Missouri University of Science and Technology Satellite Team Document Title: Propulsion Conceptual Design Document Name: Propulsion Conceptual Design Documentation Number: 04-001 Status: Current Date: 1-19-10 Revision: -
REV RELEASE DATE Revision Summary BRIEF DESCRIPTION/REASON FOR CHANGE EFFECTIVE PAGES - 1/19/2010 Initial draft. All
Signature Page REV - Prepared by: Jason Thrasher, Subsystem Member jstq58@mst.edu, (314) 610-9311 Date Prepared by: Christopher Tutza, Subsystem Member cptv22@mst.edu, (618) 616-5020 Date Approval: Dr. Henry Pernicka, M-SAT PI pernicka@mst.edu, (573) 341-6749 Date Approval: Mike Dancer, M-SAT Chief Engineer mdancer@mst.edu, (573) 429-6931 Date
Table of Contents 4 Propulsion Subsystem... 6 4.1 Introduction... 6 4.2 Propellant Investigation... 6 4.3 Propulsion Tank Design... 7 4.4 Propellant Tank Spacecraft Integration... 7 4.5 Propellant Tank Heating... 9 4.6 Pressure Transducers... 9 4.7 Pressure Regulation... 10 4.8 Propellant Distributor... 10 4.9 Propellant Lines and Connections Design... 10 4.10 Nozzle and Valve Design... 11 4.11 Nozzle and Thruster Spacecraft Integration... 12 4.12 Theory and Testing... 12 4.13 References... 13
Conceptual Design of the MR SAT Formation Flight Satellite Project at the Missouri University of Science and Technology Updated December 9, 2009 4.0 Propulsion Subsystem
4 Propulsion Subsystem 4.1 Introduction This document is the Conceptual Design Document (CDD) report for the Propulsion subsystem. The intent of this report is to highlight and document the current propulsion system design being integrated on the Missouri Rolla Satellite (MR SAT). It not only includes the status and details of the current design, but also provides insights to future plans and expectations of the Propulsion subsystem. The propulsion system design is not complete and there are still key factors that require further research and development. This report details the analysis and findings of the propulsion design process as well as the hardware product specifications. There is also information regarding planned and completed prototype testing. The propulsion designed for MR SAT is a cold gas thruster system intended to provide orbital and three axis attitude control. The propulsion system is being designed to incorporate two propellant options which are detailed in this report. The two options are a refrigerant, R-134a, stored as a saturated liquid and a more common compressed gas propellant, Xenon. The preferred propulsion system for MR SAT is the R-134a design. Due to the limitations of the University Nanosat Program (UNP), specifically the use of pressure vessels and propellants capable of phase changes, the use of R-134a requires more in-depth and specific analysis and testing prior to it being accepted by the AFRL as a safe propellant option. Hopefully this will lead to a refrigerant being accepted as a cold gas propellant and have it become a standard solution in providing small satellite propulsion systems. 4.2 Propellant Investigation When potential spacecraft propellants were being analyzed the selection process was broken down into two categories: those that could be stored as compressed gases as saturated liquid vapors. The compressed gases are analyzed in this section of the report with the saturated liquid vapor analysis detailed in a later section. It was decided that any gas options must be chemically stable and inert when exposed to a standard atmosphere, inert at temperatures and pressures in excess of the maximum expected launch temperatures and pressures, and that they are non-toxic. This suggested the noble gasses and two common stable gases already present in the atmosphere, nitrogen and carbon dioxide, both common propellants. The gases that were finally considered for analysis and comparison were: nitrogen, N2; carbon dioxide, CO2; helium He; neon, Ne; argon, Ar; and xenon, Xe. While these gasses cannot sustain human life, which makes them simple asphyxiates, they are not dangerous in a well-ventilated area and hence can be considered non-toxic. Due to its ability to be stored in a two-phase state, the extra mass of R-134a that can be stored compared to xenon, in an equivalent tank volume, results in greater ΔV capabilities. Also, other types of refrigerants, including R-12, are not considered because of their negative environmental
impact. For this reason, the use of a saturated liquid R-134a propellant is the primary propulsion design that is documented in this report. 4.3 Propulsion Tank Design In order to provide for propellant storage, the Marotta BS25-001 propellant tank was selected and purchased. An internal Propellant Management Device (PMD), consisting of filters and screens, is integrated into this tank to provide liquid propellant control (slosh control) during maneuvers. This particular PMD prevents the ingestion of liquid into the propulsion system feed lines. The PMD only restricts liquid movement by surface tension reactions on the screens and by partitioning the tank into smaller volumes. The shell is manufactured with Stainless Steel 316L with the internal filters manufactured with Stainless Steel 304L/316L. The internal screens are of expanded aluminum 901A. Table 4.3.1 highlights the specifications of this tank. Table 4.3.1: Specifications of Marotta PMD Tank Operational Temperature -40 ºC to 65 ºC -40 ºF to 150 ºF Operating Pressure 0.4 MPa 58 psi Maximum Expected Operating 1.3790 MPa 200 psi Pressure (MEOP) Minimum Burst Pressure 9.7975 MPa 1421 psi (MBP) Volume Capacity 2500 cm 3 153 in 3 Mass < 1.5 kg < 3.3 lb Maximum Body Length 32.6 cm 12.83 in Outside Diameter 11.4 cm 4.5 in Safety Margin (MEOP:MBP) 7:1 External Linkage < 1 x 10-6 scc/sec GHe max at 0.8 MPa Outlet 7/16-20 UNJF 37º STD Flare Male Inlet Threading Fitting M12 x 1.25 - Male 4.4 Propellant Tank Spacecraft Integration For the design configuration, the length of the tank is placed perpendicular to the z-axis of the satellite, as seen in Figure 4.4.1 and Figure 4.4.2. It is required that the tank be placed at the center of mass (cm) of MR SAT when MRS SAT is undocked. This placement will minimize the cm movement effects on attitude determination and control as the propellant is expelled and MR SAT s mass decreases. This exact placement is being coordinated with the Structure subsystem. Structural integration is achieved through the selected mount system. The current mounting solution is the use of two custom offset mounts with a diameter of 11.4 cm (4.488 in.). These mounts have been modeled on offset pipe clamps. The mounts will be constructed in the Missouri S&T machine shop out of aluminum 6061. Each mount consists of two parts joined by two fasteners. The mount is secured to the structure using two additional fasteners.
Figure 4.4.1: Example of Propellant Tank Integration Figure- 4.4.2: Example of Propellant Tank Layout
4.5 Propellant Tank Heating Thermal control of the propellant is required for this system because evaporation is a cooling process. Direct heating of the tank will be utilized for warming the R-134a propellant. Additional heating through cartridge heaters in the distributer is being investigated. Heating will be achieved through two flexible heaters mounted around the center circumference of the tank. Each heater consists of the polyimide film insulator, Kapton, with a heating element of either copper, nickel (Ni) or nickel iron (NiFe). The variation in element material is required to achieve the desired resistance. Leads with a length of 12 in. are provided standard for the heaters. Acrylic pressure-sensitive adhesive (PSA), which is factory applied to the aluminum backing of the heater, is used for mounting. The aluminum backing provides additional stability for curved surface applications. The adhesive is NASA approved for outgassing. The heaters are constructed of NASA approved materials (NASA-RP-1124; NASA GSFC S-311-P-079). Table 4.5.1 highlights the specification of the proposed heaters from Minco. The temperature sensors provided by the MR SAT Thermal subsystem will perform the necessary temperature monitoring for precise control of the heaters. To regulate the temperature of the tank and control heat radiation to other components, Multilayer Insulation is also under review. A full thermal analysis is needed at this time to accurately select the materials. The selection process is being conducted with regards to the NASA/TP-1999-209263 document. A fit to drawing design is being pursued through current contacts providing completed MLI blankets. Table 4.5.1: Heater Specifications Thermal Heaters Heater Location Dimensions cm (in) Resistance (Ω) Output Wattage (W) Lead Gauge Tank 1 27.178 x 27.940 28 3.57 AWG24 (10.70 x 11.00) Tank 2 7.620 x 25.400 (3.00 x 10.00) 103 0.97 AWG24 Outgassing Properties Material TML% CVCM% Kapton 0.25 0.01 PSA 0.02 0.00 Mounting Adhesive Item 0.002 (50 μ) PSA Film (Minco #10) Operating Temperature Range ( C) Operating Temperature Range ( F) TML% CVCM% (-32 to 150) (-25 to 302) 1.06 0.02 4.6 Pressure Transducers Pressure monitoring of the system is achieved through the use of two pressure transducers. The transducers are placed after the tank and after the regulator. These transducers are the AS17A model from Sensotec. The specifications for this model are listed in Table 4.6.1. Stainless Steel is used in the construction of the case of the transducers. The pressure port could have been configured to match the 1/4 in. connections previously specified for the 1/4 in. aluminum propellant lines; however, additional engineering and validation testing would have been
required for this configuration adding significant expense and time. The decision was made to use the standard 7/16-20 pressure port for this reason. Swagelok adaptors are available for the connection between the pressure port and the selected 1/8 in. Stainless Steel propellant lines with marginal increases to the mass budget. Though not designed for space application as previously selected models, the AS17A is designed to be lightweight and rugged per MIL-45208. Table 4.6.1: Pressure Transducer Specifications Pressure Range 68.9476 kpa to 68.9476 MPa 10 psia to 10,000 psia Operating Temperature -53.8889 ºC to 148.8889 ºC -65 ºF to 300 ºF Mass 141.7476 g Pressure Port 7/16-20 per MS33656E-4 Excitation 10 VDC Bridge Resistance 350 ohms 4.7 Pressure Regulation Pressure regulators increase the predictability of the system. They are valves downstream of the propellant tank. The pressure downstream of the regulator opposes the elastic force of a spring trying to open the valve. When the regulated pressure is achieved, the valve is sealed shut. When the pressure drops below the regulated pressure, the valve is forced open by the spring. This allows propellant to fill the volume between the regulator and the next valve and repressurize the region. This section of tubing can become depressurized due to the engine firing, leakage, or cooling. Having a regulator simplifies the firing process by creating a nearly constant pressure at firing. This is at the expense of extra mass and complications with design and integration, but the regulator does not require electrical power. 4.8 Propellant Distributor In order to minimize the number of fittings on the propellant lines, a distributor will be used to assist in propellant flow. Each thruster will have a dedicated line from the distributor to the corresponding thruster. (1) By minimizing the number of fittings, the number of possible leak sources is decreased. 4.9 Propellant Lines and Connections Design Providing a reliable method for the transfer of propellant from the storage tank to each thruster without leaks is challenging. Propellant lines and fittings must remain light weight while being capable of maintaining high pressures without leaking, which is a common problem for cold gas systems. The following section discusses the current design of the propellant lines and many design considerations that must be balanced for the system to be flight ready. Tube bending will be used in order to minimize the number of fittings required for the propulsion system. This is important for two reasons: fittings have a much greater mass than that of the tubing, and each connection increases the likelihood of leaks forming in the propellant line. However, tube bending is not without its design challenges. To ensure the structural properties of the tubing are not compromised during the bending process, strict adherence to manufacturer
guidelines is imperative. These guidelines include a minimum centerline radius of a bend, which will guarantee the structural integrity of the tubing (2). For the stainless steel tubing being considered, the minimum centerline bend radius is ½ in. 4.10 Nozzle and Valve Design External vendors were sought that could provide space nozzle/valve hardware that would satisfy the MR SAT requirements prior to designing an in house system. This may be cost ineffective; however, the advantages of space qualified hardware from a reputable vendor are worthwhile. Numerous large and small space propulsion development companies were contacted to find a provider of a nozzle/valve system. Micro Aerospace Solutions (MAS), of Melbourne, Florida, who has a history of developing thruster systems for micro satellites, will manufacture the MR SAT designed nozzle/valve assemblies. This system can be used with either the xenon or R-134a propellants, which are under consideration. The details of the nozzle and thruster design for MR SAT are detailed below. The MR SAT nozzle design is displayed in Figure 4.10.1. Figure 4.10.1 MR SAT Nozzle Design The valves will serve two primary functions in the propulsion system design. The first will be to hold and release the propellant through the nozzle with the required timing. The second use will be a safety feature providing a physical interruption between the propellant tank and nozzles inhibits. The valves, as stated by the Safety Policy and Requirements document NSTS 1700.7B, will act as the inhibitors within the propulsion system. The internal workings of the valve are displayed in Figure 4.10.2.
Figure 4.10.2: Lee Co INKX0512050A Valve - Internal Design It is required that three mechanically independent flow control devices be implemented in series to prevent a catastrophic hazard in the case of an attempt of premature firing. The valve and nozzle schematic is shown in Figure 4.10.3. Figure 4.10.3 Propulsion Valve Inhibit Design Schematic 4.11 Nozzle and Thruster Spacecraft Integration An effective configuration of twelve thrusters has been designed for MR SAT. Discussions are being held with the Structures subsystem as well as Power and Communications to determine the finalized positioning and mounting of the nozzle/valve system, but it appears at this early stage that the current design can be implemented. The final thruster positioning will be modified depending on the finalized cm location of MR SAT with MRS SAT detached, but the proposed preliminary layout can be seen in Figure 4.8.3.1. This setup will provide the mission the necessary detumble and the three-axis attitude control when MR SAT is in formation flight phase. 4.12 Theory and Testing The theoretical values of thrust and specific impulse only consider ideal conditions (3). The thrust value at 20 ºC (68 ºF) and 170.3 kpa (24.7 psia) was calculated as ideal, then a correction factor was multiplied through to account for the actual case. The final thrust calculated was 62.79 mn and, compared to the actual testing values, is much greater than what the tests show. The tests were performed in a vacuum chamber with the thruster mounted on a scale so the value of the thrust achieved could be noted. Several variations of temperature and pressure were tested
and the thrust force as well as the temperature and pressure of each of these trials were noted. The thrust value at 20 ºC and 20 psia (the pressure was varied by 10 psia increments) was about 20-30 mn. This is an initial testing value and the exact value will be documented when the testing is completed. 4.13 References [1] Missouri University of Science and Technology Satellite Team, 04-011 Propellant Distributor, Rev -. Jan 2010. [2] Missouri University of Science and Technology Satellite Team, 16-04-001 Propulsion Tube Manufacturing-Assembly, Rev -. Feb 2009. [3] C. R. Seubert, Refrigerant-Based Propulsion System for Small Spacecraft, Master s Thesis. University of Missouri-Rolla. May 2007.