Miniaturized green end-burning hybrid propulsion system for CubeSats

A hybrid propulsion system includes a housing, at least two electrodes, a solid-grain fuel material, a combustion chamber, an oxidizer port, and a nozzle. The housing has a first end and a second end and defines a cavity. The electrodes extend into the cavity. The fuel material is free of an oxidizer and is positioned in the cavity. The fuel material has a combustion surface and is exposed to the electrodes. The combustion chamber is defined between the combustion surface and the second end. The oxidizer port provides a flow of oxidizer to the combustion chamber. The nozzle is positioned at the second end. Combustion of the fuel material in the combustion chamber may be dominated by radiative heat transfer. Combustion of the fuel material in the combustion chamber may generate thrust of no more than 5 N at an oxidizer flow rate of no more than 5 g/s.

TECHNICAL FIELD

The present invention relates generally to hybrid rocket systems and, more specifically, to devices, systems and methods of an ignition and combustion portion of a hybrid rocket system, and particularly a small-scale hybrid rocket system.

BACKGROUND

The current state of the art for hybrid rocket ignition systems is largely based on pyrotechnic ignition methods. These methods have serious shortcomings including the inability to initiate multiple restarts using a single device, thus, limiting the applicability of the hybrid rocket. Other shortcomings include significant physical and environmental hazards. For example, making rockets safer, less toxic, and less explosive comes at a significant cost. As the propellant materials become less volatile, they also become increasingly difficult to ignite. Combustion of hybrid propellant must be initiated by an igniter that provides sufficient heat to cause pyrolysis of the solid fuel grain at the head end of the motor, while simultaneously providing sufficient residual energy to overcome the combustion activation energy to initiate combustion. Thus, hybrid rockets have typically used large, high output pyrotechnic charges to initiate combustion. Such igniters are capable of producing very high-enthalpy outputs, but are extremely susceptible to hazards of electromagnetic radiation and present significant operational hazards. Most importantly, such pyrotechnic igniters are designed as “one-shot” devices that do not allow multiple restart capability.

Due to the lack of a reliable on-demand ignition system, hybrid rockets have never been seriously considered as a viable alternative for in-space propulsion. Advancements related to3D printable plastics as alternatives to legacy solid rocket binders like HTPB has made it possible to manufacture a structural matrix with unique electrical breakdown properties. This discovery has allowed for the development of a unique on-demand ignition technology for hybrid rockets.

In the current CubeSat market, there does not exist a commercial off the shelf (COTS) propulsion system with flight heritage that has thrust levels greater than tens of millinewtons or specific impulse (Isp) levels greater than 70 seconds. Electric propulsion systems provide very high Isp but low thrust levels. The kinetic power per unit thrust surrounding electric propulsion systems are typically in the range of 10-100 W/mN3depending on selected type. These large power requirements, in addition to the low thrust of electric propulsion, either limit the mission CONOPS or extend the mission timeframe incurring addition financial expenditures. Cold gas propulsion systems are the opposite of electric systems in terms of performance with Isp levels less than 70 seconds. The low Isp of cold gas systems directly results in large form factors with poor volumetric performance.

The majority of hybrid rockets developed to date are traditional core-burning designs with significantly higher thrust levels than desired for small spacecraft. These designs also require a high length-to-diameter ratio, resulting in a form factor difficult to use in CubeSats or small satellites. Accordingly, there is a technology gap in the current CubeSat market. The gap exists in the “high” thrust and “moderate” specific impulse category. This category is traditionally filled by mono- or bi-propellants.

The technology gap in the CubeSat propulsion market leads to mission limitations including spacecraft attitude control difficulty due to centrally located thrusters, low performing cold gas systems requiring large tanks, and electric propulsion systems with large power requirements. These limitations make several mission types extremely difficult or, in some cases, unsupportable. Such missions include rendezvous, proximity operations, formation flying, and cluster management. A successful demonstration of a system in this technology gap would lead to a compact, highly efficient, and market disrupting propulsion system that that would enable CubeSats for use in critical and far-reaching missions.

Dedicated launches of Smallsats have increased recently, thus creating an industry need for a cost effective and high-performance propulsion system. Many satellite developers continue to use multi-million-dollar propulsion systems based on a propellant known as hydrazine. The hybrid propulsion system in development at the Space Dynamics Laboratory aims to address this prohibitive cost requirement to allow more satellite developers an opportunity to incorporate propulsion into their design. Current efforts are being made by AFRL, NASA, and ECAPS to develop non-hydrazine “green” propellants. However, the costs are nearly identical to conventional hydrazine systems. The cost associated with these propulsion systems has led to many Smallsats launching with no propulsion; often restricting the mission CONOPS. In addition to the costs surrounding hydrazine and other “green” propellants are the hazards they pose to personnel.

Hybrid rockets typically consist of moderately benign gaseous or liquid oxidizer and an inert solid fuel. Hybrid rockets possess operational safety and handling advantages when compared to liquid and solid propellent systems. The U.S. Department of Transportation concluded that most hybrid rocket motor designs could be safely stored and operated without risk of explosion or detonation, potentially offering significantly lower operating and integration cost. The inherent design safety of hybrid rockets increases the potential for rideshare opportunities when compared to traditional monopropellant systems.

In order for hybrid rockets to fill the technology gap discussed in the previous section there are two key disadvantages that must be addressed. The first is the difficulty surrounding motor ignition. A key reason why hybrid rockets are considered to be safe is due to the stability of their fuel. This stability makes hybrid rockets difficult to ignite. The hybrid rocket ignition source must provide sufficient heat to vaporize the solid fuel grain at the head end of the motor while simultaneously providing sufficient residual energy to overcome the activation energy of the oxidizers. Hybrid rockets typically use a pyrotechnic device called a “squib” that ignites a secondary solid fuel motor to initialize the combustion of the hybrid motor. These squibs are often susceptible to the hazards of electromagnetic radiation and can present a potential explosion hazard. Also, pyrotechnic ignitors can only be fired once, thus limiting their application as an in-space propulsive device.

The second disadvantage is that the internal motor ballistics of hybrid combustion produce regression rates that are typically 25-30% lower when compared to solid fuel motors of the same thrust and impulse. To make up for the lower regression rate, a higher oxidizer flow rate is required to maintain the same thrust level. This increases the systems oxidizer-to-fuel ratio (O/F) and ultimately results in poor mass impulse performance, erosive fuel burning, nozzle erosion, reduced motor duty cycles, potential combustion instability, and poorer overall performance of the system. To overcome this O/F problem, hybrid rockets are traditionally designed with cylindrical fuel ports that have long length-to-diameter ratios. This high aspect ratio can result in poor volumetric efficiency and thus limit a hybrid motor's application to a customarily volume constrained small satellite.

BRIEF SUMMARY OF THE INVENTION

The inventors of the present disclosure have identified that it would be advantageous to provide a hybrid rocket ignition system that has restart capability that also is safe, less toxic, and less explosive than the current state-of-the-art rocket systems.

Embodiments of the present invention are directed to various devices, systems and methods of providing a restartable ignition device for a hybrid rocket system. For example, in one embodiment, an ignition device includes a housing and at least two electrodes. The housing includes a first side and a second side and defines a bore with an axis extending therethrough between the first and second sides, the bore defining an internal surface of the housing. The at least two electrodes extend through the housing to the internal surface. The at least two electrodes are configured to be spaced apart so as to provide an electrical potential field along the internal surface between the at least two electrodes. Such housing is formed with and includes multiple flat layers such that the multiple flat layers provide ridges along the internal surface. With this arrangement, the internal surface with the ridges are configured to concentrate an electrical charge upon being subjected to the electrical potential field.

In one embodiment, the ridges, under the electrical potential field, act as miniature electrodes to arc the electrical charge. In another embodiment, the internal surface includes grooves, each groove extending between two adjacently extending ridges. In still another embodiment, each of the ridges are a periphery of each of the multiple flat layers.

In another embodiment, the multiple flat layers each define a plane oriented transverse relative to the axis of the bore. In another embodiment, the at least two electrodes define a line therebetween, the line being generally parallel with a plane defined by each of the multiple flat layers.

In another embodiment, the internal surface defines a step configuration such that the bore at the first side is larger than the bore at the second side, the step configuration exhibiting a shelf extending to a shelf notch, the shelf notch having the at least two electrodes. In another embodiment, the bore at the first side defines a first width and the bore at the second side defines a second width, the first width greater than the second width. In yet another embodiment, the bore includes a convergent configuration extending from the first side to the second side.

In another embodiment, the multiple layers are an Acrylonitrile Butadiene Styrene (ABS) material. Such multiple layers may be formed with a fused deposition modeling process or three-dimensional printing.

In accordance with another embodiment of the present invention, a method of forming an ignition device, is provided. The method includes: forming a housing with multiple flat layers, the housing having a first side and a second side defining a bore with an axis extending through the housing and between the first and second sides such that the bore is defined by an internal surface of the housing; and positioning at least two electrodes to extend through the housing to the internal surface such that the at least two electrodes are spaced and configured to provide an electrical potential field along the internal surface between the at least two electrodes; wherein the forming the housing with multiple flat layers step includes forming the internal surface to include ridges, the ridges along the internal surface being configured to concentrate an electrical charge upon being subjected to the electrical potential field.

In one embodiment, the method step of forming the internal surface to include ridges includes the step of forming multiple miniature electrodes configured to arc the electrical charge between the at least two electrodes. In another embodiment, the method step of forming the housing with multiple flat layers includes the step of forming the multiple flat layers in a plane oriented transverse relative to the axis of the bore. In still another embodiment, the method step of positioning the at least two electrodes includes the step of positioning the at least two electrodes to define a line therebetween such that the line is generally parallel with a plane defined by each of the multiple flat layers.

In another embodiment, the method step of forming the housing with multiple flat layers includes the step of forming the bore of the housing to include a step configuration to define a shelf. In another embodiment, the method step of forming the housing with multiple flat layers includes forming the bore of the housing to include a convergent configuration extending from the first side to the second side of the housing.

In another embodiment, the method step of forming the housing with multiple flat layers includes the step of forming the housing with layers of a solid grain fuel material. In another embodiment, the method step of forming the housing with multiple flat layers includes forming the housing with layers of an Acrylonitrile Butadiene Styrene (ABS) material.

In accordance with another embodiment of the present invention, a hybrid rocket system is provided. The hybrid rocket system includes a container, an ignition portion, a solid-grain combustion portion, a post combustion portion, and a nozzle. The container is sized to contain liquid or gaseous fuel. The ignition portion includes a first side and a second side, the ignition portion defining a bore with an axis extending therethrough between the first and second sides, the bore defined by an internal surface and the bore configured to receive the fuel from the container. The ignition portion includes at least two electrodes configured to provide an electrical potential field along the internal surface between the at least two electrodes. The solid grain combustion portion defines a combustion chamber such that the combustion chamber corresponds with the bore of the ignition portion. The post combustion portion is coupled to the solid grain combustion portion. The nozzle is coupled to the post combustion portion and is configured to manipulate thrust to the rocket system. The ignition portion is formed with and includes multiple flat layers such that the multiple flat layers provide ridges along the internal surface. With this arrangement, the internal surface with the ridges are configured to concentrate an electrical charge generally between the at least two electrodes upon being subjected to the electrical potential field.

In one embodiment, the multiple flat layers each define a plane oriented transverse relative to the axis of the bore.

Another embodiment relates to a hybrid propulsion system that includes a housing, at least two electrodes, a solid-grain fuel material, a combustion chamber, an oxidizer port, and a nozzle. The housing has a first end and a second end and defines a cavity. The at least two electrodes extend into the cavity. The solid-grain fuel material is free of an oxidizer and is positioned in the cavity. The fuel material also has a combustion surface and is exposed to the at least two electrodes. The combustion chamber is defined between the combustion surface and the second end. The oxidizer port is arranged to provide a flow of oxidizer to the combustion chamber. The nozzle is positioned at the second end. Combustion of the fuel material in the combustion chamber is dominated by radiative heat transfer.

The hybrid propulsion system may generate no more than about 5 N of thrust and may have oxidizer flow of no more than about 5 g/s. The fuel material may include a plurality of flat layers that provide ridges along the combustion surface, the electrodes may be configured to concentrate an electrical charge on the ridges, which may act as micro-electrodes that produce localized electrical arcing thereon and ignite the combustion surface of the fuel material. The fuel material at the combustion surface may be initially consumed or removed through combustion of the fuel material, a newly exposed internal surface of the fuel material may have newly exposed ridges that act as newly exposed micro-electrodes that produce localized electrical arcing thereon and may re-ignite the newly exposed combustion surface. The fuel material may include a plurality of flat layers formed by an additive manufacturing process. The oxidizer may be gaseous oxygen (GOX), hydrogen peroxide (H2O2), or nitrous oxide (N2O). The fuel material may include at least one of Acrylonitrile Butadiene Styrene (ABS), Polymethyl methacrylate (PMMA), Polyvinyl Chloride (PVC), and Nylon-12. The hybrid propulsion system may generate thrust in the range of about 0.1 N to about 1 N. The hybrid propulsion system may use no more than about 1 g/s oxidizer.

In another embodiments, a hybrid propulsion system includes a housing having a first end and a second end and defining a cavity, at least two electrodes extending into the cavity, a solid-grain fuel material free of an oxidizer and positioned in the cavity and exposed to the at least two electrodes, the fuel material having a combustion surface, a combustion chamber defined between the combustion surface and the second end, and an oxidizer port arranged to provide a flow of oxidizer to the combustion chamber, a nozzle positioned at the second end. Combustion of the fuel material in the combustion chamber generates thrust of no more than about 5 N at an oxidizer flow rate of no more than about 5 g/s.

The combustion may be dominated by radiative heat transfer. The fuel material may include a plurality of flat layers that provide ridges along the combustion surface, and the electrodes may be configured to concentrate an electrical charge on the ridges, which act as micro-electrodes that produce localized electrical arcing thereon and ignite the combustion surface of the fuel material. The fuel material at the combustion surface may be initially consumed or removed through combustion of the fuel material, and a newly exposed internal surface of the fuel material may have newly exposed ridges that act as newly exposed micro-electrodes that produce localized electrical arcing thereon and re-ignite the newly exposed combustion surface. The fuel material may include a plurality of flat layers formed by an additive manufacturing process. The oxidizer may be gaseous oxygen (GOX), hydrogen peroxide (H2O2), or nitrous oxide (N2O). The fuel material may include at least one of Acrylonitrile Butadiene Styrene (ABS), Polymethyl methacrylate (PMMA), Polyvinyl Chloride (PVC), and Nylon-12. The hybrid propulsion system may generate thrust in the range of about 0.1 N to about 5 N. The hybrid propulsion system may use no more than about 5 g/s oxidizer.

In a further embodiment, a method of operating a low-thrust hybrid propulsion system includes providing a housing having first and second ends, a solid-grain fuel material positioned in the housing and having a combustion surface and being free of an oxidizer, at least two electrodes positioned in the housing, a combustion chamber defined between the combustion surface and the second end, an oxidizer port, a nozzle positioned at the second end, delivering a flow of oxidizer through the oxidizer port and into the combustion chamber, and igniting the combustion surface with the at least two electrodes to generate a hot-gas, fuel-rich flow through the nozzle to generate thrust. Combustion of the fuel material in the combustion chamber is dominated by radiative heat transfer.

The hybrid propulsion system may generate thrust in the range of about 0.1 N to about 5 N. The hybrid propulsion system may generate no more than 5 N of thrust and has oxidizer flow of no more than 5 g/s. The fuel material may include a plurality of flat layers that provide ridges along the combustion surface, and the electrodes may be configured to concentrate an electrical charge on the ridges, which act as micro-electrodes that produce localized electrical arcing thereon and ignite the combustion surface of the fuel material. The fuel material at the combustion surface may be initially consumed or removed through combustion of the fuel material, and a newly exposed internal surface of the fuel material may have newly exposed ridges that act as newly exposed micro-electrodes that produce localized electrical arcing thereon and re-ignite the newly exposed combustion surface.

DETAILED DESCRIPTION OF THE INVENTION

Referring toFIGS. 1 and 2, a simplified view of a hybrid rocket system10or motor that includes an ignition system12or pre-combustion portion, according to the present invention, is provided. Referring toFIGS. 3A, 3B, and 6, in one embodiment, the ignition system12or pre-combustion portion may include a housing20formed of multiple flat layers18by employing fused deposition modeling (FDM) or three-dimensional printing. Such FDM process provides an internal surface14with ridges16formed from the multiple flat layers18deposited upon each other (SeeFIG. 6). The ignition system12may also include electrodes86and88spaced from each other and positioned adjacent the internal surface14. Upon an oxidizer or oxidizer being injected into the system and activating an electrical potential field between the electrodes86and88, the ridges16along the internal surface14may concentrate an electrical charge which seeds combustion of the solid-grain fuel material.

As will be described herein, the unique structural characteristics of the material and structure of the internal surface14and housing20provide an ignition system12that is restartable. For example, multiple re-starts have been implemented with the ignition system12set forth herein. The inventors have found that the only limitation to the number of allowable restarts is the quantity of solid fuel grain material contained within the ignition system12. Such ignition system12may require small input energy and may use only non-toxic and non-explosive with the simplicity and reliability of a monopropellant system, but with the output enthalpy equivalent to a bi-propellant igniter. As such, the restartable ignition system12may have applicability to military aircraft, missile systems for post-stall maneuvering, emergency gas generation cycles, and many other applications relating to systems that may benefit from the restartable ignition system.

With reference toFIG. 1, the basic components of the hybrid rocket system10may include a gaseous or liquid oxidizer container22or tank, a combustion portion24, and a nozzle26. The gaseous oxidizer may be nitrous oxide or gaseous oxygen or any other suitable gaseous or liquid oxidizer. The gaseous or liquid oxidizer container22may be disposed between a gas pressurization element28and a gas feed system30. The gas feed system30may feed an injector portion32, which in turn controllably injects oxidizer into the combustion portion24of the hybrid rocket system10. The combustion portion24of the system may include multiple portions, such as, the ignition system12or pre-combustion portion, a main combustion portion34, and a post combustion portion36. The main combustion portion34may be formed of one or more solid-grain fuels, such as acrylic or hydroxyl-terminated polybutadiene (HTPB), or any other suitable solid fuel grain known in the art. In one embodiment, the solid fuel grain for the main combustion portion34and post combustion portion36may be acrylonitrile butadiene styrene (ABS) or combinations of other known solid fuels. The combustion portion24and, more particularly, the post combustion portion36may be coupled to the nozzle26or other similar structure. The nozzle26may include various nozzle configurations, depending upon the application of a particular rocket system or the like. With this arrangement, the ignition system12of the present invention may be employed with the other components of the hybrid rocket system10to facilitate multiple restarts with one device, i.e., without replacing parts.

Now with reference toFIGS. 2, 3 and 3A, various views of an ignition system12or pre-combustion portion are provided. As set forth, the ignition system12or pre-combustion portion may be directly coupled to the main combustion portion34. The ignition system12, as depicted in the illustrated example, is directly coupled to a shortened minimal portion of the main combustion portion34. More important to this description is that the ignition system12or pre-combustion portion may include the housing20and first and second electrode components40,42.

In one embodiment, the housing20may include a sleeve like structure with various ports and notches therein and further, the sleeve like structure may include the internal surface14with a step configuration. For example, the housing20may include a first side44and a second side46with a bore48extending through and between the first and second sides44,46of the housing20. The second side46is illustrated as an interface surface between the housing20and main combustion portion34. The bore48may define a centrally extending axis50along a length52of the housing20. Further, the housing20may include an external surface54and the before mentioned internal surface14. The external surface54may include cylindrical shape or any another suitable structure.

The internal surface14may define the bore48of the housing20, the bore48defining a radial component such that a cross-section of the bore48may be defined as generally circular or any other suitable structure. Further, as set forth, the internal surface14may define a step configuration so as to include a shelf56. In this manner, the bore48may include a first radius58and a second radius60, the first radius58and the second radius60extending laterally from the axis50to the internal surface14of the housing20. Such first radius58may extend along the length of the bore48from the first side44of the housing20to the shelf56. The second radius60may extend along the length from the shelf56to the second side46of the housing20. With this arrangement, the first radius58may be larger than the second radius60such that the bore48exhibits a larger opening on the first side44of the housing20than on the second side46of the housing20.

With respect toFIGS. 2, 3, 3A, and 4, as set forth, the housing20may include various ports and/or notches therein. For example, in one embodiment, the bore48of the housing20may also include a first notch62and a second notch64, each defined by the internal surface14. The first and the second notches62,64may be positioned on opposite sides of the bore48so as to face each other. Each of the first and second notches62,64may extend between the shelf56and the first side44of the housing20such that the shelf56extends further at the notch to define a third radius66or a third dimension, the third radius66or dimension being larger than the first radius58and being defined from the axis50to the internal surface14at the first and second notches62,64. At least one of the first and second notches62,64may be sized and configured to exhibit electrodes86,88at, for example, base corners of the at least one of the first and second notches62,64and adjacent the shelf56, discussed in further detail herein.

Further, the housing20may include one or more ports for the electrode components. For example, the housing20may include a first port68and a second port70. The first and second ports68,70may be positioned opposite each other on the first side44of the housing20. The first port68may define a first port cavity72(shown in outline form) extending from the first port68to a first port outlet74. The first port outlet74may be disposed at a first base corner76of the first notch62on the shelf56and adjacent to the internal surface14having the third radius66. Similarly, the second port70may extend with a second port cavity78to a second port outlet80at a second base corner82of the first notch62on the shelf56. In this manner, the first port outlet74and the second port outlet80of the first notch62may be disposed at opposite first and second base corners76,82of the first notch62. A similar arrangement may be employed for the second notch64defining first and second outlets of port cavities extending to the first and second ports. In this manner, the ports and cavities extending to the first notch and/or the second notch may be sized and configured for positioning electrodes86,88of the first and second electrode components40,42. In another embodiment, one or both of the notches,62or64, or other port may include a pressure sensor configured to measure the pressure of the oxidizer at the shelf56.

With respect toFIGS. 3A and 4, as set forth, the ignition system12or pre-combustion portion includes first and second electrode components40,42. The first and second electrode components40,42may each include at least a conductive electrical wire that serves as an electrode at the end of the wire. Such electrode components may be embedded and positioned within the first and second ports68,70so that respective first and second electrodes86,88are exposed within the bore48and, more particularly at the first and second port outlets74,80defined in, for example, the first notch62. Within the bore48, the first and second electrodes86,88may be spaced a distance from each other so that, upon being electrically activated, the first and second electrodes86,88provide a voltage potential or an electrical field potential adjacent the internal surface14between the first and second electrodes86,88. As depicted, such distance or spacing between the first and second electrodes86,88may be defined by the first and second base corners76,82in, for example, the first notch62in the bore48. Further, the first and second electrodes86,88may be exposed at and flush with the internal surface14of the bore48. In another embodiment, the first and second electrodes86,88may protrude from the internal surface14of the bore48. Similar to that set forth above, another set of first and second electrodes86,88may be positioned and spaced at the second notch64.

As set forth, the housing20and bore48of this embodiment may include a step configuration to define the shelf56. The shelf56may be sized and configured to act as an impingement to the oxidizer or an impingement shelf to slow the oxidizer from moving down stream so as to increase the pressure of the oxidizer at the shelf56. The increase in pressure of the oxidizer at the shelf56may provide sufficient oxidizer for a combustion reaction of a solid fuel grain material on the internal surface14. Suitable oxidizers may include gaseous oxygen, liquid oxygen, nitrous oxide, hydrogen peroxide, hydroxylammonium nitrate, ammonium dinitramide, or air. The oxidizer pressure increase at the impingement shelf56may enable the first and second electrodes86,88to be minimally spaced (or minimally charged) to provide a charge concentration or voltage potential on the internal surface14of the bore48between the first and second electrodes86,88.

With respect toFIGS. 3B, 5, and 6, the housing20of the ignition system12may be formed from a solid-grain fuel material. In one embodiment, the solid-grain fuel material may be high or low density Acrylonitrile Butadiene Styrene (ABS) or any other suitable solid-grain fuel material that holds similar electro-mechanical, combustion, and structural properties. As set forth, the housing20may be formed with multiple flat layers18deposited upon each other, employing the Fused Deposition Modeling (FDM) method or three-dimensional printing or any other suitable process for layering a fuel grain. Upon employing the FDM method, ABS possesses a very unique electro-mechanical property such that additive manufacturing results in a distinctive surface structure that is different than the surface of a monolithically fabricated (e.g., a molded or machined) ABS structure. In particular, this surface structure, such as the internal surface14defining the bore48, is the surface structure that is transverse to a plane defined by any one of the multiple flat layers18. Such surface structure or internal surface14has the effect of concentrating electrical charges locally when the surface14of the ABS material is subjected to an electrical potential field. These high-charge concentrations produce localized electrical arcing such that the ABS material breaks down at voltages significantly lower than that of a monolithically fabricated ABS structure. Described another way, the voltage potential created between the first and second electrodes86,88, when electrically activated, causes the unique features (the ridges16formed in the multiple flat layers18shown inFIG. 6) of the surface14to act as micro-electrodes which ignites the solid-grain fuel material in the presence of an oxidizer.

In one embodiment, the multiple flat layers18may be deposited so that any one of the flat layers18define a plane that is transverse or perpendicular with the axis50of the housing20. In another embodiment, the first and second electrodes86,88(seeFIG. 4) may define a line therebetween that may be generally parallel with a plane defined by each of the multiple flat layers18. In still another embodiment, each of the flat layers18may define a plane that is substantially parallel with the axis50of the housing20. In any one of these embodiments, the multiple flat layers18, deposited upon each other, form the internal surface14with ridges16or ridged layering. The ridges16or ridged layering may be defined by peripheral ends90of the multiple flat layers18. As set forth, the unique mechanical structure (e.g., the surface characteristics created by the FMD layering) of the ridges16and multiple flat layers18, in conjunction of the material being a solid-grain fuel, such as ABS material, act as multiple micro-electrodes when subjected to an electrical potential field. Such unique mechanical structure facilitates the ignition system12to implement multiple restarts. For example, even as material from the internal surface14is initially consumed or removed through combustion, a newly exposed internal surface14maintains similar surface characteristics or surface roughness that act as micro-electrodes when exposed to an electrical potential field.

With respect toFIG. 6, an enlarged view of the multiple flat layers18and ridges of the fuel grain material are depicted. As set forth, the internal surface14defines ridges or ridged layering formed between each of the multiple flat layers18. Each of the flat layers18may include a peak92with a small radius at its peripheral end such that the structure may also include a slope extending to the peak that may be substantially linear or radial. Although depicted as uniform ridges16, such ridges may not be uniform along the internal surface14of the housing20. In this manner, the internal surface14may exhibit a rough, coarse or scratched surface. The ridges may exhibit a nodal configuration or exhibit a protruding structure that may continue or discontinue along the peripheral end90of each of the multiple flat layers18. Likewise, the internal surface14may exhibit grooves94formed between each of the multiple flat layers18. In other words, each groove94extends between adjacently extending ridges16. With this arrangement, the FDM technique of forming the housing, preferably with ABS material, provides for a unique electro-mechanical structure such that the flat layers18that exhibit the ridges16and/or grooves94therein reacts to an electrical potential field. In this manner, the structure and material itself act as multiple micro-electrodes, thereby, facilitating electrical breakdown to facilitate a restartable ignition for a hybrid rocket system.

With respect toFIGS. 7, 8, and 8A, another embodiment of an ignition system110for a hybrid rocket system10(FIG. 1) is provided. This embodiment is similar to the previous embodiment, except this embodiment exhibits a bore112, defined by an internal surface114, with a convergent or conical configuration. For example, the ignition system110may include a housing116and first and second electrode components118,119. The housing116may include a first side120and a second side122with the bore112extending through and between the first and second sides120,122. The bore112may define a centrally located axis124extending along the length of the housing116. The housing116may include first and second electrode ports126,128that may extend from the first side120to a convergent portion of the bore112so that a first and second electrode130,132may be exposed within the bore112. The housing116may also include a pressure port134with a corresponding pressure sensor136so that a pressure within the bore112may be determined upon receiving the oxidizer. Similar to that described and depicted inFIG. 6of the previous embodiment, the housing116of this embodiment may be formed with multiple flat layers18that exhibit a roughened surface or ridges16that provide the before-discussed unique structural characteristic along the internal surface114of the conical bore112. In this manner, upon the first and second electrodes130,132being activated to provide an electrical potential field, the multiple flat layers18deposited upon each other and exhibiting the ridges16and/or grooves94react and concentrate a charge, thereby, acting as multiple micro-electrodes at the internal surface114of the bore112.

As set forth in this embodiment, the bore112in the housing116is convergent. The bore112may be sized and configured to converge so as to increase the pressure of the oxidizer as it moves downstream through the bore112. The increase in pressure of the oxidizer as it moves downstream through the bore112may provide sufficient oxidizer for a combustion reaction of a solid fuel-grain material on the internal surface114. Suitable oxidizers may include gaseous oxygen, liquid oxygen, nitrous oxide, hydrogen peroxide, hydroxylammonium nitrate, ammonium dinitramide, or air. The oxidizer pressure increase at the narrower portion of the bore112may enable the first and second electrodes130,132to be minimally spaced (or minimally charged) to provide a charge concentration or voltage potential on the internal surface114of the convergent portion of the bore112between the first and second electrodes130,132.

Similar to previous embodiments, the multiple flat layers18, deposited upon each other, form the internal surface114with ridges16or ridged layering. The unique mechanical structure (e.g., the surface characteristics created by the FMD layering) of the ridges16and multiple flat layers18, in conjunction of the material being a solid-grain fuel, such as ABS material, act as multiple micro-electrodes on the internal surface114when subjected to an electrical potential field. Such unique mechanical structure facilitates the ignition system116to implement multiple restarts. For example, even as material from the internal surface114is initially consumed or removed through combustion, a newly exposed internal surface114maintains similar surface characteristics or surface roughness that act as micro-electrodes when exposed to an electrical potential field from charged electrodes130,132.

Conventional hybrid rocket motors with thrust levels greater than 5 N rely on forced convection within the boundary layer as the primary heat transfer mechanism for fuel regression. Because of convective heat transfer, the rate of fuel regression is proportional to the oxidizer mass flux through the fuel port. As the fuel port burns radially outwards along the motors longitudinal axis, the rate of fuel regression results from oxidizer diffusion into the combustion layer between the inward blowing vaporized fuel and the axially flowing oxidizer. For a constant oxidizer mass flow, the associated mass flux and the fuel regression rate decrease with burn time. As the fuel port opens up, more surface burn area is exposed, and the resulting oxidizer-to-fuel ratio shifts during the burn from a balance between the fuel regression rate drop, and the increased surface burn area. Generally, the effects of forced convection tend to make hybrid motors burn from rich-to-lean over the burn lifetime.

For Hybrid rockets with thrust levels less than 5 N, oxidizer mass flow levels are sufficiently small that the rate of convective heat transfer is significantly reduced, and radiative heat transfer dominates the fuel regression mechanism.FIG. 9shows how hybrid rockets with lower oxidizer mass flow levels are dominated by radiative heat transfer as compared to higher oxidizer mass flow levels being dominated by convective heat transfer. This issue is associated with the miniaturization of hybrid rockets and has yet to be fully understood, but generally the fuel regression rate tends to grow with time. These small hybrid motors have exhibited a lean-to-rich burn behavior.

This fuel-rich tendency leads to combustion inefficiencies in the scales necessary to achieve useful thrust levels for small satellites (e.g., less than 5 N). As a solution to provide constant fuel regression rate at these desired low thrust levels, an example hybrid rocket device was redesigned to be end-burning, resulting in a constant regression rate and oxidizer-to-fuel ratio throughout the burn lifetime. This device provides an end-burning hybrid thruster in the sub-Newton scale. The description below relates to the design and testing results of this hybrid propulsion system at various thrust levels (e.g., 0.5 N and 1 N).

The hybrid propulsion system may use various plastics such as, for example, ABS, PMMA, PVC, or Nylon-12 as fuel, and gaseous oxygen (GOX,) as the oxidizer. Other high-density oxidizers such as hydrogen peroxide (H2O2) and nitrous oxide (N2O) may also be used. These oxidizers provide several advantages, including, for example, benign handling properties, simplified plumbing, and greater characteristic velocities over traditional monopropellants.

A high voltage arc, when placed on the surface of Acrylonitrile butadiene styrene (ABS) plastic, may result in fuel vaporization along the conduction path. This principle led to the development of several lab-weight thrusters varying in thrust from approximately 4-900 N. A 5 N system was successfully demonstrated on a suborbital sounding rocket. During this flight demonstration, the system was fired 5 times with each firing lasting 5 seconds. These successful tests led an investigation of the possibility of filling the current CubeSat market technology gap with this same technology by miniaturizing the ABS-based hybrid rockets to levels less than or equal to 5 N.

An end burn configuration of a hybrid rocket system200(shown inFIG. 10), deemed the Augmented Swirling Injection (ASI) end burn motor, was designed to increase the flow of oxidizer across a fuel surface212of a fuel gran214and to maintain a constant O/F ratio. The combustion chamber216is located at the end of the fuel grain214, and lies in between the fuel grain214and the nozzle218. The electrodes220,222run through the length of the fuel grain214and terminate flush with the combustion chamber216surface allowing for fuel vaporization. The oxidizer is injected at four ports224to induce a swirling internal flow, effectively increasing the mixing of the oxidizer with the vaporized fuel in the housing226. A side view of the end burn design is shown inFIG. 9.

An alternative configuration, referred to as a “sandwich” configuration of a hybrid rocket system300(shown inFIG. 10) was designed to increase the flow of oxidizer across the fuel surface312of a pair of fuel grains314a,314band to maintain a constant O/F ratio. The combustion chamber316is located between the two fuel grains314a,314b, with one fuel grain314abumping up against the nozzle318and the other fuel grain314bcontaining the electrodes320,322. Similar to the ASI end-burn configuration200, the oxidizer is injected at eight ports324around the housing326to induce a swirling internal flow. The electrodes320,322once again run through the length of the fuel grains314a,314band terminate flush with the combustion chamber316surface allowing for fuel vaporization. A side view of the end burn design is shown inFIG. 10.

One object of the hybrid rocket system300is to minimize the area of the thruster housing326exposed to the combustion chamber316with the goal of minimizing heat loss to the housing326. The reduction in heat loss increased combustion efficiency and fuel packing efficiency beyond what was demonstrated by the rocket system200.

An issue found during testing of the system200was the reliability of the arc ignition system, including the electrodes220,222. While using ABS as the fuel, the arc path was repeatedly covered with the vaporized but unburned fuel, leading to subsequent ignition failure. The unreliable ignition of ABS required an investigation into other plastics to use as the fuel, including, for example, Nylon-12, PMMA (Acrylic), and PVC. These fuels were selected based on an initial analysis using the NASA software package Chemical Equilibrium with Applications (CEA). The CEA analysis indicated favorable combustion properties over the use of ABS. Table I shows each fuel type and the key performance metrics used to select the final fuel.

Testing of ABS showed that in at least some situation it lacked the ability to provide repeatable ignitions. The motor had to be disassembled after each ignition to clean the carbon from the electrodes. Due to the low observed variable C* and the carbon present on the electrodes after a test, it was determined that the ABS combustion was incomplete in at least some tests.

Nylon-12 was used in several burn tests and was initially found to have better ignition reliability than ABS. However, the Nylon fuel would soften and flow during a test, which caused ignition failure in some scenarios and, in some cases, blocked the injector port. The regression rate in Nylon-12 was also found to have a major dependence on the bulk fuel temperature, leading to a long motor rise time as shown inFIG. 12(Thrust profile of 1 N End-Burn motor using GOX and Nylon 12).

Testing with PVC revealed a material compatibility issue between the exhaust products and the graphite nozzle, leading to nozzle erosion and failure. A single 15 second test would produce sparks in the exhaust plume and erode the throat of a new graphite nozzle. Testing with PVC indicated that no further evaluation was required to determine its merit as a desirable fuel.

PMMA exhibited a reduced regression rate as compared to ABS, and therefore had a reduced tendency to coat the electrodes in soot. Table I shows that PMMA had the greatest C* value, which indicates a more efficient combustion process when compared to ABS and Nylon-12. Because of the reduced deposition of soot on the electrodes during each test, a greater number of consecutive ignitions was achieved. PMMA also has a higher melting point so it did not deform during extended duration burns. These favorable traits surrounding PMMA led to it being selected as another desirable fuel.

Testing the arc ignition system with these plastics found that ABS was probably the most reliable of the plastics tested. The other plastics usually initiated an arc path but failed soon after in subsequent ignition attempts. This led to a common design of having a relatively small segmented portion of the primary fuel grain containing ABS as the ignition system. This mitigated the above-mentioned ignition failure using only ABS. As a result, PMMA with an ABS ignition segment was one of the preferred motor configuration.

Two motors with different diameters of the ASI end-burning motor were tested.FIGS. 13A-D(0.5 N×4.1 cm Diameter End-Burning Motor Test Data from twenty-seven 5-second burns with GOX and PMMA) shows the results from over twenty burns of the 4.1 cm diameter 0.5 N motor, whileFIGS. 14A-D(1 N×7.5 cm Diameter End Burning 1N Motor Test Data from twenty-seven 5-second burns with GOX and PMMA) shows the results from over twenty burns of the 7.5 cm diameter 1 N motor. The chamber pressure, thrust, and ignition power were directly measured by the test stand, while vacuum thrust was calculated from these and the measured oxidizer mass flow rates.

Table II shows the calculated average Isp and the extrapolated average Vacuum Isp for each ASI end-burn configuration. The Isp was calculated over the entire duration of the burn accounting for both the rise time and the tail-off time of the motor. The low Isp of the ASI configurations is a direct result of an incorrect oxidizer-to-fuel ratio, which is controlled through both the geometry of the motors and the total injected oxidizer.

The ignition characteristics are similar for each test as shown inFIGS. 13 and 14. The chamber pressure, thrust, and Isp are not as similar between each burn, and may vary widely. These performance variations relate to the variation in the oxidizer to fuel (O/F) ratio.FIG. 15(Comparison of O/F ratios at different fuel grain diameters) shows the O/F ratio for each end burning motor diameter.

The O/F ratio tended to slightly increase with the diameter of the fuel grain. However, the overall inconsistencies in the O/F ratio and Isp suggest that combustion may be incomplete in these end burning configurations.

A comparison with a more traditional core burn configuration is shown inFIGS. 16A-B(Comparison of thrust profiles of core burn (a) and end-burn (b) motors). The core and end burning motors performed similarly to each other at the 1 N scale. A key difference is the change in the O/F ratio. The core burn motors were extremely fuel rich, while the end burning motors were oxidizer rich. The end burning motors successfully produce a more constant burn area and thrust level during steady state when compared to the core burning motors. There was limited notable difference in the Isp between the core burn and end-burn configurations.FIGS. 16A-Bshow a comparison of the thrust profiles of the core burn and end burn motors.

It was also discovered that a large amount of heat loss to the motor case occurred on a burn-to-burn basis. This heat loss directly stripped away from the combustion enthalpy, which explained much of the difference between the performance predicted by CEA and performance actually measured. Initial tests tracking the motor case temperature estimated that about 20%-40% of the enthalpy released in combustion was present as heat in the case at the end of a test, rather than being used to increase the chamber stagnation temperature. Two methods to try to mitigate this heat loss were developed and reached initial testing:

Method 1: Insulate the case from the combustion chamber with a high temperature ceramic.

Method 2: Change to a “sandwich” geometry, with the combustion chamber located between two end-burn surfaces.

Insulating the case brought the challenge of determining a suitable insulation material. Several ceramics were investigated with the common properties of withstanding high temperatures, resisting oxidizing environments, and having relatively low thermal conductivity. Some initial testing was done by modifying the end burn motors to include a ceramic insulator.

A “sandwich” end burn motor was also built and tested. One object of this approach was to minimize the area of the thruster case exposed to the combustion chamber and increase the fuel packing efficiency. Initial tests of this motor were promising showing an increase in Isp by about 23% over an end-burn configuration of the same thrust level.FIGS. 17A-D(“Sandwich” configuration showing 13×15-second burns with GOX and PMMA) show the results from the sandwich motor testing.

The heat present in the case was also reduced, averaging about 12% of the combustion energy over the tests performed. The thrust rise time also improved from greater than 1 second (e.g.,FIG. 13A,FIG. 14A) to about 350 msec. These changes also improved the ignition reliability. Table III shows the average calculated Isp and the extrapolated vacuum Isp for the “Sandwich” configuration. Again, the specific impulse was calculated over the total duration of the burn including the rise time and tail-off time of the motor.

The propulsion systems disclosed herein take advantage of radiative heating, which traditionally had impeded the use of hybrid rockets as a propulsion system in the 1 N class. The analysis of various benign fuel grains paired with gaseous oxygen showed that these systems are capable of producing a vacuum Isp of greater than 150 seconds and a thruster rise time of less than 350 msec. Other high density oxidizers such as hydrogen peroxide and nitrous oxide are expected to further increase the performance of these end-burning hybrid motors. In at least some embodiments, the ratio between the thrust and oxidizer flow rate is fixed for any given propellant combination and ambient conditions (i.e., vacuum optimized nozzle verses atmospheric optimized nozzle), but other operating conditions and losses change the ratio. For example, for ABS and GOX in vacuum, one scenario produces about 5 N using about 0.97 g/s Oxygen, or about 1 N using about 0.20 g/s Oxygen. There may be similar values for PMMA/GOX. When using HTP or N2O instead of GOX, there may be an increase in the oxidizer flow for a given thrust. In another example, the 1 N PMMA/GOX motor consumes about 0.6-0.7 g/s of gaseous oxygen.

FIG. 18is a flow chart showing steps of an example method400in accordance with the present disclosure. The method400may represent one method of use of one or more of the systems200,300and represent a method that is operable to generate any of the test data shown inFIGS. 12-17and/or Tables I-III disclosed herein.

The method400may include, at402, providing a housing having first and second ends, a solid-grain fuel material positioned in the housing and having a combustion surface and being free of an oxidizer, at least two electrodes positioned in the housing, a combustion chamber defined between the combustion surface and the second end, an oxidizer port, and a nozzle positioned at the second end. At404, the method400includes delivering a flow of oxidizer through the oxidizer port and into the combustion chamber. At406, the method includes igniting the combustion surface with the at least two electrodes to generate a hot-gas, fuel-rich flow through the nozzle to generate thrust. At408, the method includes combusting of the fuel material in the combustion chamber being dominated by radiative heat transfer.

The method400may further include generating thrust in the range of about 0.1 N to about 5 N. The method400may include generating no more than about 5 N of thrust and having oxidizer flow of no more than about 5 g/s. The fuel material may include a plurality of flat layers that provide ridges along the combustion surface, and the electrodes may be configured to concentrate an electrical charge on the ridges, which act as micro-electrodes that produce localized electrical arcing thereon and ignite the combustion surface of the fuel material. The fuel material at the combustion surface may initially be consumed or removed through combustion of the fuel material, and a newly exposed internal surface of the fuel material may have newly exposed ridges that act as newly exposed micro-electrodes that produce localized electrical arcing thereon and re-ignite the newly exposed combustion surface.