Integrated Heat Load Research Articles

Experimental Investigations on High-Speed Aerodynamic Heating over Inflatable Decelerators

Heat transfer measurements and high-speed schlieren flow visualization were performed over a cylinder with an aft-attached inflatable decelerator configuration at Mach 3.9. Back wall temperature measurements using thermocouples were used to derive the heat flux distribution over the decelerator surface. A strong nonlinear rise in peak heat flux with angle of attack is observed and windward–leeward effects are quantified. Using appropriate flow and geometric symmetries, heat maps are generated over the regions of the decelerator, which are critical for its thermal design. The effect of fore-end bluntness in terms of peak heat fluxes and integrated heat loads is investigated. The results indicate a possibility of flying at a small angle of attack to reduce the integrated heat loads while simultaneously generating some lift. Tests up to an angle of attack of 20 deg show that even though the peak heating rises strongly, the overall integrated heat loads on the surface of the decelerator remain manageable. A local minimum is seen in peak heat flux levels on the surface at angles of attack of 12–16 deg.

Numerical Simulation of Transpiration Cooling for a High-Speed Vehicle with Substructure

This paper presents a numerical model that assesses the effect of applying transpiration cooling to both the outer wall and the substructure of a high-speed flight vehicle. The porous impulse response analysis for transpiration cooling evaluation (PIRATE) code has been extended and validated to account for quasi-two-dimensional lateral heat conduction effects, thereby allowing for analysis of more complex geometries. This enables very fast calculations of the two-dimensional transient temperature response of a transpiration-cooled thermal protection system suitable for first-order systems studies. To solve for the transpiration-cooled outer wall and a two-dimensional solid substructure, PIRATE has been coupled with the commercial finite element package COMSOL. This enables modeling of the longer-duration thermal effects of the integrated heat load over a flight trajectory. Transpiration cooling using helium coolant has been applied to a wing leading-edge model with an aluminum substructure. Carbon–carbon ceramic composite and the ultra-high-temperature ceramic Zirconium diboride () are chosen as candidate materials. Results for the substructure temperature history for the space shuttle reentry trajectory are obtained, showing that transpiration cooling can lead to a 35% reduction in peak substructure temperature and a 65% reduction in thermal gradients.

Minimum-Fuel Low-Earth-Orbit Aeroglide and Aerothrust Aeroassisted Orbital Transfer Subject to Heating Constraints

A numerical optimization study of minimum-fuel low-Earth-orbit aeroglide and aerothrust aeroassisted orbital transfer of a small spacecraft subject to constraints on heating rate and heating load with a required inclination change is considered. The aeroassisted orbital transfer is formulated as a three-phase optimal control problem consisting of an exoatmospheric deorbit phase, an atmospheric flight phase that may or may not include thrust, and a second exoatmospheric flight phase. The two different variations of the three-phase optimal control problem that arise from the trajectory design are solved using a high-accuracy adaptive Gaussian quadrature collocation method. The fuel consumption of both the aeroglide and aerothrust aeroassisted maneuvers are assessed as functions of the maximum allowable heating rate, the maximum allowable integrated heat load, the maximum lift-to-drag ratio of the vehicle, and the initial mass of the vehicle. Finally, the key features of the optimal aeroglide and aerothrust maneuvers are identified.

Theoretical and numerical treatment of modal instability in high-power core and cladding-pumped Raman fiber amplifiers.

Raman fiber lasers have been proposed as potential candidates for scaling beyond the power limitations imposed on near diffraction-limited rare-earth doped fiber lasers. One limitation is the modal instability (MI) and we explore the physics of this phenomenon in Raman fiber amplifiers (RFAs). By utilizing the conservation of number of photons and conservation of energy in the absence of loss, the 3 × 3 governing system of nonlinear equations describing the pump and the signal modal content are decoupled and solved analytically for cladding-pumped RFAs. By comparing the extracted signal at MI threshold for the same step index-fiber, it is found that the MI threshold is independent of the length of the amplifier or whether the amplifier is co-pumped or counter-pumped; dictated by the integrated heat load along the length of fiber. We extend our treatment to gain-tailored RFAs and show that this approach is of limited utility in suppressing MI. Finally, we formulate the physics of MI in core-pumped RFAs where both pump and signal interferences participate in writing the time-dependent index of refraction grating.

Improved Perturbative Solution of Yaroshevskii's Planetary Entry Equation

An improved approximate analytical solution is developed for Yaroshevskii's classical planetary entry equation for the ballistic entry of a spacecraft into planetary atmospheres at circular speed. Poincaré's method of small parameters is used to solve for the altitude and flight path angle as a function of the spacecraft's speed. From this solution, other important expressions are developed including deceleration, stagnation-point heat rate, and stagnation-point integrated heat load. The accuracy of the solution is assessed via numerical integration of the exact equations of motion. The solution is also compared to the classical solutions of Yaroshevskii and Allen and Eggers. The new second-order analytical solution is more accurate than Yaroshevskii's fifth-order solution for a range of shallow (−3 deg) to steep (up to −90 deg) entry flight path angles, thereby extending the range of applicability of the solution as compared to the classical Yaroshevskii solution, which is restricted to an entry flight path of approximately −40 deg.

Optimal trajectory and heat load analysis of different shape lifting reentry vehicles for medium range application

Abstract The objective of the paper is to compute the optimal burn-out conditions and control requirements that would result in maximum down-range/cross-range performance of a waverider type hypersonic boost-glide (HBG) vehicle within the medium and intermediate ranges, and compare its performance with the performances of wing-body and lifting-body vehicles vis-a-vis the g-load and the integrated heat load experienced by vehicles for the medium-sized launch vehicle under study. Trajectory optimization studies were carried out by considering the heat rate and dynamic pressure constraints. The trajectory optimization problem is modeled as a nonlinear, multiphase, constraint optimal control problem and is solved using a hp-adaptive pseudospectral method. Detail modeling aspects of mass, aerodynamics and aerothermodynamics for the launch and glide vehicles have been discussed. It was found that the optimal burn-out angles for waverider and wing-body configurations are approximately 5° and 14.8°, respectively, for maximum down-range performance under the constraint heat rate environment. The down-range and cross-range performance of HBG waverider configuration is nearly 1.3 and 2 times that of wing-body configuration respectively. The integrated heat load experienced by the HBG waverider was found to be approximately an order of magnitude higher than that of a lifting-body configuration and 5 times that of a wing-body configuration. The footprints and corresponding heat loads and control requirements for the three types of glide vehicles are discussed for the medium range launch vehicle under consideration.

Optimal trajectory analysis of hypersonic boost-glide waverider with heat load constraint

Purpose – The purpose of the paper is to study the variation of optimal burnout angle at the end of the ascent phase and the optimal control deflection during the glide phase, that would maximize the downrange performance of a hypersonic boost-glide waverider, with variation in heat rate and integrated heat load limit. Design/methodology/approach – The approach used is to model the boost phase so as to optimize the burnout conditions. The nonlinear, multiphase, constraint optimal control problem is solved using an hp-adaptive pseudospectral method. Findings – The constraint heat load results for the waverider configuration reveal that the integrated heat load can be reduced by more than half with only 10 per cent penalty in the overall downrange of the hypersonic boost-glide vehicle, within a burnout speed range of 3.7 to 4.3 km/s. The angle-of-attack trim control requirements increase with stringent heat rate and integrated heat load bounds. The normal acceleration remains within limits. Practical implic...

Feasibility of Guided Entry for a Crewed Lifting Body Without Angle-of-Attack Control

The feasibility of flying a crewed lifting body, such as the HL-20, during entry from low Earth orbit without the steady-state body flap deflections required for angle-of-attack control was evaluated. This entry strategy mitigates the severity of the aerothermal environment on the vehicle’s body flaps and reserves control power for transient steering maneuvers. A real-time numeric predictor–corrector entry guidance algorithm was developed to accommodate the range of vehicle trim angle-of-attack profiles possible in the absence of angle-of-attack control. Results show that it is feasible to steer the vehicle from low Earth orbit to a desired target with real-time guidance while satisfying a reasonable suite of trajectory constraints, including limits on peak heat rate, peak sensed deceleration, and integrated heat load. Uncertainty analyses confirm this result and show that the vehicle maintains significant performance robustness to expected day-of-flight uncertainties. Additionally, parametric scans over mission design parameters of interest indicate that a high level of flexibility is available for the low-Earth-orbit return mission. Together, these results indicate that the proposed entry strategy is feasible: Crewed lifting bodies may be effectively flown without steady-state body flap deflections.

Radiative heat transfer analysis in modern rocket combustion chambers

Radiative heat transfer is analyzed for subscale and fullscale rocket combustion chambers for H2/O2 and CH4/O2 combustion using the P1 radiation transport model in combination with various Weighted Sum of Gray Gases Models (WSGGMs). The influence of different wall emissivities, as well as the results using different WSGGMs, the size of the combustion chamber and the coupling of radiation and fluid dynamics, is investigated. Using rather simple WSGGMs for homogeneous systems yields similar results as using sophisticated models. With models for nonhomogeneous systems the radiative wall heat flux (RWHF) decreases by 25–30 % for H2/O2 combustion and by almost 50 % for CH4/O2 combustion. Enlarging the volume of the combustion chamber increases the RWHF. The influence of radiation on the flow field is found to be negligible. The local ratio of RWHF to total wall heat flux shows a maximum of 9–10 % for H2/O2 and 8 % for CH4/O2 combustion. The integrated heat load ratio is around 3 % for H2/O2 and 2.5 % for CH4/O2 combustion. With WSGGMs for nonhomogeneous systems, the local ratio decreases to 5 % (H2/O2) and 3 % (CH4/O2) while the integrated ratio is only 2 % (H2/O2) and 1.3 % (CH4/O2).

Precision Landing at Mars Using Discrete-Event Drag Modulation

An entry, descent, and landing architecture capable of achieving Mars Science Laboratory-class landed accuracy (within 10 km of target) while delivering a Mars Exploration Rover-class payload to the surface of Mars is presented. The architecture consists of a Mars Exploration Rover-class aeroshell with a rigid, annular drag skirt. Maximum vehicle diameter, including drag skirt, is limited to be compatible with current launch-vehicle fairings. A single drag-skirt jettison event is used to control range during entry. Three-degree-of-freedom trajectory simulation is used in conjunction with Monte Carlo techniques to assess the flight performance of the proposed architecture. Results indicate that landed accuracy is competitive with preflight Mars Science Laboratory estimates, and peak heat rate and integrated heat load are significantly reduced relative to the Mars Exploration Rover entry system. Modeling parachute descent within the onboard guidance algorithm is found to remove range error bias present at touchdown; the addition of a range-based parachute deploy trigger is found to significantly improve landed accuracy.

Guided Entry Performance of Low Ballistic Coefficient Vehicles at Mars

Current Mars entry, descent, and landing technology is near its performance limit and may be unable to land payloads on the surface that exceed 1 metric ton. One option for increasing landed payload mass capability is decreasing the entry vehicle’s hypersonic ballistic coefficient. A lower ballistic coefficient vehicle decelerates higher in the atmosphere, providing the additional timeline and altitude margin necessary to land more massive payloads. This study analyzed the guided entry performance of several low ballistic coefficient vehicle concepts on Mars. A terminal point controller guidance algorithm, based on the Apollo Final Phase algorithm, was used to provide precision targeting capability. Terminal accuracy, peak deceleration, peak heat rate, and integrated heat load were assessed and compared with a traditional Mars entry vehicle concept to determine the effects of lowering the vehicle ballistic coefficient on entry performance. Results indicate that, while terminal accuracy degrades slightly with decreasing ballistic coefficient, terminal accuracy rivals the performance of current entry systems for ballistic coefficients as low as . These results demonstrate that guided entry vehicles with low ballistic coefficients (large diameters) may be feasible on Mars. Additionally, the flight performance determined by this investigation may be improved through the use of new guidance schemes designed specifically for low ballistic coefficient vehicles, as well as novel terminal descent systems designed around low ballistic coefficient trajectories.

Integrated decay heat load method to analyze repository capacity impact of a fuel cycle

Assessing the needs for repository capacity from nuclear waste disposal is essential for fuel cycle development or repository development planning. As the repository capacity is mainly constrained by thermal design limits on the repository rocks, a detailed mountain-scale heat transfer calculation is needed for repository capacity impact analysis. In this paper, a simplified repository capacity impact analysis method is proposed as an alternative to performing repository scale heat transfer analysis. The method is based on the use of integrated decay heat load (IDHL) limits. The derived integrated decay heat loads were found to appropriately represent the drift wall temperature limit (200 °C) and the midway between adjacent drifts temperature limit (96 °C) under the high temperature operating mode as long as the wastes are uniformly loaded into the repository. Results indicated that the long-term integrated decay heat load (IDHL L) and the short-term integrated decay heat load (IDHL S) can be effectively used to represent the repository capacity impact for SNFs and HLWs, respectively. Comparisons indicated good agreement between the proposed IDHL method and the repository heat transfer analysis-based approach.

Aeroheating Environments for a Mars Smart Lander

A proposed Mars Smart Lander is designed to reach the surface via lifting-body atmospheric entry (alpha = 16 deg) to within 10 km of the target site. CFD (computational fluid dynamics) predictions of the forebody aeroheating environments are given for a direct entry from a 2005 launch. The solutions were obtained using an 8-species gas in thermal and chemical nonequilibrium with a radiative-equilibrium wall temperature boundary condition. Select wind tunnel data are presented from tests at NASA Langley Research Center. Turbulence effects are included to account for both smooth body transition and turbulence due to heatshield penetrations. Natural transition is based on a momentum-thickness Reynolds number value of 200. The effects of heatshield penetrations on turbulence are estimated from wind tunnel tests of various cavity sizes and locations. Both natural transition and heatshield penetrations are predicted to cause turbulence prior to the nominal trajectory peak heating time. Laminar and turbulent CFD predictions along the trajectory are used to estimate heat rates and loads. The predicted peak turbulent heat rate of 63 W/sq cm on the heatshield leeward flank is 70% higher than the laminar peak. The maximum integrated heat load for a fully turbulent heat pulse is 38% higher than the laminar load on the heatshield nose. The predicted aeroheating environments with uncertainty factors will be used to design a thermal protection system.

Numerical Simulation of Metallic Thermal Protection System Panel Bowing

Numerical simulation of the thermoelastic response of metallic thermal protection panels is presented. The panels, which arebeing designed foruseon the windward surface of theX-33 e ighttest vehicle, deform into convex and concavebowedsurfacesdueto thermalgradientscausedbyaerodynamicheating.Threenumerical models, for the e owe eld, forthein-depth heat transfer, and for the thermoelasticdeformation, are coupled in sequence to yield the transient response of the metallic panel. The aerothermal loads are derived from computational e uid dynamic solutionsandareprescribedasadistributionfunctionwithmaximumbowheightasthegoverningparameter.Finite element models are used to simulate the thermal and structural responses. The coupled simulation is compared to a single-pass uncoupled solution. Results show negligible feedback between the structural deformation and the deformation-induced perturbation of the aerothermal heat load. Nevertheless, signie cant temperature variations on the surface of the panel are produced. The deformations induce lateral temperature gradients that increase the thermal stress within the panel. Finally, it is shown that panel bowing does not appreciably alter the trajectory integrated heat load.

Trajectory-Based Validation of the Shuttle Heating Environment

Chemically reacting,three-dimensional,fullNavier ‐Stokescalculationsaregenerated around theshuttleorbiter and are compared with the STS-2 e ight database at eight trajectory locations. Numerical estimates of quantities necessary for thermal protection system design, surface temperature and heating proe les, integrated heat load, bond-line temperatures, and thermal protection system thicknesses arecompared with theSTS-2 shuttledata. The effects of surface kinetics, turbulence, and grid resolution are investigated. It is concluded that trajectory-based thermal protection system sizing, the use of a Navier ‐Stokes e ow solver combined with a conduction analysis applied over an entry trajectory, is a benee cial tool for future thermal protection system design. This conclusion is based on a reasonable agreement between the e ight data and numerical predictions of surface heat transfer and temperature proe les, integrated heat loads and bond-line temperatures at most of the wind-side thermocouples. Theeffectsofturbulentheating on thermalprotection system designareillustrated.Forfuturelargeentry vehicles, it is concluded that the prediction of turbulent transition will be a major driver in the thermal protection system design process. Finally,onepotential payoff ofusingtrajectory-based thermalprotection system sizing,a reduction in thermal protection system mass, is illustrated. Nomenclature CT = heat transfer coefe cient, W/m 2 -K cp = heat capacity of solid cs

Factors influencing the design and performance of mars entry vehicles

reference area planetary miss distance, rp (1 + 2/i/rpFco), in Fig. 2, ft (or km) receiver bandwidth [Eq. (1) and Fig. 4], cps drag coefficient lift curve slope pitch damping coefficient diameter, ft receiver noise figure [Eq. (1)] frequency, cps total antenna gain gravitational acceleration; gc at Mars surface = 12.3 ft/sec atmosphere scale height, ft (or km) noise power density, w-sec/cycle atmospheric losses (Earth + Mars, if direct link) total circuit loss free space loss reference length, ft molecular weight mass, slug transmitted power, w pressure; psurf = surface pressure on Mars, psf (or mb, if stated) stagnation pressure, psf integrated heat load during entry, Btu/ft heat flux; Qc = convective flux at stagnation point; Qr = radiative flux, Btu/ft-sec q R Re RX Rsnt rp S