Request for Information (RFI)
NASA is currently developing the propulsion system concepts for future exploration missions including a near term human return to the lunar surface. Studies have identified high performance, non-toxic, cryogenic methane (LCH4) and oxygen (LO2) as a propellant combination for consideration for the main and reaction control system (RCS) lunar surface ascent propulsion system. LDAC-1, a NASA study, depicted a lunar ascent vehicle concept consisting of two sets of spherical propellant tanks, each set containing an oxidizer tank and a fuel tank and located diametrically opposite on either side of the crew ascent cabin. The spherical oxidizer and fuel tank sets contained the propellant for the main ascent engine and the RCS engines required for lunar orbit maneuvering and docking with the Crew Exploration Vehicle (CEV). The concept assumed the use of pressure fed main and RCS engines operating at approximately 350 psia. Two sets of spherical propellant tanks were selected for maintaining a controlled center of gravity (CG) for the lunar ascent vehicle.
For a lunar outpost exploration mission at the lunar South Pole, NASA has planned a surface stay duration of approximately 210 days for the crew habitat and lunar surface ascent stage. NASA system trade studies have shown that LCH4 ascent stage propellant tank fluid venting can be eliminated on the lunar surface for the 210 day mission with
1.A passive Multi-layer Insulation (MLI) system consisting of approximately 60 layers to protect the propellant tanks from the lunar surface and solar environmental heating;
2.Loading the LCH4 propellant tanks with densified LCH4 at 165 oR and starting with a tank ullege of approximately 15% at the KSC launch pad.
In order to verify the trade study results, a NASA inter-Center team consisting of Ames Research Center (ARC), Glenn Research Center (GRC), Kennedy Space Center (KSC) and Marshall Space Flight Center (MSFC) has proposed a series of baseline passive thermal control technology tests of a flight-representative, spherical LCH4 ascent stage propellant tank in a simulated lunar thermal environment at the GRC Creek Road Cryogenic Complex, SMiRF facility. This activity is sponsored by the NASA Cryogenic Fluid Management (CFM) Project Office.
The primary objectives of the planned test series are:—
1. To demonstrate the viability of a LCH4-LO2 ascent stage for the given lunar surface thermal environments and mission duration. This requires a reliable, high performance passive thermal isolation system conforming to the constraints outlined in the following section.
2. To provide a testing platform for lunar surface thermal control techniques, including in-tank equipment and instrumentation. Modification of the in-tank hardware will require periodic disassembly and reassembly of the MLI system. In order to enable direct comparisons of test data before and after modifications, this procedure must be accomplished with minimal effect on overall MLI performance.
3. To establish a well-defined set of baseline measurements for comparison with MLI model predictions, including the performance-lowering effects of penetrations and seams (for which reliable correlations are lacking). Proper instrumentation (primarily thermometry) of the MLI system is therefore imperative. The ultimate goal is to minimize uncertainty in MLI performance prediction.
This RFI seeks to determine whether these objectives are feasible using current techniques and commercially available materials.
The following list summarizes the salient features and constraints of the NASA test hardware and test facility to which the MLI supplier must conform.
1.The spherical test tank will not be part of the solicitation. It will be GFE. Its design is complete and fabrication drawings are available.
2.The anticipated delivery date of the test tank to GRC is no later than April 1, 2009. If necessary, the test tank could be shipped to the selected contractor for MLI installation after June 1, 2009.
3.The test tank material is 304 stainless steel, with a surface area of 7239 in2, and a nominal 48 inch diameter.
4.The thickness of the MLI system shall be no greater than 4 inches, as that is the clearance between the tank wall and the cryo-shroud.
5.The MLI system shall consist of 60 shield/spacer layers.
6.With the cryo-shroud at 700 oR (its maximum attainable temperature), the total heat leak through the (evacuated) MLI assembly (including degradations due to supports, tank penetrations, seams, etc., but not including conduction heat leaks through supports, penetrations, instrumentation wiring) shall not exceed 1 W.
7.Two circular man-way covers are located at the top and bottom of the test tank. The MLI system shall be designed to allow for ready access to the man-ways. The MLI system must therefore be at least partially removable and replaceable, with minimal degradation in performance.
8.Both man-way covers have piping penetrations that are representative of the flight tank and must be accommodated by the MLI design.
9.The test tank has two independent support methods:
•For testing, the tank will be suspended from the vacuum chamber lid by three rods threaded into the tank structure. The MLI system must accommodate the support rods.
•For test preparation outside the vacuum chamber, the tank is to be rigidly held using two removable support brackets, located 180 degrees apart on the tank’s equator. The MLI system must be removable/replaceable, with minimal degradation in performance, at least over the two support bracket mounting flats.
10.The “simulated lunar thermal environment” mentioned above is a vacuum chamber with a nominal vacuum level 10-6 torr and a nitrogen-cooled cryo-shroud with top, cylindrical, and bottom sections, with an adjustable temperature range from 200 oR to 700 oR. The inner surface of the shroud is coated with black paint.
11.For thermal characterization of the MLI system, the test facility can accommodate up to a maximum of 70 thermocouples.
A Request for Proposal for the design, fabrication and installation of a robust, predictable, low heat leak, tank-applied MLI system is anticipated to be released in early February 2009, with an anticipated contract award date is March 20 and completion in mid-July, 2009.
Prior to completing the statement of work, NASA hereby solicits information from MLI industry sources regarding industry experiences, current capabilities, and/or knowledge of flight-applicable state-of-the-art approaches to any or all of the following relevant areas:
1.Thick MLI assemblies (50 or more layers);
2.Design, fabrication and assembly techniques appropriate for spherical tanks (e.g., interleaved gore sections, overlapping sub-blanket assemblies;
3.Accommodation of tank supports and penetrations with minimal impact on MLI system performance;
4.Reliability and reproducibility of MLI assembly procedures;
5.Reproducibility of MLI performance with repeated disassembly/reassembly;
6.Tank attachment techniques;
7.Inter-layer attachment methods (e.g., tape, Velcro, stitching, bonding);
8.Inter-layer spacing techniques and materials for minimizing conduction heat-leak;
9.Strength and optical properties of commercially available outer coverings (e.g., Teflon fabric);
10.High emissivity materials (e.g., metalized mylar or kapton);
11.Materials handling and assembly procedures for minimizing contamination (e.g., emissivity degradation);
12.Optimization of constant or variable layer density;
13.Interstitial gas venting techniques (e.g., perforations);
14.Performance prediction, modeling, and analysis of MLI systems;
15.Performance verification instrumentation (e.g., measurement of temperature profiles, interstitial pressure);
16.Expected variation of key performance-impacting properties (especially emissivity) in commercially available materials;
17.Quality assurance and acceptance criteria for commercially procured materials or assemblies; and,
18.Other factors that might impact MLI system performance or performance predictability.
Responses to the RFI should also include; rough order of magnitude cost and schedule estimates; a company profile, including in-house facilities, experience of key personnel, and business size status; and a summary of previous designs for similar applications or flight systems, along with overviews of performance data.
All responses should be submitted electronically in Portable Document Format (PDF), with a practical size limit of 6 MB, via email. The due date for responses is Jan. 9, 2009.
FAR 52-215-3 Request for Information or Solicitation for Planning Purposes (Oct 1997)
(a) The Government does not intend to award a contract on the basis of this solicitation or to otherwise pay for the information solicited except as an allowable cost under other contracts as provided in subsection 31.205-18, Bid and proposal costs, of the Federal Acquisition Regulation.
(b) Although “supplier” is used in this Request for Information, your response will be treated as information only. It shall not be used as a proposal.
(c) This solicitation is issued for the purpose of a request for developmental information.
