Passive Thermal Protection System Research Articles

Parametric Study of Transpiration Cooling Using Oxides for Sharp Hypersonic Leading Edges

Recent escalating interest in the development of highly maneuverable hypersonic vehicles demands sharp leading edges. However, sharp leading edges induce severe aerothermal conditions where conventional passive or ablative thermal protection systems fail to protect the leading edge. Here, we numerically demonstrate transpiration cooling employing oxide coolants as a new alternative system to thermally protect sharp leading edges. We parametrically characterize the performance of transpiration cooling for various coolant properties, flight conditions, and leading edge radii using a semi-analytic boundary-layer model validated with third-order direct numerical simulations. We further utilize direct numerical simulation to examine the impact of the thermochemical behavior of oxide vapors with the external hypersonic flow on transpiration cooling. Our findings do not readily align with an optimal set of material properties for transpiration cooling. Instead, certain coolant properties are more appropriate for various flight conditions and leading edge sizes. Our results also demonstrate that the thermochemical interactions between the oxide vapors and the external hypersonic flow have a negligible impact on the performance of transpiration cooling. Our study provides numerical frameworks to evaluate the performance of transpiration cooling and optimize the coolant properties for various flight conditions to protect sharp leading edges, which are paramount across hypersonic applications.

Preliminary study on the thermal insulation of a multilayer passive thermal protection system with carbon-phenolic composites in a combustion chamber

This paper is a case study on a thermal protection system in a combustion chamber of a ramjet. A multilayer passive thermal protection system (MPTPS) with carbon-phenolic material was established, and the temperature distribution and insulation performance of MPTPS were studied. The problem is that the carbon-phenolic material is pyrolyzed when subjected to high temperatures, leading to the changes in its thermal properties (e.g., density, specific heat capacity, and thermal conductivity.). Moreover, the carbon-phenolic material absorbs heat during the pyrolysis procedure of phenolic resin. Therefore, in this paper, several experimental studies were conducted on carbon-phenolic composites to investigate its pyrolysis characteristics, including tube furnace heating experiment and Py-GC-MS experiment. On the other hand, a one-dimensional multilayer transient heat conduction approach was developed, many calculations were numerically performed to study the effects of negative heat sources and the thickness of carbon-phenolic composites on the thermal insulation performance of MPTPS. As a result, the carbon-phenolic material was pyrolyzed from 200 °C, and the mass-loss rate was 25% when the heating temperature increased up to 900 °C; the negative heat source delayed the temperature rising point and decreased the outer wall temperature; the thickness of the carbon-phenolic composites affected the final outer wall mainly by conduction.

Optimization of Active and Passive Thermal Protection Systems for a Hypersonic Vehicle

The hypersonic vehicle model Michigan–AFRL Scramjet in Vehicle (MASIV) is used to optimize both the active and the passive thermal protection systems on a scramjet-powered generic X-43 waverider. The passive thermal protection system is the NASA ARMOR design; a silicon dioxide insulation layer is sandwiched between a radiation shield and the vehicle titanium skin. The active thermal protection system consists of a heat exchanger on the combustor wall; the coolant is the liquid hydrogen fuel. The portion of the fuel that is used for cooling (and not for propulsion) is recirculated back into the fuel tank, which causes the temperature of fuel in the tank to rise. Gradient-based optimization is performed to determine 1) the minimum insulation thickness distribution required and 2) the optimal coolant mass flow rate and its variation in time. Constraints are that surface temperatures cannot exceed the failure temperatures, and the volume of fuel in the fuel tank cannot exceed the tank volume. The hypersonic vehicle heat transfer model is adaptable and can handle new materials for thermal protection, as they are developed.

The heat dissipation, transport and reuse management for hypersonic vehicles based on regenerative cooling and thermoelectric conversion

The thermal energy management (TM) of a hypersonic vehicle should concern the full process of the heat dissipation, transport and reuse. In this paper, the aerodynamic heat of a hypersonic cruiser is dissipated by passive thermal protection system (TPS), transported by regenerative cooling (RC) network, and reused by RC network and thermoelectric (TE) conversion component. The TM system accordingly includes three subsystems of TPS, RC network and TE component. An equivalent thermal equilibrium model and an overall equivalent heat transfer coefficient are developed to build up the mutual correlation between the aerodynamic heating and TM system instead of the one-way influence, and account for the coupling design rationale of the TM subsystems. For passive TPS, the distribution, area and weight of relevant concepts are obtained; for RC network, the determination method of the heat capacity and coolant mass-flow-rate is developed, and the heat transport performance at specific vehicle regions is numerically analyzed by using hydrocarbon and liquid hydrogen fuel as coolants; for TE component, a TE-AFRSI concept is established by integrating mid- and low-temperature TE stages into the advanced flexible reusable surface insulation (AFRSI), and the concept is optimized by considering the thermal protection, weight increment and heat reuse performance. The design roadmap of TM system is finally proposed and the influence of the overall equivalent heat transfer coefficient is clarified. The results show that the aerodynamic heat and the transported or reused heat proportion will increase, while the scale of passive TPS will be reduced by the increase of overall equivalent heat transfer coefficient.

Thermal conductivities of a needle-punched carbon/carbon composite with unbalanced structures

The XYZ 3-dimensional thermal conductivities of the C/C up to 2000 °C were measured by a laser flash method. Carbon fiber-reinforced carbon composites (C/C) were generally developed for aerospace missions due to their excellent thermal resistivity at ultrahigh temperature. C/C must endure harsh environments such as thousands of degrees Celsius without degradation of its mechanical properties. To solve this problem, among the passive thermal protection system, we suggest a method of conducting more heat through the mono-axial direction, which resulted in ease of the thermal rise in the heat receiving part. For example, the X-43A flight applied unbalanced C/C (UCC) with different carbon fiber orientation ratios according to the XY direction in the leading edge part. To investigate the difference in thermal conductivity between unbalanced C/C (UCC) and balanced C/C (BCC), unbalanced and balanced preforms were prepared by a needle punching process, and then they were densified by pitch infiltration and a carbonization process. We compared and analyzed the effects of unbalanced C/C(UCC) and balanced C/C (BCC) structures on the thermal conductivity. We also designed the “rule of mixtures” equation for calculating thermal conductivities of each C/C using reported data of carbon fiber and graphite matrix. Our calculations of thermal conductivity ratio match the ratio of real data.

Trends in the development of flexible thermal protection systems for modern aircraft

The emergence of new types of space and aviation technology necessitates the development of new types of thermal protection systems capable of operating at high temperature and long operating times. There are several types of thermal protection systems for different operating conditions: active thermal protection systems using forced supply of coolant to the protected surface, passive thermal protection systems using materials with low thermal conductivity without additional heat removal, high-temperature systems, which are simultaneously elements of the bearing structure and provide thermal protection, ablation materials. Heat protection systems in the form of rigid tiles and flexible panels, felt and mats are most common kind of heat protecting systems. This article examines the trends of development of flexible reusable heat protection systems intended for passive protection of aircraft structural structures from overheating.

Оценка эффективности пассивной системы тепловой защиты породного массива

Оценка эффективности пассивной системы тепловой защиты породного массива

The design of thermal management system for hypersonic launch vehicles based on active cooling networks

In this paper, a design method of thermal management system (TMS) for hypersonic vehicles is developed. The system consists of a passive thermal protection system (TPS) and an active cooling network (ACN) with coolant of kerosene. In most previous studies, the passive TPS and ACN are always designed separately and thus leads to over-conservative results that deviating from real engineering conditions. A coupled design method is developed in present work, and the process includes calculation of aerodynamic heat, determination of passive TPS concept distribution, computation of TPS and ACN scales, and iterative design. The coupled design is realized based on two keys, the active cooling is coupled in aerodynamic heating and heat transfer in TPS by an equivalent thermal equilibrium model, and the overall capacity of active cooling is indicated by an equivalent heat transfer coefficient. The model and the coefficient act as the rationale and the equivalent parameter of the whole process, respectively. The TMS of a reusable launch vehicle is established under a typical trajectory. The influences of equivalent heat transfer coefficient on aerodynamic heating, passive TPS and ACN are studied. The results show that the weight of passive TPS decreases, while the coolant mass flow rate increases with the enhancement of active cooling.

Thermal and thermomechanical performance of actively cooled pyramidal sandwich panels

Previous research has proved that pyramidal core sandwich panel is an excellent kind of passive Thermal Protection Systems (TPSs) for hypersonic aerospace aircrafts suffered from in-service environment. To explore its further insulation performance and load-bearing capability, four kinds of actively cooled pyramidal core sandwich panels are designed and investigated in this paper. Convection flow mechanism for cooling channels is described in details. Influences of the coolant velocity, channel height ratio and flow direction of the two layers on thermal performance are firstly studied and discussed in an appropriate way. Analysis of thermomechanical performance is then performed. The effect of the cover board thickness on thermomechanical performance is also studied. In consequence, a light-weight, structural and thermal integrated thermal protection system (ITPS) has been designed to protect hypersonic aircrafts from intense radiation heat flux in service.

Low-energy helium plasma effects on textured micro-porous tungsten

The effects of low-energy helium plasma on surface evolution and restructuring of micro-porous tungsten and tungsten-rhenium micro-engineered materials was experimentally investigated. Tungsten foam samples of 45 and 80 Pores Per Inch (PPI) of three types of textured surface architecture (nodular W, dense Re-pillars, and W-coated Re-pillars) were exposed to low-energy helium plasma at 80 and 200 eV to fluence levels of 1, 2, and 7 × 1025 m−2 at 1123 K. The sidewalls of Re pillars exhibited surface holes due to trapped helium while their tips were found to form complex nano-scale web structures, 100's of microns across, rather than nano-tendrils (fuzz) that are often observed in flat W samples. A nano-necklace bead/string structure was observed for the first time on Re pillars, likely driven by a high density of nano-meter size helium bubbles crowded near pillar tips. W-coated Re pillars exhibited typical fuzz nanostructures seen on flat and untextured W-foam, albeit much shorter. While the height of W-fuzz was found to depend on the square root of ion fluence, the nano-web height was found to be linearly dependent on ion fluence. The fuzz was found to grow much slower on nodular foams as compared to planar tungsten, by as much as 25% slower in the W-coated Re pillar samples. Weight loss due to ion sputtering was found to be significantly reduced in nodular foam as compared to planar tungsten, but was found to be more pronounced with Re pillars. Surface coverage by nano-webs may provide insulation against plasma disruptions and act as a passive Thermal Protection System (TPS) that may be ablated during severe plasma disruptions, thus protecting the underlying surface. W-pillars or W-coated Re-pillars are found to allow trapped helium to diffuse out of the surface instead of building up sub-surface stress.

Thermal performance of multilayer insulation: A review

A multilayer insulation (MLI) is a passive thermal protection system used in cryogenics and space exploration programs as a thermal insulator. It is very thin and light weight. Due to less weight and higher thermal performance of MLI, it found application in space programs to store cryogenic liquid propellant. The effective thermal conductivity of MLI is in the order of 10−5 W/mK. The prediction of heat transfer in MLI is very complex due to the anisotropic conductivity and combination of radiation, gas conduction and solid conduction modes of heat transfer. This work is an attempt to gather some significant research outcome in MLI performance prediction and the factors to be considered while modeling the MLI.

Development and characterization of a contoured passive thermal protection system

AbstractPresent work deals with process refinement, fabrication, and testing of a novel contoured passive thermal protection system (TPS), based on aluminosilicate fiber reinforced silica matrix composite. Taguchi method‐based statistical approach was employed to understand the effect of process parameters such as thickness, layer density, stacking sequence and firing temperature on dielectric properties, flexural strength, and bulk density of composites. Analysis of variance was performed to ensure the statistical significance. The developed TPS was characterized for thermal insulation properties by testing under infrared heating conditions. The science behind the insulation characteristics of the TPS was analyzed using FTIR, XRD, and SEM to understand the associated high‐temperature phase transitions. Present study has shown that, the back‐face temperature exponentially decays with TPS thickness. For TPS having 9 mm thickness, the back‐face temperature was found to be below 100°C when the front face was exposed to 1000°C. The insulation index was found to be increasing with increase in temperature and TPS thickness. Improved thermal insulation characteristics of TPS are due to energy consuming processes such as crystalline changes and micro cracks formation.

Storey-based stability of unbraced structural steel frames subjected to variable fire loading

Structural steel has poor fire resistance properties and often requires thermal protection. Passive thermal protection systems such as insulation for steel members by application of spray-applied fire-resistance materials (SFRM) on surfaces of structural steel members are expensive and represent a significant portion of building costs for steel structures. Also, design codes around the world are progressing toward performance-based approaches rather than prescriptive-based approaches in design for fire safety. However, it is difficult to account realistically for all probable fire scenarios, and the actual fire resistance of a structure may vary significantly depending on the nature of the fire, location of origin and characteristics of the building. A means to define and identify the maximum and minimum fire loading scenarios causing instability of a given structure is therefore desirable. Presented in this paper is a novel global optimization approach for determining the highest temperature, lowest temperature, most localized, and most distributed fire scenarios causing instability for an unbraced structural steel frame. The investigation assumes that the columns are fire protected but the beams are unprotected, and may be extended to apply in other configurations and framing materials.

Design of Thermal Protection System for Reusable Hypersonic Vehicle Using Inverse Approach

In this study, an inverse heat conduction analysis based on the Levenberg-Marquardt method is applied to reconstruct the aerothermal heating and thermal protection material response of a reusable hypersonic vehicle (RHV) from the knowledge of temperature measurements taken within the thermal protection system (TPS). In addition to reconstruction of TPS material response, a lightweight passive TPS for the RHV based on these predictions is designed. In this analysis, the experimental temperature data are generated by a direct approach; this is due to the lack of real experimental setup that is used to measure the temperature. The direct analysis has performed by an explicit one-dimensional finite difference code to analyze the in-depth temperature response of the TPS. The temperature data determined from the direct approach are used to simulate the temperature measurements. The influence of measurement errors upon the accuracy of the inverse results is also investigated. Finally, results show that the Levenberg-Marquardt method is an effective technique to design a lightweight passive TPS for an RHV by using the inverse approach.

Selection of Materials and Design of Multilayer Lightweight Passive Thermal Protection System

A methodology has been established aiming to design a lightweight thermal protection system (TPS), using advanced lightweight ablative materials developed at the NASA Ames Research Center. An explicit finite-difference scheme is presented for the analysis of one-dimensional transient heat transfer in a multilayer TPS. This problem is solved in two steps, in the first step, best candidate materials are selected for TPS. The selection of materials is based mainly on their thermal properties. In the second step, the geometrical dimensions are determined by using an explicit finite-difference scheme for different combinations of the selected materials, and these dimensions are optimized for the design of lightweight TPS. The best combination of material employs silicone impregnated reusable ceramic ablator (SIRCA), Saffil, and glass-wool for the first, second, and third layer, respectively.

Design and heat transfer characteristics analysis of combined active and passive thermal protection system for hydrogen fueled scramjet

Many methods have been widely carried out to develop the thermal protection system of hydrogen fueled scramjet to meet the cooling requirements of engine. A combined thermal protection system is designed and developed by combing active and passive thermal protection systems and proposed to endure the long duration safe operation of a scramjet at high flight Mach number. The combined thermal protection system proposed is designed in the light of the dangerous conditions for an active thermal protection system, and its performance is evaluated with a one-dimensional coupled heat transfer model of a hydrogen fueled scramjet at different Mach numbers and equivalence ratios. Test results indicate that the thermal protection system proposed could be used to expand the flight range of a scramjet by properly selecting conductivity and thickness of passive layer.

Transient aero-thermal mapping of passive Thermal Protection system for nose-cap of Reusable Hypersonic Vehicle

The temperature field history of passive Thermal Protection System (TPS) material at the nose-cap (forward stagnation region) of a Reusable Hypersonic Vehicle (RHV) is generated. The 3-D unsteady heat transfer model couples conduction in the solid with external convection and radiation that are modeled as time-varying boundary conditions on the surface. Results are presented for the following two cases: (1) nose-cap comprised of ablative TPS material only (SIRCA/PICA), and (2) nose-cap comprised of a combination of ablative TPS material with moderate thermal conductivity and insulative TPS material. Comparison of the temperature fields of SIRCA and PICA [Case (1)] indicates lowering of the peak stagnation region temperatures for PICA, due to its higher thermal conductivity. Also, the use of PICA and insulative TPS [Case (2)] for the nose-cap has higher potential for weight reduction than the use of ablative TPS alone.

Projectile Nosetip Thermal Management for Railgun Launch to Space

A major issue in achieving direct launch to space using electromagnetic railguns from the surface of the earth is that we are at the bottom of a gravity well, requiring very high launch velocities to achieve orbit insertion. The atmosphere is most dense at the earth's surface-this, combined with high velocity, leads to very stressing thermal conditions for the launch aerobody. To minimize aerodynamic losses, a typical hypersonic flight body is a blunt slender cone. The thermal loads are greatest on the nosetip of this cone, and for small bodies (<10 kg) the thermal loads may exceed the capabilities of present passive thermal protection systems, such as dense, ablative carbon phenolics. Even for launch from a high-altitude airborne platform, aerothermal loads can be considerable unless rocket assist is provided. This paper discusses aspects of the launch package and characteristics of the railgun needed to achieve orbit insertion velocities

A reusable space-rescue vehicle: re-entry simulation

This paper presents the conceptual design of a re–entry vehicle suitable for the emergency return of personnel from the International Space Station. For injured occupants, or those with emergency medical conditions, a re–entry environment with a maximum g loading of less than 1.1 would be necessary. This has been met by using a relatively conventional design employing leading–edge bluntness and a passive thermal–protection system. The paper addresses the guidance and control required to meet these conditions and thermal modelling is used to determine the cabin temperature rise during re–entry. In comparison with vehicles employing greater aerodynamic sophistication, such as wave–riders, the cross–range is inferior, but in other respects it is shown that the requirements of the re–entry environment can be met.

Heat shield design for re-entry and launch. The use of conduction-assisted radiation on sharp-edged wings

Restricted accessMoreSectionsView PDF ToolsAdd to favoritesDownload CitationsTrack Citations ShareShare onFacebookTwitterLinked InRedditEmail Cite this article Nonweiler Terence R. F. 1999Heat shield design for re-entry and launch. The use of conduction-assisted radiation on sharp-edged wingsPhil. Trans. R. Soc. A.3572197–2225http://doi.org/10.1098/rsta.1999.0428SectionRestricted accessHeat shield design for re-entry and launch. The use of conduction-assisted radiation on sharp-edged wings Terence R. F. Nonweiler Terence R. F. Nonweiler Victoria University, Wellington, New Zealand and APECS Ltd, 3 Hawkley Hurst, Hawkley, Hants GU33 6NS, UK Google Scholar Find this author on PubMed Search for more papers by this author Terence R. F. Nonweiler Terence R. F. Nonweiler Victoria University, Wellington, New Zealand and APECS Ltd, 3 Hawkley Hurst, Hawkley, Hants GU33 6NS, UK Google Scholar Find this author on PubMed Search for more papers by this author Published:01 August 1999https://doi.org/10.1098/rsta.1999.0428AbstractThe conduction of heat within the material of a wing leading edge can be treated both simply and (usually) accurately by ‘conducting–plate theory’. The theory leads to a concise ‘reference solution’ for the variation of temperature over a solid wedge–shaped nose that is heated by a boundary layer. This indicates the best choice of material, and we look at ways a given mass of the material may be best be installed to reduce the nose temperature, without sacrificing the benefits that a sharp–nosed wing may have to offer. It is shown that although these materials may be dense, a heavy mass is not required to achieve acceptable temperatures at the edge in hypersonic flight. Something perhaps between 5 and 10 kg per metre of edge (or say 3 − 7 lb ft−1 is usually enough. However, the nose temperature depends on sweepback and surface pressure. To avoid temperatures of 1600 K or more at a highly swept sharp edge in hypersonic flight, it is necessary that neither the wing loading nor the surface pressure exceed more than 2 kPa (40 lbf ft−2). We cite values of less than this that relate to the design of a reentry vehicle with a wing loading of only 680 Pa (14 and a ¼ lbf ft7minus;2). However, rounding the nose (with a radius usually of just a few millimetres) can provide reductions of up to perhaps 20 % in the nose temperature. This allows this form of temperature control to be extended to wings of higher loading and to regions of lower sweep, including at or near the wing apex where the heating rates are most intense. Previous ArticleNext Article VIEW FULL TEXT DOWNLOAD PDF FiguresRelatedReferencesDetailsCited by Larrimbe L, Pettinà M, Nikbin K, Jones E, Katz A, Hawkins C, DeCerbo J, Brown P, Vandeperre L and White M (2016) High Heat Flux Laser Testing of HfB 2 Cylinders , Journal of the American Ceramic Society, 10.1111/jace.14474, 100:1, (293-303), Online publication date: 1-Jan-2017. Kumar S and Mahulikar S (2015) Selection of Materials and Design of Multilayer Lightweight Passive Thermal Protection System, Journal of Thermal Science and Engineering Applications, 10.1115/1.4031737, 8:2, Online publication date: 1-Jun-2016. East R (2010) Re-Entry Vehicle Flight Profiles Encyclopedia of Aerospace Engineering, 10.1002/9780470686652.eae318 Squire T and Marschall J (2010) Material property requirements for analysis and design of UHTC components in hypersonic applications, Journal of the European Ceramic Society, 10.1016/j.jeurceramsoc.2010.01.026, 30:11, (2239-2251), Online publication date: 1-Aug-2010. Savino R, De Stefano Fumo M, Paterna D and Serpico M (2005) Aerothermodynamic study of UHTC-based thermal protection systems, Aerospace Science and Technology, 10.1016/j.ast.2004.12.003, 9:2, (151-160), Online publication date: 1-Mar-2005. Townend L (2001) Domain of the Scramjet, Journal of Propulsion and Power, 10.2514/2.5865, 17:6, (1205-1213), Online publication date: 1-Nov-2001. This Issue01 August 1999Volume 357Issue 1759Theme Issue ‘Hypersonic aircraft: lifting re–entry and launch’ compiled by L. H. Townend Article InformationDOI:https://doi.org/10.1098/rsta.1999.0428Published by:Royal SocietyPrint ISSN:1364-503XOnline ISSN:1471-2962History: Published online01/08/1999Published in print01/08/1999 License: Citations and impact Keywordsleading edge designeffect of roundingconducting–plate theoryconduction–assisted coolingre–entry heatingsharp–edge temperature