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Technology Development Project Selections
September 4, 2013

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SMALL SPACECRAFT PROPULSION

Operational Demonstration of the MPS-120 Cubesat High-impulse Adaptable Modular Propulsion System
PI: Christian Carpenter
Aerojet General Corporation
Redmond, WA

The cubesat platform has enabled a new paradigm of low-cost space missions featuring readily available launch opportunities and low cost satellites with rapid development and production schedules using commercial-of-the-shelf components. Due to a lack of propulsive capabilities, cubesat missions are confined to their dispersal orbits, limiting their mission applicability. Current state-of-the-art cold-gas propulsion systems provide a delta-velocity of approximately 10 meters per second, but a greater delta-velocity capability will benefit NASA and the commercial market.

Aerojet has made significant investments, over the last three years, to develop the MPS-100 Cubesat High-impulse Adaptable Modular Propulsion System (CHAMPS) product line to satisfy the cubesat market need for commodity propulsion systems capable of providing a delta-velocity greater than 200 meters per second. The MPS-100 CHAMPS is a one-cubesat-unit (1U) miniature hydrazine propulsion system. The proposed project will demonstrate an MPS-120 CHAMPS with additive manufacturing to enable a twenty-fold delta-velocity increase over state-of-the-art cold gas propulsion systems for cubesats. The project will perform the first operation of Aerojet’s MR-142 miniature hydrazine rocket engine to demonstrate a four-fold reduction in engine size. The project will also perform the first operation of an additive manufactured piston propellant tank and miniature isolation system that demonstrates safe storage of hydrazine on cubesats and a credible path to low cost, short lead-time propulsion systems that are not burdened by the operational costs usually associated with ground handling of hazardous fluids.

Advanced Hybrid Rocket Motor Propulsion Unit for Cubesats
PI: John DeSain
The Aerospace Corporation
El Segundo, CA
Partner: Pennsylvania State University, University Park, PA

Most current cubesat propulsion techniques utilize cold gas thrusters with low specific impulse (less than 80 seconds for nitrogen gas). Current large-scale satellites use chemical propulsion systems with toxic hydrazine for monopropellant and bipropellant thrusters. Cubesats, as secondary payloads, are generally precluded from using hazardous material such as hydrazine. The proposed research will demonstrate a novel propulsion unit that utilizes a hybrid rocket motor. Hybrid rocket motors are an attractive bipropellant for use with cubesats because, unlike solid rocket motors, hybrid motors are not necessarily single-use thrusters and might be used for short duration thrust events like those needed for station-keeping or orientation control, when a suitable igniter system is employed. Hybrid motors can also be used in longer duration burns such as deorbit maneuvers or orbit insertion. Since one of the two propellants is a solid material, hybrid rockets have much greater simplicity and lower mass than bipropellant options using all liquid propellants.

The propulsion unit will be designed to be compatible with the existing cubesat footprint. Robotic manufacturing techniques such as rapid prototyping are capable of producing structures out of novel binder materials not readily producible by casting, molding, or machining. Portions of the propulsion system can be designed to serve as structure, reducing or eliminating the need for a completely separate cubesat frame. The fabricated polymer material used in rapid prototyping is suitable for use as a fuel grain in a hybrid motor. Previously, the Aerospace Corporation, with Pennsylvania State University, has tested hybrid rocket motor fuel grains produced by rapid prototyping. The major disadvantage of many hybrid motors made of printable plastic is that their performance is not as high as with other materials or liquid bipropellant systems. However, recent experiments that incorporate novel flow structures have demonstrated increased performance. The combination of state-of-the-art rapid prototyping of hybrid motors with current manufacturing of propellant systems could conceivably produce a high-thrust bipropellant motor in a compact space that could outperform existing cold gas thrusters on cubesats.

Iodine RF Ion Thruster Development
PI: Kurt Hohman
Busek Company. Inc.
Natick, MA

Busek proposes to develop a unique miniature radio-frequency (RF) ion propulsion system where iodine replaces the typical xenon propellant. Integration of the high-density, low-pressure-stored iodine greatly enhances the performance of a cubesat propulsion system, while eliminating the hazards and accompanying pressurized-tank safety issues and stringent procedures. In this project we will leverage our experience with miniature RF ion engines and our more recent iodine electric propulsion research.

The ability to precisely control the low thrust and high specific impulse lend the system to several applications including escape trajectory missions, orientation and pointing control, orbital maintenance, and de-orbit. Due to its low thrust, the system is less likely to be suitable for rendezvous or docking, where the need to react quickly may be required. The module will require low power and will be contained in a volume typical of a standard cubesat, (10 centimeter per side).

Depending on the required total impulse, the module could fit within one-half (0.5U) to one-cubesat-unit (1U). At 0.5 U, the estimated propellant stored is 250 grams and leads to a propulsion module that can deliver 4,900 Newton-seconds of total impulse over the course of a little more than 1 year. At 1U, the total amount of propellant is likely to be more than the thruster can process in its lifetime. At the conclusion of this project, Busek will be prepared to advance into a sub-orbital demonstration of the propulsion module.

1U Cubesat Green Propulsion System with Post-Launch Pressurization
PI: Michael Tsay
Busek Company. Inc.
Natick, MA
Partner: NASA Goddard Space Flight Center, Greenbelt, MD

Busek proposes to develop a one-cubesat-unit (1U) integrated propulsion system utilizing a non-toxic, “green” propellant. The system will also feature an innovative post-launch pressurization scheme. With the ability to launch completely unpressurized, the system will pose minimum hazards to the spacecraft integrator, primary payload, and launch vehicle. It will be a candidate for rideshare opportunities because of its low toxicity, safety, and minimum need for launch waivers. The proposed 1U propulsion system will only require a communication port and 15 watts of power at any spacecraft bus voltage. Estimated total impulse capability is on the order of 600 Newton-seconds, which can be increased by expanding the propellant tank volume. In the first phase of the project critical components of the 1U system will be developed, culminating in a system demonstration test.

The NASA Goddard Space Flight Center has agreed to support the development of the post-launch pressurization device as well as to provide mission infusion analysis. From the infusion of existing technologies, future flight prototype systems can be procured with minimum additional cost. A potential test platform for the prototype system is the SpaceLoft XL suborbital reusable launch vehicle developed by UP Aerospace.

Inductively Coupled Electromagnetic Thruster System Development for Small Spacecraft Propulsion
PI: John Slough
MSNW LLC
Redmond, WA

The thruster development program proposed is based on recent discoveries and advances in the development of an electromagnetic plasma thruster operating at very small scale and power. In the Inductively Coupled Electromagnetic (ICE) thruster, plasma generation is achieved with a small (30 mm diameter), spiral-wound coil, driven by an integrated radio frequency (RF) oscillator in a mode commonly referred to as an inductively coupled plasma (ICP). These devices have been employed for commercial plasma generation and have operated reliably for many years. Unlike conventional ICPs, the ICE thruster coil driver, as well as all other circuit elements are immersed in the fluid propellant providing for a power processing unit energy transfer efficiency of near unity.

The use of a porous glass frit as the interface between the driver coil and plasma generation zone at the thruster exit eliminates the need for a complex, miniature high-pressure gas feed and valve system. More importantly, it provides for efficient conversion of propellant liquid to gas with the gas delivered in the most optimal position for coupling to the RF fields. When operated at lower flow the predominance of the axial j x B electromagnetic force on the plasma/neutral fluid mixture provides for a means to achieve high specific impulse at low power for maximum thruster efficiency and propellant utilization. The induction coil and integrated circuit drive can be made quite small. Micro-electro-mechanical system ICPs have been made on a single printed circuit board that can operate at less than a watt to tens of watts. The total volume of the ICE thruster plus power processing unit is anticipated to be less than 1/8th of a cubesat unit (0.125U).

The ability to run on virtually any liquid propellant provides for much higher propellant density than gas with no significant tankage mass or volume. Operating with ionic or corrosive liquids is feasible as the system is electrode-less. No flow or evaporation issues were observed with the water-fed ICE thruster, and more than adequate flow was obtained with backing pressures less than one atmosphere.

Due to the high circuit quality factor (Q) of RF generators (Q ~ 300-500) it is possible to operate at the anticipated power (10-50 watts) with power transfer efficiencies to the plasma of 90%. Miniature RF drivers that are typically powered with a 12.8-volt supply are commercially available. Operation at other bus voltages can be easily accommodated.

While the essential physics of the device has been elucidated and the critical functions required for operation of the thruster have been demonstrated, there is a need to bring all components together in a manner that will provide for direct evaluation as a thruster. This will be accomplished with thrust stand measurements and in-situ thruster optimization. This technology development will bring the ICE thruster through technology readiness level 4. In a parallel effort the power processing unit and other flight hardware will be designed, prototyped, and tested, taking the technology readiness level of the ICE thruster system and subsystems to technology readiness level 5. With success, a follow-on effort would provide for the construction of a full ICE thruster system to technology readiness level 6 (prototype demonstration in an operational environment, i.e. flight) as work to be accomplished in year two.
 


SMALL EARTH RETURN VEHICLES

Technology Development for the Maraia Earth Return Capsule
PI: Alan Strahan
NASA Johnson Space Center
Houston, TX

A sounding rocket flying out of Spaceport America in New Mexico, will be used to help advance the understanding of the high-speed performance and active control of a candidate small atmospheric entry capsule that may one day return scientific samples from the International Space Station or demonstrate entry technologies for later use at Mars. The SpaceLoft XL rocket, built and operated by UP Aerospace, will carry a 7-kilogram, 25-centimeter diameter subscale prototype of the capsule, to an altitude greater than 110 kilometers, where it will deploy the capsule via spring ejection for a high-altitude descent towards the earth, aided only by gravity. After ascent motor burn out and prior to capsule deployment, the rocket will de-spin from 2,500 degrees per second to approximately 20 degrees per second, and then eject its nose cone to provide a sufficiently large opening for the capsule to exit.

Once free of the rocket, the spin-stabilized capsule will descend and reach a peak descent velocity greater than Mach 3.5, at an altitude of approximately 75 kilometers. The capsule will decelerate to subsonic velocity by the time it reaches an altitude of 30 kilometers and it will deploy a parachute at approximately 15 kilometers (50,000 feet) altitude. During its supersonic, transonic, and subsonic flight, which approximates the lower portion of a return from orbit, the capsule will exercise active control with a small cold-gas thruster system responding to software commands and inertial measurement unit-sensed attitude and rates. Its aerodynamic and stability characteristics will be observed via on-board measurements of attitude, rate, and acceleration. An upward looking wide-angle lens camera will also observe capsule attitude based on the position of the horizon.

A unique design aspect of the capsule recovery system will be demonstrated with the use of compressed gas to deploy the prototype recovery parachute, at the desired altitude.

This particular capsule being investigated is unique in that it started as a Soyuz shape, but was shortened to maximize internal volume for a given height and diameter, increasing its potential to return samples, though potentially reducing its stability. Other capsules of a somewhat similar shape have shown stability issues in the low-supersonic and transonic velocity ranges. Acquiring dynamic stability data and demonstrating the actual control of such a capsule is very difficult using wind tunnels or other ground-based facilities, which is where this flight opportunity will prove most valuable.

Overall this test will investigate and inform the design of the entry, descent, and landing subsystems for a small earth return capsules. Such a capsule design would then allow for the on-demand return of small samples from the ISS and provide for the advancement of other exploration related entry, descent, and landing technologies, by acting as an atmospheric entry testbed.

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Page Last Updated: September 4th, 2013
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