Patent Publication Number: US-2017363044-A1

Title: Small satellite propulsion system utilizing liquid propellant ullage vapor

Description:
FIELD 
     The present invention generally pertains to a small satellite propulsion system, and more specifically, a propulsion system that generates thrust for a small satellite using natural vapor from stored liquid propellant. 
     BACKGROUND 
     A typical chemical thruster provides liquid propellant to an injector where the propellant is effectively sprayed into a combustion chamber to generate thrust. In such a design, a pressurizing gas is typically used to force the liquid propellant through to the injector. In other words, in addition to the liquid propellant, a gas is required. This gas is stored in a tank classified as a pressure vessel. The pressure vessel is generally a safety concern and can make obtaining a launch slot for a small satellite more difficult. 
     Accordingly, an improved approach to provide thrust for a small satellite may be beneficial. 
     SUMMARY 
     Certain embodiments of the present invention may be implemented and provide solutions to the problems and needs in the art that have not yet been fully solved by conventional small satellite propulsion systems. For example, some embodiments of the present invention generally pertain to a small satellite propulsion system that uses vapor pressure that naturally sits on top of liquid propellant to generate thrust in a small satellite. 
     In one embodiment, an apparatus includes vapor that naturally sits on top of liquid propellant stored within a propellant tank. The vapor flows from the propellant tank through a membrane, which is used for metering flow, to interact with a reacting surface to generate thrust. 
     In another embodiment, an apparatus may include a propellant tank and a valve. The propellant tank includes vapor that naturally sits on top of liquid propellant. The valve, when open, causes the vapor to flow from the propellant tank and through a membrane to interact with a reacting surface to generate thrust. 
     In yet another embodiment, an apparatus includes a first propellant tank comprising vapor that naturally sits on top of liquid propellants, and a second propellant tank comprising a second propellant. The vapor and the second propellant is mixed prior to passing through a membrane. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       In order that the advantages of certain embodiments of the invention will be readily understood, a more particular description of the invention briefly described above will be rendered by reference to specific embodiments that are illustrated in the appended drawings. While it should be understood that these drawings depict only typical embodiments of the invention and are not therefore to be considered to be limiting of its scope, the invention will be described and explained with additional specificity and detail through the use of the accompanying drawings, in which: 
         FIGS. 1A and 1B  are diagrams illustrating a small satellite propulsion system, according to an embodiment of the present invention. 
         FIG. 2A  is a diagram illustrating a behavior of vapor flowing through a membrane and reacting surface, according to an embodiment of the present invention. 
         FIG. 2B  is a diagram illustrating a behavior of vapor flowing through a thermal transpiration membrane and reacting surface, according to an embodiment of the present invention. 
         FIG. 3  is a graph illustrating theoretical performance comparison with other currently available small satellite propulsion systems, according to an embodiment of the present invention. 
         FIG. 4  is a graph illustrating a change in orbital velocity and burn time based on a specific thermal transpiration propulsion system using hydrogen peroxide, according to an embodiment of the present invention. 
         FIG. 5  is a graph illustrating thrust and Knudsen number based on membrane pore size in a small satellite propulsion system, according to an embodiment of the present invention. 
         FIG. 6  is a graph illustrating thrust and flowrate based on membrane area, according to an embodiment of the present invention. 
     
    
    
     DETAILED DESCRIPTION OF THE EMBODIMENTS 
     Some embodiments generally pertain to a novel small satellite propulsion system (hereinafter “propulsion system”). The propulsion system may use natural vapor (or gas) from liquid propellant(s), a membrane, and a reacting surface to generate thrust in a small satellite. The thrust may be on the order of milli-Newtons to Newtons. The membrane in some embodiments may use small pores or tubes to meter the flowrate of the vapor and/or create additional pressure (e.g., as would be the case for a thermal transpiration membrane). 
     In some embodiments, this propulsion system may not substantially add to the system requirements of the small satellite. More specifically, the propulsion system may not dictate the design of the small satellite. Furthermore, the propulsion system takes into account that the small satellites are generally not the primary payload on a launch vehicle and are designed to abide by the launch vehicle code. Because small satellites are small in nature, e.g.,  10  cm by  10  cm, the propulsion system in certain embodiments may not encompass a majority of the real estate on the small satellite. 
     In some embodiments, the propulsion system may be utilized in a monopropellant design ( FIG. 1A ), and in other embodiments, the propulsion system may be utilized in a bipropellant design ( FIG. 1B ). The propulsion system does not require complex moving parts or expensive electric components that would otherwise be in, for example, an electrical propulsion system. The propulsion system does not require a separate pressure vessel for transmission of a high pressuring gas. More specifically, high pressuring gas is not required to push the vapor through the propulsion system. 
     In certain embodiments, the propulsion system may allow for green propellant options. This way, traditional hazardous propellants may not be required. These chemicals may include hydrazine, monomethyl hydrazine, and dinitrogen tetroxide. Hydrazine is a monopropellant currently employed on many satellites. While hydrazine has good performance, it presents a multitude of safety issues to personnel due to its corrosivity, toxicity, and carcinogenic properties. Monomethyl hydrazine (MMH) and dinitrogen tetroxide (N2O4) are often used in satellite bipropellant systems and have similar safety issues. Green propellants often have slightly lower performance but are normally much safer to handle and are normally better for the environment. Green propellants allow for more satellite manufacturers, without the funds, equipment, or skills to handle dangerous chemicals like hydrazine, to enter the market. 
       FIGS. 1A and 1B  are diagrams illustrating a propulsion system  100 A,  100 B, according to an embodiment of the present invention. In some embodiments, propellant system  100 A may include a monopropellant design, as shown in  FIG. 1A . The monopropellant design may include a propellant tank  105 , which includes liquid propellant  115  and natural vapor (hereinafter “vapor”)  120  from liquid propellant  115 . In other embodiments, small satellite propellant system  100 B may include a bipropellant design, as shown in  FIG. 1B . In the bipropellant design, an additional propellant tank  165  may be used. In such an embodiment, one tank would hold an oxidizer and the other tank would hold fuel. Propellant tank  165  in certain embodiments includes second propellant  170  that may be mixed with vapor  120  before passing through membrane  145 . 
     Propellant tank  105  and propellant tank  165  may be printed in plastic or metal in some embodiments. Printing the tank may allow for more complex geometries that might be more difficult or expensive with traditional machining processes. In other embodiments, propellant tank  105  and propellant tank  165  may be constructed using typical machining processes. Propellant tanks  105 ,  165  might also be shaped to fit the available space on the spacecraft or be integrated with other components. Since propellant tanks  105 ,  165  may not hold more than atmospheric pressure, a specific shape may not be a requirement for the small satellite. 
     It should be appreciated that liquid propellant  115  may be stored in a liquid phase at ambient conditions, i.e., for simple, low pressure, and high density tanking. However, in other embodiments, liquid propellant  115  may be stored in cooler or warmer conditions. It might be beneficial to store liquid propellant  115  at cooler conditions to lower the vapor pressure or freeze the propellant for storage when not in use. Heating the propellant might prove beneficial for increasing vapor pressure. 
     For purposes of simplicity, the operation of the monopropellant thrust design will be explained below with the understanding that the bipropellant thrust design would operate in a similar manner. 
     As discussed above, the monopropellant thrust design includes a propellant tank  105 . Propellant tank includes liquid propellant (or in certain embodiments a liquid oxidizer or fuel)  115  and vapor  120  that naturally sits on top of liquid propellant  115 . In some embodiments, liquid propellant  115  may include, but is not limited to, hydrogen peroxide, hydrazine, isopropyl nitrate, nitromethane, nitrous oxide, and any other type of propellant that would be readily appreciated by a person of ordinary skill in the art. Liquid propellant  115  should have a reactive vapor phase. This means that salt-based propellants, such as hydroxylannonium nitrate (HAN) and AF-315, would not work with propulsion system  100 A. 
     Generally, in a traditional propulsion system, pressurant gas pushes the liquid to, for example, a catalyst to create thrust. However, in propulsion system  100 A, rather than pushing liquid to a catalyst, vapor  120  that naturally sits on top of liquid propellant  115  may flow from propellant tank  105  and through propulsion system  100 A. Vapor  120  may react spontaneously on the reacting surface  150  or may require preheating reacting surface  150  before a reaction occurs. Reacting surface  150  in some embodiments may include a catalyst. As vapor  120  flows from propellant tank  105 , liquid propellant  115  may continue to regenerate additional vapor  120  until all liquid propellant  115  is used. 
     It should be appreciated that a vacuum of space downstream of nozzle  160  may cause the flow of vapor  120 . This is due to the pressure difference between propellant tank  105  and space. In some embodiments, the flow of vapor  120  may be viewed as vapor  120  being pulled from propellant tank  105  through propulsion system  100 A for generating thrust. In other embodiments, the flow of vapor  120  may be viewed as vapor  120  being expanded from propulsion system  100 A for generating thrust after interacting with reacting surface  150 . 
     During operation, when a latching valve (or valve)  130  is opened, vapor  120  may begin to flow from propellant tank  105 . Valve  130  in some embodiments is not only used for controlling the flow of vapor  120 , but also used for safety on the launch pad, as well as for starting and stopping propulsion system  100 A while in orbit. 
     To prevent liquid propellant  115  from flowing through propulsion system  100 A, a liquid/vapor separation membrane  125  may be used. Liquid/vapor separation membrane  125  may separate liquid propellant  115  and the vapor  120 . Should liquid propellant  115  flow through the propulsion system, liquid propellant  115  may block the flow of vapor  120  through membrane  145  and to reacting surface  150 . For example, if the liquid were to flow through propulsion system  100 A, and membrane  145  is completely blocked by liquid, negligible thrust would be generated. It should be appreciated that the liquid may not flow successfully through propulsion system  100 A, due to the liquid&#39;s wetting action in membrane  145 . This contact may inhibit flow of vapor  120  through small path areas of membrane  145 . 
     In certain embodiments, small satellite propulsion system  100 A may also include a propulsion management device  110  that guides liquid propellant  115  in one direction and vapor  120  in another direction. 
       FIG. 2A  is a diagram  200  illustrating a behavior of vapor flowing through a membrane and a reacting surface, according to an embodiment of the present invention. More specifically,  FIG. 2A  shows the behavior of vapor downstream from valve  130  of  FIG. 1A . In an embodiment,  FIG. 2A  shows a first chamber  205 , which is the cold side, a membrane  210 , a reacting surface  215 , a second chamber  220 , which is the hot side, and a converging, diverging nozzle  225 . Membrane  210  may include a plurality of small pores or tubes. The diameter for each pore or tube may vary depending on the gas that is being pumped through, for example, propulsion system  100 A. The size of the pore may control the flow rate of vapor to reacting surface  215 . Simply put, the size of the pore may dictate the amount of vapor that flows through membrane  210 . In other words, membrane  210  meters the flow rate of vapor to reacting surface  215 , where vapor is decomposed resulting in hot gas downstream in chamber  220 . In certain embodiments, membrane  210  may be made of a low thermal conductivity material to prevent heat flow back to first chamber  205 . 
     In certain embodiments the membrane may be a thermal transpiration membrane, which would result in an increase in pressure of the vapor. See, for example,  FIG. 2B . The thermal transpiration membrane would in some embodiments have pore diameters on the order of the gas mean free path. 
     Referring to  FIG. 1A , for example, when propulsion system  100 A is activated, the flow of vapor  120  through membrane  145  may be very slow, because there is not yet a temperature gradient within membrane  145 . This may not occur in embodiments that do not use a thermal transpiration membrane. For example, in order for a thermal transpiration membrane to pump gas, a temperature gradient may be required in some embodiments. In those embodiments, the temperature gradient is established with vapor  120  reacting on the reacting surface. 
     As shown in  FIG. 2B , temperature may begin to increase due to reaction between the vapor and reacting surface  215 . When this occurs, the flow rate of vapor through transpiration membrane  210  begins to increase due to thermal transpiration induced pumping and becomes steady. The steady flow of vapor through transpiration membrane  210  and the subsequent reaction on reacting surface  215  produces hot decomposition gases in second chamber  220 . The hot decomposition of gases can then be forced out a nozzle, as shown in  FIG. 2A , to generate thrust. It should be appreciated that unless valve  130  of  FIG. 1A  is closed, the flow of vapor  120  may continue to exist due to the vacuum near nozzle  160 . 
     Using  FIG. 1A  as an example, if a higher flowrate is desired, valve  130  may be used instead of a membrane to meter the flowrate of vapor. In such an embodiment, membrane  145  of  FIG. 1A  may be removed and vapor  120  may interact directly with reacting surface  150 . This might be suited for short pulses and higher vapor pressure propellants. 
     In certain embodiments, thermal management may be necessary to regulate heat transfer through the system. In those embodiments, thermal guard  140  may be placed between membrane  145  and the cold propellant gas (or vapor) in chamber  135 . Thermal guard  140  may keep the propellant gas in chamber  135  thermally separated from the hot reacting surface  150 . Thermal guard  140  may also be useful for maintaining a larger change in temperature, which is beneficial for embodiments utilizing a thermal transpiration membrane. Thermal guard  140  may be made of aluminum in some embodiments. To allow the propellant to flow easily through thermal guard  140 , thermal guard  140  may be perforated or made of foam, for example. 
     To maintain a minimum temperature of the liquid propellant  115 , a thermal connector  165  may be used to transfer heat from the hot side of the system, i.e., from chamber  155 , back to propellant tank  105 . Although  FIG. 1A  shows a thermal connector  165  connecting the hot side of the system to propellant tank  105 , one of ordinary skill in the art would readily appreciate that other configurations and/or techniques may be used to transfer heat. This temperature regulation may be used to control the pressure of vapor  120 . It should be appreciated that chamber  135  may be referred to as the “cold chamber” and chamber  155  may be referred to as “hot chamber”. 
       FIG. 3  is a graph  300  illustrating a theoretical performance comparison with other currently available small satellites propulsion systems and this system utilizing hydrogen peroxide and a thermal transpiration membrane, according to an embodiment of the present invention. Graph  300  shows a thrust over power performance comparison for propulsion systems with thrust/power&gt;1. Therefore, extremely low thrust, high power systems are not shown since the high power systems are outside the bounds of the chart. The high power systems would mostly include electric systems. As can be seen, propellant system described herein can generate a much higher thrust/power than most available systems for small satellites, and also has the flexibility to perform within their ranges if desired. 
       FIG. 4  is a graph  400  illustrating a theoretical orbital velocity change and burn time based on hydrogen peroxide loaded for a specific thermal transpiration propulsion system, according to an embodiment of the present invention. The amount of propellant loaded, and therefore, the mission orbital velocity change, can be altered as it is not fixed for some of the embodiments described herein. This allows for more mission control by the satellite manufacturer. 
       FIG. 5  is a graph  500  illustrating theoretical thrust and Knudsen number based on membrane pore size in a small satellite propulsion system, according to an embodiment of the present invention. Graph  500  also shows the thrust benefits from using pumping membrane verses a non-pumping membrane of the same pore size (thermal transpiration verses non-thermal transpiration). The dotted line represents the membrane pore size generating the maximum difference in thrust between a non-pumping membrane and a pumping membrane. 
       FIG. 6  is a graph  600  illustrating a theoretical thrust and flowrate based on membrane area, according to an embodiment of the present invention. In this illustration, graph  600  shows that as the area of the membrane is increased, more vapor is allowed to flow through the system. A higher flowrate can lead to higher thrust in certain embodiments. 
     One or more embodiments generally discuss a novel approach to generating thrust by using vapor that naturally sits on top of liquid propellant. In some embodiments, the liquid propellant may be stored in a liquid phase. However, in other embodiments, the liquid propellant may be stored in a frozen state, and may be heated at a later time or allowed to be thawed or sublime over time. 
     The propulsion system may enable a small satellite, such as a CubeSat, to perform orbit changes, extend the life of the mission, perform rendezvous, perform formation flying, and any other movement once placed in orbit. The propulsion system may have low pressure, low power, and low complexity. The propulsion system may also work with a variety of propellants that have a reactive vapor phase and may be stored in a liquid or solid state. 
     The propulsion system may include a transpiration membrane having a pore size that can be manipulated to provide a range of thrust options. Multiple stages of thermal transpiration may lead to a higher chamber pressure, leading to a higher thrust and performance The propulsion system may incorporate solid oxide fuel cells (SOFC) that can lead to power generation. SOFCs may cause the system to generate small amounts of power while running, which can be stored for later use. The SOFC(s) in some embodiments would be placed in the propellant(s) vapor path. 
     It will be readily understood that the components of various embodiments of the present invention, as generally described and illustrated in the figures herein, may be arranged and designed in a wide variety of different configurations. Thus, the detailed description of the embodiments of the systems, apparatuses, methods, and computer programs of the present invention, as represented in the attached figures, is not intended to limit the scope of the invention as claimed, but is merely representative of selected embodiments of the invention. 
     The features, structures, or characteristics of the invention described throughout this specification may be combined in any suitable manner in one or more embodiments. For example, reference throughout this specification to “certain embodiments,” “some embodiments,” or similar language means that a particular feature, structure, or characteristic described in connection with the embodiment is included in at least one embodiment of the present invention. Thus, appearances of the phrases “in certain embodiments,” “in some embodiment,” “in other embodiments,” or similar language throughout this specification do not necessarily all refer to the same group of embodiments and the described features, structures, or characteristics may be combined in any suitable manner in one or more embodiments. 
     It should be noted that reference throughout this specification to features, advantages, or similar language does not imply that all of the features and advantages that may be realized with the present invention should be or are in any single embodiment of the invention. Rather, language referring to the features and advantages is understood to mean that a specific feature, advantage, or characteristic described in connection with an embodiment is included in at least one embodiment of the present invention. Thus, discussion of the features and advantages, and similar language, throughout this specification may, but do not necessarily, refer to the same embodiment. 
     Furthermore, the described features, advantages, and characteristics of the invention may be combined in any suitable manner in one or more embodiments. One skilled in the relevant art will recognize that the invention can be practiced without one or more of the specific features or advantages of a particular embodiment. In other instances, additional features and advantages may be recognized in certain embodiments that may not be present in all embodiments of the invention. 
     One having ordinary skill in the art will readily understand that the invention as discussed above may be practiced with steps in a different order, and/or with hardware elements in configurations which are different than those which are disclosed. Therefore, although the invention has been described based upon these preferred embodiments, it would be apparent to those of skill in the art that certain modifications, variations, and alternative constructions would be apparent, while remaining within the spirit and scope of the invention. In order to determine the metes and bounds of the invention, therefore, reference should be made to the appended claims.