Wrapped multilayer insulation design and testing

New vehicles need improved cryogenic propellant storage and transfer capabilities for long duration missions. Multilayer insulation (MLI) for cryogenic propellant feedlines is much less effective than MLI tank insulation, with heat leak into spiral wrapped MLI on pipes 3–10 times higher than conventional tank MLI. Better insulation for cryogenic feed lines is an important enabling technology that could help NASA reach cryogenic propellant storage and transfer requirements. Improved insulation for Ground Support Equipment could reduce cryogen losses during launch vehicle loading. Wrapped-MLI (WMLI) is a high performance multilayer insulation using innovative discrete spacer technology specifically designed for cryogenic transfer lines and Vacuum Jacketed Pipe (VJP) to reduce heat flux.The poor performance of traditional MLI wrapped on feed lines is due in part to compression of the MLI layers, with increased interlayer contact and heat conduction. WMLI uses discrete spacers that maintain precise layer spacing, with a unique design to reduce heat leak. A Triple Orthogonal Disk spacer was engineered to minimize contact area/length ratio and reduce solid heat conduction for use in concentric MLI configurations.A new insulation, WMLI, was developed and tested. Novel polymer spacers were designed, analyzed and fabricated; different installation techniques were examined; and rapid prototype nested shell components to speed installation on real world piping were designed and tested. Prototypes were installed on tubing set test fixtures and heat flux measured via calorimetry. WMLI offered superior performance to traditional MLI installed on cryogenic pipe, with 2.2 W/m² heat flux compared to 26.6 W/m² for traditional spiral wrapped MLI (5 layers, 77–295 K). WMLI as inner insulation in VJP can offer heat leaks as low as 0.09 W/m, compared to industry standard products with 0.31 W/m. WMLI could enable improved spacecraft cryogenic feedlines and industrial hot/cold transfer lines.     

Theory to boil-off gas cooled shields for cryogenic storage vessels

An intermediate refrigeration with boil-off gas cooled shields using the boil-off gas stream is an alternative method to the conventional intermediate refrigeration with a cryogenic liquid.By using an analytical calculation method relations are derived, which enable complete predictions about the effectiveness of an intermediate refrigeration with boil-off gas cooled shields as a function of the number of shields for the different stored cryogenic liquids. For this theoretical derivation however, the restrictive assumption must be made that the thermal conductivity of the used insulation material has a constant value between the considered temperature boundaries.For purposes of a more exact calculation a numerical method is therefore suggested, which takes into consideration that the thermal conductivity is temperature-dependent. For a liquid hydrogen storage vessel with a perlite-vacuum insulation e.g., the effectiveness of one shield and its equilibrium temperature are given as a function of the position of the shield in the insulation space.     

A coupled numerical analysis of shield temperatures, heat losses and residual gas pressures in an evacuated super-insulation using thermal and fluid networks: Part II: Unsteady-state conditions (cool-down period)

This paper analyses the cool-down period of a 300 L super-insulated cryogenic storage tank for liquid nitrogen. Storage tank and evacuated shields are the same as described in part I of this paper where stationary states were investigated. The aim of the present paper is to introduce thermal resistance networks as a tool to quantitatively understand and control also unsteady-states like cool-down of super-insulations. Numerical simulations using thermal resistance networks have been performed to determine time dependence of local shield temperatures and heat loss components. Coupling between radiation and solid conduction is investigated under these conditions. Using the numerical results, we have checked an experimental method suggested in the literature to separate heat losses through the insulation from losses through thermal bridges by measurement of unsteady-state evaporation rates. The results of the simulations confirm that it takes the outer shields much longer to reach stationary temperature; cool-down does not proceed uniformly in the super-insulation. Coupling between different heat transfer modes again is obvious. Thermal emissivity is important also during the early phase of cool-down. Using the obtained numerical results, the experimental method to separate heat loss components could only roughly been confirmed for thick metallic foils.     

Novel load responsive multilayer insulation with high in-atmosphere and on-orbit thermal performance

Aerospace cryogenic systems require lightweight, high performance thermal insulation to preserve cryopropellants both pre-launch and on-orbit. Current technologies have difficulty meeting all requirements, and advances in insulation would benefit cryogenic upper stage launch vehicles, LH₂ fueled aircraft and ground vehicles, and provide capabilities for sub-cooled cryogens for space-borne instruments and orbital fuel depots. This paper reports the further development of load responsive multilayer insulation (LRMLI) that has a lightweight integrated vacuum shell and provides high thermal performance both in-air and on-orbit.LRMLI is being developed by Quest Product Development and Ball Aerospace under NASA contract, with prototypes designed, built, installed and successfully tested. A 3-layer LRMLI blanket (0.63 cm thick, 77 K cold, 295 K hot) had a measured heat leak of 6.6 W/m² in vacuum and 40.6 W/m² in air at one atmosphere. In-air LRMLI has an 18× advantage over Spray On Foam Insulation (SOFI) in heat leak per thickness and a 16× advantage over aerogel. On-orbit LRMLI has a 78× lower heat leak than SOFI per thickness and 6× lower heat leak than aerogel.The Phase II development of LRMLI is reported with a modular, flexible, thin vacuum shell and improved on-orbit performance. Structural and thermal analysis and testing results are presented. LRMLI mass and thermal performance is compared to SOFI, aerogel and MLI over SOFI.     

低温推进剂贮箱大面积冷屏热分析及成本优化

运载火箭低温贮箱采用大面积冷屏与多层隔热材料组成的复合结构可以有效减少低温推进剂蒸发损耗,延长低温推进剂在轨贮存时间。通过建立多层隔热材料耦合90 K大面积冷屏的传热模型,获得了引入大面积冷屏后对多层隔热材料层间温度分布及热流密度影响的变化规律,对比了采用冷屏技术和直接对液氢采用主动制冷两种方式,同等条件下采用冷屏在主动制冷系统重量和功耗方面可分别节省60%和64%。研究了低温推进剂不同在轨贮存时间和冷屏安装在多层隔热材料中不同位置时热管理系统重量和功耗成本,以成本最小为目标获得了90 K冷屏布局最优化设计方法。     

有效导热系数对低温容器日蒸发率的影响

低温容器是气体液化分离加工工业中的重要设备，绝热设计是低温容器设计的重要组成部分，它直接影响低温容器的日蒸发率。在不同的绝热材料和一定漏热温差的条件下，给出了普通绝热和真空粉末绝热型低温容器的日蒸发率与热材料的有效导热系数以及绝热层厚度之间的关系曲线。为简化设计和比较低温容器的特性提供方法和依据。     

Cost effective honeycomb and multi-layer insulation debris shields for unmanned spacecraft

Ways to improve the tolerance of unmanned spacecraft to hypervelocity impact are presented. Two new honeycomb and multi-layer insulation (MLI) shields were defined: (1) double honeycomb, and (2) enhanced or toughened MLI (with additional Kevlar 310 and/or Betacloth layers). Following hypervelocity impact testing, a new ballistic limit threshold was defined, based on rear facesheet perforation and witness plate damage characteristics. At 12 km/s, the ballistic limit of single honeycomb was 0.58 mm (aluminium sphere), rising to 0.91 mm for double honeycomb, 1.00 mm for double honeycomb with MLI and 1.17 mm for double honeycomb with toughened MLI. A damage equation, based on the modified Cour-Palais equation with ESA constants, was compared with the data and found to be conservative. The impact angle exponent was increased in order to reduce the equation under-prediction for the oblique incidence data. An equivalent rear wall thickness was defined in order to distinguish between shield types above 7 km/s. The spacecraft survivability analysis showed that the double honeycomb and toughened MLI significantly reduced the number of perforating particles over the baseline single honeycomb design. The mass increase of these shields is approximately 1.2 kg/m² for double honeycomb and 0.8 kg/m² for toughened MLI.     

A coupled numerical analysis of shield temperatures, heat losses and residual gas pressures in an evacuated super-insulation using thermal and fluid networks: Part I: Stationary conditions

This paper describes numerical simulations, using thermal networks, of shield temperatures and radiative and conductive heat losses of a super-insulated cryogenic storage tank operating at 77 K. Interactions between radiation and conductive heat transfer modes in the shields are investigated, by calculation of local shield temperatures. As a new method, fluid networks are introduced for calculation of stationary residual gas pressure distribution in the evacuated multilayer super-insulation. Output from the fluid network is coupled to the iterative thermal network calculations. Parameter tests concern thickness and emissivity of shields, degree of perforation, residual gas sources like desorption from radiation shields, spacers and container walls, and permeation from the inner container to the evacuated insulation space. Variations of either a conductive (thickness of Al-film on Mylar) or a radiative parameter (thermal emissivity) exert crosswise influences on the radiative or conductive heat losses of the tank, respectively.     

Multiple vapour cooled shield position tolerances

Vapour cooled shields are used to intercept part of a cryostat's heat leak. This paper is concerned with the positional tolerances for multiple shields and two and three shield cases are analysed in detail.     

Impact damage on sandwich panels and multi-layer insulation

Most spacecraft rely intensively on sandwich construction for external structures with multi layer thermal insulation where appropriate. Experience gained in ESA with various spacecraft (ROSETTA, METOP, ATV,…) covers a substantial range of materials and configurations. In this work, the applicability of simple damage equations (e.g. those presently used for single or Whipple shield ballistic limits) to more complex configurations (e.g. sandwich plates with and without MLI) is analyzed. The different sandwich configurations which were submitted to testing are reviewed, impact test results are presented and compared with impact reference data on single plates and Whipple shields. It has been found that sandwich panels have a better tolerance to hypervelocity impacts than monolithic structures. MLI placed in front of the sandwich panels contributes significantly to the overall protection performance in the range of the projectile diameters tested. The complexity of the sandwich structure is responsible for a considerable scatter in the test results. The predictors for Whipple shields applied to sandwich panels with and without MLI can only be considered on a case by case basis for risk assessment analysis.     

低温真空多层绝热结构热阻的理论分析

低温下真空多层绝热是非常有效的绝热方式，广泛应用于众多领域的科研和工程实践中。低温下真空多层绝热的影响因素很多，具体包括反射屏之间隔层的材料特性、反射屏的层数、层密度、真空度、捆扎的松紧程度等。运用热阻网络分析了间隔物为纤维材料的低温真空多层绝热结构的热阻组成，对总热阻中固体导热热阻Rconduction(solid)、辐射换热热阻Rradiation、残余气体导热热阻Rconduction(gas）分别建立了理论模型，并进行了理论计算推导，得到了低温下真空多层绝热结构总热阻的计算公式。     

Launch ascent testing of a representative Altair ascent stage methane tank

In order to support long duration cryogenic propellant storage, the NASA is investigating the long duration storage properties of liquid methane. The Methane Lunar Surface Thermal Control (MLSTC) testing is using a tank of the approximate dimensions of the Altair lunar ascent propellant tanks. The tank was insulated with multilayer insulation and placed inside of a vacuum chamber to simulate the various environments that would be encountered during launch and travel from the earth to the lunar surface, including long duration stays on the lunar surface. One of these environments to be studied is the launch and ascent environment; while all the effects of this mission phase cannot be simulated at the same time, an effort was made to simulate as many as possible. Boil-off testing included ambient pressure ground hold testing followed by a rapid depressurization of the vacuum chamber during which the liquid methane tank was allowed to come to steady state condition in the high vacuum environment. The data gathered from the series of tests fit with-in pre-test predictions and yielded much needed test data for rapid depressurization using liquid methane.     

Cryogenic propellant liquefaction and storage for a precursor to a human Mars mission

An analysis of cryogenic liquefaction and storage methods for in-situ produced propellants (oxygen and methane) on Mars is presented. The application is to a subscale precursor sample return mission, intended to demonstrate critical cryogenic technologies prior to a human mission. A heat transfer analysis is included, resulting in predicted cryogenic tank surface temperatures and heat leak values for different conditions. Insulation thickness is traded off against cryocooler capacity to find optimum combinations for various insulation configurations, including multilayer insulation and microspheres. Microsphere insulation is shown to have promise, and further development is recommended.     

Cryogenic propulsion for the Titan Orbiter Polar Surveyor (TOPS) mission

Liquid hydrogen (LH2) and liquid oxygen (LO2) cryogenic propellants can dramatically enhance NASA’s ability to explore the solar system due to their superior specific impulse (I_sp) capability. Although these cryogenic propellants can be challenging to manage and store, they allow significant mass advantages over traditional hypergolic propulsion systems and are therefore enabling for many planetary science missions. New cryogenic storage techniques such as subcooling and the use of advanced insulation and low thermal conductivity support structures will allow for the long term storage and use of cryogenic propellants for solar system exploration and hence allow NASA to deliver more payloads to targets of interest, launch on smaller and less expensive launch vehicles, or both. These new cryogenic storage technologies were implemented in a design study for the Titan Orbiter Polar Surveyor (TOPS) mission, with LH2 and LO2 as propellants, and the resulting spacecraft design was able to achieve a 43% launch mass reduction over a TOPS mission, that utilized a traditional hypergolic propulsion system with mono-methyl hydrazine (MMH) and nitrogen tetroxide (NTO) propellants. This paper describes the cryogenic propellant storage design for the TOPS mission and demonstrates how these cryogenic propellants are stored passively for a decade-long Titan mission that requires the cryogenics propellants to be stored for 8.5 years.     

空间液氦制冷器隔热屏蔽系统设计及优化

以空间液氦制冷器为研究对象,介绍了空间液氦制冷器隔热屏蔽系统的设计思路以及方法,设计了一套以气冷屏为主、包括真空多层绝热和隔热支撑结构在内的隔热屏蔽系统,并编制数值计算程序对其进行了优化。计算结果显示了气冷屏尺寸是影响制冷器内部温度场以及内胆热负荷的关键因素,提出了气冷屏尺寸的优化结果。     

高真空多层绝热低温液体运输半挂车的开发研制

简述高真空多层绝热低温液体运输半挂车的基本结构、绝热特性 ,并阐述了该车在低温液体运输的应用前景。     

Development of No-Vent™ liquid hydrogen storage system for space applications

A number of technological advances required to store and maintain normal-boiling-point and densified cryogenic liquids, including liquid hydrogen, under zero boil-off conditions in-space, for long periods of time, have been developed. These technologies include (1) thermally optimized compact cryogen storage systems that reduce environmental heat leak to the lowest-temperature cryogen, which minimizes cryocooler size and input power, and (2) actively-cooled shields that surround the storage systems and intercept heat leak. The processes and tools used to develop these technologies are discussed. A zero boil-off liquid hydrogen storage system technology demonstrator for validating the actively-cooled shield technology is presented.     

Passive ZBO storage of liquid hydrogen and liquid oxygen applied to space science mission concepts

Liquid hydrogen and oxygen cryogenic propulsion and storage were recently considered for application to Titan Explorer and Comet Nuclear Sample Return space science mission investigations. These missions would require up to 11 years of cryogenic storage. We modeled and designed cryogenic propellant storage concepts for these missions. By isolating the propellant tank’s view to deep space, we were able to achieve zero boil-off for both liquid hydrogen and oxygen propellant storage without cryocoolers. Several shades were incorporated to protect the tanks from the sun and spacecraft bus, and to protect the hydrogen tank from the warmer oxygen tank. This had a dramatic effect on the surface temperatures of the propellant tank insulation. These passive storage concepts for deep space missions substantially improved this application of cryogenic propulsion. It is projected that for missions requiring larger propellant tank sizes, the results would be even more dramatic.     

Cryogenic leak test procedure for a large superconducting cyclotron helium vessel

A procedure for testing the vacuum integrity of very large stainless steel weldments used at cryogenic temperatures has been developed at Michigan State University. This development, which uses large quantities of liquid nitrogen, is a modification of a technique commonly applied to small devices and involves cooling the cryostat's liquid helium vessel (bobbin) to liquid nitrogen temperature, and then proceeding immediately with leak testing. This method was applied to the K800 superconducting magnet helium vessel, which seemed leak tight at room temperature, but was found to have an easily detectable helium leak when cooled. After repairing the leak, retesting revealed no leaks, where upon the K800 cryostat construction was completed; i.e. the bobbin was wrapped with superinsulation, a liquid nitrogen radiation shield was added, and the assembly was inserted into the vacuum jacket. The final leak test occurred when the cryostat was cooled to liquid helium temperature and was found to be helium leak tight.