Abstract:
The disclosed subject matter relates to a radiation shielding apparatus including a cryogenic vessel and a cryogenic hydrogen radiation shielding material capable of providing a radiation shield, the cryogenic hydrogen radiation shielding material including cryogenic hydrogen.

Description:
FIELD OF THE INVENTION 
       [0001]    This invention relates to protecting a manned spacecraft from radiation in space. 
       BACKGROUND 
       [0002]    Human susceptibility to the harsh space radiation environment has been identified as being a major hurdle for exploration beyond low Earth orbit (LEO). High energy protons and nuclei ions from Solar Energetic Particles (SEPs) and Galactic Cosmic Rays (GCRs) can result in radiation doses that are dangerous to astronaut health and even survivability if the astronauts are not adequately shielded. These high energy particles also cause significant amounts of secondary radiation when they impinge on spacecraft structure. The secondary neutron radiation may cause human radiogenic cancers. Hydrogen or hydrogen rich materials are ideal materials for radiation shielding because hydrogen does not easily break down and become a source for secondary radiation. 
         [0003]    When a spacecraft is positioned in LEO, the Earth&#39;s magnetic field provides some radiation protection to the spacecraft and the astronauts occupying it. Radiation protection for astronauts is critical for the future of human space flight since conventional spacecraft construction materials such as aluminum are susceptible to secondary radiation when SEPs or GCRs impinge on them. Because of the size of an aluminum nucleus, the secondary radiation produced while shielding space radiation can be just as damaging as the primary radiation and this secondary radiation contributes to the total ionizing dose received by the astronauts. Other types of hydrogen-rich materials, such as polyethylene, have been tested to determine their effectiveness at reducing the dose received from all sources of radiation. Such shielding materials do not produce the same level of damaging secondary radiation, however, the presence of carbon atoms in polyethylene means that there is less hydrogen shielding material per unit of shielding material mass than there would be if hydrogen itself is used as the shielding material. However, hydrogen is a challenging substance to store and manage and, therefore, has not been considered as a viable shielding material for spacecraft. 
         [0004]    Developing a system using cryogenic material, hydrogen, that is maintained at, for example, 10-12 K (“K” here and throughout refers to “° K” or “degrees Kelvin”), for radiation shielding presents several challenges. Thermal challenges include, for example, heat leak from the space environment into cryogenic hydrogen shielding due to, for example solar irradiation, planetary albedo, heat leak from the crew capsule that is maintained at room temperature of about 300 K, power system, propulsion, etc. into the cryogenic hydrogen shield. It is also challenging to process the cryogenic hydrogen on the ground, prior to launch, and bring it to a frozen temperature of 10 K while the hydrogen is contained in a tank that is in an ambient approximately 300 K environment. 
       BRIEF DESCRIPTION 
       [0005]    In one embodiment, a radiation shielding apparatus is provided. The radiation shielding apparatus includes a cryogenic vessel and a cryogenic hydrogen radiation shielding material capable of providing a radiation shield, the cryogenic hydrogen radiation shielding material includes hydrogen at a temperature of less than or equal to about 20 K, wherein the cryogenic hydrogen radiation shielding material is contained in the cryogenic vessel. 
         [0006]    In another embodiment, a spacecraft is provided. The spacecraft includes a radiation shielding apparatus and a crew module. The radiation shielding apparatus includes a cryogenic vessel and a cryogenic hydrogen radiation shielding material capable of providing a radiation shield, the cryogenic hydrogen radiation shielding material includes hydrogen at a temperature of less than or equal to about 20 K, wherein the cryogenic hydrogen radiation shielding material is contained in the cryogenic vessel. The crew module includes a walled enclosure with an exterior surface and a hatch to permit access and egress to an internal area within the walled enclosure, wherein the radiation shielding apparatus is disposed adjacent to the exterior surface of the crew module. 
         [0007]    In another embodiment, a spacecraft is provided. The spacecraft includes a fuselage, a radiation shielding apparatus, a crew module and a radiator system. The fuselage defines an internal volume within the spacecraft. The radiation shielding apparatus is disposed in the internal volume of the fuselage and includes a cryogenic vessel, insulation material and a cryogenic hydrogen radiation shielding material capable of providing a radiation shield, the cryogenic hydrogen radiation shielding material including solid hydrogen, subcooled solid hydrogen or a mixture thereof, wherein the cryogenic hydrogen radiation shielding material is contained in the cryogenic vessel. The crew module is disposed in the internal volume of the fuselage and includes a walled enclosure with an exterior surface and a hatch to permit access and egress to an internal area within the walled enclosure, the internal area of the crew module being substantially maintained at about room temperature. The radiator system is to remove heat emitting from the crew module. The radiation shielding apparatus is disposed between the fuselage and the exterior surface of the crew module. The radiator system is disposed between the exterior surface of the crew module and radiation shielding apparatus. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0008]    These and other features, aspects, and advantages of the present invention will become better understood when the following detailed description is read with reference to the accompanying drawings in which like characters represent like parts throughout the drawings, wherein: 
           [0009]      FIG. 1A  is a top perspective view of a shielded capsule; 
           [0010]      FIG. 1B  is a cross-sectional view of the shielded capsule of  FIG. 1A ; 
           [0011]      FIG. 2  is a cross-sectional view of a shielded capsule; 
           [0012]      FIG. 3  is a schematic representation of a ground cooling system and section view of a spacecraft, radiation shield and space thermal system; 
           [0013]      FIG. 4  is a sectional view of a radiation shield; 
           [0014]      FIG. 5  illustrates a comparison of radiation shielding provided by different materials; and 
           [0015]      FIG. 6  illustrates a comparison of radiation shielding provided by different materials. 
       
    
    
     DETAILED DESCRIPTION 
       [0016]    Studies have shown there will be a need to protect astronauts during, for example, interplanetary missions (e.g., Mars) from deep space radiation with an annual allowable radiation dose less than 500 mSv. For a typical crew module that is 4 meter in diameter and 8 meter in length, the mass of polyethylene radiation shielding required would be more than 17,500 kg at a needed shielding a real density of approximately 140 kg/m 2 . By comparison, the requirement for hydrogen shielding is 70 kg/m 2 , much less than polyethylene shielding. Vapor hydrogen has a very low density, and the storage tank can&#39;t fit into a 5 meter payload fairing for a rocket that might launch the crew module. Liquid and solid hydrogen have much higher densities and are preferable to vapor hydrogen for the purpose of packaging the required hydrogen areal density in a reasonable volume. For example, the thickness of solid hydrogen needed to shield astronauts is about 0.43 m and the combined diameter of the crew module with shielding is about 4.86 m. However, a challenge with using either liquid or solid hydrogen as shielding material is that the hydrogen has to be stored at cryogenic temperatures. The Cryogenic Hydrogen Radiation Shielding (CHRS) requires a thermal system to prevent heat leak into cryogenic tank from the crew module (substantially maintained at room temperature, for example, about 300 K) to avoid phase change of the cryogenic hydrogen. However, even after accounting for the mass of the thermal and containment system for CHRS, CHRS may halve the mass of a radiation shield when compared to polyethylene shields. The crew module is intended to be suitably maintained in temperature and atmosphere to adequately support life and provide an environment in which astronauts could live. 
         [0017]    CHRS material includes liquid hydrogen, subcooled liquid hydrogen, solid hydrogen and subcooled solid hydrogen or a mixture thereof, preferably solid hydrogen, subcooled solid hydrogen or a mixture thereof and more preferably subcooled solid hydrogen. Liquid hydrogen at a pressure of  1  atm can be stored at a maximum temperature of about 20 K. Subcooled liquid hydrogen can be stored at a temperature from about 14 K to about 20 K. Solid hydrogen can be stored at a maximum temperature of about 14 K (the triple point of hydrogen). Subcooled solid hydrogen can be stored at a temperature of less than about 14 K, preferably from about 10 K to about 12 K. Subcooled solid hydrogen may have an advantage in that it can absorb more heat without changing phase. CHRS material has a lower mass density compared to other radiation shielding materials, such as aluminum and polyethylene. The degree of radiation shielding provided by a CHRS material depends on the mass of hydrogen per unit surface area. 
         [0018]    One embodiment includes a shielded capsule  100  including cryogenic hydrogen radiation shielding as shown in  FIG. 1A  and  FIG. 1B .  FIG. 1A  and  FIG. 1B  include a crew module  102  defined by a circumferential side wall  104  and end walls  106  and  108  and having an inner volume  110  which the crew may inhabit. The crew module  102  is protected by an annular cryogenic vessel  112  adjacent the circumferential side wall  104  and by toroidal cryogenic vessels  114  and  116  adjacent end walls  106  and  108 , respectively. When the capsule is in operation, the annular cryogenic vessel  112  and toroidal cryogenic vessels  114  and  116  contain CHRS material to provide radiation shielding to the crew module  102 . A bore  118  in the middle of toroidal cryogenic tank  114  leads to hatch  120  to allow for crew access and egress from inner volume  110  of the crew module  102 . Bore  118  and/or the hatch  120  can be closed with a suitable radiation shielding hatch cover in order to minimize radiation from reaching the inner area of the crew module through bore  118  and/or hatch  120 . The suitable radiation shielding hatch cover may be movable and constructed of a suitable radiation shielding material, such as for example, polyethylene. 
         [0019]    Another embodiment includes a shielded capsule  200  including cryogenic hydrogen radiation shielding as shown in  FIG. 2 .  FIG. 2  includes a crew module  202  having an inner volume  204  which the crew may inhabit, circumferential side wall  206  and end walls  208  and  210 . The crew module  202  is surrounded by a cryogenic vessel  212 . When the capsule is in operation, the cryogenic vessel  212  contains CHRS material, such as solid hydrogen, to provide radiation shielding to the crew module  202 . The cryogenic vessel  212  includes a circumferential vessel portion  214  adjacent circumferential side wall  206  and vessel end portions  216  and  218  adjacent end walls  208  and  210 , respectively. A bore  220  in the middle of vessel end portions  216  leads to hatch  222  which allows for crew access and egress from the crew module  202  of the shielded capsule  200 . Bore  220  and/or the hatch  222  can be sealed with a suitable radiation shielding hatch cover  224  that may be movable and constructed of a suitable radiation shielding material, such as for example, polyethylene. A passive thermal management system, such as a 100 K thermal shield, includes end sections  228  and  230  adjacent end walls  208  and  210 , respectively, and a side wall section  232  adjacent the circumferential side wall  206  and is positioned between the crew module  202  (circumferential side wall  206  and end walls  208  and  210 ) and the cryogenic vessel  212  (circumferential vessel portion  214  and vessel end portions  216  and  218 ). Conduits  225 ,  226  and  227  may provide a thermal link between a radiator  233  and the exemplified 100 K thermal shield insulation material. The radiator system (thermal management system end sections  228  and  230  and side wall section  232 , conduits  225 ,  226  and  227  and radiator  233 ) rejects heat into deep space, the latter existing at a temperature of about 7 K. The radiator system removes heat emitting from the inner volume  204  of the crew module  202  in order to insulate and minimize heat transfer to the cryogenic vessel  212  from the inner volume  204  being maintained at about room temperature (about 300 K). Such heat transfer from inner volume  204  can affect and be problematic to the maintenance of the low temperature of the CHRS in the cryogenic vessel  212 . 
         [0020]    The CHRS system components including a cryogenic tank or vessel and insulation material and their design and materials should be selected based on mechanical and fluid engineering criteria including thermal performance (e.g., insulation) and structural performance (e.g., ability to maintain integrity &amp; internal pressure) experienced in the various rigors of space as well as in a gravitational environment, such as, on a planet (e.g., Earth). The tank or vessel may be suitably constructed of, for example, metal, such as aluminum, as well as composite or composite overwrapped tank skins. For example, the cryogenic tank or vessel that contains the CHRS material should be able to withstand some pressure increase. As a result, suitable tank or vessel specification should be determined, including, for example, proper material and wall thickness. The cryogenic system components may include various conduits to supply material to and vent material from the cryogenic system including the cryogenic tank or vessel as well as sensors to monitor the cryogenic system including the cryogenic tank or vessel. 
         [0021]    Insulation of the tank is important to maintain the temperature of the hydrogen contained therein. Such a change in temperature can be affected by various factors including convection (caused by, for example, heat flowing from the ambient atmosphere to the tank at the launch pad), conduction (caused by, for example, heat flowing from spacecraft components through the support structure to the tank) and radiation (caused by, for example, heat transmitting by solar irradiation, or planetary albedo impinging on the tank surface). 
         [0022]    The cryogenic tank or vessel (the terms “tank” or “vessel” may be used interchangeably any where herein) may include design features and components to maintain the CHRS material therein. For example, when the CHRS material includes liquid hydrogen, low or zero gravity fluid management using screen channel and/or vane systems are two possible options for the fluid management system. Such fluid management systems may be needed to provide the required fluid distribution in the cryogenic tank or vessel, and suppress the formation of large gas bubbles therein. A vane system may also be used in several locations of the cryogenic system in order to create enough surface tension force to move gas present in the cryogenic tank or vessel to a vent location or a cooler location for recondensation. 
         [0023]    The cryogenic system supplying, supporting and maintaining the cryogenic tank or vessel and the CHRS material therein may be active or passive and include a space thermal system and ground cooling system. The ground cooling system may be utilized to supply, support and maintain the cryogenic system and cryogenic tank or vessel aboard a spacecraft prior to launch, including, for example, on Earth. Such a ground cooling system may, for example, utilize a cryogenic hydrogen subcooler to cool hydrogen close to triple point temperature within a day and a helium cooler to freeze and subcool the hydrogen to 10 K. Such a ground cooling system may be included in a spacecraft or separate there from, preferably it is housed at a launch facility separate from the spacecraft and located on or close to the launch pad. In the latter preferred embodiment, the ground cooling system is connected to the spacecraft and disconnected at or before launch. 
         [0024]    The space thermal system (thermal management system) may be utilized to supply, support and maintain the cryogenic system and cryogenic tank or vessel aboard a spacecraft after launch or once a separate ground cooling system is disconnected from the spacecraft. In one embodiment, the CHRS including the solid hydrogen, the cryogen thermal and storage system would have an areal mass density of 70 kg/m 2 . Such a system could utilize, for example, a passive thermal control system including solar shields, load responsive multilayer insulation (LRMLI), multilayer insulation (MLI), aluminum foam (for example, 3% density), and 100 K shield cooled by a 4 meter diameter radiator. Such a design may utilize the benefit of the 7 K temperature of deep space (for example, when the spacecraft is not in planetary orbit) by pointing the radiator towards deep space. The preliminary thermal analysis results show that the heat leak from a crew module is 50 Watt, which can be easily compensated with a small heater, such as radiator  233  shown in  FIG. 2 . As a further example, the CHRS can absorb 1 Watt of heat in deep space from the Sun and 130 Watt of heat from the Earth and Sun over a couple of orbits in LEO. For a one year mission to Mars, for example, a spacecraft may stay in LEO for a few hours. The overall heat leak could be about 32,500 kJ for the whole mission, which could increase the temperature of, for example, solid hydrogen from 10 K to close to 14 K (the triple point of hydrogen). In this example, with CHRS, the mass of crew module with radiation shielding could be reduced from more than 26,500 kg to less than 17,800 kg. CHRS could save nearly 8,800 kg for a 4 m diameter and 8 m long cylindrical crew module and halves the required shielding mass when compared with polyethylene shields. Such could, for example, save close to 44 million dollars in launch cost, based on $5000/kg estimate for SpaceX Felcon 9. In another embodiment, the space thermal system may also include a cryocooler, for example, a 14K cryocooler, in the design to actively store the hydrogen at a desired cryogenic temperature and in, for example, solid form. Such a cryocooler, for example, a 14K cryocooler, may be beneficial on space missions lasting more than 1 year. 
         [0025]      FIG. 3  illustrates an embodiment showing a space thermal system and a ground cooling system, around a crew module. It shows half of a section view of the spacecraft, since the spacecraft is reasonably symmetric about the bottom horizontal edge of the schematic. Spacecraft  300  includes a fuselage  302  with an internal volume (area)  303 , a crew module  304  and space thermal system  306 . The crew module  302  includes a walled enclosure  308  with an exterior surface  309  and an internal volume (area)  310  within the walled enclosure  308 . Space thermal system  306  has a tank  312  including, for example, a metal skin, for example, aluminum, and a foam insert, for example, aluminum foam, preferably about 1% to about 3% density aluminum foam, more preferably about 1% density aluminum foam. In the embodiment, tank  312  is encased with several exemplary layers of insulation materials. Encasing tank  312  is a tank integrated multilayer insulation (IMLI-a product of Quest Thermal Group)  314  composed of, for example, layers of multilayer insulation (MLI) with polymer spacers. Encasing the tank IMLI  314  is a 100 K thermal shield  316  composed of, for example, aluminum. Encasing the 100 K thermal shield  316  is a 100 K thermal shield load responsive multilayer insulation (LRMLI-a product of Quest Thermal Group)  318  composed of, for example, layers of MLI supporting a lightweight metallic vacuum shell with polymer spacers. Encasing the 100 K thermal shield LRMLI  318  is 100 K thermal shield IMLI  320  composed of, for example, layers of MLI with polymer spacers. Spacecraft  300  also includes low thermal conductivity support structure  322  (for example, T300) and thermal connections  324  and  326  that provide heat sinks for heat interception. 
         [0026]    Ground cooling system  328  includes liquid hydrogen supply cluster  330 , hydrogen freezing cluster  332 , hydrogen tank fill and vent cluster  334 , LRMLI vent cluster  336  and liquid hydrogen ground subcooling return cluster  338 . Liquid hydrogen subcooling supply cluster  330  is connected at  340  to a hydrogen subcooler and includes conduit system  342 , burst disk/relief valve  344 , seal-off valve  346  and thermal acoustic oscillation damper  348 . Conduit system  342  is connected to tank  312  at thermal connection  350 . Hydrogen freezing cluster  332  performs a freeze and subcooled freezing operation on hydrogen in tank  312  and includes an inlet and outlet for the hydrogen freezing coolant in a conduit  352  that runs from seal-off valve  354  to seal-off valve  356  through thermal connection  364 , section  358  that passes through tank  312  and thermal connection  366 . Hydrogen tank fill and vent cluster  334  includes conduit system  370  and burst disk/relief valve  372 . Conduit system  370  is connected to a hydrogen source at  374  and provides hydrogen to tank  312  via thermal connection  376 . Hydrogen tank fill and vent cluster  334  also includes conduit system  378  with pyro valve  380  connected to vent  385 , seal-off valve  382  and burst disk/relief valve  383  connected to vent  386 . Conduit system  378  is connected to tank  312  via thermal connection  384  to vent hydrogen from tank  312 . Conduit systems  370  and  378  are connected via conduit system  387  that includes thermal acoustic oscillation damper  388 . LRMLI vent cluster  336  is used to vent the LRMLI for convection insulation while the radiation shield is on the ground in an environment with an atmosphere and includes conduit system  389 , burst disk/relief  390  and seal-off valve  391 . Conduit system  389  is connected to tank  312  at  392  and vents through a vacuum pump at  393 . Liquid hydrogen ground subcooling return cluster  338  is connected at  394  to the return side of the hydrogen subcooler and includes conduit system  395 , burst disk/relief valve  396 , seal-off valve  397  and thermal acoustic oscillation damper  398 . Conduit system  395  is connected to tank  312  at thermal connection  399 . 
         [0027]    Using the CHRS system, the crew module can be substantially maintained at about room temperature with a 50 Watt heater, while keeping the CHRS temperature at the desired low temperature, for example, below 14 K. 
         [0028]    The mass and the power requirements of the CHRS system aboard a spacecraft should be determined and incorporated into the overall spacecraft design. For example, the mass of solar panels that may be needed for the power requirement should be calculated and added to the CHRS system when liquid hydrogen is used. 
         [0029]      FIG. 4  illustrates a CHRS tank and insulation  400  embodiment including LRMLI  402  and  404 , MLI or IMLI  406 ,  408 ,  410  and  412 , 90-100 K shield  414 , gaps  416  and  418 , CHRS tank walls  420  and  422  and solid hydrogen and aluminum foam  424 . A crew module is positioned in this embodiment closest to LRMLI  402 . 
         [0030]      FIG. 5  illustrates comparisons on depth-effective dose estimates versus shielding thickness using the ICRP definition of quality factors for several materials.  FIG. 6  illustrates comparisons on depth-effective dose estimates versus shielding thickness using the NASA Solid cancer definition of quality factors for several materials. Calculations For both  FIG. 5  and  FIG. 6  are for 1-year GCR exposures at solar minimum of a human behind each of the shielding materials. The shielding materials in  FIG. 5  and  FIG. 6  are aluminum (graphs  501  and  601 , respectively), epoxy (graphs  502  and  602 , respectively), water (graphs  503  and  603 , respectively), polyethylene (graphs  504  and  604 , respectively), and liquid hydrogen (graphs  505  and  605 , respectively). The horizontal axis indicates the g/cm2 of each of the materials and the vertical axis indicates the radiation dose (exposure) in millisievert (mSv). 
         [0031]    Another benefit of an embodiment utilizing, for example, CHRS material could be used for other mission purposes, such as fuel for a final burn that could help capture the spacecraft into low Earth orbit on a return trajectory or even be used for a burn on a lunar ascent vehicle. Such a dual use could further increase the mass advantage of such embodiments. 
         [0032]    This written description uses examples as part of the disclosure, including the best mode, and also to enable any person skilled in the art to practice the disclosed implementations, including making and using any devices or systems and performing any incorporated methods. The patentable scope is defined by the claims, and may include other examples that occur to those skilled in the art. Such other examples are intended to be within the scope of the claims if they have structural elements that do not differ from the literal language of the claims, or if they include equivalent structural elements with insubstantial differences from the literal languages of the claims.