Thermal Capacitor Block with Integrated Fluidic Channels

A spacecraft propulsion system comprises an attitude adjustment thruster system with multiple thrusters (488a-d) receiving heated propellant via a shared thermal capacitance block (275). The thermal capacitance block (275) with integrated fluidic channels receives energy from a solar concentrator (320) and stores the heat.

FIELD OF THE DISCLOSURE

The disclosure generally relates to operating a spacecraft and more specifically to operating an attitude adjustment thruster system with a common heater implemented as a thermal capacitance block.

BACKGROUND

With increased commercial and government activity in the near space, a variety of spacecraft and missions are under development. For example, some spacecraft may be dedicated to delivering payloads (e.g., satellites) from one orbit to another. In such orbital transfer missions, thruster systems, including attitude adjustment thrusters, may be designed for reliability to consistently achieve mission objectives.

Current attitude adjustment thruster systems may include multiple thrusters (e.g., electro-jets) that receive propellant from a tank and have electrical components to add energy to the propellant. These systems may have a number of possible failure modes.

Furthermore, managing and distributing energy in the spacecraft (including to the attitude adjustment thrusters) remains a challenge. There are existing inefficiencies in collecting solar energy and storing it. Additionally, there are heat sources in various spacecraft subsystems that need to be removed from the spacecraft and radiated into space.

SUMMARY

This disclosure generally relates to improving reliability of a maneuvering or attitude adjustment thruster system . . . .

In one embodiment, a device for heating a propellant in a spacecraft includes a thermal capacitor block of a certain volume, the thermal capacitor block configured to operate in a low-pressure environment and including at least one material to store thermal energy. The device further includes one or more integrated fluidic channels traversing the thermal capacitor block and configured to carry the propellant so as to transfer heat from the at least one material to the propellant, wherein the one or more fluidic channels occupy a minority of the volume, and the material occupies a majority of the volume.

In another embodiment, a spacecraft includes a propellant tank, at least one thruster, and a heating device configured to receive propellant from the propellant tank and supply the propellant to the thruster. The heating device includes a thermal capacitor block of a certain volume, the thermal capacitor block configured to operate in a low-pressure environment and includes at least one material to store thermal energy. The heating device further includes one or more fluidic channels traversing the thermal capacitor block and configured to carry the propellant so as to transfer heat from the thermal capacitor block to the propellant, wherein the one or more fluidic channels occupy a minority of the volume, and the material occupies a majority of the volume.

In yet another embodiment, a spacecraft maneuvering system includes a propellant tank and a shared thermal capacitor including a material configured to store thermal energy, the shared thermal capacitor configured to (i) receive, via a plurality of ingress ports, a propellant for a plurality of respective fluidic channels, (ii) transfer the stored thermal energy to the propellant in the plurality of fluidic channels, and (iii) output the propellant from the plurality of fluidic channels via a plurality of a respective egress ports. The system further includes a plurality of valves configured to restrict an amount of propellant directed to the respective ingress ports; and a controller configured to control the plurality of the valves to change flow rates through the plurality of the fluidic channels in response to signals indicative of intended spacecraft maneuvers.

In yet another embodiment, a method of maneuvering a spacecraft includes directing a propellant from a propellant tank to a plurality of fluidic channels in thermal communication with a shared thermal capacitor, via a plurality of respective valves, including controlling, by a controller, flow rates through the valves in accordance with intended spacecraft maneuvers. The method further includes transferring thermal energy stored in a material of the thermal capacitor to the propellant in the plurality of fluidic channels and directing the propellant from the plurality of fluidic channels to respective ones of a plurality of thrusters.

In yet another embodiment, a solar concentrator device configured to operate in a spacecraft and collect solar energy includes: a reflecting side configured to focus solar radiation on a thermal target disposed at the spacecraft, and a radiating side configured to receive thermal energy from the spacecraft and to radiate the received thermal energy. The device further includes a heat transfer component to transfer thermal energy from the spacecraft to the radiating side.

In yet another embodiment, a method of managing thermal energy in a spacecraft includes: focusing, using a reflective side of a solar concentrator, solar radiation on a thermal target; transferring, using a heat transfer component, thermal energy from the spacecraft to the radiating side of the solar concentrator; and radiating, using the radiating side of the solar concentrator, the transferred thermal energy.

In yet another embodiment, a thermal system for use in a spacecraft includes a thermal target including an absorbing surface configured to absorb solar radiation, and a solar concentrator configured to direct solar radiation toward the thermal target, via an optical path. The thermal system is configured to reflect at least a portion of thermal radiation emitted by the thermal target, in the optical path, back to the thermal target.

In yet another embodiment, a method of managing thermal energy in a spacecraft includes directing, using a solar concentrator, solar radiation on a thermal target via an optical path, and causing at least a portion of thermal radiation emitted by the thermal target to be reflected back to the thermal target, in the optical path.

DETAILED DESCRIPTION

Spacecraft may be configured for transferring a payload from a lower energy orbit to a higher energy orbit according to a set of mission parameters. The mission parameters may include, for example, a time to complete the transfer and an amount of propellant and/or fuel available for the mission. Generally, spacecraft may collect solar energy and use the energy as well as the stored propellant to generate thrust using one or more thrusters. The spacecraft may use distinct sets of thrusters for changing orbit energy and maneuvering to change spacecraft orientation with respect to an orbit. In some implementations, a spacecraft maneuvering system, or, equivalently, an attitude adjustment thruster system, may be used for docking with other spacecraft, payloads, or fuel depots in space. Different thruster types and/or operating modes may trade off the total amount of thrust with the efficiency of thrust with respect to fuel or propellant consumption, defined as a specific impulse. Different thruster types or thruster systems may have different levels of complexity and reliability. Generally, optimizing and/or simplifying thruster systems may improve spacecraft efficiency and/or reliability.

A disclosed spacecraft includes a system of multiple attitude adjustment, otherwise referred to as maneuvering, thrusters sharing energy system components to kinetically energize supplied propellant. In some implementations, the attitude adjustment (i.e., maneuvering) system may be a passive structure expelling the energized propellant through nozzles. Such thruster configuration may increase thruster system reliability by reducing possible failure modes, for example, by obviating individual heating elements in the maneuvering thrusters. Furthermore, the disclosed enabling energy system components increase energy efficiency of the spacecraft by efficiently converting sunlight into propellant energy. Still furthermore, the disclosed maneuvering thruster system is scalable because a robust common heater may replace individual in-thruster electrical heating elements (e.g., in electro-jets) that can have power limitations.

FIG.1is a block diagram of a spacecraft100configured for transferring a payload between orbits. The spacecraft100may include the disclosed attitude adjustment thruster system (referred to as a subsystem in the context of a thruster system that includes other thrusters) and supporting energy system components. The spacecraft100includes a number of systems, subsystems, units, or components disposed in or at a housing110. The subsystems of the spacecraft100may include sensors and communications components120, mechanism control130, propulsion control140, a flight computer150, a docking system160(for attaching to a launch vehicle162, one or more payloads164, a propellant depot166, etc.), a power system170, a thruster system180that includes a primary propulsion (main) thruster subsystem182and an attitude adjustment thruster subsystem184, and a propellant system190. Furthermore, any combination of subsystems, units, or components of the spacecraft100involved in determining, generating, and/or supporting spacecraft propulsion (e.g., the mechanism control130, the propulsion control140, the flight computer150, the power system170, the thruster system180, and the propellant system190) may be collectively referred to as a propulsion system of the spacecraft100.

The sensors and communications components120may include a number of sensors and/or sensor systems for navigation (e.g., imaging sensors, magnetometers, inertial motion units (IMUs), Global Positioning System (GPS) receivers, etc.), temperature, pressure, strain, radiation, and other environmental sensors, as well as radio and/or optical communication devices to communicate, for example, with a ground station, and/or other spacecraft. The sensors and communications components120may be communicatively connected with the flight computer150, for example, to provide the flight computer150with signals indicative of information about spacecraft position and/or commands received from a ground station.

The flight computer150may include one or more processors, a memory unit, computer readable media, to process signals received from the sensors and communications components120and determine appropriate actions according to instructions loaded into the memory unit (e.g., from the computer readable media). Generally, the flight computer150may be implemented using any suitable combination of processing hardware, that may include, for example, applications specific integrated circuits (ASICS) or field programmable gate arrays (FPGAs), and/or software components. The flight computer150may generate control messages based on the determined actions and communicate the control messages to the mechanism control130and/or the propulsion control140. For example, upon receiving signals indicative of a position of the spacecraft100, the flight computer150may generate a control message to activate one of the thruster subsystems182,184in the thruster system180and send the message to the propulsion control140. The flight computer150may also generate messages to activate and direct sensors and communications components120.

The docking system160may include a number of structures and mechanisms to attach the spacecraft100to a launch vehicle162, one or more payloads164, and/or a propellant refueling depot166. The docking system160may be fluidicly connected to the propellant system190to enable refilling the propellant from the propellant depot166. Additionally or alternatively, in some implementations at least a portion of the propellant may be disposed on the launch vehicle162and outside of the spacecraft100during launch. The fluidic connection between the docking system160and the propellant system190may enable transferring the propellant from the launch vehicle162to the spacecraft100upon delivering and prior to deploying the spacecraft100in orbit.

The power system170may include components for collecting solar energy, generating electricity and/or heat, storing electricity and/or heat, and delivering electricity and/or heat to the thruster system180. To collect solar energy, the power system170may include solar panels with photovoltaic cells, solar collectors or concentrators with mirrors and/or lenses, or a suitable combination of devices. In the case of using photovoltaic devices, the power system170may convert the solar energy into electricity and store it in energy storage devices (e.g., lithium ion batteries, fuel cells, etc.) for later delivery to the thruster system180and other spacecraft components. In some implementations, the power system180may deliver at least a portion of the generated electricity directly (i.e., bypassing storage) to the thruster system180and/or to other spacecraft components. When using a solar concentrator, the power system170may direct the concentrated (having increased irradiance) solar radiation to photovoltaic solar cells to convert to electricity. In other implementations, the power system170may direct the concentrated solar energy to a solar thermal receiver or simply, a thermal receiver, that may absorb the solar radiation to generate heat. The power system170may use the generated heat to power a thruster directly, as discussed in more detail below, and/or to generate electricity using, for example, a turbine or another suitable technique (e.g., a Stirling engine). The power system170then may use the electricity directly for generating thrust or storing electrical energy.

The thruster system180may include a number of thrusters and other components configured to generate propulsion or thrust for the spacecraft100. Thrusters may generally include main thrusters in the primary propulsion subsystem182that are configured to substantially change speed of the spacecraft100, or as attitude control thrusters in the attitude control thruster subsystem184that are configured to change direction or orientation of the spacecraft100without substantial changes in speed.

One or more thrusters in the primary propulsion subsystem182may be a microwave-electro-thermal (MET) thrusters. In a MET thruster cavity, an injected amount of propellant may absorb energy from a microwave source (that may include one or more oscillators) included in the thruster system180and, upon partial ionization, further heat up, expand, and exit the MET thruster cavity through a nozzle, generating thrust.

Another one or more thrusters in the primary propulsion subsystem182may be solar thermal thrusters. In one implementation, propellant in a thruster cavity acts as the solar thermal receiver and, upon absorbing concentrated solar energy, heats up, expands, and exits the nozzle generating thrust. In other implementations, the propellant may absorb heat before entering the cavity either as a part of the thermal target or in a heat exchange with the thermal target or another suitable thermal mass thermally connected to the thermal target. In some implementations, while the propellant may absorb heat before entering the thruster cavity, the primary propulsion thruster subsystem182may add more heat to the propellant within the cavity using an electrical heater or directing a portion of solar radiation energy to the cavity.

Thrusters in the attitude adjustment subsystem184may use propellant that absorbs heat before entering the cavities of the attitude adjustment thrusters in a heat exchange with the thermal target or another suitable thermal mass thermally connected to the thermal target. In some implementations, while the propellant may absorb heat before entering thruster cavities, the thrusters of the attitude adjustment thruster subsystem184may add more heat to the propellant within the cavity using corresponding electrical heaters.

The propellant system190may store the propellant for use in the thruster system180. The propellant may include water, hydrogen peroxide, hydrazine, ammonia or another suitable substance. The propellant may be stored on the spacecraft in solid, liquid, and/or gas phase. To that end, the propellant system190may include one or more tanks, including, in some implementations, deployable tanks. To move the propellant within the spacecraft100, and to deliver the propellant to one of the thrusters, the propellant system190may include one or more pumps, valves, and pipes. The propellant may also store heat and/or facilitate generating electricity from heat, and the propellant system190may be configured, accordingly, to supply propellant to the power system170.

The mechanism control130may activate and control mechanisms in the docking system160(e.g., for attaching and detaching a payload or connecting with an external propellant source), the power system170(e.g., for deploying and aligning solar panels or solar concentrators), and/or the propellant system190(e.g., for changing configuration of one or more deployable propellant tanks). Furthermore, the mechanism control130may coordinate interaction between subsystems, for example, by deploying a tank in the propellant system190to receive propellant from an external propellant source connected to the docking system160.

The propulsion control140may coordinate the interaction between the thruster system180and the propellant system190, for example, by activating and controlling electrical components (e.g., a microwave source) of the thruster system140and the flow of propellant supplied to thrusters by the propellant system190. Additionally or alternatively, the propulsion control140may direct the propellant through elements of the power system170. For example, the propellant system190may direct the propellant to absorb the heat (e.g., at a heat exchanger) accumulated within the power system170. Vaporized propellant may then drive a power plant (e.g., a turbine, a Stirling engine, etc.) of the power system170to generate electricity. Additionally or alternatively, the propellant system190may direct some of the propellant to charge a fuel cell within the power system190. Still further, the attude adjustment thruster subsystem184may directly use the heated propellant to generate thrust.

The subsystems of the spacecraft may be merged or subdivided in different implementations. For example, a single control unit may control mechanisms and propulsion. Alternatively, dedicated controllers may be used for different mechanisms (e.g., a pivot system for a solar concentrator), thrusters (e.g., a MET thruster), valves, etc. In the following discussion, a controller may refer to any portion or combination of the mechanism control130and/or propulsion control140.

FIGS.2A and2Billustrate the distinction between a known attitude control thruster system200aand an example disclosed attitude control thruster system200b, each exemplifying the attitude control subsystem184. The systems200a,binclude propulsion controls240a,b, power systems270a,b, thruster systems284a,b, and propellant systems290a,b, exemplifying, respectively, propulsion control140, power system170, attitude control thruster subsystem184, and propellant system190. The power system270bmay include a thermal block275, described below. The thruster system284aincludes four thrusters286a-d. The thruster systems284b, analogously, includes four thrusters288a-d. Generally, the systems200a,bmay include any suitable number (2, 2, 4, 5, 6, 7, 8, 9, 10, 12, etc.) of thrusters. The thrusters288a-dmay be different from the thrusters286a-d. For example, the thrusters288a-dmay operate without the need for electrical components, enabling a more robust and scalable maneuvering system.

The thruster system200a, is configured to use thrusters286a-dthat each receive a supply of propellant from the propellant system290aand a supply of power from the power system270a. Each of the thrusters286a-dis configured to individually convert the supplied power into energy of the supplied propellant. Such thrusters may be, for example, electro-jet or micro-plasma thrusters.

In contrast, the thruster system200bis configured to use thrusters288a-d that each receive high-temperature (e.g.,100,150,250,400,600,900,1200, etc. ° C.) propellant pre-heated by the power system270b. To that end, the power system170bmay include the thermal capacitor block275, that may be referred to as a thermal capacitance block, thermal block or a thermal capacitor. The thermal block275may act as a thermal target receiving energy from a solar collector, as described below. In some implementations, the thermal block275may include an electrical heater to raise the block temperature through Joule heating. The heater may be embedded in a ceramic material to electrically isolate the heater from a potentially conductive material of the block275. The thermal block275may be a device that combines functionalities of a heat storage device and a thermal exchanger by combining thermal mass having substantial heat capacity with a number of integrated fluidic channels. The number of the integrated fluidic channels may correspond to the number of thrusters (e.g., four fluidic channels for the thrusters288a-d). The thrusters288a-dmay run through the thermal capacitor block275so as to heat the working fluid, which, in the case of the system200b, is the propellant for the thrusters288a-d. The thermal capacitor block is described in more detail below, with reference toFIGS.6A, B.

It should be noted that a thruster system with a thermal capacitance block (e.g., block275) may be configured to operate with a single thruster, using a dedicated fluidic channel. Furthermore, the thermal capacitor block may include a secondary fluidic channel, coupled to a non-thruster component. The secondary channel may carry a working fluid used, for example, for power production (e.g., with a turbine).

FIG.3schematically illustrates an example implementation300of the thruster system200bwith the thermal block275receiving solar energy (designated by rays310) collected and directed by a solar collector320disposed outside of a satellite housing330and controlled by a pointing system340. The solar collector320may be referred to as the solar concentrator320. The pointing system340is mechanically attached to the solar collector320and may include a portion of the mechanism control130along with a mechanism (e.g., actuators, beams, gears, tethers, etc.) for deploying and/or moving the solar collector320. The pointing system may move the solar collector to create an angle with respect to the rays of the sun that would guide the solar radiation toward the thermal block275. The implementation300is described further with reference toFIG.4.

FIG.4is a perspective illustration of a spacecraft400in which the implementation300of the system200bmay operate. A housing401(e.g., housing330) of the spacecraft400has one side removed to show internal components. The system200bmay operate within the spacecraft400in conjunction with the solar collector320, as illustrated inFIG.3. InFIG.4, a propellant tank405included in the propellant system290bis fluidicly connected via conduits408a,bto thrusters488a-b(which may be two of the thrusters288a-d). Thrusters488c-dconnect to the tank405in a similar manner, although the corresponding conduits are obscured by the housing401inFIG.4. The spacecraft400may include a main thruster489. The main thruster may share propellant with the maneuvering thrusters488a-d. The main thruster489may be a MET thruster and use electrical energy. The electrical energy may be collected using solar cells with or without the use of the solar collector320or generated using a thermoelectric plant that may include using the heat accumulated in the thermal block275.

The conduits408a,bmay include pipes or tubes made out of metal or another suitable material and connected in series with suitable connectors. In some implementations, the conduits408a,bmay run through the thermal block275, the walls of the conduits408a,bin thermal contact with the thermal block275. In other implementations, the material of the block275may form walls of channels integrated into the block275. These integrated channels may connect to incoming and outgoing conduits408a,bto form continuous fluidic channels connecting the tank405with the thrusters488a,b. The integrated channels may help ensure rapid heat transfer from the heat stored in block275to the working fluid. One or more pumps, one or more valves, and/or one or more splitters/combiners may be configured to control the flow through the conduits408a,band the corresponding fluidic channels. Such pumps, valves and/or splitters are not shown inFIG.4to avoid clutter, but discussed below with reference toFIGS.5and6.

The solar concentrator320inFIGS.3and4is illustrated as a curved mirror configured to focus onto the thermal block275solar rays approaching the mirror, substantially in parallel, from the direction of the sun. Generally, the solar concentrator320may include one or more mirrors, lenses, and/or fiber-optic guides to collect and guide solar radiation toward the block275serving as a thermal target. The block275is configured to absorb the radiant solar energy, converting it to heat. In some implementations, optics of the solar collector320may divert a portion of solar energy to another thermal target or a solar cell array, for example.

As discussed above, a mechanism (e.g., included in the pointing system340ofFIG.3) may move the solar collector320so as to guide solar radiation toward the block275. In some implementations, maneuvers to orient the spacecraft400as a whole may contribute to positioning of the solar collector320.

FIGS.5A, B illustrate example implementations500a,bof the attitude adjustment thruster system200bwith particular emphasis on enabling thruster system functionality by controlling the propellant system290b. The implementations500a,binclude controllers510a,b, implementing at least portions of the propulsion control140. The controllers510a,bcontrol a propellant system590(example implementations of the propellant system200b).

The propellant system590may include a tank592in fluidic connection with a pump594, a splitter596, and valves598a-d. The tank592and the pump594may be configured to convert a multiphase microgravity mixture of a propellant to a liquid propellant downstream of the pump596. The splitter596may be configured to split the liquid stream from the pump596into four portions, each directed to one of the valves598a-d. In some implementations, the splitter596may split the incoming propellant into equal portions under the condition that the valves598a-dare fully open. In other implementations, the splitter may have asymmetry in the splitting of the incoming stream, even when the valves598a-dare fully open.

The valves598a-dmay be configured to determine, at least in part, flow rates of the propellant toward each of the thrusters288a-d. For example, a partial restriction in valve598a(or598b-d) may be configured to reduce the flow rate to thruster288a(or, respectively, thrusters288b-d). A full closing of one of the valves598a-dmay be configured to fully stop the flow of the propellant to the corresponding thruster.

In operation, the controller510a(or the controller510b) may activate the pump596when a flight computer (e.g., flight computer150) signals that attitude adjustment thrust is required to maneuver the spacecraft (e.g., spacecraft400). The controller510a(or the controller510b) may compute the degree to which each of the valves598a-dis open based on the amount and direction of required thrust. The propellant flows through the valves598a-dand, via ingress ports572a-d, into the thermal capacitor block275heated to a high temperature. Passing through the block275, the propellant may vaporize. For example, when water serves as the propellant, the block275may convert the propellant into superheated steam. Fluidic conduits (e.g., conduits408a-b) may then direct to heated gas propellant ejected out of the egress ports574a-dinto each of the thrusters288a-d. The pressure of the gas in each thruster may depend on the flow rate out of the pump594and the degree to which each of the corresponding valves598a-dis opened. The pressurized propellant gas generates thrust as it exits the thrusters288a-d. To maximize thrust, the thrusters288a-dmay include suitable expansion nozzles. The thrusters288a-ddisposed at different locations with respect to the center of the spacecraft (e.g., as thrusters488a-dinFIG.4) and producing different amounts of thrust, generate torque to turn or maneuver the spacecraft.

The example implementation500bincludes a sensor599, configured to measure the temperature of the block275. The sensor may be a thermocouple, a fiberoptic sensor, an IR sensor or any other suitable contact or contactless sensor. The controller510bmay control the pump594and/or the valves598a-din view of the temperature of the block275. For example, the controller may increase the power to the pump594(and, thereby, the flow rate of the propellant given constant valve configuration) as the temperature of the block275decreases. The temperature of the block275may decrease, for example, due to radiation losses or due to the heat removed by the propellant flowing through the block275.

In some implementations, an imaging thermal sensor or a configuration of sensors may measure temperature non-uniformity in the block275. The controller510bmay control the valves288a-din view of the measured temperature non-uniformity. Additionally or alternatively, the controllers510a,bmay use any other suitable inputs, including, for example, motion of the spacecraft (e.g., as measured by accelerometers, gyroscopic, and/or imaging sensors) to control the pump594and the valves598a-d.

FIGS.6A, B illustrate example thermal capacitor blocks600aand600b. The blocks600a,bmay be made of metal (e.g., copper, brass, aluminum, steel, beryllium, or any other suitable metal or allow). In some implementations, the block may be made of ceramic, or any other suitable crystalline or amorphous material. Furthermore, the block material need not be uniform. Portions of the blocks600a,bmay be made of different materials. Generally, tradeoffs between high thermal conductivity, high specific heat, high melting point, low density, low cost, good manufacturability, etc. may dictate the choice of block material.

Although the blocks600a,bare illustrated inFIGS.6A,B as parallelepipeds, the blocks600a,bmay have any suitable shape (e.g., cubes, cylinders, cones, truncated cones, toroids, etc.). The blocks600a,bmay be of any suitable dimensions, with linear dimensions of 5, 10, 20, 30, 40, 50 cm or any other suitable value.

In some implementations, a block may include an inner portion and an outer portion of different materials. For example, block600bmay include a cylindrical core608. The core608may include a phase-changing (e.g., two-phase) material that may melt, at least partially, at an operating temperature of the block600bto store heat in a phase change (as latent heat). Heat storage in a phase change may allow storing heat while maintaining substantially constant temperature that may be an operating temperature of the thermal block600b. The phase-changing material may be a salt (e.g., chlorides, nitrides, or nitrates of sodium and/or potassium) or a metal (e.g., allows of tin, indium, lithium, lead, etc.) with lower melting point that the outer portion of the block600b. In some implementations, the phase-changing material may be a material configured to change between a liquid and a gas phase. In some implementations, the thermal block may include a three-phase material, configured to store heat in successive phase changes.

The blocks600a,bmay include respective absorptive surfaces605a,bconfigured to absorb a substantial portion (e.g., 50, 60, 70, 80, 90, 95% or more) of the impinging solar radiation (e.g., focused by a solar collector). To that end, in some implementations, the absorptive surfaces605a,bmay be covered with a black paint or another suitable coating. In other implementations, the blocks605a,bmay include plates of absorptive (for a suitable portion of solar spectrum) thermally conductive material (e.g., anodized aluminum).

In some implementations, a thermal block (e.g., blocks600a,b) may include an electrical heater configured to heat the block. The electrical heater may be configured to carry electrical current to heat the block through Joule heating. The electrical heater may be imbedded in a ceramic (or another dielectric) portion of the thermal block, the ceramic portion conductive thermal coupling to the portion of the thermal block with integrated fluidic channels.

The block600aincludes integrated fluidic channels612a-dthat traverse the block600a. The block600bincludes integrated fluidic channels614a,bthat traverse the block600b. Generally, a thermal capacitor block (e.g., blocks600a,b) may include any suitable number of channels (1, 2, 5, 10, etc.). The channels (e.g.,612a-dand614a,b) may occupy only a small volumetric portion of the block (e.g., 50, 20, 10, 5, 2, 1% or smaller). The channels may have any suitable shape. The channels may meander through the block and have length substantially greater than any linear dimension of the block. Furthermore, cross-section area of the channels may change along the length of the channels. The increasing cross-section downstream with respect to the flow direction of the propellant may accommodate an increasing volume flow rate as the propellant absorbs heat.

The fluidic channels612a-dmay meander through the thermal block600a, each channel (612a-d) substantially constrained to a respective one of the four parallel planes. The number of turns may be any suitable number. Fluidic channels614a-b, on the other hand, have helical portions spiraling through the thermal block600b. The helical portions of the channels614a-bmay maintain constant distance from the block core608. Generally, a thermal block may include any suitable number (e.g., 3, 4, 5, 6, 8, 10, etc.) of helical channels. Furthermore, the angular distance between any two channels may stay constant along the length of the block in the direction of propellant flow. In this manner, the helical channels may have substantially equivalent (i.e., symmetrical) relationship to the geometry of the thermal block.

A variety of manufacturing methods may be used to build thermal capacitor block (e.g., blocks275,600a,b). In some implementations, a subtractive manufacturing method step may include machining at least a portion of an integrated channel in at least a portion of the thermal block. Another manufacturing step may include assembling portions of the thermal capacitor block with at least partial integrated channels. For example, manufacturing the thermal capacitor block600amay include steps for machining portions of the block600a, the portions separated by the planes of the meandering channels612a-d.

In other implementations, a manufacturing method for a thermal capacitor block (e.g., blocks275,600a,b) may include an additive manufacturing step. For example, a three-dimensional (3D) printing step may produce a monolithic thermal block with integrated helical channels. In other implementations, a 3D printing step may produce a solid material portion of a thermal capacitor block (e.g.,600b), while leaving a void for a two-phase portion.

Additionally or alternatively, manufacturing a thermal capacitor block portion with integrated channels may include a combination of additive and subtractive steps. For example, a layer of material of the thermal block may be created (e.g., deposited, fused, etc.) using an additive method. Subsequently, a subtractive manufacturing stem (e.g., mechanical or laser drilling) may create portions of the channels within the layer. The portions of the channels in the layer may traverse the layer at any suitable angle. Thus, for example, the manufacturing method, may create, layer by layer, a thermal capacitor block with integrated helical channels (e.g., block600b). An additive step of a manufacturing method may deposit subsequent layers in a liquid state, and the subtractive step may create holes in each layer once the corresponding layer solidifies. Thus, integrated channels of a suitable shape may be created.

A device for heating propellant may use a number of techniques for minimizing heat loss of a thermal capacitor block (e.g., block275) included in the device. The device may include one or more structures for mitigating conductive, convective, and/or radiant heat losses of the thermal capacitor block.

FIG.7illustrates configurations700a-dof a for managing radiant energy transfer between the thermal block275and the environment surrounding the thermal block275in a device for heating propellant. In configurations700a-dthe device includes a reflector710configured to reflect radiant energy emitted by the block275back to the block275. The reflector710may surround the thermal block275to minimize radiation heat loss. Fluidic conduits connected to ingress and egress ports of the thermal block275may penetrate the reflector710.

The reflector710may be attached to the block by stand-offs712. The stand-offs712may minimize conduction of heat from the thermal block275to the reflector, while maintaining structural integrity of the device. To that end, the cross-section area of the stand-offs may be minimized and the material of the stand-offs may have suitably low thermal conductivity.

Using the stand-offs712, the reflector710may be substantially separated from the thermal capacitor block275by a gap714. The gap714may be substantially vacated of matter (i.e., a vacuum or very low-pressure gap). For example, the gas pressure within the gap may be less than 1/100, 1/1000, or 1/10000 or any other suitable fraction of atmospheric pressure. In a manner of speaking, operating a thermal capacitor block with a reflector separated by the gap714advantageously uses the vacuum in the space environment to minimize convective heat loss. Thus, the design of the device for heating the propellant using the thermal capacitance block275, reflects that the device is configured to operate in the low pressure of the space environment.

The block275is configured to receive radiant energy of the sun collected and focused by solar concentrators720a-c(e.g., solar concentrator320). The collected and focused solar energy may impinge on the thermal block275through a window722. The window722opens an optical path for the solar radiation to reach the thermal block275. Conversely, the thermal block275may lose some heat by radiating through the same optical path (in reverse direction) opened through window722, particularly when the solar concentrators720a-care not collecting solar energy (e.g., when spacecraft does not have a line of sight to the sun). One or more techniques may mitigate radiative heat loss through the window722. In some implementations, an optical element may mitigate the radiative heat loss by reflecting radiation emitted by the block275back to the block275along substantially the same optical path as used by the solar radiation focused on the block275.

In configuration700a, the solar concentrator720amay refract the solar rays onto the thermal block275. A surface of the solar concentrator may be coated to reflect a suitable portion of the infrared (IR) spectrum (e.g., in near- and mid-IR). The IR reflective surface may reflect the radiation emitted by the thermal block275back onto the thermal block275. The coated refractive element of the solar concentrator may act as a concave IR mirror from the perspective of the thermal block275. ThoughFIG.7illustrates the solar concentrator720aas a single refractive element, generally, a solar concentrator may combine a suitable number of refractive and/or reflective elements. At least one of the refractive elements in the combination may include a reflective IR coating.

It should be noted that a reflective IR coating may reflect a portion of the impinging solar radiation, reducing the rate of energy collection. However, the spectral shape difference between the solar radiation, centered in the visible portion of the spectrum, and the energy radiated by a heated thermal block (e.g., block275), centered in IR, contributes to the benefit of the coating for preventing thermal loss. The reflective IR coating may be a transparent heat-reflective (THR) coating, and may include silver, gold, copper, and/or TiO2, for example.

Configuration700bmay include a filtering element730, configured to reflect most of the energy emitted by the thermal block275back to the thermal block275, while transmitting the substantial portion of solar radiation collected by the solar concentrator720. The filtering element may include a coating as described above. In some implementations, the filtering element730may be included in the optics of the solar concentrator720. For example, the filtering element may be a diffractive optical element (e.g., a Fresnel lens) that contributes to focusing the solar radiation onto the thermal block275.

Configurations700cand700dillustrate operational states in a technique of using the solar720concentrator to prevent radiative heat loss of the thermal block275without relying on a coating or another filtering technique. In configuration700c, representing the first operational state, the solar concentrator720points (e.g., as positioned by the pointing system340) so as to reflect the radiation emitted by the block275back to the block275. In configuration700d, representing the second operational state, the solar concentrator720points (e.g., as positioned by the pointing system340) so as to reflect and focus solar radiation onto the block275. A controller may be configured to receive a signal from a sensor sensing availability of solar energy and position the solar concentrator accordingly. In some implementation, the solar concentrator may be configured with a different focal length depending if the solar concentrator720is operating in a collecting (i.e., configuration700dor retro-reflecting (i.e., configuration700c) mode. To that end, a mechanical system (e.g., the pointing system340) may change the curvature of at least one optical element of the solar concentrator720in response to a change in the operating mode.

In some implementations, a solar concentrator (e.g., solar concentrator320,720) may be configured to help manage excess heat within a spacecraft (e.g., heat generated when operating a main thruster). To that end, the solar concentrator may include a radiating side configured to receive thermal energy from the spacecraft and to radiate the received thermal energy into space.

FIG.8illustrates an example implementation800of the thruster system200b. The implementation800may include a solar concentrator820configured to operate in a spacecraft (e.g., spacecraft400) and collect solar energy. The solar concentrator device820may include a reflecting side822configured to focus solar radiation on a thermal target disposed at the spacecraft and a radiating side824configured to receive thermal energy from the spacecraft and to radiate the received thermal energy. The solar concentrator device820may include a heat transfer component840configured to transfer heat from the spacecraft to the radiating side824. The surface of the radiating side824may be configured (e.g., painted, coated, anodized, etc.) to have low reflectance, and, consequently, high emissivity. For example, the radiating side824may be painted black. Regardless of the apparent color in the visible spectrum, the radiating surface824may have a suitably low reflectance (e.g., less than 0.3, 0.2, 0.1 or another suitable value) in the mid-IR spectral region.

A substantial portion of the radiating side824may be separated from the reflecting side822by a gap, providing thermal insulation between the two sides822,824. To that end, the two sides822,824may be separated by stand-offs.

In some implementations, the heat transfer component840may conductively transfer the heat from the spacecraft to the radiating side824. To that end, the heat transfer component may be made of metal (e.g., copper, aluminum, etc.).

In some implementations, the heat transfer component840may include a heat pipe to transfer the heat. The heat pipe may be configured to carry hot fluid from a subsystem in the spacecraft to the radiating side824of the solar concentrator820. After transferring heat to the radiating side824of the solar concentrator820, the cooled working fluid may return along the heat pipe to cool the subsystem.

The heat pipe of the heat transfer component840may be configured as an active heat pipe equipped with one or more pumps. The pump or pumps may transfer the working fluid between the spacecraft subsystem in need of cooling and the radiating side824of the solar concentrator820. Additionally or alternatively, the heat transfer component840may include a passive heat pipe. The passive heat pipe may transfer a condensed working fluid from the radiating side824(the cool end of the heat pipe) to the subsystem in need of cooling (the hot end of the heat pipe) using capillary channels of the heat pipe. The working fluid may evaporate at the hot end of the heat pipe, and the evaporated working fluid may expand to the cool end.

The heat transfer component may include rigid fluidic ducts and flexible fluidic ducts. The flexible fluidic ducts may accommodate deployment and movement of the solar concentrator820.

In some implementations, the working fluid of the heat pipe may be a propellant for one or more thrusters. The propellant may be water, hydrazine, hydrogen peroxide, or any other suitable propellant.

The following aspects are explicitly considered.

Aspect 1. A method of manufacturing a heat exchanger block, the method including: depositing a first layer of block material in liquid form; cooling the first layer of block material until solid; creating holes through the first layer; depositing a second layer of block material in liquid form along a direction of deposition; cooling the second layer of block material until solid; creating holes through the second layer, shifted from the holes in the first layer along the direction transverse to the direction of deposition.

Aspect 2. The method of aspect 1, further comprising using a three-dimensional (3D) printing technique.

Aspect 3. The method of aspect 1, further comprising using a subtractive method to create the holes.