Description
NASA/GRC has a requirement for the Scaling Relationships Study for Large Propellant Tanks NASA intends to purchase this service from Dr. Franklin T. Dodge at the Southwest Research Institute of San Antonio, Texas. SOW for the Scaling Relationships Study for Large Propellant Tanks. Introduction- NASA is currently developing the propulsion system concepts for exploration missions including a near term human return to the lunar surface. The propulsion concept defined in the NASA Exploration System Architecture Study (ESAS) considered the use of cryogenic propellants ? liquid oxygen (LO2), liquid methane (LCH4) and/or liquid hydrogen (LH2) for the low gravity portion of the mission. These exploration propulsion system concepts included a large LO2/LH2 pump fed propulsion second stage capable of LEO orbit insertion of a lunar surface payload, of LEO loiter and of a LEO engine restart for the trans-lunar injection of a crewed capsule, service module and the lunar surface payload required for lunar exploration missions. The LH2 propellant storage tank of this large propulsion stage is cylindrical and has been estimated to be 10 m (~30 feet) in diameter, 10 meters (30 feet) long, and to have an approximate volume of 620 m3 (22,000 ft3). The LO2 propellant storage tank is ellipsoidal and has been estimated to be 8 m (~24 feet) in diameter, approximately 6 m (18 feet) long and to have an estimated volume of 190 m3 (6,700 ft3).This stage is comparable in size to the Saturn V S-II stage (~ 9 m in diameter); however, the S-II did not have an orbital engine restart requirement, i.e., was a single burn. The S-IVB stage was about 20 ft in diameter and the engine was re-started after a LEO coast period of up to 4.5 hours. Similarly, the 10 ft. diameter Centaur upper stages have restarted after orbital coast periods of up to 9 hours. Thus an earth departure stage will be the largest stage with an orbital engine restart requirement after extended coast periods in low earth orbit (LEO) of up to 14 days. Cryogenic fluid management subsystems will be required to control the propellant tank pressure and assure that the tank pressure and bulk fluid temperatures support the propulsion system engine start box after the pre-pressurization sequence in preparation for engine restart. The current method for tank pressure control during orbital coast periods is to position the tank fluid with a settling thrust provided by a separate ancillary propulsion system then venting the ullage. Tank venting and pressure control can be accomplished without propellant settling by using a thermodynamic vent system (TVS). De-stratification of the bulk fluid temperature can be accomplished with a mixer and the thermal energy removed through utilization of Joule-Thomson refrigeration techniques. Although a cryogenic TVS has not been demonstrated in low-gravity, extensive testing of TVS?s has been accomplished in normal gravity. All of the previous experimentation with fluid mixing and TVS pressure control systems has been with very small tanks using simulant fluids in drop tower tests or with tank volumes of less than 1000 cubic feet using liquid hydrogen (LH2) or liquid nitrogen (LN2) in a 1-g environment. Objective- The objective of this study is to 1. Determine the risk to NASA with extrapolating results from prior ground tests of fluid mixer and TVS pressure control components to the proposed large propellant tanks of an earth departure stage. 2. Provide NASA an understanding of the need of near full scale tests to verify the performance of the fluid mixers and TVS pressure control systems required for an earth departure stage of the size under consideration. Task Description- 1.) Determine the thermodynamic transient response scaling relationships for both LH2 and LO2 for mixer size and power, and destratification time to eliminate fluid stratification in a LEO loiter. ? Using the developed scaling relationships and existing, reported NASA and contractor 1-g LH2 and LO2 (LN2) stratification test data predict the mixer size and power, and destratification time for an earth departure stage to eliminate fluid stratification for a 50 % fill level in LEO loiter. ? NASA ground test data are available for the following tank volumes: 18 m3 (639 ft3); MSFC ? MHTB and 5 m3 (175 ft3); GRC ? K-Site. 2.) Determine the thermodynamic transient response scaling relationships for both LH2 and LO2 for sizing a TVS heat exchanger and its J-T valve, and for determining the fluid losses required to maintain the tank pressure within +/- 10 % of the propellant tank operating pressure (< 138 kPa (20 psia)) for tank fill levels from 50 to 100% in LEO loiter ? Using the developed scaling relationships and existing, reported NASA and contractor 1-g LH2 and LO2 (LN2) TVS test data predict the resultant size of the TVS heat exchanger and its J-T valve, and the fluid losses required to maintain an earth departure stage tank pressure within +/- 17 kPa (2.5 psi) for a 50% tank fill level in LEO loiter ? NASA ground test data are available for the following tank volumes: 18 m3 (639 ft3); MSFC ? MHTB, 5 m3 (175 ft3); GRC ? K-Site, and 2 m3 (71 ft3); MSFC ? STUSTD 3.) Establish a qualitative value of the confidence level to use the scaled one-g data, combined with available flight data and CFD modeling results, to predict an earth departure stage low-g behavior. Study Assumptions- ? Tank geometry: ? LH2 is cylindrical with sq. root of 2 ellipsoidal domes; approximately 10 m diameter and 10 m long with an approximate volume of 620 m3. ? LO2 is ellipsoidal; approximately 8 m diameter and 6 m long with an approximate volume of 190 m3. ? Tank heat leak ranges: ? LH2 from 185 to 200 watts ? LO2 from 110 to 130 watts ? Nominal tank pressure during orbital coast is 138 kPa (20 psia) ? Approximate tank conditions required to support engine restart: ? LH2 - 138 kPa (20 psia) at 20 K (37 R) ? LO2 - 138 kPa (20 psia) at 91 K (164 R) ? Both axial jet and spray bar mixer systems are to be considered ? Assume acceleration environment of 10-6 g?s due to aerodynamic drag during LEO coast periods Deliverable- Final report of the three tasks described in the Task Description section above. Period of Performance- A draft version of the final report as denoted in the Deliverable section must be submitted for NASA GRC review three months after contract award. The final version of the report must be submitted one month after the completion of the NASA GRC review. The NASA GRC review will not exceed two weeks. The study key personnel shall participate via telecom during the NASA GRC review. Presently a concept for an earth departure stage requires that it be the largest propulsion stage ever developed with an orbital engine restart requirement after extended coast periods in low earth orbit (LEO). This large size presents a challenge to the NASA Cryogenic Fluid Management (CFM) community to validate the tank pressure control systems using existing but substantially smaller scale ground test hardware and test facilities. It is recommended that this study be performed by Dr. Franklin T. Dodge, a Senior Technical Advisor of the Southwest Research Institute (SwRI). SwRI is an independent, non-profit research institution located in San Antonio, Texas. Rational- Dr. Dodge is the most qualified individual to perform this study because he is recognized by his peers in academia, industry and NASA as both a national and international expert in the field of cryogenic fluid management for low gravity applications. He has spent over 43 years in this field, authored over 100 publications, has been the Principal Investigator for a successfully flown low gravity flight experiment, holds several patents in this field and has consulted as a Co-Investigator for over six flight experiments. He is an expert in the field of fluid thermodynamic behavior in low gravity, is knowledgeable in the use of thermodynamic vent systems to control propellant tank pressures in low gravity and has acted as a consultant on several proposed NASA cryogenic fluid flight experiments in the 1990?s. Although currently retired from SwRI, Dr. Dodge maintains an office at the SwRI San Antonio campus and has availability to all the SwRI resources. Recommendation- It is recommended that this study be performed by Dr. Franklin T. Dodge, a Senior Technical Advisor of the Southwest Research Institute (SwRI). SwRI is an independent, non-profit research institution located in San Antonio, Texas. The Government intends to acquire a commercial item using FAR Part 12. Interested organizations may submit their capabilities and qualifications to perform the effort in writing to the identified point of contact not later than 4:30 p.m. local time on or before 4/1/08 to: ( NASA Glenn Research Center, Attn: Dennis D.Pehotsky, 21000 Brookpark Rd., MS 500-319, Cleveland, OH 44135.) Such capabilities/qualifications will be evaluated solely for the purpose of determining whether or not to conduct this procurement on a competitive basis. A determination by the Government not to compete this proposed effort on a full and open competition basis, based upon responses to this notice, is solely within the discretion of the government. Oral communications are not acceptable in response to this notice. All responsible sources may submit an offer which shall be considered by the agency. An Ombudsman has been appointed. See NASA Specific Note "B". Any referenced notes may be viewed at the following URLs linked below.