propulsion gt3 final submission

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  1. 1. Propulsion System Aaron Spanner, Annie Lin, Edwin Romero, Diana Alsindy, Jae Oh, Jansen Quiros i. Introduction The Students for the Exploration and Development of Space at the University of California, San Diego (SEDS@UCSD) Chapter is researching additively manufactured propulsion thrusters. The first monopropellant engine, named Callan, was designed and tested by SEDS@UCSD. This engine includes an additively manufactured diffuser section, reaction chamber, and nozzle module, which are printed in separate pieces to be bolted together. The catalyst pack is not additively manufactured and is assembled through traditional manufacturing processes. Presented in this document is the ground testing of Callan engine in Purdue University prior to Ground Tournament 3, the Propellant Feed system Ground testing at Open Source Maker Labs in Vista, CA, as well as the design changes of the first iteration of the thruster. The 3D printed thruster aims to promote the Cube Quest Challenge and its mission of completing a lunar orbit in December 2018. (SEDS@UCSD) will be approaching this challenge by designing Triteia: a 6U configuration CubeSat designed to achieve a polar lunar orbit from a trans-lunar injection trajectory through the SLS EM-1 secondary payload deployment sequence. Triteia transforms from an unassuming, ordinary CubeSat to an autonomous and intelligent power management system with a state-of-the-art additively manufactured high test hydrogen peroxide (H2O2) propulsion unit that allows for extraordinarily fast in-space translational speeds. This is an unprecedented level of detail in design as Triteias propulsion system is built with direct metal laser sintering (DMLS) techniques that manufacture the thruster as 3 separate modules: the diffuser plate, reaction chamber, and the nozzle, thereby allowing for unlimited customization and total aesthetic control. SEDS@UCSD has embarked on designing an entirely new, Hydrogen Peroxide (H2O2) monopropellant propulsion system with never-before seen Delta-V and thrust capabilities onto the 6U Triteia CubeSat. Upon anticipation for securing the Lunar Derby prize in NASAs Cube Quest Competition, this sophisticated propellant structure will be one of the pioneers of its kind to evolve away from the conventional electric propulsion thrusters. 1
  2. 2. Table of Contents i. Introduction 4. Subsystem Verification 1. System Requirements 4.1 Callan Testing 1.1 Volume 4.1.1 Theoretical Results 1.2 Safety 4.1.2 Experimental Results 1.3 Material 4.2 Brassboard Testing 1.4 Thermal 4.2.1 Theoretical Results 1.5 Storage and Handling 4.2.2 Experimental Results 1.6 Electrical Power 2. Subsystem Design 2.1 Nomenclature 2.2 Propulsion System Configuration 2.3 Component List 2.4 Thruster Design Changes 3. Subsystem Analysis 3.1 Delta-V Analysis/Propellant Mass Budgets 3.1.1 Constant Mass Flow Rate 3.1.2 Variable Mass Flow Rate 3.2 Propellant Bladder 3.3 Pressure Vessel Sizing and Stresses 3.3.1 Structural Analysis 3.4 Lines and Fittings 3.5 Flammability of Materials 2
  3. 3. I. System Requirements Category Requirement: Verification Method: 1.1 Volume The propulsion system was allotted 3300 of volume within the 10cm xcm3 20 cm x 30 cm structure. The CAD of our system fits within the chassis that is designed to fit the allotted volume. The assembled system will be measured to verify. [SPS.SPL.005] Dispenser/Payload Cleanliness. Payloads shall comply with the GSDO-RQMT-1080, Cross-Program Contamination Control Requirements document for visibly clean standard level. The system meets the payload cleanliness design challenge, because all fittings, valves, tubes, and tanks that make up the system will be subject to oxidizer cleaning by AstroPak and subsequent inspection. [SPS.SPL.006] Payload Storage: Payloads shall be storable up to 6 months under conditions listed in Table 3-2. The system meets the storage design challenge based on the analysis performed because the materials chosen for the propellant storage are class 1 compatible with hydrogen peroxide. Aclar 22C was chosen for the fuel bladder material because it has a low AOL with higher concentrations of HTP. 1.2 Safety [IDRD.3.4.4.3] Pressurized systems with lines and fittings less than 1.5 inches diameter (outside diameter (OD)) must have a Factor of Safety (FOS) for Pressure of 2.0x MDP for proof and 4.0x MDP for ultimate. All lines and fittings meet the minimum proof FoS of 2x MDP and ultimate FoS of 4x MDP based on the analysis performed, and will undergo acceptance and qualification tests during GT4. [IDRD.3.4.4.3] Pressurized systems with reservoirs / pressure vessels must have a FOS of 1.5 x MDP for proof and a 2.0x MDP for ultimate. All reservoirs/pressure vessels meet the minimum proof FoS of 1.5x MDP and ultimate FoS of 2x MDP based on the analysis and simulations performed, and will undergo acceptance and qualification tests during GT4. [IDRD.3.4.4.3] Pressurized systems for other components and their internal parts which are exposed to system pressure must have a FOS of 1.5x MDP for proof and 2.5x MDP for ultimate. All other components ie solenoid and service valves meet the minimum proof FoS of 1.5x MDP and ultimate FoS of 2.5x MDP based on vendor data, and will undergo acceptance and qualification tests during GT4. 1.3 Material All material and components in contact with H2O2, must not decompose with H2O2 unless otherwise intended. The tanks and tubing are made out of Al 6063 and the bladder is made of Aclar 22C, and these have great compatibility with 90% HTP. Furthermore, all the valves in the system are made of stainless 316, which also has good compatibility with peroxide. Only the catalyst bed, made of silver and nickel, is meant to decompose the peroxide. The amount of decomposition is verified by our analysis. 3
  4. 4. [IDRD 3.4.8.5] Materials and processes shall be in accordance with NASA-STD-6016. For materials that create potential hazardous situations as described in the paragraphs below and for which no prior NASA test data or rating exists, the payload developer will present other test results for SLS Program review or request assistance from the MSFC in conducting applicable tests. All materials and processes are in accordance with NASA-STD-6061. An oxygen compatibility assessment was made for all flammable materials. A materials offgassing assessment was performed for all materials using the MAPTIS database. For those materials where no flammability or offgassing data was provided, we will request additional assistance from NASA MSFC. 1.4 Thermal [Secondary Payload Users Guide] The payload must be able to endure surface temperatures ranging from 200F with direct Sun on one side to -143F with deep space on the other side. The valves, pressure transducers, and thermocouples were selected with vendor data that show an operating range encapsulating 200F to -143F. This will be verified by testing in an environmental chamber with simulated temperature cycling. The secondary payload integrated with the deployer inside the MSA is not expected to radiate heat or contribute to the thermal loading for the SLS vehicle. The propulsion system has three valves in series, giving it triple redundancy to ensure that no peroxide will leak out of the system nor will it allow the thruster to fire. No other part of our sub-system can radiate heat. 1.5 Storage and Handling Secondary payload design must be compatible with storage of up to six months under launch site environments while awaiting integration into the vehicle. Storage temperatures can range from 65-85F. Other environmental conditions are discussed in the following sections. The system meets the storage design challenge based on the analysis performed because the materials chosen for the propellant storage are class 1 compatible with hydrogen peroxide. Aclar 22C was chosen for the fuel bladder material because it has a low AOL with higher concentrations of HTP. This amount of decomposition is verified by our analysis. 1.6 Electrical Power The valves and data acquisition instruments of the propellant feed system were chosen to keep the electrical power consumption below 8 watts due to battery and solar power constraints. The safety / redundancy valves can not draw extra power. Vendor data for the PTs and Thermocouples show a maximum power draw of 0.8W. The vendor for solenoid valves will consume a maximum of 0.25W each. This falls well below the 8W allotted. During final system testing the power draw will be verified in an environmental testing chamber. 4
  5. 5. II. Subsystem Design A subsystem testing on component was performed on the first iteration of Callan at Purdue University in West Lafayette, Indiana using a test system provided by the Maurice J. Zucrow Propulsion Laboratory. The purpose for conducting testing on the thruster was to capture data about the chamber pressure, temperature, heat flux across the chamber wall and the resulting pressure drop across the catalyst pack. This suite of information from the extensive testing of Callan during startup and steady state operation not only helps prove the flight technology level of additively manufactured thrusters, but also helps with the analysis of other systems on-board the Triteia CubeSat as a result of thruster operation in space. The test matrix outlining the experiments required to obtain the data mentioned included a series of short pulses to characterize and determine the ideal startup procedure for the thruster, as well as a series of long duration burns to mimic the burn operations while the thruster is on the mission. A system level testing was performed at Open Space Maker Labs (OSML) for the propellant feed system to verify the design of the Triteia 6U Cube Satellite propellant feed system, an assembled and tested brass board system for Callan. Callan was subjected to iterative water flow testing to simulate the blowdown system that is projected

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