An approximate method of calculating the minimum total weight of the two-insulation--two-coolant thermal protection system is developed. The equations derived in the development of the approximate method enable insight into the parameters that control the system weight. Two cases are considered: the case where the outer coolant location is unrestricted within the insulating wall and the case where the outer coolant location is restricted within the insulating wall. The effects on system weight of material properties and the outer coolant location within the insulating wall are discussed. A comparison of weights predicted by the approximate method and numerical solutions is shown. / Master of Science

No description available.

No description available.

A new gage which can measure skin friction and heat flux simultaneously was designed, constructed, and tested. This gage is the combination of a non-nulling type skin friction balance and a heat flux microsensor. By mounting the heat flux microsensor directly on the surface of the floating element of the skin friction balance, it was possible to perform simultaneous measurements of the skin friction and the heat flux. The total thickness of the heat flux microsensor is less than 2 μm, so the presence of this microsensor creates negligible disruption on the thermal and the mechanical characteristics of the air flow. Tests were conducted in the Virginia Tech supersonic wind tunnel. The nominal Mach number was 2.4, and Reynolds number per meter was 4.87 x 10⁷ with total pressure of 5.2 atm and total temperature of 300 °K. Results of the tests showed that this new gage was quite reliable and could be used repeatably in the supersonic flow. This gage also has an active heating system inside of the cantilever beam of the skin friction balance so that the surface temperature of the floating element can be controlled as desired. With these features, the effects of a temperature mismatch between the gage surface and the surrounding wall on the measurements of the skin friction and the heat flux were investigated. An infrared radiometer was used to measure the surface temperature distributions. Without the active heating, the amount of temperature mismatch generated by the gage itself was from 2.5 °K to 4.5 °K. The active heating produced the temperature mismatch of 18.7 °K. The largest temperature mismatch corresponds to the levels typically found in high heat flux cases when it is expressed in dimensionless terms. This temperature mismatch made sizable effects — a 24 % increase in the skin friction measurement and a 580 % increase in the heat flux measurements. These experimental results were compared with the computational results using the Computational Fluid Dynamics code GASP. The input flow conditions were obtained from the boundary layer measurements. The temperature mismatch was input by specifying the density and the pressure at each grid point on the wall. The Baldwin-Lomax algebraic turbulence model was used with the thin layer approximations. The comparison showed that the difference in the skin friction and heat flux was less than 10 % of the measured data when the temperature mismatch was less than 8.5 °K, but the difference was increased as the amount of the temperature mismatch increased. It is presumed that the disagreement between the measurements and the calculations was caused mainly by deficiencies in the turbulence model for this complex, developing viscous flow, because the Baldwin-Lomax model cannot account for the multiple length scale in this flow. / Ph. D.

Hypersonic flight at sea-level conditions induces severe thermal loads not seen by any other type of current hypersonic system. Appropriate design of the hypersonic round requires a solid understanding of the thermal environment. Numerous codes were obtained and assessed for their applicability to the problem under study, and outside of the GASP Conjugate Heat Transfer module, Navier-Stokes code from Aerosoft, Inc., no efficient codes are available that can model the aerodynamic heating response for a fully detailed projectile, including all subassemblies, over an entire trajectory. Although the codes obtained were not applicable to a fully detailed thermal soak analyses they were useful in providing insight into ablation effects. These initial trade studies indicated that ablation of up to 1.25 inches could be expected for a Carbon-Carbon nosetip in this flight environment. In order to capture the thermal soak effects a new methodology (BMA) was required. This methodology couples the Sandia aerodynamic heating codes with a full thermal finite element model of the desired projectile, using the finite element code ANSYS from ANSYS, Inc. Since ablation can be treated elsewhere it was not included in the BMA methodology. Various trajectories of quadrant elevations of 0.5, 10, 30, 50, and 80 degrees were analyzed to determine thermal time histories and maximum operating temperatures. All of the trajectories have the same launch condition, Mach 8 sea-level, and therefore will undergo the same initial thermal spike in temperature at the nose-tip of approximately 3,100 K (5600R). Of the five trajectories analyzed the maximum internal temperatures experienced occurred for the 50 degree quadrant elevation trajectory. This trajectory experienced temperatures in excess of 1,000 K (1800R) for more than 80% of its flight time. The BMA methodology was validated by comparisons with experiment and computational fluid solutions with an uncertainty of 10% at a cost savings of over three orders of magnitude. / Ph. D.

The analytical theory of the “passive thermal design of spacecraft" can be divided into two parts. The first part is concerned with the description of the radiant heat transfer to spacecraft external surfaces. The second part is concerned with calculating temperature over a spacecraft when the radiant heat incident, on the spacecraft's wall, is known. The first part, the calculation of the heat incident on a spacecraft's external surfaces, has been investigated in the literature. References one, two, and three are examples of such papers. Unfortunately, the results of auch papers are either numerical or else too specialized to be of general interest for the analytical study of the thermal design of spacecraft. The second part, the calculation of temperatures over a spacecraft when the incident radiant heat is known, is also dealt with in the literature. References four and five are examples of such papers. The heat flow, occurring in the walls of spacecraft, is nonlinear because of thermal radiation and few exact solutions are known. This problem is usually attacked by "linearizing'' the nonlinear term or by directly employing power aeries. The solution of the nonlinear heat equation by the linearization process is valid only for small temperature variations. When temperature differences are large, the linearized solutions do not properly account for the nonlinear radiation terms and series error can result. When power series are employed directly to solve the nonlinear heat flow equation, the labor required to solve the time dependent problem is generally excessive because the elementary functions cannot be used efficiently. In this thesis, the radiant heat transferred to spacecraft is found by the use of Fourier series. The resulting solutions are simple and are valid for spacecraft of very general geometry. Heat transfer calculation which previously required extensive integration on electronic computers can be calculated by the results of this thesis with only trivial labor. Also, the results have the advantage of being well suited for use in the solution of the nonlinear heat transfer equation. The problem of heat flow including nonlinear radiation is also attached in this thesis. The method of solution used is closely related to the well known method of successive approximations and allows solution of nonlinear equations which do not have the classical “Small perturbation parameter.” Also, the method of solution used makes good use of the elementary functions so that time dependent problems can be solved without excessive labor. The problems solved in this thesis includes: the temperature time history of a body at uniform temperature but exposed to periodic radiative heating, the temperature time history of a body having nonuniform temperatures and exposed to periodic radiative heating, and finally the problem of linear heat flow with nonlinear boundary conditions. In each case it is shown how linearized solutions neglect the important results of nonlinear radiation heat transfer. / Master of Science

A variation of the boundary element method is developed to determine the distribution of frictional heat and the ensuing surface or subsurface temperature rise caused by frictional heating between sliding solids. The theoretical model consists of two semi-infinite substrates each coated with a film of arbitrary thickness and thermal properties. A three dimensional transient analysis is developed which involves the thermal coupling of the two sliding solids at the true contact areas. The boundary element solution is based on a moving Green's function which naturally incorporates the combined conduction and convection effects due to sliding. Results are presented to display some of the important numerical characteristics of the boundary element solution method. Results are also presented that show the sensitivity of surface temperature rise to contact area evolvement, geometry and subdivision. The effects of surface film thickness and thermal properties on surface temperature rise are presented for a range of Peelet numbers. Lastly, a comparison of theoretical predictions and experimental measurements for surface temperature rise of a graphite epoxy ball loaded against a rotating sapphire disk is presented. / Master of Science

No description available.

The emerging and competitive environment in the space technology requires the improvements in the capability of aerodynamic vehicles. This leads to the analysis in drag reduction of the vehicle along with the minimized heat transfer rate. Using forward facing solid aerospike is the simplest way among the existing drag reduction methodologies for hypersonic blunt cone bodies. But the flow oscillations associated with this aerospike makes it difficult to implement. When analyzing this flow, it can be understood that this oscillating flow can be compared to conical cavity flow. Therefore in the spiked flows, it is decided to implement the technique used in reducing the flow oscillation of the cavities. Based on this method the shallow conical cavity flow generated by the aerospike fixed ahead of a 120o blunt cone body is fissured as multiple cavities by so many disks formed from 10o cone. Now the deep conical cavities had the length to mean depth ratio of unity; this suppresses the unnecessary oscillations of the shallow cavity. The total length of the telescopic aerospike is fixed as 100mm. And one another conical tip plain aerospike of same length is designed for comparing the telescopic spike’s performance at hypersonic flow Mach numbers of 5.75 and 7.9. A three component force balance system capable of measuring drag, lift and pitching moment is designed and mounted internally into the skirt of the model. Drag measurement is done for without spike, conical tip plain spiked and telescopic spiked blunt cone body. The three configurations are tested at different angles of attack from 0 to 10 degree with a step of 2. A discrete iterative deconvolution methodology is implemented in this research work for obtaining the clean drag history from the noisy drag accelerometer signal. The drag results showed the drag reduction when compared to the without spike blunt cone body. When comparing to the plain spiked, the telescopic spiked blunt cone body has lesser drag at higher angles of attack. Heat transfer measurements are done over the blunt cone surface using the Platinum thin film gauges formed over the Macor substrate. These results and the flow visualization give better understanding of the flow and the heat flux rate caused by the flow. The enhancement in the heat flux rate over the blunt cone surface is due to the shock interaction. And in recirculation region the heat flux rate is very much lesser when compared to without spike blunt cone body. It is observed that the shock interaction in the windward side is coming closer towards the nose of the blunt cone as the angle of attack increases and the oscillation of the oblique shock also decreases. Schlieren visualization showed that there is dispersion in the oblique shock, particularly in the leeward side. In the telescopic spike there are multiple shocks generated from each and every disk which coalesces together to form a single oblique shock. And the effect of the shock generated by the telescopic spike is stronger than the effect of the shock generated by the conical tip plain spike.

The recent resurgence in hypersonics is centered around the development of SCRAMJET engine technology to power future hypersonic vehicles. Successful flight trials by Australian and American scientists have created interest in the scramjet engine research across the globe. To develop scramjet engine, it is important to study heat transfer effects on the engine performance and aerodynamic forces acting on the body. Hence, the main aim of present investigation is the design of scramjet engine configuration and measurement of aerodynamic forces acting on the model and heat transfer rates along the length of the combustor. The model is a two-dimensional single ramp model and is designed based on shock-on-lip (SOL) condition. Experiments are performed in IISc hypersonic shock tunnel HST2 at two different Mach numbers of 8 and 7 for different angles of attack. Aerodynamic forces measurements using three-component accelerometer force balance and heat transfer rates measurements using platinum thin film sensors deposited on Macor substrate are some of the shock tunnel flow diagnostics that have been used in this study.