Attachment B RST-PROP-SPEC-0124-.pdf

Effective Date: July 21, 2020

National Aeronautics and Space Administration

Goddard Space Flight Center Greenbelt, Maryland

RST-PROP-SPEC-0124, Revision -

Roman Space Telescope (RST), Code 448

Attachment B, Propellant Tank Specification

GSFC RST CMO

July 21, 2020

Released

RST Propellant Tank Spec RST-PROP-SPEC-0124, Revision -ii

Propellant Tank Specification Review/Signature/Approval Page

Prepared by:

Daniel Ramspacher

Approved by:

Jason Hylan

Electronic Approval available on-line in the RST CM Tool file://gs448share/WFIRST/01.0%20Project%20Management/1.4%20Configuration-Data%20Mgmt/Templates/Document%20Templates/TDMS/Templates%20(boilerplates)/Historical/in iii

Preface This document is a Roman Space Telescope (RST) Configuration Management (CM)-controlled document.

Note: Prior to May 20, 2020, the project name was Wide Field Infrared Survey Telescope

(WFIRST).

For the purposes of configuration management, the prefixes “WFIRST” and “RST” are completely interchangeable. For example, RST-MGMT-PROC-0024 is the same as WFIRST-

MGMT-PROC-0024.

Changes to this document require prior approval of the applicable Configuration Control Board (CCB) Chairperson or designee. Proposed changes shall be submitted to the RST CM Office (CMO), along with supportive material justifying the proposed change.

In this document, a requirement is identified by “shall,” a good practice by “should,” permission by “may” or “can,” expectation by “will,” and descriptive material by “is.”

Questions or comments concerning this document should be addressed to:

RST Configuration Management Office Mail Stop 448 Goddard Space Flight Center Greenbelt, Maryland 20771 iv

Change History Log

Revision Effective Date Description of Changes (Reference the CCR & CCB/ERB Approval Date)

Revision - July 10, 2020 Initial Release per CCR-0222 v

Table of TBDs/TBRs/TBSs [optional]

Item No. Location Summary Individual/ Organization

Actionee

Due Date vi

Table of Contents

1 INTRODUCTION

1.1 Purpose

1.2 Scope

1.3 Related Documentation

1.3.1 Applicable Documents [and Forms]

2 CONTRACT DESCRIPTION

2.1 Propellant Tank Description

3 FUNCTIONAL/PERFORMANCE REQUIREMENTS

3.1 Propellant Tank Functional/Performance Requirements

3.1.1 Internal Pressure

3.1.1.1 MEOP

3.1.1.2 Proof Pressure

3.1.1.3 Burst Pressure

3.1.1.4 Minimum Internal Pressure

3.1.1.5 Differential Pressure

Differential Pressure Capability Differential Pressure During Nominal Expulsion

3.1.1.6 Interface Tube Design Pressure

Proof Pressure Burst Pressure

3.1.2 Leakage

3.1.2.1 External Leakage

3.1.2.2 Internal Leakage

3.1.3 Expulsion Efficiency

3.1.4 Leak Before Burst

3.1.5 Diaphragm Cycles

3.2 Resource Allocations

3.2.1 Mass Allocation

3.3 Electrical Grounding

3.3.1 Mechanical Contact Resistance

4 PHYSICAL REQUIREMENTS

4.1 Interface Documentation

4.1.1 Mechanical Interface

4.2 Mass Properties

4.2.1 Component Masses

4.2.2 Center of Mass Location

4.2.3 Center of Mass Accuracy

4.2.4 Determination of Moments and Products of Inertia

4.3 Physical Envelope

4.4 Fluid Interfaces

4.4.1 Bonded Brackets

5 ENVIRONMENTAL REQUIREMENTS

5.1 Quasi-Static Acceleration

vii

5.2 Frequency Requirement

5.2.1 Stowed Fundamental Launch Frequencies

5.3 Vibration

5.3.1 Sinusoidal Vibration

5.3.2 Random Vibration

5.4 Shock

5.5 External Pressure

5.5.1 External Operating Pressure Range

5.6 Ground Environments

5.7 Thermal Requirements

5.7.1 Flight Interface Design Temperature Limits

6 CLEANLINESS

6.1 Surface Contamination

6.1.1 Surface Contamination Levels at Delivery

6.1.1.1 External Cleanliness

Particulate Contamination Molecular Contamination – Exposed Surfaces

6.1.1.2 Internal Cleanliness

6.1.2 Surface Contamination Generation

6.1.2.1 Particulate Generation

6.1.2.2 Molecular Generation

Material Selection Material Selection - Silicones

6.1.3 Cleanability

6.1.3.1 Cleanability – Sensitive Surfaces

6.1.3.2 Cleanability – Sensitive Surface Cleaning Methods

6.1.4 Transportation Cleanliness

7 DESIGN & CONSTRUCTION REQUIREMENTS

7.1 Parts, Materials & Processes (PMP)

7.1.1 Materials

7.1.1.1 Material Conductivity

7.1.1.2 Fluid Compatibility

7.2 Safety

7.3 Identification and Marking

7.4 Workmanship

7.4.1 Workmanship Standards

7.5 Reliability and Mission Lifetime

7.5.1 Mission Life

7.5.2 Shelf Life

7.6 Ground Handling

7.6.1 Propellant Tank Lifting Point

7.6.1.1 Lift Point Locations

7.6.1.2 Lift Point Capability

7.6.1.3 Lift Point Proof Testing

7.6.2 GSE Cleanliness

viii

8 MECHANICAL DESIGN REQUIREMENTS

8.1 Mechanical Factors of Safety

8.2 Fracture Control Requirements

8.3 Metallic Pressurized Lines and Fittings Burst Factor

8.4 Joints

8.4.1 Joint Edge Distance

8.4.2 Bonded Joint Allowable

8.4.3 Reserved

8.4.4 Fastened Joints

8.4.4.1 Fastener Locking

8.4.4.2 Fastened Joint Margin of Safety

8.4.4.3 Installation Torque Documentation

APPENDIX A ABBREVIATIONS AND ACRONYMS

List of Figures Figure 4-1: Tank Envelope Figure 4-2: Pressurant Tube Dimensions, all dimensions in inches Figure 4-3: Propellant Tube Dimensions, all dimensions in inches

List of Tables Table 5-1 Propellant Tank Generic Sine Vibration Environment Table 5-2 Random Vibration Limit Levels for Components Weighing Less Than 22.7 kg Table 5-3 ASD Limit Load Reduction Table 5-4 Minimum Workmanship ASD Level for Components Weighing Less than 45.4 kg. . 12 Table 5-5: Minimum Workmanship Reduction Table 5-6 Shock Limit Levels Table 5-7 Temperature Limits of the Loaded Propellant Tank Table 8: Temperature Limits of the Unloaded Propellant Tank Table 8-1 Design Factors of Safety(1) Table 8-2 Minimum Edge Distance

1 INTRODUCTION

1.1 Purpose

The Nancy Grace Roman Space Telescope (RST) is a mission responding to the 2010 National Research Council New Worlds, New Horizons (NWNH) Astronomy and Astrophysics Decadal Survey top priority recommendation in the large space mission category. The science program includes two dedicated investigations to tackle outstanding questions in dark energy research and exoplanet exploration, and includes a substantial General Observer program to enable targeted investigations of astrophysical phenomena to advance other goals from the Decadal Survey. A coronagraph instrument is included in the payload for purposes of advancing the present state of the art of coronagraph technology. This document defines the requirements for the Propellant Tank used for propellant storage for the RST Propulsion System.

1.2 Scope

This specification describes the mechanical, environmental, and verification testing requirements for space-qualified Propellant Tanks for the NASA Goddard Space Flight Center (GSFC) RST Mission. The RST Propulsion System requires four (4) propellant tanks.

1.3 Related Documentation

The latest versions of all documents below should be used. RST documents can be obtained from URL: https://ipdtdms.gsfc.nasa.gov.

1.3.1 Applicable Documents [and Forms]

The following documents and drawings in effect on the day this specification was signed shall apply to the fabrication and to the mechanical and environmental requirements of the Propellant Tank to the extent specified herein. In the event of conflict between this specification and any referenced document, this specification will govern, with the exception of the RST Propellant Tank Statement of Work (RST-PROP-SOW-0043), in which case the Statement of Work takes precedence.

Document Number Title

ASTM E-595-07

Standard Test Method for Total Mass Loss and Collected Volatile Condensable Materials from Outgassing in a Vacuum Environment

IEST-STD-CC1246E Product Cleanliness Levels And Contamination Control Program

MSFC-STD-3029A

Guidelines for the Selection of Metallic Materials for Stress Corrosion Cracking Resistance in Sodium Chloride Environments

NASA-STD-5019A Fracture Control Requirements for Spaceflight Hardware NASA-STD-5020A –w/ Change 1

Requirements for Threaded Fastening Systems in Spaceflight Hardware

NASA-STD-8719.24 NASA Expendable Launch Vehicle Payload Safety Requirements https://ipdtdms.gsfc.nasa.gov/

Document Number Title GSFC-STD-7000A General Environmental Verification Standard (GEVS) RST-SYS-PLAN-0051 WFIRST Contamination Control Plan MIL-PRF-26536G Performance Specification Propellant, Hydrazine JSC-SPEC-C-20C Water, High Purity, Specification for MIL-PRF-27401G Propellant, Nitrogen, Pressurizing Agent MIL-PRF-27407D Propellant, Helium, Pressurizing Agent MIL-PRF-27415C Propellant Pressurizing Agent, Argon

AWS D17.1 Specification for Fusion Welding of Aerospace Applications

MSFC-STD-3029

Guidelines for the Selection of Metallic Materials for Stress Corrosion Cracking Resistance in Sodium Chloride Environments

NASA-STD-5001B Structural Design And Test Factors Of Safety For Spaceflight Hardware

NASA-STD-6016A Standard Materials and Processes Requirements for Spacecraft NASA-HDBK-7005 Dynamic Environment Criteria NASA-STD-7001 Payload Vibroacoustic Test Criteria RST-PROP-SOW-0043 RST Propellant Tank Statement of Work

2 CONTRACT DESCRIPTION

2.1 Propellant Tank Description

The RST Propulsion Subsystem will use four (4) propellant tanks. Each propellant tank will have a separate pressurant and propellant compartments, and will store 284 kg of propellant to feed continuous gas free propellant to the thrusters.

Please refer to the Statement of Work (SOW, RST-PROP-SOW-0043) for a complete list of all hardware, verification testing, test reports, and analyses that should be delivered with or before the Propellant Tank flight hardware.

3 FUNCTIONAL/PERFORMANCE REQUIREMENTS

This section defines the functional and performance requirements for the Propellant Tanks as defined in Section 3.1.

3.1 Propellant Tank Functional/Performance Requirements

3.1.1 Internal Pressure

3.1.1.1 MEOP

The Maximum Expected Operating Pressure (MEOP) shall be 400 psia.

3.1.1.2 Proof Pressure

Proof pressure capability of the propellant tank shall be at least 500 psia.

3.1.1.3 Burst Pressure

Burst pressure of the propellant tank shall be at least 600 psia.

3.1.1.4 Minimum Internal Pressure

The propellant tank shall be capable of withstanding an internal vacuum differential of 14.7 psi without any material yielding.

3.1.1.5 Differential Pressure

Differential Pressure Capability

The propellant tank shall be capable of at least 150 psi differential between the pressurant and propellant ports at End of Life (EOL).

Differential Pressure During Nominal Expulsion The propellant tank shall have less than 15 psid pressure differential between the static pressures at the pressurant and propellant ports while expelling at a hydrazine flow rate of 0.1904 lbm/s.

3.1.1.6 Interface Tube Design Pressure

Proof Pressure

The interface tube design proof pressure capability of the propellant tank shall be at least 600 psia.

Burst Pressure The interface tube burst pressure of the propellant tank shall be at least 1600 psia.

3.1.2 Leakage

3.1.2.1 External Leakage

The external leakage rate of the propellant tank shall not exceed 10-6 sccs of gaseous helium across the entire operational pressure range.

3.1.2.2 Internal Leakage

Internal leakage across diaphragm shall not exceed 0.13 sccm of gaseous helium with the helium applied to the pressurant side of the tank and a pressurant differential of 100 ± 10 psid.

3.1.3 Expulsion Efficiency

The expulsion efficiency shall be greater than 99.2% of the volume of the propellant mass in 4.2.1.

3.1.4 Leak Before Burst

The tank shall comply with Safe-Life design methodology based on fracture mechanics techniques per NASA-STD-8719.24.

3.1.5 Diaphragm Cycles

The diaphragm shall be qualified for at least 60 cycles.

3.2 Resource Allocations

3.2.1 Mass Allocation

Each propellant tank shall have a total empty mass less than or equal to 32.7 kg (72.0 lbm).

3.3 Electrical Grounding

3.3.1 Mechanical Contact Resistance

The DC resistance of the mechanical contact between two conductive mating surfaces (internal to the component, and at the spacecraft interface) shall be less than or equal to 2.5 mΩ DC resistance.

4 PHYSICAL REQUIREMENTS

The propellant tank shall be installed with the pressurant side on the top and the propellant side on the bottom.

4.1 Interface Documentation

The contractor shall use metric units when interfacing with NASA GSFC including any drawings, documents, models, except for the following cases:

• Heritage Hardware: Hardware that has been previously qualified, or of similar design heritage, may be specified in English units where use of metric equivalents would lead to additional cost to the program.

• Fasteners: Although bolt patterns will be defined using metric dimensioning, use of English fasteners (with hole dimensioning and tolerancing) is permitted.

• Angular Measurement: Angular measurement may be expressed in degree of arc or in an appropriate subdivision of degree of arc such as second of arc (arc-sec) when advantageous to application.

4.1.1 Mechanical Interface

The mounting interface shall be defined in the Mechanical Interface Control Drawing (MICD), which will be developed between the contractor and NASA GSFC.

4.2 Mass Properties

4.2.1 Component Masses

a. See Section 3.2.1 for total empty mass of each propellant tank (diaphragm in the expelled position).

b. A single propellant tank shall be able to contain a minimum of 284 kg of Hydrazine at

Beginning Of Life (BOL) at a fill fraction of 77% at the high temperature condition,

c. The total empty mass of each propellant tank (diaphragm in the expelled position) shall be measured to within 50 grams.

4.2.2 Center of Mass Location

The contractor shall define the center of mass for an empty tank (diaphragm in the expelled position) in the Mechanical ICD.

4.2.3 Center of Mass Accuracy

The center of mass for an empty tank (diaphragm in the expelled position) shall be determined to within ±3 mm relative to an external reference.

4.2.4 Determination of Moments and Products of Inertia

The final moments of inertia of each propellant tank shall be calculated to within 1%.

The final products of inertia of each propellant tank shall be calculated to within 5%.

4.3 Physical Envelope

The outside dimensions of each propellant tank, including mounting flanges and interface tubes, shall not exceed the volume envelope per Figure 4-1 under all pressurized and unpressurized conditions. The propellant tank shall fit within and mount to the tank stand described in the Interface Control Document (ICD) Drawing #2222818 in its flight configuration. This drawing is available upon request by the vendor. The gas and liquid feed tubes shall be aligned with each other and a single mounting hole when projecting their centerlines around tangential around the tank shell. All dimensions in inches and considered nominal unless otherwise noted.

Figure 4-1: Tank Envelope

4.4 Fluid Interfaces

The pressurant fluid interface shall be 0.25”+0.004/-0.000 inch (6.35 mm ±0.1016 mm) outer diameter with a 0.028”±0.002 (0.7112 mm ±0.1016 mm) wall thickness, and the material shall be SS-304L. The dimensions of the interface tubes shall meet the dimensions shown in Figure 4-2.

Figure 4-2: Pressurant Tube Dimensions, all dimensions in inches

The propellant fluid interface shall be 0.375”±0.004 (9.525 mm ±0.1016 mm) inch outer diameter with a 0.028”±0.004 (0.7112 mm ±0.1016 mm) wall thickness, and the material shall be SS-304L. The dimensions of the interface tubes shall meet the dimensions shown in Figure 4-3.

Figure 4-3: Propellant Tube Dimensions, all dimensions in inches

4.4.1 Bonded Brackets

Each propellant tank shall have one (1) pressurant tube bracket bonded to the tank shell in accordance with the location to be determined by COR once the tank design and geometry has been completed. The bonding process shall be in accordance with the SOW (RST-PROP-SOW- 0043).

5 ENVIRONMENTAL REQUIREMENTS

Environmental design requirements are specified in this section.

Each propellant tank shall meet its performance requirements defined in Section 3.0 during and after exposure to the environments specified in this section.

5.1 Quasi-Static Acceleration

Quasi-static acceleration represents the combination of steady-state accelerations and the low frequency mechanically transmitted dynamic accelerations that occur during launch.

Each propellant tank, when loaded to its worst-case fill fraction and hard-mounted at its interface, shall be designed to withstand the quasi-static Design Limit Loads (DLL) of 9.5g.

5.2 Frequency Requirement

5.2.1 Stowed Fundamental Launch Frequencies

Each propellant tank, loaded to its worst-case fill fraction, shall have a fundamental frequency greater than 50 Hz when hard mounted at its interface.

Note:

1. Requirements are met assuming rigid stiffness to restrained Degree of Freedom (DOF).

2. Components with first fundamental frequency greater than 75 Hz may verify through low level signature tests. Components with a frequency less than 75 Hz are required to perform modal testing and provide a test correlated FEM for dynamic analysis. Requirements for the FEM are in the SOW.

3. Frequency requirements apply to all modes with modal effective mass fraction (MEMF) >5% of FEM mass. MEMF is compared to the total mass of the element/subsystem/component, not against higher levels of assembly. e.g. OBS modes MEMF is compared to OBS mass

4. For modes with MEMF between 5% and 10% that do not meet the frequency requirement(s), additional assessments can be made by the observatory systems group to determine the effects of the mode on observatory dynamics and subsystem/element design limit loads.

5. Components/subsystems verification is required through low level signature tests.

6. Failure to meet frequency requirement requires a test correlated Finite Element Model

(FEM) be supplied for dynamic analysis. Requirements for the FEM are in the SOW, Section 3.5.

7. Failure to meet frequency requirements may result in higher DLL than specified herein.

5.3 Vibration

5.3.1 Sinusoidal Vibration

Each propellant tank, loaded to its worst-case fill fraction and mounted to the government-furnished tank stand specified in RST-PROP-SOW-0043, shall withstand the sinusoidal vibration levels defined in Table 5-1.

Note:

1. Levels defined are acceptance levels. Analysis is conducted against acceptance with the safety factors defined in Table 8-1. Testing factors are defined in SOW.

2. Peak levels at the low end of the frequency range (5 - 20 Hz) may be ramped up as needed to accommodate table limitations.

3. The sine sweep vibration levels shown in Table 5-1 are defined at the interface of the element/subsystem.

4. Input levels may be notched to limit the test specimens CG response to 1.25 times its

Design Limit Load outlined in Section 5.1.

5. Test verification is required over the range 5 to 50 Hz. Analytical verification is required from 50 to 100 Hz.

6. Elements/Subsystems/components with a first fundamental frequency equal to or greater than 75 Hz can forgo sine vibration testing from 5 to 50 Hz, while verification through analysis of the 50 to 100 Hz is still required.

7. Subsystems/components with a fundamental frequency equal to or greater than 150 Hz can forgo the sine vibration environment.

Table 5-1 Propellant Tank Generic Sine Vibration Environment

Axial Direction Sine Vibration Spectrum (Limit level)

Transverse Direction Sine Vibration Spectrum (Limit level)

Freq (Hz) g's Freq (Hz) g's

5 1.3 5 2.2

10 1.3 20 2.2 19 2 25 6.5 28 7.5 40 6.5 100 7.5 50 1 100 1

Levels may be notched such that the interface forces do not exceed 1.25 times the design limit load outlined in Section 5.1. Peak levels at the low end of the frequency range (5 – 20 Hz typically) may be ramped up as needed to accommodate shaker table displacement limitations.

5.3.2 Random Vibration

The propellant tank, loaded to its worst case fill fraction and mounted to the government-furnished tank stand specified in RST-PROP-SOW-0043, shall withstand the random vibration environment in Table 5-2 applied at the interface to the propellant tank.

The random vibration inputs are based on GEVS and may be notched to limit the test article’s cg response to 1.25 times DLL. Notched inputs must envelope minimum workmanship levels in order to screen for design or manufacturing flaws. A delicate optic, detector, sensor, etc., which is shown analytically to have negative margin at minimum workmanship levels may notch further at those frequency bands with project concurrence.

Table 5-2 shows the limit level random vibration environment for components with mass less than

22.7 kg. For components weighing more than 22.7 kg, relief can be provided in accordance with Table 5-3. Levels defined are acceptance levels. Analysis is conducted against acceptance with the safety factors defined in Table 8-1. Testing factors are defined in the SOW. Table 5-4 defines minimum workmanship for components weighing less than 45.4 kg. For components weighing more than 45.4 kg, relief can be provided in accordance with Table 5-5. Random vibration specs may be updated based on acoustic analysis.

During testing, force limiting control is recommended and the control method must be approved by the NASA/GSFC COR.

Table 5-2 Random Vibration Limit Levels for Components Weighing Less Than 22.7 kg

Frequency (Hz) ASD Limit Level (g2/Hz)

20 0.013 20 - 50 +6 dB/oct

50 - 800 0.08 800 - 2000 -6 dB/oct

2000 0.013 Overall 10.0 grms

Table 5-3 ASD Limit Load Reduction

ASD Limit Load Reduction for Flight Hardware Weighing More Than 22.7 kg(1, 2, 3) dB reduction =10 log(W/22.7) ASD (50 - 800 Hz) = 0.08 x (22.7/W)

(1) W represents weight.

(2) Slopes are maintained at plus and minus 6dB/oct for components weighing up to 59.0 kg.

Above 59.0 kg, slopes are adjusted to maintain an ASD level of 0.01 g2/Hz at 20 and 2000Hz.

(3) For components weighing more than 182.0 kg, levels are maintained at the 182 kg level.

Table 5-4 Minimum Workmanship ASD Level for Components Weighing Less than 45.4 kg.

Frequency (Hz) ASD Level (g2/Hz)

20 0.01

20 - 80 +3 dB/oct 80 - 500 0.04

500 - 2000 -3 dB/oct 2000 0.01

Overall 6.8 grms

Table 5-5: Minimum Workmanship Reduction

ASD Minimum Workmanship Reduction for Flight Hardware Weighing More Than 45. kg(1, 2, 3) dB reduction =10 log(W/45.4) ASD (50 - 800 Hz) = 0.04 x (45.4/W)

(1) W represents weight.

(2) Sloped portions of the spectrum are maintained at plus and minus 3/dB/oct.

(3) Workmanship spectrum must be equal to or greater than 0.01 g2/Hz, i.e., for components whose weight is greater than 182.0 kg, the workmanship spectrum is 0.01g2/Hz from 20 to 2000 Hz.

5.4 Shock

Each propellant tank, loaded to its worst case fill fraction and mounted to the government-furnished tank stand specified in RST-PROP-SOW-0043, shall be designed to meet its performance requirements after being subjected to the shock environment in Table 5-6, applied at the propellant tank interface to the RST spacecraft structure. Levels defined are acceptance levels. Testing factors are defined in Table 5-6.

Table 5-6 Shock Limit Levels

Frequency (Hz) Limit Level 100 100 1435 1750 10000 1750

Note:

A shock susceptibility and attenuation assessment is to be performed on the propellant tank.

This assessment can be based on past shock tests or other relevant information. The shock assessment will be used to determine whether the shock environment is low risk for the propellant tank. If the environment is low risk, then the shock test may be deferred to the Observatory level with concurrence from the NASA/GSFC COR. Levels defined represent limit level. Analysis is conducted against limit level with FS defined in section 8.1. For testing factors see section 9.4.2

5.5 External Pressure

5.5.1 External Operating Pressure Range

Each propellant tank, loaded to its worst case fill fraction, shall be designed to meet all performance requirements while operating over a pressure range of 1.08 x 105 N/m2 (813 Torr) to 1.3 x 10-12 N/m2 (1 x 10-14 Torr).s

5.6 Ground Environments

Each propellant tank shall meet all of their performance requirements during exposure to air temperatures between +5 and +30 degrees C and relative humidity between 30% and 70%.

5.7 Thermal Requirements

5.7.1 Flight Interface Design Temperature Limits

Each propellant tank shall be capable of surviving indefinitely when its temperatures are within the limits shown in Table 5-7 without damage or permanent performance degradation.

Table 5-7 Temperature Limits of the Loaded Propellant Tank

Minimum Temperature (ºC) Maximum Temperature (ºC) Operational (In Spec) 12 40 Protoflight/Qualification 12 40 Survival 8 45

Table 8: Temperature Limits of the Unloaded Propellant Tank

Minimum Temperature (ºC) Maximum Temperature (ºC) Operational (In Spec) 10 50 Protoflight/Qualification 8 60 Survival 5 65

6 CLEANLINESS

The requirements in this section ensure the cleanliness of each propellant tank at delivery, so as not to adversely affect its own performance, as well as not be a source of contamination to other items, including not generating contaminants following delivery in excess of that permitted below by virtue of its design, materials of construction, or operation.

6.1 Surface Contamination

6.1.1 Surface Contamination Levels at Delivery

6.1.1.1 External Cleanliness

Particulate Contamination If each propellant tank has surfaces exposed to the space environment, all such surfaces shall meet a particulate surface cleanliness level 300 per IEST-STD-CC1246E on all external surfaces, or equivalent.

Molecular Contamination – Exposed Surfaces If each propellant tank has surfaces exposed to the space environment, all such surfaces shall meet a molecular surface cleanliness level R1 of IEST-STD-CC1246E on all external surfaces, or equivalent.

6.1.1.2 Internal Cleanliness

The propulsion components, lines, and fittings shall be cleaned, and verified internally clean, to level 50R1 per IEST-STD-CC1246E prior to integration in the system.

6.1.2 Surface Contamination Generation

6.1.2.1 Particulate Generation

The propellant tank contractor shall not employ any of the following particle generating materials or processes into the propellant tank design or construction without prior approval by the NASA/GSFC COR:

• Paints prone to shedding due to large paint pigment molecules, overspray, poor adhesion, etc.

• Surfaces prone to corrosion or oxides because of a lack of corrosion protection or dissimilar metals in close contact.

• Fabrics with brittle constituents (e.g., composites, graphite or glass).

• Perforated materials when material is highly susceptible to tear propagation (e.g., multi-layer insulation (MLI)).

• Metal oxides (bare [untreated] aluminum and magnesium, iron, non-corrosion resistant steel, etc.).

• Braided metallic or synthetic wires, ropes, slings, etc. unless measures have been taken to contain any broken filaments or fibers (sheathing, sealing with polymers, covering, etc.).

• Woven materials especially cut or unfinished ends (metal braid, EMI shielding, lacing cord, expando sleeving), unless measures have been taken to prevent fraying or generation of particles (cut with a hot knife, seal with polymer, bag, etc.).

• Materials with thin films known to erode or crack or flake when subjected to normal handling (e.g., indium tin oxide [ITO] or other rigid or brittle semiconductor or ceramic coating on flexible substrates, Teflon, MLI, etc.).

• Foams, highly textured materials.

• Trapped debris in holes.

6.1.2.2 Molecular Generation

Material Selection

The propellant tank materials shall have a total mass loss (TML) less than 1.00% and a collected volatile condensable mass (CVCM) less than 0.10%, when measured in accordance with ASTM E-595 unless a materials usage agreement has been generated and approved by the NASA/GSFC

COR.

Material Selection - Silicones Silicones on external (to the spacecraft) surfaces shall not be exposed to the space environment unless approved by the NASA/GSFC COR. Silicones should be avoided or minimized. It is highly recommended that silicones be baked out at a high temperature prior to integration into the system to prevent extended bakeouts of the entire assembly.

6.1.3 Cleanability

6.1.3.1 Cleanability – Sensitive Surfaces

If any surfaces are not cleanable with Isopropyl alcohol and polyester wipes or light vacuuming, they shall be identified on the MICD.

6.1.3.2 Cleanability – Sensitive Surface Cleaning Methods

Alternate cleaning methods shall be identified and appropriate documentation provided for any surfaces that are not cleanable with Isopropyl alcohol.

6.1.4 Transportation Cleanliness

Materials and enclosure used for transportation and storage shall not generate molecular or particle contaminants or degrade the surface cleanliness of the item or adjacent items

7 DESIGN & CONSTRUCTION REQUIREMENTS

7.1 Parts, Materials & Processes (PMP)

7.1.1 Materials

Each propellant tank will be comprised of materials and processes in accordance with the requirements in the SOW, RST-PROP-SOW-0043

7.1.1.1 Material Conductivity

All parts should be passivated and mounting surfaces on the propellant tank shall be conductive as defined in Section 3.3.

7.1.1.2 Fluid Compatibility

All wetted materials shall be compatible with prolonged exposure to the following fluids:

• De-ionized water per JSC-SPEC-C-20

• Hydrazine per MIL-PRF-26536G, High Purity

• GHe per MIL-PRF-27407D

• Grade A, Ar per MIL-PRF-27415C

• Grade A, GN2 per MIL-PRF-27401G

7.2 Safety

All flight hardware shall meet the applicable sections of NASA-STD-8719.24 ELV Safety Requirements.

For pressure vessels and pressurized structures to be analyzed with linear elastic fracture mechanics, fracture properties shall be accounted for in material selection. These properties include fracture toughness; threshold values of stress intensity under sustained loading; sub-critical crack-growth characteristics under sustained and cyclic loadings; the effects of fabrication and joining processes**; the effects of cleaning agents, dye penetrants, coatings, and proof test fluids;

and the effects of inspection couplants or materials, temperature, load spectra, and other environmental conditions.

**- Instrument the welding process fixturing forces used to hold parts together for fusion welds, so that these residual stresses (when fixturing is removed) are properly accounted for in subsequent stress analysis.

7.3 Identification and Marking

Each propellant tank shall be permanently marked with the part number and a unique sequential serial number in the area designated on the interface control drawing in a manner to be approved by the NASA/GSFC COR.

All markings shall use alcohol proof ink.

7.4 Workmanship

7.4.1 Workmanship Standards

The workmanship standards and processes outlined in the SOW will be used.

7.5 Reliability and Mission Lifetime

7.5.1 Mission Life

Each propellant tank shall meet all performance specifications through three (3) years of ground testing and five (5) years of operation in space, following a three (3) month commissioning period.

7.5.2 Shelf Life

The propellant tank, when loaded to its worst-case fill fraction, shall not suffer any degradation in performance when stored for five (5) years either on the S/C or in bonded storage.

7.6 Ground Handling

7.6.1 Propellant Tank Lifting Point

7.6.1.1 Lift Point Locations

There shall be a minimum of one (1) lift point, with the exact location and number approved by the NASA/GSFC COR and documented in the ICD.

7.6.1.2 Lift Point Capability

Each Propellant Tank shall include an analysis showing that the lift points are capable of lifting the Propellant Tank mass (while empty) with the appropriate factors of safety (see Table 8-1).

7.6.1.3 Lift Point Proof Testing

Lift points and supporting structure comprised of composite materials, Beryllium, bonded joints and/or bonded inserts shall be proof tested with the appropriate factors of safety (see Table 8-1).

7.6.2 GSE Cleanliness

All Ground and Test support equipment shall be compatible with the flight component and the environment where the flight component or test component will reside (cleanroom, thermal vacuum chamber, vibration cell, etc.)

8 MECHANICAL DESIGN REQUIREMENTS

For additional guidance on the design and analysis of threaded fastening systems in NASA spaceflight hardware, consult NASA-STD-5020A, “Requirements for Threaded Fastening Systems In Spaceflight Hardware”.

8.1 Mechanical Factors of Safety

Positive Margin of Safety (MS) shall be demonstrated analytically for the Propellant Tank, including interfaces to MGSE, using the MS formula below with the appropriate Factors of Safety (FS) defined in Table 8-1 applied to flight limit loads.

Note:

1. MS is defined as follows:

𝑀𝑀𝑀𝑀 =

𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝑒𝑒 𝐿𝐿𝐿𝐿𝐿𝐿𝐿𝐿𝐿𝐿 𝐿𝐿𝐴𝐴𝐴𝐴𝐿𝐿 𝐴𝐴𝑜𝑜 𝑀𝑀𝐿𝐿𝑜𝑜𝑒𝑒𝑆𝑆𝑆𝑆

𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐴𝐿𝐿𝑒𝑒𝐿𝐿 𝐿𝐿𝐿𝐿𝐿𝐿𝐿𝐿𝐿𝐿 𝐿𝐿𝐴𝐴𝐴𝐴𝐿𝐿 𝐴𝐴𝑜𝑜 𝑀𝑀𝐿𝐿𝑜𝑜𝑒𝑒𝑆𝑆𝑆𝑆 ∗ 𝐹𝐹𝑀𝑀

− 1

2. Positive MS in joints that are both bonded and fastened is demonstrated assuming only the bond carries the load, only the bolt carries the load, or if the capability of the joint requires both the bond and fasteners, a detailed analysis is performed.

a. The adhesive portion of the joint is usually the stiffest load path and thus carries most of the load. In this scenario, the full strength of the bolted interface is not realized until after the failure of the adhesive.

3. Positive MS is demonstrated for adhesively bonded joints for all potential failure modes under all loading conditions and temperatures.

a. Failure modes may include cohesive failure in the adhesive, failure at the adhesive to adhered interface, and adherent failure. Composite adherent failure modes may include in-plane tension/compression, in-plane shear, peel, interlaminar shear, and interactions between failure modes.

b. Fiber orientation should be considered in the computation of strength for a bonded joint where one or more of the adherends are made of composite material.

c. Interaction between adhesively bonded joints in close proximity should be included when analyzing bonded joints.

d. Temperature stresses in bonded joints and sandwich panels should consider bond and consolidation cure temperatures.

e. Adhesive bond fillets and bondline thickness should be considered in bonded joint design and analysis under temperature loading.

Table 8-1 Design Factors of Safety(1)

Type of Hardware Static /Sine Random/Acoustic (4,5) Tested Metallic Yield 1.25(3) 1.6 Tested Metallic Ultimate 1.4(3) 1.8 Stability/Buckling Ultimate 1.4 1.8 Beryllium Yield 1.4 1.8 Beryllium Ultimate 1.6 2.0 Composite Ultimate(1) 1.5 1.9

Type of Hardware Static /Sine Random/Acoustic (4,5) Bonded inserts/Joints Ultimate 1.5 1.9

Glass/Ceramic 3.0(6) 3.8 Bond in Glass/Ceramic 1.5 1.9

(1) Factors of Safety for pressurized systems to be compliant with NASA-SDT-8719.24 (with Annex)

(NASA EXPENDABLE LAUNCH VEHICLE PAYLOAD SAFETY REQUIREMENTS).

(2) All composite structures must be tested to 1.25 x limit loads.

(3) For qualification by analysis only, positive margin must be shown with FS of 2.0 and 2.6 for metallic yield and ultimate, respectively.

(4) Factors shown are applied to statistically derived peak response based on acceptance RMS level. As a minimum, the peak response must be calculated as a 3-sigma value.

(5) Factors shown assume that qualification/prototype testing is performed at 3dB above limit/acceptance limits. If the difference between acceptance and qualification is less than 3dB then the above factors may be applied to qualification minus 3dB.

(6) For qualification by analysis only, a no-test safety factor of 5.0 may be used for glass (not ceramic), per NASA-STD-5001B Table 3.

8.2 Fracture Control Requirements

The vendor shall evaluate all hardware to determine fracture criticality. A four-tier system is used to evaluate fracture criticality of hardware; each tier has associated mitigation requirements that shall be followed, as specified below.

Tier 1 fracture classification: Light-weighted optics, ceramic structural elements used to transfer load from one component to another, and any glass/ceramic hardware designated high risk by RST shall mitigate risk of catastrophic failure due to cracks/flaws through the following requirements:

a. Proving design is fail-safe or,

a. Industry standard fracture analysis using NASGRO, Flaw Growth Analysis per NASA-STD- 5018, or other fracture analysis software demonstrating safe-life for 4 times number of mission cycles.

and,

b. Positive MS using B-basis Weibull accounting for Weibull scale effect per ASTM C1683 and,

c. Slow crack growth analysis of glass/ceramics made of materials with SiO bonds or,

a. Positive MS using B-basis Weibull accounting for Weibull scale effect per ASTM C1683

b. Proof testing with 2.0 proof test factor

For all flight hardware designated Tier 1 fracture critical, a Fracture Control Plan (FCP) shall be developed and approved by the RST FCB.

Tier 2 fracture classification. All glass/ceramics not classified as Tier 1 shall mitigate risk of catastrophic failure due to cracks/flaws through the following requirements:

a. Proving design is fail-safe or,

a. Positive MS using B-basis Weibull accounting for Weibull scale effect (ASTM 1683) or

a. Industry standard fracture analysis using NASGRO, Flaw Growth Analysis per NASA-STD- 5018, or other fracture analysis software demonstrating safe-life for 4 times number of mission cycles

Tier 3 fracture classification: Metals with fracture toughness to yield strength ratio less than

0.33 (KIc/Fty < 1.66 mm1/2 (0.33 in1/2)); thin walled metallic components serving as springs or flexures; and fastened joints identified as critical, that experience unconcentrated stress at greater than 30% of Ultimate strength shall mitigate risk of catastrophic failure due to cracks/flaws through the following requirements:

a. Proving design is fail-safe or,

a. Demonstrate sufficient fatigue strength for 4 times the number of mission cycles or,

a. Prove safe-life design through industry standard linear-elastic fracture analysis for 4 times the number of mission cycles.

Tier 4 fracture classification: Fracture control requirements (per NASA-STD-5019) shall apply to the following elements only: (1) Pressure vessels, dewars, lines, and fittings (per NASA-STD 8719.24);(2) Castings (unless hot isostatically pressed and the flight article is proof tested to 1.25 times limit load);(3) Weldments;(4) Parts made of materials in Tables II or III of MSFC-SPEC- 3029A if under sustained tensile stress;(5) Parts made of materials susceptible to cracking during quenching;(6) Nonredundant, mission-critical preloaded springs loaded to greater than 25 percent of ultimate strength.

Notes for fracture/fatigue:

1. Fracture analysis is performed assuming an initial crack size. NDE or proof testing is required to verify this initial crack size assumption. If proof testing, the proof tests must be at load sufficient enough to screen for any crack size above the assumed. This often results in over-test compared to DLL.

2. If NASGRO is used for brittle component fracture analysis, the NASGL module is used.

3. If the max principle stress in the glass/ceramic containing SiO bonds in the glass with a minimum detectable crack size (a) is calculated (from KIc =σ√πa) to be within 25% of fracture toughness, then slow crack growth is not a concern.

4. A fail-safe component is a structural component in a redundant load path which after the loss of the component, the remaining load path(s) has sufficient structural capability to withstand the redistributed loads, and the loss of the component will not cause a catastrophic hazard.

Examples of redundant load paths are multiple fasteners in bolted joints, stitched welded structures, and riveted and bolted structures. A factor of safety of 1.0 shall be applied to the remaining load path(s) when demonstrating fail-safe.

5. Critical fastened joints are joints whose failure results mission ending failure. Class-C fasteners that are contained within Class-C hardware are not critical fasteners.

6. Mission cycles include ground testing and handling, transportation, launch, and on orbit loading. The load spectrum will include the load level and the accompanying number of cycles and duration (for static loading) at each level during the hardware service life. Mission cycle load spectrum is included in the FCP.

7. In cases where fracture toughness of a metal is not known, contact RST project for help in fracture classification.

8.3 Metallic Pressurized Lines and Fittings Burst Factor

Metallic pressurized lines and fittings shall be designed to sustain the design Burst Pressure (BP) as defined in notes without leakage or rupture.

Rationale: To ensure the safety of pressurized lines and fittings.

Notes:

1. Burst Pressure is defined as follows:

𝐵𝐵𝐵𝐵 = 𝐵𝐵𝐵𝐵𝑜𝑜𝑆𝑆𝐿𝐿 𝐹𝐹𝐴𝐴𝐹𝐹𝐿𝐿𝐴𝐴𝑜𝑜(𝐵𝐵𝐹𝐹) 𝑥𝑥 𝑀𝑀𝐴𝐴𝑥𝑥𝐿𝐿𝐿𝐿𝐵𝐵𝐿𝐿 𝐸𝐸𝑥𝑥𝐴𝐴𝑒𝑒𝐹𝐹𝐿𝐿𝑒𝑒𝐿𝐿 𝑂𝑂𝐴𝐴𝑒𝑒𝑜𝑜𝐴𝐴𝐿𝐿𝐿𝐿𝑂𝑂𝑂𝑂 𝐵𝐵𝑜𝑜𝑒𝑒𝑆𝑆𝑆𝑆𝐵𝐵𝑜𝑜𝑒𝑒(𝑀𝑀𝐸𝐸𝑂𝑂𝐵𝐵)

Burst Factor (BF) for diameter ≥ 1.5 in = 2.5, BF for diameter < 1.5 in = 4.0

8.4 Joints

8.4.1 Joint Edge Distance

The minimum bolt edge distance shall be as specified in Table 8-2 unless a detailed analysis is performed to justify otherwise.

Notes

1. Use appropriate bearing value from MMPDS for metals

2. D = Bolt hole or insert sleeve diameter, e = edge distance from center of bolt to edge of panel

Table 8-2 Minimum Edge Distance

Minimum Edge Distance (e/D)

Composite/Ceramic 2.5 min Metallic 1.5 min

8.4.2 Bonded Joint Allowable

Bonded joint allowables shall be B-basis produced through coupons representative of flight hardware, tested under their appropriate environments, and taking into account strength degradation due to thermal cycles.

Rationale: Bond strength is dependent on materials and workmanship, e.g. surface prep. Due to scatter from the workmanship process and material variability, B basis allowables are developed for structural analysis.

Note: Bonded joints on flight hardware are proof tested to 1.25 times limit load.

8.4.3 Reserved

8.4.4 Fastened Joints

8.4.4.1 Fastener Locking

Each threaded fastening system in spaceflight hardware shall incorporate a minimum of one locking feature that does not depend upon preload to function.

Notes:

1. Locking features and their installation processes, including verification methods, should be specified in the engineering documentation.

2. Locking features should be verifiable per Section 4.6 of NASA-STD-5020A.

8.4.4.2 Fastened Joint Margin of Safety

Positive MS for all threaded fastening systems shall be demonstrated for:

a) Ultimate design loads under all structural failure modes in conjunction with the applicable maximum expected range of environmental conditions without failure per the following:

1) The ultimate design load, Pu, is calculated using the following formula.

2) 𝐵𝐵𝑈𝑈 = 𝐹𝐹𝐹𝐹 ∗ 𝐹𝐹𝑀𝑀𝑈𝑈 ∗ 𝐵𝐵𝐿𝐿

i. where FF is the fitting factor, FSu is the ultimate factor of safety, and PL is the limit load.

3) The fitting factor (FF) defined in NASA-STD-5020A Section 4.2.2 is used.

4) Analysis for ultimate design loads will address potential rupture in all elements of the threaded fastening system, including the fastener, the internally threaded part, such as a nut or an insert, and the clamped parts.

5) Insert pull-out strengths given in vendor documentation typically assume full bolt to insert thread engagement. An appropriate knock down in pull-out strength is included if the bolt does not fully engage the insert.

6) Ultimate strength analysis of a fastening system under applied tensile loading is performed per NASA-STD-5020A, Section 4.4.1.

7) Ultimate strength analysis of a fastening system under applied shear loading is performed per NASA-STD-5020A, Section 4.4.1.

8) Ultimate strength analysis of bolts under shear loading is based on the assumption that no shear load is carried by friction between the faying surfaces.

9) For fasteners under simultaneous applied tensile and shear loads, along with any applicable bending analysis will account for interaction of the combined loading per NASA-STD-5020A, Section 4.4.4.

10) When possible, joints should be designed for bearing failure rather than bolt tension or shear failure.

b) Yield design loads in conjunction with the applicable maximum expected range of environmental conditions without detrimental yielding per the following:

1. If one or more of the following applies:

i. Fastener yielding causes the joint to separate under an applied tensile load that is less than the design separation load;

ii. Fastener yielding causes the joint to suffer detrimental slip under an applied shear load that is less than the applicable design shear load;

iii. Some other design-specific reason exists for why fastener yielding is detrimental (e.g., any fastener yielding that adversely affects the form, fit, or function of the design);

then fastener yielding is detrimental and analysis is performed per NASA-STD- 5020A, Section 4.4.2 to show the fastener’s total tensile load, when accounting for maximum preload and the yield design tensile load, does not exceed the allowable yield tensile load.

2. The yield design load, PY, is calculated using the following formula.

𝐵𝐵𝑌𝑌 = 𝐹𝐹𝐹𝐹 ∗ 𝐹𝐹𝑀𝑀𝑌𝑌 ∗ 𝐵𝐵𝐿𝐿

where FF is the fitting factor, FSY is the yield factor of safety, and PL is the limit load.

3. The fitting factor (FF) defined in NASA-STD-5020A section 4.2.2 will be used.

4. If MS for joint slip is shown to be positive, friction may be included as a shear load path for yield strength analysis of joint members (e.g. bearing). NASA-STD-5020A Appendix A.10 is used for joint slip analysis.

c) Gapping load in conjunction with applicable maximum or minimum temperatures without gapping per the following:

1) Margin of safety for gapping is defined as follows:

i) 𝑀𝑀𝑀𝑀𝑠𝑠𝑠𝑠𝑠𝑠 = 𝑃𝑃𝑝𝑝−𝑚𝑚𝑚𝑚𝑚𝑚

𝐹𝐹𝐹𝐹∗𝐹𝐹𝐹𝐹𝑠𝑠𝑠𝑠𝑝𝑝∗𝑃𝑃𝐿𝐿

- 1 where FF is the fitting factor, and Pp-min is the minimum preload.

2) Gapping factor of safety (FSsep) of 1.0 for non-gapping critical joints, FSsep=1.25 for non-catastrophic failure gapping critical joints, and FSsep=1.4 for catastrophic failure gapping critical joints is used for margin calc.

3) The fitting factor (FF) defined in NASA-STD-5020A Section 4.2.2 will be used.

4) NASA-STD-5020A, Appendix A.12 is used to guide gapping analysis to show no separation for each threaded fastening system that is subject to applied tensile loading, with the assumption of minimum preload.

5) Analysis of threaded fastening systems will address nominal, maximum, and minimum preloads per NASA-STD-5020A Section 4.3.

See NASA-STD-5020A, section 4. and Appendices A.10 – A.12 as reference.

8.4.4.3 Installation Torque Documentation

Engineering documentation shall specify the installation torque range or specify an applicable standard that defines the installation torque range.

Note:

1. The engineering documentation should clearly identify when the installation torque is the torque above running torque.

APPENDIX A ABBREVIATIONS AND ACRONYMS

Abbreviation/ Acronym Definition

AC Alternating Current AFT Abbreviated Functional Test Al Aluminum CCB Configuration Control Board CCR Configuration Change Request CE Conducted Emissions CG Center of Gravity CIR Configuration Initial Release CM Configuration Management CMO Configuration Management Office CMOS Complementary Metal Oxide Semiconductor COR Contracting Officer's Representative CPT Comprehensive Performance Test CS Conducted Susceptibility CVCM Collected Volatile Condensable Mass DA Double Amplitude DC Direct Current DDD Displacement Damage Dose DILS Deliverable Items List and Schedule EED ElectroExplosive Actuators EEE Electrical, Electronic, and Electromechanical ELDR Enhanced Low Dose Rate EMC Electromagnetic Compatibility EMI Electromagnetic Interference EOL End of Life ESD Electrostatic Discharge FS Factor of Safety GeBK Germanium Black Kapton GSE Ground Support Equipment GSFC Goddard Space Flight Center ICD Interface Control Document I&T Integration and Test ITO Indium Tin Oxide LET Linear Energy Transfer LPT Limited Performance Test LVDS Low Voltage Differential Signal MBU Multi Bit Upset MICD Mechanical Interface Control Document

Abbreviation/ Acronym Definition

MLI Multi-Layer Insulation Mohms Megaohms MOP Maximum Operating Pressure MOSFET Metal Oxide Semiconductor Field-Effect Transistor MS Margin of Safety NASA National Aeronautics and Space Administration NEA Non-Explosive Actuators NIEL Non-Ionizing Energy Loss OSR Optical Solar Reflector PDL Product Design Lead QCM Quartz Crystal Monitor RE Radiated Emissions RF Radio Frequency RS Radiated Susceptibility RST Roman Space Telescope S/C Spacecraft SEB Single Event Burnout SEGR Single Event Gate Rupture SEE Single Event Effects SEFI Single Event Functional Interrupt SEL Single Event Latchup SEM Scanning Electron Microscope SEU Single Event Upset SHE Single Hard Error SOW Statement of Work SPL Sound Pressure Level TBD To Be Defined TBR To Be Reviewed TID Total Ionizing Dose TML Total Mass Loss UUT Unit Under Test VDA Vapor Deposited Aluminum VDC Voltage, Direct Current

6.1.1.2 Internal Cleanliness