate and reliable data from most aircraft configurations. Additional suggestions for testing models in ground effect are also given in this chapter. Finally, a few shortcomings that still need to be investigated are mentioned at the end of Chapter 6. AFRIKAANS : Hierdie verhandeling ondersoek tegnieke om vliegtuigmodelle in grondeffek in 'n laespoed-windtonnel te toets. Alhoewel hierdie tipe van toetse al voorheen gedoen is, is die huidige studie uniek deurdat 'n oorhoofse balans eerder as 'n interne balans gebruik word. Die voordeel hiervan is dat modelle wat moeilik op 'n naald- of "sting"-balans monteer kan word, baie keer makliker monteer kan word met 'n stang wat deur die bokant van die model steek. Posisioneering van 'n naald-gemonteerde model naby aan die vloer van die tonnel by hoe invalshoeke is gewoonlik ook baie moeilik indien nie onmoontlik nie. Laastens is sleurkrag-metings wat met 'n oorhoofse balans gemeet is gewoonlik meer akkuraat as sleurkrag-metings wat met 'n interne naald-tipe balans gedoen is. Die nadele wat met hierdie toetsmetode geassosieer kan word, word geïdentifiseer en moontlike oplossing word voorgestel. Hierdie sluit die berekening in van akkurate moment-transformasies en monteersleureffekte en -steurings. Die verhandeling ondersoek ook algemene prosedures wat met grondeffektoetse geassosieer kan word, byvoorbeeld akkurate wandkorreksies en die nodigheid van die rolvloer. 'n Vereenvoudigde vooraf-toetsreeks was uitgevoer om moontlike tekortkominge in die bestaande toerusting en prosedures te identifiseer. Hierdie toetsreeks word in Hoofstuk 2 bespreek. Hoofstuk 4 en 5 verduidelik die veranderinge wat aan die bestaande toerusting gemaak is na aanleidng van hierdie toetsreeks. Hierdie veranderinge sluit 'n teleskopiese windskerm in om die monteerstang te isoleer van die wind, sowel as 'n interne heimeganisme om die invalshoek van die model te verstel. Die korreksieprosedures en algemene teorie word in Hoofstuk 3 opgesom. Die outeur se gevolgtekking in Hoofstuk 6 stel dat die toetsingenieur, met behulp van die gebruik van die tegnieke in hierdie verhandeling beskryf, in staat behoort te wees om betroubare metings te kan neem van meeste vliegtuigkonfigurasies. Verdere voorstelle vir die toets van modelle in grondeffek word ook in hierdie hoofstuk gemaak. Uiteindelik word 'n paar tekortkominge genoem wat moontlik in 'n toekomstige studie ondersoek kan word. / Dissertation (MEng)--University of Pretoria, 1999. / Mechanical and Aeronautical Engineering / unrestricted

No description available.

One of the characteristics of the high speed ows over blunt bodies is the detached shock formed in front of the body. The distance of the shock from the stagnation point measured along the stagnation streamline is termed as the shock stand o distance or the shock detachment distance. It is one of the most basic parameters in such ows. The need to know the shock stand o distance arises due to the high temperatures faced in these cases. The biggest challenge faced in high enthalpy ows is the high amounts of heat transfer to the body. The position of the shock is relevant in knowing the temperatures that the body being subjected to such ows will have to face and thus building an efficient system to reduce the heat transfer. Despite being a basic parameter, there is no theoretical means to determine the shock stand o distance which is accepted universally. Deduction of this quantity depends more or less on experimental or computational means until a successful theoretical model for its predictions is developed. The experimental data available in open literature for spherical bodies in high speed ows mostly lies beyond the 2 km/s regime. Experiments were conducted to determine the shock stand o distance in the velocity range of 1-2 km/s. Three different hemispherical bodies of radii 25, 40 and 50 mm were taken as test models. Since the shock stand o distance is known to depend on the density ratio across the shock and hence gamma (ratio of specific heats), two different test gases, air and carbon dioxide were used for the experiments here. Five different test cases were studied with air as the test gas; Mach 5.56 with Reynolds number of 5.71 million/m and enthalpy of 1.08 MJ/kg, Mach 5.39 with Reynolds number of 3.04 million/m and enthalpy of 1.42 MJ/kg Mach 8.42 with Reynolds number of 1.72 million/m and enthalpy of 1.21 MJ/kg, Mach 11.8 with Reynolds number of 1.09 million/m and enthalpy of 2.03 MJ/kg and Mach 11.25 with Reynolds number of 0.90 million/m and enthalpy of 2.88 MJ/kg. For the experiments conducted with carbon dioxide as test gas, typical freestream conditions were: Mach 6.66 with Reynolds number of 1.46 million/m and enthalpy of 1.23 MJ/kg. The shock stand o distance was determined from the images that were obtained through schlieren photography, the ow visualization technique employed here. The results obtained were found to follow the same trend as the existing experimental data in the higher velocity range. The experimental data obtained was compared with two different theoretical models given by Lobb and Olivier and was found to match. Simulations were carried out in HiFUN, an in-house CFD package for Euler and laminar own conditions for Mach 8 own over 50 mm body with air as the test gas. The computational data was found to match well with the experimental and theoretical data

The hypersonic laminar to turbulent transition problem above Mach 10 is addressed experimentally in the short duration VKI Longshot gun tunnel. Reentry conditions are partially duplicated in terms of Mach and Reynolds numbers. Pure nitrogen is used as a test gas with flow enthalpies sufficiently low to avoid its dissociation, thus approaching a perfect gas behavior. The stabilizing effects of Mach number and nosetip bluntness on the development of natural boundary layer disturbances are evaluated over a 7 degrees half-angle conical geometry without angle of attack. <p><p>Emphasis is initially placed on the flow characterization of the Longshot wind tunnel where these experiments are performed. Free-stream static pressure diagnostics are implemented in order to complete existing stagnation point pressure and heat flux measurements on a hemispherical probe. An alternative method used to determine accurate free-stream flow conditions is then derived following a rigorous theoretical approach coupled to the VKI Mutation thermo-chemical library. Resulting sensitivities of free-stream quantities to the experimental inputs are determined and the corresponding uncertainties are quantified and discussed. The benefits of this different approach are underlined, revealing the severe weaknesses of traditional methods based on the measurement of reservoir conditions and the following assumptions of an isentropic and adiabatic flow through the nozzle. The operational map of the Longshot wind tunnel is redefined accordingly. The practical limits associated with the onset of nitrogen flow condensation under non-equilibrium conditions are also accounted for. <p><p>Boundary layer transition experiments are then performed in this environment with free-stream Mach numbers ranging between 10-12. Instrumentation along the 800mm long conical model includes flush-mounted thermocouples and fast-response pressure sensors. Transition locations on sharp cones compare favorably with engineering correlations. A strong stabilizing effect of nosetip bluntness is reported and no transition reversal regime is observed for Re_RN<120000. Wavelet analysis of wall pressure traces denote the presence of inviscid instabilities belonging to Mack's second mode. An excellent agreement with Linear Stability Theory results is obtained from which the N-factor of the Longshot wind tunnel in these conditions is inferred. A novel Schlieren technique using a short duration laser light source is developed, allowing for high-quality flow visualization of the boundary layer disturbances. Comparisons of these measurement techniques between each other are finally reported, providing a detailed view of the transition process above Mach 10. / Doctorat en Sciences de l'ingénieur / info:eu-repo/semantics/nonPublished