1,721,007 research outputs found
Rotational Temperature Measurement in Hypersonic Shock Tunnel using Tunable Diode Laser Absorption Spectroscopy
No full textAtmospheric entry of a spacecraft is considered as one of the critical phases in any planetary mission. Although it lasts only for a few minutes, rapid heating and deceleration can cause severe problems to the spacecraft. Hence, the understanding of the re-entry flow field around the vehicle and design of an optimized thermal protection system (TPS) to it is important, which depends on the reliable data on aerodynamic heating. Shock tunnels are proven ground test facilities for generating a hypersonic flow to investigate the high-temperature effects around the test model. As per the shock tunnel design, the shock compressed reservoir gas at the end of shock tube will expand through a nozzle to simulate the specific kinetic energy of a flow and this may be non-equilibrium in both chemically and thermally at high enthalpy flow conditions. So, the flow generated by shock tunnels differ fundamentally from the atmosphere air flow in a real flight. A better understanding of a freestream temperature generated by a shock tunnel is strongly needed for aerothermodynamic analysis.
In this work, hypersonic shock tunnel HST-2 and free piston driven shock tunnel (FPST) HST-3 in the laboratory for hypersonics and shockwave research (LHSR) were used to generate a hypersonic freestream. The rotational temperature of the freestream was measured using tunable diode laser absorption spectroscopy (TDLAS) technique by probing water vapor (2) absorption lines near 1392 nm using a vertical cavity surface emitting laser (VCSEL) scanned at 25000 Hz. Experiments were performed at different locations from the nozzle exit and different enthalpy conditions in a flow Mach number, M≈8 in HST-2 and M≈10 in HST-3.
The effect of vibrational non-equilibrium over rotational/translational temperature through V-R,T relaxation process of 2 and 2 in a gas flow has been observed at low enthalpy conditions. The effect of thermal and pressure collisions in the hypersonic flow have been analyzed from the measured rotational temperature and full width at half maximum (FWHM) of the absorption line. The establishment of thermal equilibrium between the rotational and translational temperature in the flow has been analyzed and validated the assumption of rotational/translational temperature equilibrium in the hypersonic flow
Experimental Investigation Of Hypersonic Boundary Layer Modifications Due To Heat Addition And Enthalpy Variation Over A Cone Cylinder Configuration
No full textDespite years of research in high speed boundary layer flow, there is still a need for insightful experiments to realize key features of the flow like boundary layer response to different conditions and related transition mechanisms. Volumes of data on the these problems point to the fact that there is still much to be understood about the nature of boundary layer instability causing transition and growth of boundary layer in different conditions. Boundary layer stability experiments have been found to be more useful, in which the boundary layer is perturbed and its behavior observed to infer useful conclusions. Also, apart from the stability part, the effect of various changes in boundary layer due to the perturbation makes interesting observation to gain more insight into the understood and the not so understood facets of the same.
In view of the above, the effect of a steady axisymmetric thermal bump is investigated on a hypersonic boundary layer over a 60º sharp cone cylinder model. The thermal bump, placed near tip of the cone, perturbs the boundary layer, the behavior of which is observed by recording the wall heat flux on the cone and cylinder surface using platinum thin film sensors. The state of the boundary layer is qualitatively assessed by the wall heat flux comparisons between laminar and turbulent values. The same thermal bump also acts as a heat addition source to boundary layer in which case this recorded data provides a look into the effect of the heat addition to the wall heat flux. To gain a larger view of heat addition causing changes to the flow, effects of change in enthalpy are also considered.
Experiments are performed in the IISc HST2 shock tunnel facility at 2MJkg−1 stag-nation enthalpy and Mach number of 8,with and without the thermal bump to form comparisons. Some experiments are also performed in the IISc HST3 free piston driven shock tunnel facility at 6MJkg−1, to investigate the effect of change in stagnation enthalpy on the wall heat flux. To support the experimental results theoretical comparisons and computational studies have also been carried out.
The results of experiments show that the laminar boundary layer over the whole model remains laminar even when perturbed by the thermal bump. The wall heat flux measurements show change on the cone part where there seems to be fluctuation in the temperature gradients caused by the thermal bump, which decrease at first and then show an increase towards the base of the cone. The cylinder part remains the same with and without the thermal bump, indicating heavy damping effects by the expansion fan at cone cylinder junction. A local peak in wall heat flux is observed at the junction which is reduced by 64% by the action of the thermal bump. The possible reason for this is attributed to the increased temperature gradients at the wall due to delayed dissipation of heat that is accumulated in the boundary layer as a result of the thermal bump action. The comparison of data for enthalpies of 2MJkg−1 and 6MJkg−1 show that there are negligible real gas effects in the higher enthalpy case and they do not affect the wall heat flux much. Also it is found that the thermal bump fails to dump heat into the flow directly though it creates heat addition virtually by mere discontinuity in the surface temperature and causes temperature gradients fluctuation in the boundary layer. Considering the thermal bump action and the change in stagnation enthalpy of the flow, there seems to be no change in both cases that can be attributed to a common observation resulting from the factor of change in heat inside the boundary layer
Development of Novel Heat Transfer Gauges Based on Large Carbon Clusters to Measure Total as well as Radiative Heat Flux for Planetary Entry Configurations in Hypersonic Shock Tunnels
No full textThe quest of travelling beyond earth, preludes with ground based experimental studies, detailed analysis and accurate calculations in the aspects of having a safer design of flying vehicles. As the vehicles plunge into the dense atmosphere with greater velocities to hypersonic Mach numbers, the shockwave produced ahead of the aerodynamic body becomes highly intense producing volatile conditions at a temperature of several thousands of Kelvins. Predominantly the unsteady effects are dominated by radiations in the velocities which are greater than 6km/s. During such high enthalpy flows, the atmospheric molecules which cross the strong shockwave are excited to higher energy states. Therefore the shocked gas ahead of the space vehicle is at a state of chemical and thermal non-equilibrium. To attain equilibrium condition, energy is released from high enthalpy fluid to surroundings. The aerodynamic body which faces this energy release is heated by all modes of heat transfer. Behind the normal shock, excited molecules relax to lower levels by energy releasing mechanisms including emission of radiation. In this process, initial photons emitted are absorbed by other molecules further raising its energy levels leading to dissociation and ionization. During recombination of molecular species more photons are released. Such radiations from molecules and shock wave reach the surface of the aerodynamic body. Collective absorption of all incident radiation heats up the surface of planetary entry body. In particular, the radiative heating predominates at very high velocities. Direct measurement of total radiative heating is highly challenging due to the complexity in finding out a proper measurement device. Existing literatures show that only a partial amount of radiative heating could be measured by thin film gauges, since the efficiency of thin film based measurement technique depends on the absorption of sensing element used and the wavelength range of the radiation.
In the present work, it is attempted to measure the radiative heat flux over aerodynamic body in the hypersonic flow condition. To overcome the limitations imposed by the existing measurement technique, a novel thermal sensing element based on Carbon is devised, which is denoted as Large Carbon Cluster. LCC is prepared by single step pyrolysis technique with benzene and ferrocene as precursor mixture. The ratio of precursor mixture is varied to find the proper LCC layer to be formed on a ceramic substrate to get a particular electrical resistance in order to use it as a thermal sensing element. Calibration of the devised carbon allotrope i.e. LCC is found to be having very good thermo-electric characteristics.
Several thermal gauges are developed based on LCC for aerodynamic models to test them for the total heat flux rate in Mach 8 hypersonic flow generated in hypersonic shock tunnel – HST2. The performance of the gauges is compared with the existing platinum based thin-film thermal gauges. It is found that the LCC based thin film gauges perform better than platinum thin-film heat transfer gauges. The durability of LCC is also found to be better than platinum.
The main aim of finding LCC is that it has good optical absorptivity than any other thermal sensing element; therefore it can be used in radiative heat flux measurements. Aerodynamic models are prepared with the radiative thermal gauges based on LCC and these models are tested in different atmospheric hypersonic test flows. The results reveal that radiative heat flux rate is significantly measured even at lower velocity hypersonic flow conditions. This gives a great confidence on using the LCC based thermal gauges for higher velocity flow conditions and to real time test flights
Experimental Investigation Of The Effect Of Nose Cavity On The Aerothermodynamics Of The Missile Shaped Bodies Flying At Hypersonic Mach Numbers
No full textHypersonic vehicles are exposed to severe heating loads during their flight in the atmosphere. In order to minimize the heating problem, a variety of cooling techniques are presently available for hypersonic blunt bodies. Introduction of a forward-facing cavity in the nose tip of a blunt body configuration of hypersonic vehicle is one of the most simple and attractive methods of reducing the convective heating rates on such a vehicle. In addition to aerodynamic heating, the overall drag force experienced by vehicles flying at hypersonic speeds is predominate due to formation of strong shock waves in the flow. Hence, the effective management of heat transfer rate and aerodynamic drag is a primary element to the success of any hypersonic vehicle design. So, precise information on both aerodynamic forces and heat transfer rates are essential in deciding the performance of the vehicle. In order to address the issue of both forces and heat transfer rates, right kind of measurement techniques must be incorporated in the ground-based testing facilities for such type of body configurations. Impulse facilities are the only devices that can simulate high altitude flight conditions. Uncertainties in test flow conditions of impulse facilities are some of the critical issues that essentially affect the final experimental results. Hence, more reliable and carefully designed experimental techniques/methodologies are needed in impulse facilities for generating experimental data, especially at hypersonic Mach numbers.
In view of the above, an experimental program has been initiated to develop novel techniques of measuring both the aerodynamic forces and surface heat transfer rates. In the present investigation, both aerodynamic forces and surface heat transfer rates are measured over the test models at hypersonic Mach numbers in IISc hypersonic shock tunnel HST-2, having an effective test time of 800 s. The aerodynamic coefficients are measured with a miniature type accelerometer based balance system where as platinum thin film sensors are used to measure the convective heat transfer rates over the surface of the test model.
An internally mountable accelerometer based balance system (three and six-component) is used for the measurement of aerodynamic forces and moment coefficients acting on the different test models (i.e., blunt cone with after body, blunt cone with after
body and frustum, blunt cone with after body-frustum-triangular fins and sharp cone with after body-frustum-triangular fins), flying at free stream Mach numbers of 5.75 and 8 in hypersonic shock tunnel. The main principle of this design is that the model along with the internally mounted accelerometer balance system are supported by rubber bushes and there-by ensuring unrestrained free floating conditions of the model in the test section during the flow duration. In order to get a better performance from the accelerometer balance system, the location of accelerometers plays a vital role during the initial design of the balance. Hence, axi-symmetric finite element modeling (FEM) of the integrated model-balance system for the missile shaped model has been carried out at 0° angle of attack in a flow Mach number of 8. The drag force of a model was determined using commercial package of MSC/NASTRAN and MSC/PATRAN. For test flow duration of 800 s, the neoprene type rubber with Young’s modulus of 3 MPa and material combinations (aluminum and stainless steel material used as the model and balance) were chosen. The simulated drag acceleration (finite element) from the drag accelerometer is compared with recorded acceleration-time history from the accelerometer during the shock tunnel testing. The agreement between the acceleration-time history from finite-element simulation and measured response from the accelerometer is very good within the test flow domain. In order to verify the performance of the balance, tests were carried out on similar standard AGARD model configurations (blunt cone with cylinder and blunt cone with cylinder-frustum) and the results indicated that the measured values match very well with the AGARD model data and theoretically estimated values. Modified Newtonian theory is used to calculate the aerodynamic force coefficient analytically for various angles of attack.
Convective surface heat transfer rate measurements are carried out by using vacuum sputtered platinum thin film sensors deposited on ceramic substrate (Macor) inserts which in turn are embedded on the metallic missile shaped body. Investigations are carried out on a model with and without fin configurations in HST-2 at flow Mach number of 5.75 and 8 with a stagnation enthalpy of 2 MJ/kg for zero degree angle of attack. The measured heating rates for the missile shaped body (i.e., with fin configuration) are lower than the predicted stagnation heating rates (Fay-Riddell expression) and the maximum difference is about 8%. These differences may be due to the theoretical values of velocity gradient used in the empirical relation. The experimentally measured values are expressed in terms of normalized heat transfer rates, Stanton numbers and correlated Stanton numbers, compared with the numerically estimated results. From the results, it is inferred that the location of maximum heating occurs at stagnation point which corresponds to zero velocity gradient. The heat-transfer ratio (q1/Qo)remains same in the stagnation zone of the model when the Mach number is increased from 5.75 to 8. At the corners of the blunt cone, the heat transfer rate doesn’t increase (or) fluctuate and the effects are negligible at two different Mach numbers (5.75 and 8). On the basis of equivalent total enthalpy, the heat-transfer rate with fin configuration (i.e., at junction of cylinder and fins) is slightly higher than that of the missile model without fin.
Attempts have also been made to evaluate the feasibility of using forward facing cavity as probable technique to reduce the heat transfer rate and to study its effect on aerodynamic coefficients on a 41° apex angle missile shaped body, in hypersonic shock tunnel at a free stream Mach number of 8. The forward-facing circular cavities with two different diameters of 6 and 12 mm are chosen for the present investigations. Experiments are carried out at zero degree angle of attack for heat transfer measurements. About 10-25 % reduction in heat transfer rates is observed with cavity at gauge locations close to stagnation region, whereas the reduction in surface heat transfer rate is between 10-15 % for all other gauge locations (which is slightly downstream of the cavity) compared with the model without cavity.
In order to understand the influence of forward facing cavities on force coefficients, measurement of aerodynamic forces and moment coefficients are also carried out on a missile shaped body at angles of attack. The same six component balance is also being used for subsequent investigation of force measurement on a missile shaped body with forward facing cavity. Overall drag reductions of up to 5 % is obtained for a cavity of 6 mm diameter, where as, for the 12 mm cavity an increase in aerodynamic drag is observed (up to about 10%). The addition of cavity resulted in a slight increase in the missile L/D ratio and did not significantly affect the missile lateral components. In summary, the designed balances are found to be suitable for force measurements on different test models in flows of duration less than a millisecond.
In order to compliment the experimental results, axi-symmetric, Navier-Stokes CFD computations for the above-defined models are carried out for various angles of attack using a commercial package CFX-Ansys 5.7. The experimental free stream conditions obtained from the shock tunnel are used for the boundary conditions in the CFD simulation. The fundamental aerodynamic coefficients and heat transfer rates of experimental results are shown to be in good agreement with the predicted CFD.
In order to have a feeling of the shock structure over test models, flow visualization experiments have been carried out by using the Schlieren technique at flow Mach numbers of 5.75 and 8. The visualized shock wave pattern around the test model consists of a strong bow shock which is spherical in shape and symmetrical over the forebody of the cone. Experimentally measured shock stand-off distance compare well with the computed value as well as the theoretically estimated value using Van Dyke’s theory. These flow visualization experiments have given a factual proof to the quality of flow in the tunnel test section
Investigation of Heat Transfer Rates Around the Aerodynamic Cavities on a Flat Plate at Hypersonic Mach Numbers
Full text linkAerodynamic cavities are common features on hypersonic vehicles which are caused in both large and small scale features like surface defects, pitting, gap in joints etc. In the hypersonic regime, the presence of such cavities alters the flow phenomenon considerably and heating rates adjacent to the discontinuities can be greatly enhanced due to the diversion of flow. Since the 1960s, a great deal of theoretical and experimental research has been carried out on cavity flow physics and heating. However, most of the studies have been done to characterize the effect downstream and within the cavity. In the present study, a series of were carried out in the shock tunnel to investigate the heating characteristics, upstream and on the lateral side of the cavity. Heat flux measurement has been done using indigenously developed high resistance platinum thin film gauges. High resistance gauges, as contrary to the conventionally used low resistance gauges were showing good response to the extremely low heat flux values on a flat plate with sharp leading edge. The experimental measurements of heat done on a flat plate with sharp leading edge using these gauges show good match with theoretical relation by Crabtree et al. Flow visualization using high speed camera with the cavity model and shock structures visualized were similar to reported in supersonic cavity flow. This also goes to state that in spite of the fluctuating shear layer-the main feature of hypersonic flow over a cavity ,reasonable studies can be done within the short test time of shock tunnel.
Numerical Simulations by solving the Navier-Stokes equation, using the commercially available CFD package FLUENT 13.0.0 has been done to complement the experimental studies
Analysis Of Solar Pumped Chemical Oxygen Iodine Laser
No full textChemical Oxygen Iodine Laser(COIL) is an electronic transition high energy chemical laser having a wavelength of 1.315 /mi. This is the first chemical laser to operate on an electronic rather than a rotational or vibrational transition. In principle the COIL can be operated either in pulsed or cw mode. Its interest lies in high chemical efficiency, high power and wavelength which is shortest among all the chemical lasers. COIL finds a wide range of applications as its output wavelength at 1.315/zm couples well with the surface of most metals. The applications include surface hardening and modification of metals, welding, drilling and cutting of metals, cutting of ceramics, micro machining, laser deposition of non metallic coatings on metallic surfaces, monitoring of atmospheric pollutants and solar hazardous waste detoxification. Moreover, its wavelength is suitable for fiber optic transmission.
In COIL the laser output at 1.315 /an is achieved by stimulated emission on the f (2-PL/2) -* -f (2-p3/2) magnetic dipole transition in atomic iodine. The population inversion on this transition is obtained by resonant collisions! energy transfer from metastable excited Oj^A) molecules produced by a chemical reaction of KOH, H2O? and Cl2. The chemical reaction of H2O2 and Cl2 that produces oxygen molecules is highly exothermic, and because of spin conservation considerations, channels its energy directly into the metastable electronically excited singlet delta state of oxygen molecule. Since the O2(1A) has a 45 mins lifetime and hence an extremely low small signal gain coefficient, it cannot be lased directly. Lasing can be achieved, however, if this energy is transferred to an atom or molecule which has a reasonable transition moment between its excited and ground states. The iodine 52P^2 -> 52P3/2 magnetic dipole transition has an acceptable transition moment and is nearly resonant with the 02{lA) state in oxygen. Excited iodine atoms are obtained by mixing O2(l A) and l2 molecules resulting in their dissociation and subsequent excitation.
Power levels in excess of 25 kW have been reported in COIL. Due to wide range of applications and mainly for its use as a laser weapon, efforts are being made to enhance the power to higher levels. The dissociation of I2 controls the gain of the coil and hence power. In the pure COIL scheme some of the I2 remains undissociated due to the recombination reactions. Hence if we add a mechanism to dissociate the residual I2 molecules, we can enhance the performance of the COIL. So we propose to add a solar pumping to conventional COIL, which by photo exciting the undissociated I2lead to increase in efficiency.
The thesis contains six chapters in which chapter 1 contains a general introduction and the definition of the research problem. The basic theory and the chemical reactions are discussed in chapter 2, The proposed model is discussed and the rate equations are solved in chapter 3. The numerical scheme and the computer code along-with the validation of the code are presented in chapter 4. The numerical results for the species concentrations, population inversion density and the output power for the proposed solar pumped COIL are presented in chapter 5, Final conclusions and future scope of the proposed research are presented in the final chapter 6.
(Pl refer the original document for formulas
Experimental Investigations on Hypersonic Waverider
Full text linkIn the flying field of space transportation domain, the increased efforts involving design and development of hypersonic flight for space missions is on toe to provide the optimum aerothermodynamic design data to satisfy mission requirements. Aerothermodynamics is the basis for designing and development of hypersonic space transportation flight vehicles such as X 51 a, and other programmes like planetary probes for Moon and Mars, and Earth re-entry vehicles such as SRE and space shuttle. It enables safe flying of aerospace vehicles, keeping other parameters optimum for structural and materials with thermal protection systems. In this context, the experimental investigations on hypersonic waverider are carried out at design Mach 6.
The hypersonic waverider has high lift to drag ratio at design Mach number even at zero degree angle of incidence, and this seems to be one of the special characteristics for its shape at hypersonic flight regime. The heat transfer rates are measured using 30 thin film platinum gauges sputtered on a Macor material that are embedded on the test model. The waverider has 16 sensors on top surface and 14 on bottom surface of a model. The surface temperature history is directly converted to heat transfer rates. The heat transfer data are measured for design (Mach 6) and off-design Mach numbers (8) in the hypersonic shock tunnel, HST2. The results are obtained at stagnation enthalpy of ~ 2 MJ/kg, and Reynolds number range from 0.578 x 106 m-1 to 1.461 x 106 m-1. In addition, flow visualization is carried out by using Schlieren technique to obtain the shock structures and flow evolution around the Waverider. Some preliminary computational analyses are conducted using FLUENT 6.3 and HiFUN, which gave quantitative results. Experimentally measured surface heat flux data are compared with the computed one and both the data agree well. These detailed results are presented in the thesis
Characterization of a large scale hypersonic shock tunnel and investigation of the effect of roughness on a large cone boundary layer flow
No full textHigh speed flight is ridden with engineering challenges, some of which have been studied
for decades with only minor improvements. Boundary layer transition to turbulence in
highMach number flow is one such topic, which, although has seen impressive improvements,
remains far from staple engineering prediction. The boundary layer transition at
high speeds is critical due to the high levels of heating that the turbulent boundary layer
brings, but the same turbulent layer is also sought after for its resistance towards separation
in adverse pressure gradients. This advantage is capitalized by Scramjet inlets where
boundary layer separation can lead to unstart of the engine, and so the boundary layer
must not be laminar when entering the engine. Given the hypersonic flight conditions of
low unit length Reynolds numbers at high altitudes, most high Mach number flights cannot
expect to have natural boundary layer transition and therefore must actively trip the
boundary layer towards transition. This effect of a given roughness on the boundary layer
at aMach 8 hypersonic flowover a generic axisymmetric forebody is the prime objective of
this study. The study is performed at a low Reynolds number condition, which is realistic
for a Mach 8 flight, to test an unfavorable condition for producing a turbulent boundary
layer.
The present study employs an 800mmlong sharp cone as the test body for the boundary
layer experiments. The long model is housed in a recently characterized large scale
shock tunnel. Experiments are conducted in Mach 8 flow at a low Reynolds number of
3.2 million per m. To trip the flow, roughness in the form of diamond shaped isolated
3-dimensional elements are primarily used, given the previous success from other studies
which make it the most effective trip shape in flows with such speeds. For flows with
boundary layer edge conditions that are hypersonic, the trip height requirements have
been previously found to be greater than the local velocity boundary layer thickness. For
the present study, the trip heights have been varied up to 5 times the local boundary layer
thickness. For all the cases studied, there is no clear hint of a transitional boundary layer.
The baseline case without trips, is found to be fully laminar. The cases with trips do affect
the boundary layer, and this is observed by the heat flux variations behind the trips,
but in no case does the heat flux variation clearly indicate development of a transitional
boundary layer. This result brings out the importance of Reynolds numbers at high altitude
hypersonic flight conditions, presenting a case of challenging tripping environment
that may not even lead to a turbulent boundary layer. Moreover, the axisymmetric nature
of the body makes transition further difficult, where previously obtained data in other
studies showthat the flat plate case is relatively easier to bring to boundary layer transition.
Finally, a comparison of the data is made with a few existing roughness correlations for
transition prediction. It is found that the correlations do not work well in the hypersonic
boundary layer edge flow conditions and that they over predict the effect of the Reynolds
numbers.
In carrying out this boundary layer study, a large scale shock tunnel has been calibrated
and characterized. Apart from the regular calibration, a new analysis of the shock tube
data is presented that helps to study the available test gas slug in such a facility. The new
analysis brings out the various non-ideal effects involved in the deterioration of the performance
of a facility and can be applied to any such facility without the need for any special
equipment or measurement. For the large scale shock tunnel, this analysis helps to quantify
the test gas slug as a function of operational incident shockMach number alone. This
information is further found useful in accurate test time prediction in non-tailored mode
of tunnel operation but may be applied to tailored conditions as well. Finally, using a previously
established early driver gas arrival mechanism, attempts are made to predict and
verify the driver gas arrival in the tunnel mode of operation. This process has been found
to be partially successful, with further work required in this potential application of the
presented test gas slug analysis
Going Beyond Counting First Authors in Author Co-citation Analysis
Full text linkThe present study examines one of the fundamental aspects of author co-citation analysis (ACA) - the way co-citation
counts are defined. Co-citation counting provides the data on which all subsequent statistical analyses and mappings
are based, and we compare ACA results based on two different types of co-citation counting - the traditional type that
only counts the first one among a cited work's authors on the one hand and a non-traditional type that takes into
account the first 5 authors of a cited work on the other hand. Results indicate that the picture produced through this non-traditional author co-citation counting contains more coherent author groups and is therefore considerably clearer. However, this picture represents fewer specialties in the research field being studied than that produced through the traditional first-author co-citation counting when the same number of top-ranked authors is selected and analyzed. Reasons for these effects are discussed
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