Transcript
Page 1: Propulsion GT3 Final Submission

 

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 (H 2 O2 ) propulsion unit that allows for extraordinarily fast in-space translational speeds. This is an unprecedented level of detail in design as Triteia’s 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 (H 2 O2 ) 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 NASA’s 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.

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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

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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 H2 O2 , must not decompose with H 2 O2 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.

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[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 User’s Guide] The payload must be able to endure surface temperatures ranging from 200F <TBR-001 > with direct Sun on one side to -143F < TBR-001 > 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.

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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 to occur during the actual flight.

Figure 2.1 (left) 3D Printed Thruster. (middle-left) 3D Printed Front Diffuser Plate with Truncated Nozzle. (middle-right) Purdue System Set-Up For Verification Testing. (right) Systems Verification Testing For Propellant Feed System Cold Flow

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2.1 Nomenclature

Q= Volumetric Flow Rate ṁ= Mass Flow Rate

=Dry massmf

=Propellant massmp =Wet massmo

⍴= Density R e = Reynold’s Number U= Ullage Volume P= Pressure P0n = Stagnation pressure at chamber μ= Viscosity f= Friction Factor l= Length of Straight Pipe g= Acceleration Due to Gravity k= Specific Heat Ratio E = Young’s Modulus

D= Inner Diameter of Pipe 𝜉= Coefficient of Loss A*= Throat Area t= Thickness

= Residence timetΔ σ= Stress SF= Safety Factor F= Force Isp= Specific Impulse

= First Burn Timet1 =Burned off Massm

burn,1 =Second Burn Timet2

=Burned off Massmburn,2 = Stagnation TemperatureT 0

of Chamber R= Specific Gas Constant

2.2 Propulsion System Configuration As shown in the figure below, Triteia propellant feed system repositioned the thruster to the top to conserve space margins and free up space for a potential payload.

Figure 2.2. (left) Original Propellant Feed System Configuration. (right) Modified Propellant Feed System Configuration

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2.3 Component List Table 2.1: Component List

Component

Part Number Vendor Material

NDE Techniqu

e

Acceptance Testing

Qualification Testing (for FLT HWR)

Image

Solenoid Valve

IEPA2421141H Lee Company SS 316 Ultrasonic -Proof at 1.5*MDP

-Leak Test

-API 591 1

- API 6A   1

Service Valve For Press 99-V24000-01 Valcor

Al 2024-T85

11 Ultrasonic

-Proof at 1.5*MDP

-Leak Test

-API 591   1 - API 6A   1

Service Valve For H2O2 99-V24000-01 Valcor Al 6063 Ultrasonic

-Proof at 1.5*MDP

-Leak Test

-API 591   1 - API 6A   1

Pressure Relief Valve

PRFA2819700

L Lee Company Al 6063 Ultrasonic

-Proof at 1.5*MDP -Leak Test

-API 591   1 - API 6A   1

Pressure Vessel Custom Part Shwabel Al 6063 Ultrasonic

-Proof at 1.5*MDP -Leak Test

-Vibration - Cycle(1) - Leak Test - Burst

Fuel Bladder Custom Part Aero Tec

Laboratories Inc.

Aclar 22C Liquid Dye Penetrant Inspection

-Leak Test -Expansion

-Vibration - Cycle(1)

- Leak Test - Burst

1 "Qualification Standards on Performance Type Testing for Valves Used Top Side and in Facilitie." (2012): n. pag. Cameron. Web.

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Pressure Transducer 105C12

PCB Piezotronics SS 316

Radiography

-Check against

calibrated Pressure Gauge

-Vacuum Environment

Testing -Vibration -Leak Test

Thermocouple

Custom part Thermometrics Corporation

SS 316 Radiography

-Check against

calibrated Thermomete

r

-Vacuum Environment

Testing -Vibration -Thermal Cycling

Tubing TBD Kaiser Aluminum

Al 6063 Radiography

-Proof at 1.5*MDP

-Leak Test

-Vibration - Cycle(1)

- Leak Test - Burst

Fittings Custom Part United Titanium

Al 6063 Radiography

-Proof at 1.5*MDP

-Leak Test

-Vibration - Cycle(1)

- Leak Test - Burst

Thruster Custom Part Metal

Technologies Inc. (MTI)

Inconel 718

Ultrasonic -Hot Fire Test at Purdue

-Vibration -Leak Test -Vacuum

Environment Testing

Note: (1) - The cycle testing includes the information from Table XYX. Table 2.2: Qualification Cycle Test Requirements

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2.4 Thruster Design Changes Table 2.3: Thruster Design Changes

Design Changes

Description

3D Print Reduction

The holes built in the vertical direction were more susceptible to percent reductions in area after 3D printing. The hole size also had an effect on the percent reduction in hole area, showing that holes with smaller diameters tended to “shrink” more than larger sized holes. For the 2nd print iteration, the Internal holes were over-sized in the CAD based on test data to obtain correctly sized holes after printing.

Orifice Plate Deflection

The original assembly procedure for compressing the catalyst pack had the compression take place against the front diffuser plate, which was already attached to the chamber. This compression plastically deformed the diffuser plate. For the next Callan iteration the assembly process will have the compression take place against a block of scrap metal, and then the front diffuser plate will be attached after the compression has occurred. This will avoid the deflection seen in the image to the right.

Surface Roughness

The chamber roughness and inner nozzle surface will be machined to be as smooth as possible, smooth enough to create a seal with the baffles, ~Ra=16. Blue surfaces are the ones we wish to minimize surface roughness.

O-Rings Grooves/ Gaskets

The vermiculite gaskets used during testing began to disintegrate when in contact with HTP. The vermiculite will be replaced with 316 Stainless Steel gaskets that can withstand the 1300F temperature and the highly corrosive conditions of HTP during firing.

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Nozzle Truncation

The first iteration truncated print of Callan had an expansion ratio of 4.023, not 1.8 as intended. Throat radius is 0.0396” and exit plane radius was 0.0794”, but has been adjusted to 0.0531”.

III. Subsystem Analysis

3.1 Delta-V analysis/Propellant Mass Budgets 3.1.1 Constant Mass Flow Rate The values in the Delta-V budget table were calculated assuming a constant mass flow rate. Since the configuration of the propulsion system is a blowdown configuration and will in reality have a variable mass flow rate as tank pressures decay over time, the values in the table below are rough estimates subject to change once experimental values are obtained for the solenoid valve pressure drops and the catalyst pack pressure drop. Table 3.1: Overview of Delta-V budget

Mission Phase Maneuver Name GT2

Delta-V (m/s)

GT3 Delta-V (m/s)

Comments

Cruise Trajectory Correction

Maneuver (TCM) 46

30.41

Orbit Insertion Lunar Orbit Injection (LOI) 350 350

Error Margin 56.8 72.39

To be used for additional correction maneuvers, the protection of historic lunar

sites, etc.

Total 452.8 452.8 Grand total Delta-V for the

mission

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Table 3.2: Total Mass and Volume Budget

GT2 Raw Value GT2 Marginal Value

GT3 Raw Value

GT3 Marginal Value

Delta-V, V  Δ 396 m/s 452.8 m/s 380.41 m/s

452.8 m/s

Dry mass,  mf 6.482 kg 6.799 kg 7.25 kg N/A

Propellant mass, mp 1.910 kg 2.227 kg (16.6%) 2.136kg 2.351 kg

Wet mass, mo 8.392 kg 8.709 kg 9.386 kg 9.600 kg

First Burn Time,  t1 82.143sec 85.246sec 62.357 sec. N/A

Burned off Mass, mburn,1 0.248kg 0.257kg 0.188 kg N/A

Second Burn Time,  t2 550.446sec 571.203sec 635.318 sec N/A

Burned off Mass, mburn,2 1.662kg 1.726kg 1.921 kg N/A

Mass Flow Rate, m 0.00302 kg/s 0.00302 kg/s 0.00302 kg/s 0.00302 kg/s

Specific Impulse, Isp 156.27 sec N/A 156.27 sec N/A

3.1.2 Variable Mass Flow Rate The propulsion system of Triteia is a blowdown system, and as such, the internal tank pressure decays as fuel is depleted. This means that steady-state analysis does not fully describe the system. In order to account for the unsteady flow in the system, this quasi-static iterative method accurately describes the system and allows the calculation of Thrust, , and Delta-v curves that are functions of time.Isp Assumptions:

- H202 is an incompressible fluid - Depletion of the tank is isothermal (slow depletion) - Helium pressurant behaves as an ideal gas - Mass flow rate out of tanks is equal to the mass flow rate out of engine (mass

continuity) - Quasi-static and one dimensional flow - Neglect transient startup - Chamber temperature is adiabatic flame temperature of H202 decomposition - Velocity in engine chamber can be neglected (absolute conditions are same as

stagnation) - Isentropic nozzle with choked flow - Pressure drop across system determined experimentally

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Analysis: The change in ullage volume after a small enough residence time, = :tΔ t )(

i+1 − ti U t  Δ = ⍴

ṁ ∙ Δ In combination with the ideal gas law for an isothermal process:

(1)U UPtank,i i

= Ptank,i+1 i+1

Describe the pressure of the tanks after by:tΔ

(2)Ptank,i+1 =

P Utank,i i

(U + ∙Δt)1 ⍴ṁ

This equation still depends on the mass flow rate of the system, which for choked flow through an isentropic nozzle is described by:

 ṁ = A* ∙ Pchamber,i√ k

T R0( 2k+1)−

k+12(k−1)

With an experimentally determined hydraulic pressure loss ( ) across the solenoidPHL

valves and engine catalyst pack (from cold-flow testing):

Pchamber,i = P tank,i − PHL

Such that:

(3) ṁ = A* ∙ P( tank,i − PHL)√ k

T R0( 2k+1)−

k+12(k−1)

From equations (2) and (3), the pressure of the tanks at time i+1 can be determined from the pressure and ullage volume of the tanks at time i. Equation (1) can be used to find ullage volume at time i+1. This set of equations can be iterated through our MATLAB code until the propellant mass is depleted to generate vectors of tank pressure and ullage volume data. This data is then used to calculate the transient Thrust, , and Delta-v that are characteristic of the blowdown system. See testing Isp and verification section 4.2 for an analysis of the testing performed in order to obtain the experimental hydraulic pressure loss across the solenoid valves and catalyst pack. 3.2 Propellant Bladder The fuel containment bladder was designed to account for the volume constraint of the tank profile and the amount of HTP the bladder material would cause to decompose. See figure 3.1 for a schematic of the diagram. Table 3.3: Bladder Material and Active Oxygen Loss 2

Material HTP Concentration

AOL @ 30 C/week

AOL @ 66 C/week

Decomposed Mass [kg]

Decomposition Volume [cm3 ]

Aclar 22C 90% 0.175 1.2 0.2571 183.6428571

2 Ventura, Mark. "Long Term Storability of Hydrogen Peroxide - 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit (AIAA)." Long Term Storability of Hydrogen Peroxide - 41st  

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Table 3.3 shows information about the Active Oxygen Loss of Aclar 22C, which was chosen as the ideal propellant bladder material because the amount of liquid volume that decomposed was much smaller than any other compatible materials.

Figure 3.1. (Left) Outline of bladder within the pressure vessel. (Right) Similar bladder fabricated by Aero Tec Laboratories Inc.

3.3 Pressure Vessel Sizing and Stresses The values in the pressure vessel volume requirements table were also calculated for a constant mass flow rate. After the propellant volume was found, the ullage volume was found and losses resulting from the blowdown configuration were included. The losses found included the residual propellant volume and the amount of liquid propellant lost to vapor from decomposition. The amount of propellant that changed phase from liquid to vapor was calculated using the active oxygen loss of the material . 3

Note that these values are also rough estimates since the delta-V values that were used to find the volume parameters assume a constant mass flow rate. Table 3.4: Pressure Vessel Volume Parameters

Description Value [cm^3] Assumptions

Propellant Volume 856.872 10% propellant margin

Ullage Volume 285.624 25% of total tank volume

Residual Propellant Volume 34.274 4% of the propellant volume

Volume from Propellant Decomposition See Table 3.3

3 Ventura, Mark. "Long Term Storability of Hydrogen Peroxide - 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit (AIAA)." Long Term Storability of Hydrogen Peroxide - 41st

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3.3.1 Structural Analysis The Triteia pressure vessels were designed in accordance with ASME design equations: Cylinder Section The following ASME equation taken from ASME Boiler and Pressure Vessel Code Section III, Division 1 was used to determine the thickness of the cylindrical section for the vessels:

   t =   PR

S(E)+0.4P where

- P: Internal Pressure - R: Outside radius - S: Allowable Stress - E: Weld Joint Efficiency

ASME 2:1 Elliptical Head The following ASME 2:1 elliptical head thickness equation taken from ASME Pressure Vessel Design Section VIII, Division 1 was used to determine the thickness of the heads:

   t =   PDK

2(S)(E)−0.2P

where - P: Internal Pressure - D: Internal Diameter - K: factor where K = ⅙*[2+(D/2h)2 ] - h: Internal height of the head.Typical 2:1 elliptical head with 1000mm inside

diameter will have a height of 250mm - S: Allowable Stress - E: Weld Joint Efficiency

Based on the width and length of the chassis frame, the maximum internal radius available for the pressure vessels was:

  .0455 m (1.795 in)  R =   Using this radius value as our constraint, the following table was produced:

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Table 3.5: Pressure Vessel Design Parameters

Tank Section

MDP

Material FoS ASME Equation Inputs ASME Equation Outputs

Final Thickness

Cylinder Section

4.275 E-06

(620 psi)

Aluminum 6063

2

P = 8549499.043 (1240 psi) R = .0455 m (1.795 in)

S = 213737476.089 Pa (31 ksi)

E = 0.8

Min Required Thickness =.0023 m (.0905 in)

.00317 m (.125 in)

Elliptical Head

4.275 E-06

(620 psi)

Aluminum 6063

2 P = 8549499.043 (1240 psi) R = .0455 m (1.795 in)

S = 213737476.089 Pa (31 ksi)

E = 0.8

Min Required Thickness =.0023 m (.0905 in)

.00317 m (.125 in)

Note: Since the factor of safety for the shell is , which satisfies   1240 psi  P =   oS  2  F =   and exceeds the proof requirement in table 3-9 of SLS-SPIE-RQMT-018. Since the final thickness of the pressure vessels was set to .125 inches, the pressure at which the vessels begin to yield was estimated to be P = 11.435 MPa (1658.506 psi). Since the mounting extrusions seen on the elliptical heads were expected to add stress concentrations, a simulation was run to verify how much the extrusions decrease the burst pressure of the vessels. The mounting face of the extrusions were fixed, and an internal pressure of 1400 psi was applied to the internal tank walls. Figure 3.2: Simulation of tank at 1400 psi

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As seen in the above figure, the mounting extrusions create stress concentrations on the internal surface that weaken the vessel. However, the vessel does not yield at 1400 psi, and with a design pressure of 620 psi, the vessel maintains a FoS of 2.26, well above the needed FoS of 2. 3.4 Lines and Fittings Similar to the design of the Triteia vessel cylinder section, the following ASME equation taken from ASME Boiler and Pressure Vessel Code Section III, Division 1 was used to determine the thickness of the tubing for the Triteia propulsion system:

   t =   PR

S(E)+0.4P The outer diameter of the lines and fittings will be 0.25 in (R = 0.125 in) and the material is Aluminium 6063 which has a yield strength S of 31 ksi . The design pressure through 4

the lines is 533 psi. Therefore, with a FoS of 4.0 our calculations use an Internal Pressure P of 2132 psi. We are assuming a weld efficiency of 50% (E = 0.5) for Aluminum. The minimum thickness t is calculated to be:

  .016 in t = (2132 psi) (0.125 in)(31000 psi) (0.50) + (0.4)(2132 psi) = 0

Since the minimum wall thickness that our vendor carries is t = 0.018 in , our Aluminum tubes satisfy the safety requirements. 3.5 Flammability of Materials Table 3.6: Flammability Data for Metals 5

Material Lowest Burn Pressure Highest No-burn Pressure

Mpa psia Burn Length(in) Mpa psia Burn length (in)

Silver (pure) No Data >68.9 >10,000 0 Nickel 200 mesh No Data >68.9 >10,000 N/A Inconel 718 ≤3.4 ≤500 0.5-4.3 None 300 series Stainless Steel

0.4 200 0.1-1.3 0.8 111 0.1-0.9

Aluminum 6063 No Data

Since gaseous oxygen will be present in the system due to the decomposition of hydrogen peroxide propellant during storage and the production of oxygen exhaust

4 United States. ASM International. Aluminum Alloy 6063 . Alloy Digest, Inc., n.d. Web. 21 Sept. 2016. 5 Manual 36 (Safe Use of Oxygen and Oxygen Systems: Handbook for Design, Operation, and Maintenance: 2nd Edition), pp.18­21, Table 3­1, 3­2   

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when propellant contacts the silver catalyst, flammability analysis of materials under worst operating environment is crucial. According to MAPTIS Material Selection Database, the bladder material Aclar 22C has an “I” flammability rating in a 40% oxygen concentration and 10.2 psia environment in accordance to NASA-STD-6001B rating standard based on Test Rpt No. M105258-B tested in 1997. This non-metal data should be re-tested because it exceeds 10 years (Ref). Assistance from MSFC will be requested to conduct additional test. Aluminum 6063 is a class 1 compatible material with high concentration hydrogen peroxide. It is commonly used as the material for pipe, tubing and other uses(Ref). In the Triteia propulsion system, Aluminum 6063 tanks will contain hydrogen peroxide propellant under 533 psi tank pressure. Aluminum 6063 will be assumed flammable because no data could be found on MAPTIS and ASTM manual 36. The following table summarized the flammability analysis on Aluminum 6063 tubing and tank based on Oxygen compatibility assessment. Table 3.7. Bladder,Tank,Tubing OCA

 * PCT data assuming commercially pure aluminum. Aluminum 6063 is not found on table 3­1  Table 3.8. Lee Solenoid Valve OCA

  Other materials that will be exposed to gaseous oxygen include thruster chamber material Inconel 718 and catalyst pack mesh materials silver and nickel 200. 90%

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hydrogen peroxide decomposes into 0.7076 mol fraction of H 2 O and 0.2924 mol fraction of O2 . As the liquid propellant decomposes along the catalyst pack, the concentration of gaseous oxygen increases. The chamber pressure of the Callan thruster is required to be 125 psi with estimate of 173 psi pressure drop across the catalyst pack. As shown in the table above, the lowest burn pressure of Inconel 718 is 500 psia and 385 psi is well below the lowest burn pressure of Inconel 718 in 100% oxygen concentration environment, therefore it is safe to assume Inconel 718 thruster component is non-flammable in the system. Silver and nickel 200 both have highest no-burn pressure above 10,000 psia and are considered to be non -flammable in the triteia propulsion system.

IV. Subsystem Verification 4.1 Callan Testing The objective of this testing was to obtain data about the thruster’s pressure, temperature, thrust, the heat flux across the thruster walls, and flow velocity of the propellant through the lines. Confirming the thruster design and efficiency from the thrust values at ideal steady state operation would then prove the flight technology readiness level of additively manufactured thrusters. The ideal testing of the Callan thruster was to perform a series of burn sequences or pulse tests and determine what amount of mass flow rate sprayed at the catalyst pack and at what time intervals would produce the highest temperature increase in the catalyst pack for the least amount of propellant. As a result, several pulse tests would be performed in order to compare which pulse length and at what interval would be the most efficient. After confirming the most efficient pulse, the thruster would then be tested at 15 sec and at 82 sec in order to replicate the burn times that have been calculated during mission operation. These steady state tests would help SEDS@UCSD analyze the fatigue and thermal stresses experienced by the thruster during long operation. The test matrix is shown below. Figure 4.1: Callan Engine Setup at Purdue University June 17th, 2016 (left). Callan Engine Test Matrix (right)

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4.1.1 Theoretical Results The purpose of pulsing the Callan thruster is to warm and prime the catalyst pack for sustained operation. As a result, the most important data for the pulse tests would be the temperature data at various locations on both the inside and outside walls of the thruster. The locations deemed the most critical were the top of the chamber, the mid section of the chamber, and the bottom of the chamber. Based on academic literature regarding hot fire tests and monopropellant hydrogen peroxide thrusters, the correct sequence to prime a thruster similar in size to the Callan thruster would have pulse lengths ranging from 200-600 milliseconds with 5 second intervals between each pulse and repeating this process 6 times for every pulse. Despite this research, the required number and length of pulses to successfully prime the thruster was suggested to SEDS@UCSD by Purdue personnel to be specific to every engine and every test conducted on any given day. As a result, the pulse tests conducted at Purdue University were conducted by tracking the data from the thermocouple located at the aft diffuser plate. After the first pulse test, the real time graphs from the thermocouple would rise exponentially, then approach a peak value and finally decrease. Once the graph would begin decreasing, another pulse was conducted and the graph would again rise but now reach a higher peak than before. The pulse lengths were instead ½ second pulses. This process was followed for all 3 pulse tests conducted at Purdue. Based on the research of other hydrogen peroxide monopropellant rocket engines, it was expected that the temperature profile across the catalyst pack would increase as the distance from the front diffuser increased. That is, the temperature of the catalyst pack immediately following the front diffuser plate was expected to exhibit the lowest temperature. The maximum temperature of the catalyst pack would exist in the section just upstream of the aft diffuser plate. This temperature profile would exist during the transient stage of thruster operation. Once the thruster reached steady state from the priming of the catalyst pack, the catalyst pack would reach a uniform temperature profile across the catalyst pack at the decomposition temperature of 90% hydrogen peroxide at approximately 1400F or 1033K. 4.1.2 Experimental Results The experimental results for all three pulse tests conducted show a different temperature profile for the catalyst pack than was expected. Whereas the literature explains the highest temperatures to be at the aft diffuser plate, the tests showed the highest temperatures to be in the mid section of the two diffuser plates, followed by slightly lower temperatures at the front diffuser plate, and finally the lowest temperatures at the aft diffuser plate. This temperature profile was found true for all three pulse tests. Although this temperature profile was not expected, this data was confirmed by the infrared images and videos taken for each pulse test. Both pieces of

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equipment measured the maximum temperatures at the mid section for all three pulse tests.

Pulse Test 1 The first pulse test in the series of experimental tests served as a baseline to properly pulse and improve the pulse procedure of the Callan thruster. From the data collected, it was determined that the mid-section of the catalyst pack exhibited the highest temperature, followed by the front-section of the catalyst pack, and finally followed by the aft-section of the catalyst pack. The maximum temperature as evidenced by the data shown in Figure 4.3 (a) hovers around 527K or 489F. Furthermore, based on the results of the first pulse test it can be generalized that every pulse results in an increase in the temperature of the catalyst pack. Examining the mass flow rate throughout the first pulse test in Figure 4.3 (b), it can be noticed that the flow rates are excessively high. But, looking closer at the data, it can be seen that maximum mass flow rate occurs at about 1.4 grams per second (zeroed) when a pulse is actuated and returns to 0 grams per second after flow is shut off. Although no thermocouples were used to directly measure the varying catalyst pack temperatures, this temperature for the first pulse test was confirmed by the high speed footage. The high speed footage shows there was a clear exhaust exiting the truncated nozzle. Because hydrogen peroxide decomposes into water and oxygen when in contact with a reactive catalyst, a clear exhaust indicated that superheated steam (which is transparent) and oxygen (which is also transparent) was exiting the nozzle. As a result, these visual cues were a strong indication that full decomposition was taking place inside the thruster, as the images show.

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(a) (b) Figure 4.3. Plots of (a) Temperature vs. Time at different locations on the thruster and (b) Mass Flow Rate vs. Time from Pulse Test 1.

Pulse Test 2 The second pulse test produced much different results than the first pulse test. The figures included below show that now, small droplets were exiting the truncated nozzle during the start of the pulse and a liquid stream was exiting the nozzle during the end of the pulse, unlike during the first pulse test. In the second pulse test, the same temperature profile of the catalyst pack can be observed. The mid-section of the catalyst pack exhibits the highest temperature, followed by the front catalyst pack, and finally followed by the aft catalyst pack. However, the maximum temperature of the catalyst pack dropped to 385K or 233F. The general trend in heating up the catalyst pack applied in the second pulse test where a pulse resulted in an increase in the catalyst pack’s temperature. The mass flow rates show a maximum flow rate of 3.7 grams per second which is similar to the results of the first pulse test. The mass flow rate returns to 0 grams per second once the main valve for the test system is closed.

(a) (b) Figure 4.4. Plots of (a) Temperature vs. Time at different locations on the thruster and (b) Mass Flow Rate vs. Time from Pulse Test 2.

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Pulse Test 3 The third pulse test confirms the temperature profile of the catalyst pack. The mid-section of the catalyst pack operates at the highest temperature, followed by the front section, and finally followed by the aft section. In addition, after every pulse, the catalyst pack temperature increased, similar to the first and second pulse tests. The maximum temperature is consistent with the results of the second pulse test and occurs at 385K or 233F. Similar to the second pulse test, the maximum mass flow rate occurs at approximately 3.75 grams per second and returns to 0 grams per second when the main valve is shut off.

(a) (b)

Figure 4.5. Plots of (a) Temperature vs. Time at different locations on the thruster and (b) Mass Flow Rate vs. Time from Pulse Test 3.

Thermal images of the Callan thruster were taken during all three pulse tests. Figure 12 illustrates the maximum temperature achieved during the third pulse test. Examining the images, it is possible to see a consistency between the recorded data from the thermocouples and the infrared thermometer as both sources of data measure a maximum temperature around 355K or 179F. Figure 4.6: (Left to Right): Thermal Distribution During Pulse Test 3

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4.2 Brassboard Testing (Propellant Feed System Testing) In order to further verify the design of the Triteia 6U Cube Satellite propellant feed system, SEDS@UCSD assembled and tested a brass board system for Callan. Callan was subjected to iterative water flow testing to simulate the blowdown system that is projected to occur during the actual flight. This testing was performed in order to obtain experimental hydraulic pressure loss data to be used in the iterative MATLAB code discussed in section 3.1.2. The goal was to obtain a relationship between instantaneous tank pressure and the instantaneous pressure drop across system components such as solenoid valves and the catalyst pack. This information would consider the blowdown nature of the system. The brassboard PID, shown below, followed the actual flight plumbing and instrumentation diagram with the exception of thermocouples, service valves, and the two solenoids used in ensuring triple redundancy.

Figure 4.8 Press System and Propellant Feed System Testing PID

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The team assembled the brassboard in-house at OSML and mounted it vertically for testing. This included flaring all tubing, tightening all components, machining the mounts for the system, wiring of solenoids and thermocouples, and designing and implementing the LabVIEW and MATLAB codes for data acquisition and analysis. The pressurization system was modified slightly from a previous SEDS@UCSD project, and used a K-bottle of gaseous nitrogen for the pressurant gas. This setup can and will be used for further verification testing of the system. 4.2.1 Theoretical Results An average mass flow rate of 0.00302 kg/s is expected to ensure that enough thrust will be produced to reach a delta-v of 396.0 m/s. As a blowdown system, the thrust of the system is expected to decay exponentially as opposed to a pressure fed system with a constant pressure outlet. However, assuming steady-state flow, a close approximation can be determined. The head loss through the tubing and fittings were calculated to be 10 psi. The flight solenoid valves are expected to have a pressure drop of 75 psi each, and the catalyst bed was expected to have a nominal pressure drop of 173 psi. Given the total pressure drop of 10 psi through the tubing and fittings, 225 psi through the three valves and the 173 psi through the catalyst bed, the maximum design pressure for the pressure vessels on the flight system will be 533 psi. For the valves, Lohms law for liquids was used: 6

 L =I

KV√ S

H Where H is the differential pressure loss, L is the valve’s Lohm rating, I is volumetric flow rate, S is specific gravity, V is a viscosity compensation factor, and K is a units constant. The pressure drop across the lines and fittings was estimated using the following major and minor head loss equations:

( )( )  hL,  major = f l

D 2gV2  P (ρg)  Δ = h

L, major  

ξ( )  hL, minor =   2g

V2

P ξ ρV  Δ =   21 2

The Ergun Equation was used to generate a MATLAB script to calculate pressure drop across the catalyst pack. The relevant equations are: 7

6 http://www.leeimh.com/resources/engineering/p136_LohmLaws­LiquidFlow.htm 7 Ergun, S., Orning, A.A. "Fluid Flow Through Packed Columns", Chemical Engineering Progress. 48 89­94 (1952). 

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4.2.2 Experimental Results The pressurized system tested a range of pressures between 400 and 660 psi in the pressure vessels to determine the time dependency of pressure drop across the catalyst pack and valves. Pressure transducers were placed above the tanks, above and below the solenoid, and at the engine chamber in order to measure the pressure drop across the solenoid valve and the catalyst pack and relate them to tank pressure. Symmetric filling and pressurization of propellant was tested by using a single port to fill both tanks, and a separate single port to pressurize. The simultaneous release of the propellant through the system was tested using a single high pressure solenoid valve. The brassboard was tested at MDP’s of 400, 550, and 660 psi. The following table summarizes the pressure drop data for the three tests run: Table 4.1: Operating Pressures of Propellant Feed System

GT2 GT3

Item Estimated MDP (psi) Estimated MDP (psi) Tested MDP (psi)

Pressure Vessels 390 533 400, 530, 660

Estimated (psi)P  Δ Estimated (psi)P  Δ Tested (psi)P  Δ

Solenoid Valves (3) 60 75 88-62, 96-71, 111-81

Lines 10 10 27-9, 31-10, 34-12

Catalyst Pack 75 173 120-40, 160-45, 215-55

Estimated Chamber Pressure (psi)

Estimated Chamber Pressure (psi)

Tested Chamber Pressure (psi)

Engine 125 125 10-5, 12-6, 15-8

After reviewing this data and observing a consistently minimal chamber pressure, we decided to remove the catalyst pack and test again at 530 psi in order to determine if it was actually the catalyst pack creating this pressure drop to such low chamber pressures. When we ran this test we discovered the exact same pressure drop across

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the engine and the same chamber pressure, even with the catalyst pack removed. See the below plots of this data, and notice how they are practically identical despite the catalyst pack being installed in the first set but absent in the second set. Just as a note, the pressure spike seen just before 80 seconds represents the depletion of the water in the tanks and the transition to flow of the purely gaseous nitrogen pressurant.

Figure 4.9. Cold Flow with Catalyst pack Screens @550 psi (Left) Callan Head Loss (Right) Chamber Pressure.

Figure 4.10. Cold Flow without Catalyst pack Screens @550 psi (Left) Callan Head Loss. (Right) Chamber Pressure.

We concluded that the cause of this pressure loss that was independent of the catalyst pack was the deformation of the front diffuser plate caused by compression of the catalyst pack for the Purdue testing. The deformation caused the orifices in the diffuser plate to partially and fully close, thus reducing the effective open area of the diffuser. This reduction in area meant that a very large pressure drop occurred across the

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diffuser, and thus we could not create a high enough inlet pressure to obtain useful data on catalyst pack pressure drop. To correct this issue we need to test with the second print of our engine Callan. The second print will have the the design changes discussed in an earlier section, with changed assembly process to prevent the same deformation from occurring, and is expected to ship on September 22, 2016. We did not have this engine available for testing before the GT3 deadline so we were unable to find an experimental hydraulic pressure drop relationship with tank pressure to be used in the iterative blowdown MATLAB code discussed earlier. Our intention is to use the setup we currently have to test this new, un-deformed engine, with the purpose of obtaining the experimental pressure drop. Once this data is collected, the effects of the blowdown system can be included in our delta-V analysis and mass budget analysis.

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