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Spacecraft Propellant Storage and Feed Systems

Description:

 
 

TECHNOLOGY AREA(S): Space Platforms

The technology within this topic is restricted under the International Traffic in Arms Regulation (ITAR), 22 CFR Parts 120-130, which controls the export and import of defense-related material and services, including export of sensitive technical data, or the Export Administration Regulation (EAR), 15 CFR Parts 730-774, which controls dual use items. Offerors must disclose any proposed use of foreign nationals (FNs), their country(ies) of origin, the type of visa or work permit possessed, and the statement of work (SOW) tasks intended for accomplishment by the FN(s) in accordance with section 5.4.c.(8) of the solicitation and within the AF Component-specific instructions. Offerors are advised foreign nationals proposed to perform on this topic may be restricted due to the technical data under US Export Control Laws. Please direct questions to the AF SBIR/STTR Contracting Officer, Ms. Gail Nyikon, gail.nyikon@us.af.mil.

OBJECTIVE: Develop and demonstrate decreased mass, volume and power requirements for spacecraft liquid chemical propellant storage and feed hardware.

DESCRIPTION: Typical propellant storage and feed systems for spacecraft using liquid chemical propulsion comprise compressed helium or nitrogen driving the propellant from the storage tank. Mission requirements will drive the choice of blow down or regulated pressure feed, likewise, the choice of feed will further drive the type of propellant management device. Also common for hydrazine monopropellant systems, the driving pressure gas and propellant will be within the same tank separated by an elastomeric diaphragm. Also, it is frequently necessary to provide some sort of environmental control for the propellant storage tank to ensure the propellants do not freeze or fall to sub-nominal temperature for thruster operation. These systems are proven for reliability and have long flight legacies, however, they are not free of concerns and there remains opportunity for improvement of the design.

Pressurized tankage presents a significant logistical and cost footprint in the regards to component qualification or verification, acquisition lead time, and spacecraft processing operations. Where monopropellant thrusters are used, catalyst poisoning is always of concern. Though standards for the purity of hydrazine as well as for preparation of the elastomer diaphragm materials, such as AFE-332, that the hydrazine would be continuously contacting within a diaphragm tank are well established, introduction of contaminating substances acquired from the hydrazine or diaphragm leaching may have potential to alter thruster delivered performance due to catalytic poisoning. Similar concerns are also present for emerging advanced green monopropellant formulations.

Other limitations faced with liquid propulsion systems on board spacecraft relate to impulse variability and determination of propellant remaining. In blow down systems, the change in delivered performance of the thruster due to decreasing feed pressure must be mapped in order to be able to determine thrust commands to accomplish desired maneuvers. For missions with large delta-V requirements, significant amounts of propellant will be required driving need of large compressed gas tanks reducing mass and volume available for payload. Liquid chemical thrusters that can deliver variable thrust from a compact configuration, such as combined functionality of low thrust monopropellant and high thrust bipropellant modes, for different mission phases have been developed and are commercially available.

Of interest is a liquid propellant storage and feed system that does not grow in volume and component manufacturing risk with propellant throughput (such as state of the art compressed gas approaches) that also mitigates typical concerns associated with reliability, repeatability, and contamination. Envisioned applications are for thrusters in the range of ~0.25 lbf to ~5.0 lbf, with design knowledge to scale up and be able to support to the 100 lbf level. Minimum impulse bit performance repeatability and predictability that is superior or, as a minimum, equivalent to the state of the art is desired.

Performance and capability advantages to all type of spacecraft platforms from extremely volume limited applications such as Cubesats up to large scale, long life systems such as GPS should be assessed and presented.

Developmental effort should include a physics based understanding in terms of a mathematical expression; capturing details of power requirements, geometry, material make-up, duty cycle, and propellant throughput range as a function of relevant parameters.

Energy requirements should be bounded within today’s nominal satellite bus architecture capabilities.

Additional considerations should include streamlined manufacturing process with high yield and minimal quality assurance required. Estimates of storage life and guidance to maximize storage should also be considered, minimal storage requirements are desired. Approaches with applicability to both state of the art and emerging green propellant formulations are encouraged.

The thruster technology should be capable of supporting a 15-year mission in GEO or Medium Earth Orbit (MEO) and 5 years in Low Earth Orbit (LEO) after ground storage of 5 years.

PHASE I: Demonstrate a feasibility concept and accompanying base model approach that shows path to meet manufacturability and performance metrics stated. The effort should clearly address and estimate propulsion system inert weight and overall flight system impacts as well as model and manufacturing technical challenges.

PHASE II: Demonstrate proof of concept with flight scaled components in relevant environment. Propulsion system inert weight and flight system impacts shall be optimized from those estimated in Phase I. Leading model and manufacturing technical challenges shall be retired or have a clearly defined path to mitigation.

PHASE III DUAL USE APPLICATIONS: The Offeror shall develop viable demonstration cases in collaboration with the government or the private sector. Follow-on activities are to be sought aggressively throughout all mission applications within DoD, NASA, and commercial space platforms by Offeror.

REFERENCES:

  • Ballinger, I.A.; Lay, W.D.; Tam, W.H., “Review and History of PSI Elastomeric Diaphragm Tanks”, AIAA 95-2534, 31st AIAA/ASME/SAE/ASEE Joint Propulsion Conference, San Diego, CA, July, 1995.
  • Ballinger, A.; Sims, D., “Development of an EPDM Elastomeric Material for use in Hydrazine Propulsion Systems”, AIAA 2003-4611, 39th AIAA Propulsion Conference, Huntsville, AL July 21, 2003.
  • Honse, J.P.; Bangasser, C.T.; Wilson, M.J., “Delta-Qualification Test of Aerojet 6 and 9 lbf MR-106 Monopropellant Hydrazine Thrusters for Use on the Atlas Centaur Upper Stage during the Lunar Reconnaissance Orbiter (LRO) and Lunar Crater Observation and Sensing Satellite (LCROSS) Missions”, AIAA 2009-5481, 45th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, Denver, CO August, 2009.
  • Owens, B.; Cosgrove, D.; Sholl, M.; Bester, M., “On-Orbit Propellant Estimation, management, and Conditioning for THEMIS Spacecraft Constellation”, AIAA 2010-2329, SpaceOps Conference, Huntsville, AL, April, 2010.
  • United States Patent 5,417,049.

KEYWORDS: Spacecraft Propulsion, Chemical Propulsion, Propellant Storage, Propellant Feed system, Pump, Blow Down, Thruster

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