Attachment B specification.pdf

Effective Date: September 9, 2020

National Aeronautics and Space Administration

Goddard Space Flight Center Greenbelt, Maryland

RST-PROP-SPEC-0053, Revision -

Roman Space Telescope (RST), Code 448

Pressure Transducer Specification

GSFC RST CMO

September 9, 2020

Released

RST Pressure Transducer Spec RST-PROP-SPEC-0053, Revision -ii

Pressure Transducer Specification

Review/Signature/Approval Page

Prepared by:

Wanyi Ng

Approved by:

Betsy Forsbacka

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 - September 4, 2020 Initial Release per CCR-0007 v

Table of Contents

1 INTRODUCTION

1.1 Purpose

1.2 Scope

1.3 Related Documentation

1.3.1 Applicable Documents [and Forms]

1.3.2 Reference Documents

2 CONTRACT DESCRIPTION

2.1 Pressure Transducer Description

3 FUNCTIONAL/PERFORMANCE REQUIREMENTS

3.1 Pressure Transducer Flight Unit Functional/Performance Requirements

3.1.1 Fluid Compatibility

3.1.2 Pressures

3.1.2.1 Operating Pressure Range

3.1.2.2 Proof Pressure

3.1.2.3 Burst Pressure

3.1.2.4 Pressure Cycle Life

3.1.2.5 Interface Tube Design Pressure

Proof Pressure Burst Pressure

3.1.3 Leakage Requirements

3.1.3.1 External Leakage

3.1.3.2 Leak Path

3.1.4 Accuracy

3.2 Resource Allocations

3.2.1 Mass Allocation

3.2.2 Nominal Power Allocation

3.2.3 Peak Power

3.3 Power

3.3.1 Primary (Unregulated) Power Input Requirements

3.3.1.1 Operating Voltage Range

3.3.1.2 Abnormal Voltages

3.3.1.3 Voltage Ripple

3.3.1.4 Sudden Removal of Power

3.3.1.5 Polarity Reversal Protection

3.3.1.6 Over-Current Protection

3.3.1.7 Primary Power Return Ground

3.3.1.8 Power Wiring Redundancy

3.3.2 Load Induced Noise Requirements

3.3.2.1 Turn-on Current Transients

3.3.2.2 Operational Current Transients (one time)

3.3.2.3 Low Frequency Repetitive Transients (during nominal operation)

3.4 Electrical Grounding

3.4.1 Primary Power DC Isolation

vi

3.4.2 Internally Generated Secondary Returns

3.4.3 Internally Generated Secondary to Primary DC Isolation

3.4.4 Mechanical Contact Resistance

3.4.5 Mating Method

3.4.6 Connector and Back-Shell DC Resistance

3.5 Signal and Data Interfaces

3.5.1 Active Analog Telemetry

3.5.1.1 Active Analog Telemetry Interface

3.5.1.2 Active Analog Telemetry Short Circuit Protected

3.5.1.3 Active Analog Telemetry Source and Load Resistance

3.5.1.4 Active Analog Telemetry Voltage Range

4 PHYSICAL REQUIREMENTS

4.1 Interface Documentation

4.1.1 Mechanical Interface

4.1.2 Electrical 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 Mounting

4.5 Interface Tube

4.5.1 Dimension

4.5.2 Material

5 ENVIRONMENTAL REQUIREMENTS

5.1 Quasi-Static Acceleration

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 Pressure

5.5.1 Operating Pressure Range

5.5.2 Maximum Depressurization Rate

5.6 Ground Environments

5.7 Thermal Requirements

5.7.1 Flight Interface Design Temperature Limits

5.8 Charged Particle Radiation Requirements

5.8.1 Definitions

5.8.2 Total Ionizing Dose

5.8.2.1 Minimum TID Tolerance for EEE Parts and Materials

5.8.3 Displacement Damage Dose

5.8.4 Single Event Effects

5.8.5 Charging Environment

vii

6 CLEANLINESS

6.1 Surface Contamination

6.1.1 Surface Contamination Levels at Delivery

6.1.1.1 External Cleanliness

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 EEE Parts

7.1.2 Materials

7.1.2.1 Material Conductivity

7.2 Electrical

7.2.1 Test Sensors

7.2.2 Interface Requirements

7.2.2.1 Connector Selection

Connector Specifications External Box Power Connectors Contact Derating Redundant Contact Derating

7.2.2.2 Signal Segregation

7.2.2.3 Test and Flight Signal Isolation

7.2.2.4 Test Interfaces

Facility-Induced Noise Facility-Induced ESD GSE Malfunction Facility-Induced GSE Malfunction

7.2.3 Dielectric Structures

7.3 Electromagnetic Compatibility

7.3.1 Conducted Emissions

7.3.1.1 Applicability of Conducted Emissions

7.3.1.2 CE101 – Differential Mode Current Emission Limits (Low Frequency)

7.3.1.3 CE102/03 – Differential Mode Current Emission Limits (Low Frequency)

7.3.1.4 Common Mode Current Emissions Limits

7.3.1.5 Conductive Emission Measurement Bandwidths

7.3.1.6 Common Mode Emission Time Domain Measurement

7.3.2 Conducted Susceptibility

7.3.2.1 Applicability of Conducted Susceptibility

7.3.2.2 CS101 – Power Leads, 30 Hz to 150 kHz Limit

7.3.2.3 CS02/CS114 – Power Leads, 150 kHz to 50 MHz Limit

viii

7.3.3 Radiated Emissions

7.3.3.1 RE102 - Electric Field Emissions

7.3.4 Radiated Susceptibility

7.3.4.1 RS103 - Electric Field

7.4 Identification and Marking

7.5 Workmanship

7.5.1 Workmanship Standards

7.5.2 Connector

7.5.2.1 GSE Cable Connectors

7.5.2.2 Prevention of Connector Mismating

Connector Uniqueness Connector Keying Accessibility Connector Gender “Scoop-Proof”

7.5.2.3 Connector Identification

7.6 Reliability and Mission Lifetime

7.6.1 Mission Life

7.6.2 Operating Time

7.6.2.1 Total Time

7.6.2.2 Failure Free Time

8 MECHANICAL DESIGN REQUIREMENTS

8.1 Mechanical Factors of Safety

8.2 Fracture and Fatigue

8.3 Metallic Pressurized Lines and Fittings Burst Factor

8.4 Materials

8.4.1 Dissimilar Metals

8.4.2 Material Allowables and Stiffness

8.4.3 CTE Data

8.4.4 Stress Corrosion Cracking

8.5 Joints

8.5.1 Joint Edge Distance

8.5.2 Bonded Joint Allowable

8.5.3 Fastened Joints

8.5.3.1 Fastener Locking

8.5.3.2 Fastened Joint Margin of Safety

8.5.3.3 Installation Torque Documentation

8.5.3.4 Seal Analysis

8.5.3.5 Critical Fasteners

9 LOGISTICS

9.1 Ground Support Equipment

9.2 Transportation Equipment

APPENDIX A ABBREVIATIONS AND ACRONYMS

List of Figures ix

Figure 7-1 CE101/CE03 Differential Mode Current Emission Limits Figure 7-2 Common Mode Current Emission Limits Figure 7-3 Component Level RE102 Radiated Electric Field Emission Limits

List of Tables Table 5-1 Item Name Design Limit Loads Table 5-2 Item Name Generic Sine Vibration Environment Table 5-3 Random Vibration Limit Levels for Components Weighing Less Than 22.7 kg Table 5-4 Minimum Workmanship ASD Level for Components Weighing Less than 45.4 kg. . 12 Table 5-5 Shock Limit Levels Table 5-6:Depressurization Profile Table 5-7 Temperature Limits at Box Mounting Interface, Unwetted Table 5-8 Propellant-Wetted In-Flight Software Temperature Limits Table 7-1 Emission and Susceptibility Requirements Table 7-2: Receiver Notch RE102 Test Limit Levels Table 7-3: RS103 Radiated Susceptibility Levels Table 8-1 Design Factors of Safety Table 8-2 Minimum Edge Distance

1 INTRODUCTION

1.1 Purpose

The 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 propulsion subsystem pressure transducers.

1.2 Scope

This specification describes the electrical, mechanical, environmental, and verification testing requirements for a space-qualified propulsion subsystem pressure transducers for the NASA Goddard Space Flight Center (GSFC) RST Mission.

1.3 Related Documentation

RST documents can be obtained in the RST CM Tool.

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 electrical, mechanical, and environmental requirements of the pressure transducers 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 Pressure Transducer Statement of Work (RST-PROP-SOW-0017) in which case the Statement of Work takes precedence.

Document Number Title ANSI/TIA/EIA-422-B Electrical Characteristics of Balanced Voltage Digital Interface

Circuits

ANSI/TIA/EIA-644-A-

Electrical Characteristics of Low Voltage Differential Signaling (LVDS) Interface Circuits

ASTM E-595-07 Standard Test Method for Total Mass Loss and Collected Volatile Condensable Materials from Outgassing in a Vacuum Environment

ECSS-E-ST-50-12A SpaceWire - Links, Nodes, Routers and Networks EEE-INST-002 Instructions for EEE Parts Selection, Screening, Qualification, and Derating IEST-STD-CC1246E Product Cleanliness Levels And Contamination Control Program MIL-STD-1553B Department of Defense Interface Standard for Digital Time

Division Command/Response Multiplex Data Bus

MIL-STD-461F Military Standard, Electromagnetic Emission And Susceptibility Requirements For The Control Of Electromagnetic Interference

(EMI)

MIL-STD-462, Notice 2 Electromagnetic Interference Characteristics, Measurement of, 1 May 1970

MSFC-STD-3029A Guidelines for the Selection of Metallic Materials for Stress Corrosion Cracking Resistance in Sodium Chloride Environments

NFPA 70 National Fire Protection Association National Electric Code NASA-STD-5017A Design and Development Requirements for Mechanisms 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

GSFC-STD-7000A General Environmental Verification Standard (GEVS) 541-WI-5330.1.41 Fastener Locking Using Arathane 5753 RST-SYS-PLAN-0051 RST Contamination Control Plan RST-PROP-SOW-0017 RST Pressure Transducer Statement of Work

1.3.2 Reference Documents

The following documents are referenced herein and amplify or clarify the information presented in this document. These documents are not binding on the content of this document.

Document Number Title AMS5569 REV. B Steel, Corrosion and Heat Resistant, Seamless and Welded Hydraulic

Tubing 19Cr - 9.5Ni - 0.03C max Cold Drawn, 1/8 Hard Temper AMS5647 REV. K Steel, Corrosion-Resistant, Bars, Wire, Forgings, Mechanical Tubing, and

Rings 19Cr - 9.5Ni Solution Heat Treated

RST Pressure Transducer Spec RST-PROP-SPEC-0053, Revision-

2 CONTRACT DESCRIPTION

2.1 Pressure Transducer Description

The pressure transducers are measurement devices that are capable of monitoring the pressure in various segments of the propulsion subsystem, beginning during ground operations and continuing throughout the mission life. The transducers convert sensed gas and liquid pressure into electrical signals for telemetry purposes. The transducers must cover a full pressure range of 0-400 psia full scale output (FSO).

The pressure transducer design must be qualified for space flight or the vendor will include one

(1) unit designated as qualification to test to the qualification levels listed in this document.

Designs that have been qualified previously must include qualification documentation as described in the Pressure Transducer Statement of Work (RST-SOW-0017).

3 FUNCTIONAL/PERFORMANCE REQUIREMENTS

This section defines the functional and performance requirements for the pressure transducers as defined in Section 3.1.

3.1 Pressure Transducer Flight Unit Functional/Performance Requirements

3.1.1 Fluid Compatibility

Each pressure transducer, including all welds, shall show no degradation or loss of accuracy due to exposure for a 99-month period (equivalent to mission life of 3 years of ground tests, 5 years of operation in space, 3 months commissioning) to any of the following fluids:

• Nitrogen per MIL-PRF-27401G

• Helium per MIL-PRF-27407D

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

• Deionized and Distilled Water per JSC-PSEC-C-20

• Isopropyl Alcohol per FED-STD-TT-I-735A

• Argon per MIL-PRF-27415C

3.1.2 Pressures

3.1.2.1 Operating Pressure Range

All four (4) pressure transducers shall have a pressure range from 0 psia to the Maximum Expected Operating Pressure (MEOP) for RST, 400 psia.

3.1.2.2 Proof Pressure

The pressure transducers shall be capable of withstanding proof pressure no less than 1.5 x MEOP, 600 psia, without any permanent physical deformation, yielding, or cracking, and without any loss of performance or accuracy as specified herein.

3.1.2.3 Burst Pressure

The pressure transducers shall demonstrate, through a qualification program, burst pressures no less than 2.5 x MEOP, 1000 psia.

3.1.2.4 Pressure Cycle Life

The pressure transducers shall maintain the operational requirements defined herein after no less than 2,000 pressure cycles, where one cycle consists of the sensed pressure changing from atmospheric pressure to MEOP and then returning to atmospheric pressure. The Contractor shall provide data to show that the transducer design has successfully operated in a manner that verifies the capability to achieve this cycle life.

3.1.2.5 Interface Tube Design Pressure

Proof Pressure

Each pressure transducer’s interface tube shall be capable of withstanding the proof pressure of 1.5xMEOP, 600 psia.

Burst Pressure The Contractor shall demonstrate through a qualification program that each pressure transducer’s interface tube has a burst pressure of no less than 4.0 x MEOP, 1600 psia.

3.1.3 Leakage Requirements

3.1.3.1 External Leakage

Each pressure transducer shall demonstrate a total leakage rate of no greater than 1 x 10-6 standard cubic centimeters per second (sccs) GHe when pressurized to MEOP.

3.1.3.2 Leak Path

All electrical circuits shall be hermetically sealed from the propellant or pressurant flow. This seal shall be tested to verify a total leak rate of no greater than 1 x 10-5 standard cubic centimeters per second (sccs) GHe when pressurized to MEOP. All internal propellant leaks shall be contained within the transducer housing.

3.1.4 Accuracy

Static Error Band accuracy shall be 0.25% at a minimum using Best Fit Straight Line – Root Sum of Squares Method (BFSL-RSS). The Static Error Band equation is also defined below.

Static Error Band = SQRT (Linearity2 + Hysteresis2 + Repeatability2)

The error value is given as percentages of full scale, with each pressure transducer’s full scale defined as the maximum pressure in the transducer’s range (FSO is 400 psia). The transducers shall maintain these accuracy requirements throughout the temperature range and input voltage range specified herein.

3.2 Resource Allocations

3.2.1 Mass Allocation

Each pressure transducer shall have a mass of less than or equal to 0.4 kg (0.88 lb).

3.2.2 Nominal Power Allocation

Each pressure transducer shall have a nominal (continuous) power consumption of less than or equal to 0.875 W for a duration of 100 minutes at a bus voltage of 32 VDC.

3.2.3 Peak Power

The peak power shall not exceed its daily average by more than 10%.

3.3 Power

3.3.1 Primary (Unregulated) Power Input Requirements

3.3.1.1 Operating Voltage Range

The pressure transducers shall be designed to operate over the bus voltage range of +24 to +35 VDC at its primary power inputs during all normal mission phases and for all expected load conditions (except when turned off).

3.3.1.2 Abnormal Voltages

The pressure transducers shall survive without electrical overstress after exposure to an anomalous voltage range of 0 to +40 VDC and input power short or open. This requirement is to be verified by analysis or test of a non-flight unit.

3.3.1.3 Voltage Ripple

The pressure transducers shall be designed to meet their performance requirements in the presence of low and RF frequency ripple voltages specified in paragraphs 7.3.2.2 and 7.3.2.3, respectively.

3.3.1.4 Sudden Removal of Power

The pressure transducers shall meet their performance requirements without degradation after exposure to an abrupt, unannounced removal of power.

3.3.1.5 Polarity Reversal Protection

The pressure transducers should have built-in protection to prevent damage due to polarity reversal at the power inputs (up to -35V), where feasible.

3.3.1.6 Over-Current Protection

The pressure transducers should not use non-resetting over-current protection (i.e., fuses) internal to the unit, but if they are required by the contractor to isolate an internal fault, prior approval shall be obtained from the NASA/GSFC Contracting Officer’s Representative (COR).

3.3.1.7 Primary Power Return Ground

The pressure transducers shall provide a dedicated Primary Power return in the same connector as the primary power.

3.3.1.8 Power Wiring Redundancy

The pressure transducers shall provide a redundant power interface. The input power should be “diode-OR” with primary power.

3.3.2 Load Induced Noise Requirements

3.3.2.1 Turn-on Current Transients

The amplitude and duration of in-rush currents, at both initial application of prime power and subsequent component turn-on, shall be per the following:

• <10 microseconds: 10 times of max steady state or 10A whichever is larger

• Between 10 microseconds and 10 milliseconds: 3 times of max steady state or 3A whichever is larger

• >10 milliseconds: not to exceed max steady state

The compliance of this requirement can be demonstrated using a solid-state power switch with the similar turn on characteristics as the flight power control switch.

3.3.2.2 Operational Current Transients (one time)

a. The pressure transducer operational current transients shall be less than or equal to 5A for a period of less than 1 millisecond.

b. The rate of change of the current shall be less than or equal to 20 milliamperes/microsecond.

3.3.2.3 Low Frequency Repetitive Transients (during nominal operation) Any repetitive current transients, in the frequency range between 1 Hz to 200 Hz, shall not exceed 0.8 A peak-to-peak.

This measurement will be performed with an oscilloscope current probe in time domain.

3.4 Electrical Grounding

3.4.1 Primary Power DC Isolation

At the pressure transducer primary power interfaces, primary power and primary power returns shall be isolated from signal grounds and from the component chassis by a DC resistance of greater than or equal to 1 MΩ.

3.4.2 Internally Generated Secondary Returns

Secondary returns (power and signal grounds) shall be referenced to the component chassis ground by connecting them at one or more places.

3.4.3 Internally Generated Secondary to Primary DC Isolation

Secondary power (or signal) inputs shall be isolated from primary power by a DC resistance of greater than 1 MΩ.

3.4.4 Mechanical Contact Resistance

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

3.4.5 Mating Method

The primary mating method for a RST component is metal-to-metal contact between component mounting feet (or base plate) and the spacecraft structure. When the use of this method is not possible as determined by the Mechanical Interface Control Document (MICD), the use of a ground strap is necessary.

3.4.6 Connector and Back-Shell DC Resistance

Component connectors including the EMI back-shells shall be electrically connected to chassis ground of the electronics box with a DC resistance less than or equal to 100 mΩ.

3.5 Signal and Data Interfaces

3.5.1 Active Analog Telemetry

3.5.1.1 Active Analog Telemetry Interface

The pressure transducers shall have a mounted circular connector/receptacle. The pressure transducers shall provide the following active analog signals to the Spacecraft via connector pins:

• Power (+)

• Power (-)

• Signal (+)

• Signal (-)

3.5.1.2 Active Analog Telemetry Short Circuit Protected

The sources of active analog signals shall be short circuit protected.

3.5.1.3 Active Analog Telemetry Source and Load Resistance

The source and load resistances (at the source and destination, respectively) shall be documented in the pressure transducer Electrical Interface Control Document (EICD) such that the overall system compatibility of the telemetry signal is reviewed and controlled.

3.5.1.4 Active Analog Telemetry Voltage Range

The active analog telemetry voltage range shall be within 0 to +5 VDC.

4 PHYSICAL REQUIREMENTS

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:

• SI units with parenthetical English units are permitted for engineering and manufacturing drawings.

• Angular measurement may be expressed in degree of arc or in an appropriate subdivision of degree of arc such as minute or second of arc (arc-min or arc-sec).

• Hardware that has been previously built, or new hardware of similar design heritage, may be specified in English units where use of metric equivalents would lead to additional risk and cost to the program, but interfaces with that hardware must use metric with parenthetical English units.

• Hybrid SI/standard astronomical units are permitted for Astronomer User’s Manuals, user interfaces (Graphical User Interfaces and files) consistent with the Astronomer's User Manuals, and requirements documents.

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.

4.1.2 Electrical Interface

The electrical interface shall be defined in the Electrical Interface Control Document (EICD), which will be developed between the contractor and NASA GSFC.

4.2 Mass Properties

4.2.1 Component Masses

Refer to Section 3.2.1, Mass Allocation.

4.2.2 Center of Mass Location

The contractor shall define the center of mass in the Mechanical ICD.

4.2.3 Center of Mass Accuracy

The center of mass of the pressure transducers shall be determined to within ±2.5 mm relative to an external reference.

4.2.4 Determination of Moments and Products of Inertia

The final moments of inertia of the pressure transducers shall be calculated to within 1%. The final products of inertia of the pressure transducers shall be calculated to within 5%.

4.3 Physical Envelope

The pressure transducers’ outside dimensions, including mounting flanges and connectors, shall not exceed the volume envelope of 18.5 cm length x 6 cm width x 4 cm height as defined in the Mechanical ICD.

4.4 Mounting

Each pressure transducer shall be capable of being bolted to the flat spacecraft mounting surface.

The mounting interface shall be defined in the MICD.

4.5 Interface Tube

Each pressure transducer shall contain a straight interface tube that can later be welded directly to the RST propulsion subsystem.

4.5.1 Dimension

The interface tube shall be no less than 1.5 +/-0.005 inches long, with a consistent outer diameter (OD) of 0.25 +/-0.003 inches and a wall thickness of 0.028 +/- 0.003 inches.

4.5.2 Material

The interface tube material shall be 304L stainless steel per AMS5647 and AMS5569.

5 ENVIRONMENTAL REQUIREMENTS

Environmental design requirements are specified in this section.

The pressure transducers shall meet their performance requirements in section 5.0 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.

The pressure transducers shall be designed to withstand the quasi-static Design Limit Loads (DLL) defined in the Mass-Acceleration Curve (MAC) shown in Table 5-1 without damage or degradation of performance.

Note:

1. For masses within the breakpoints, linear interpolation is used to determine load.

2. Loads are considered to act in any direction, one axis at a time.

Table 5-1 Item Name Design Limit Loads

Component Mass (Kg) Limit Load (g)

0.5 35.9 1 35

5.2 Frequency Requirement

5.2.1 Stowed Fundamental Launch Frequencies

The pressure transducers shall have a fundamental frequency greater than 100 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.

5.3 Vibration

5.3.1 Sinusoidal Vibration

The pressure transducers shall withstand the sinusoidal vibration levels defined in Table 5-2.

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 the SOW, Table 5-1, Test Factors and Durations.

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 5-2 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 Table 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 until higher levels of assembly upon approval by the

NASA/GSFC COR.

Table 5-2 Item Name Generic Sine Vibration Environment

Frequency (Hz) Limit Level

5 to 100 10.0 g

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

a. The pressure transducers shall withstand the random vibration environment in Table 5-3 applied at the interface to the pressure transducer.

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 5-3 shows the limit level random vibration environment for components with mass less than

22.7 kg. 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 5-1, Test Factors and Durations. Table 5-4 defines minimum workmanship for components weighing less than 45.4

kg. 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 5-3 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 5-4 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

5.4 Shock

The pressure transducers shall be designed to meet their performance requirements after being subjected to the shock environment in Table 5-5, applied at the pressure transducer interface to the RST spacecraft structure. Levels defined are acceptance levels. Testing factors are defined in the SOW, Table 5-1, Test Factors and Durations.

Table 5-5 Shock Limit Levels

Frequency (Hz) Limit Load (g)

100 75 100 – 738 6.5 dB/oct

738 858 10000 858

Note: A shock susceptibility and attenuation assessment is to be performed on the pressure transducers. This assessment can be based on past shock tests or other relevant information.

5.5 Pressure

5.5.1 Operating Pressure Range

The pressure transducers 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).

5.5.2 Maximum Depressurization Rate

The pressure transducers shall be designed to meet all performance requirements after exposure to a depressurization profile shown in Table 5-6 during launch and ascent.

Table 5-6:Depressurization Profile

Transonic Late Early sec Pa sec Pa sec Pa 0 101353 0 101353 0 101353 10 97906 10 99285 10 96527 15 94459 15 96527 15 92390 20 89288 20 92046 20 86530 25 82393 25 85840 25 78946

41.2 54469 44.8 54469 28 73361

43.2 53090 46.8 53780 30.2 67224

45.2 43024 48.8 44816 32.2 62053

46.6 37232 50.4 41369 33.5 58606

48.7 29648 52.1 37922 35 55159

50.9 22753 53.8 34474 36.4 51711 54 16548 55.3 31716 37.9 48264 58 11032 56.8 28958 39.4 44816 65 6895 58.4 26201 41 41369

70 5516 60 23443 42.7 37922 80 3448 61.4 21374 44.4 34474 110 2069 62.8 19306 46.3 31027 150 345 64.2 17237 48 27924

65.5 15514 50 25856

66.9 13790 52 19857 68 12411 57 13790

68.8 11584 65 8964 75 7585 80 4137 85 4137 110 2069 110 2069 150 345 150 345

5.6 Ground Environments

a. The pressure transducers shall meet all of their performance requirements during exposure to air temperature between +5 and +30 degrees C and relative humidity between 30% and 70%.

b. After being powered “OFF” and exposed to air temperatures of +5 to +30 degrees C and relative humidity of 0 to 70%, the pressure transducers shall meet all of their performance requirements.

5.7 Thermal Requirements

5.7.1 Flight Interface Design Temperature Limits

a. When powered “OFF”, the pressure transducers shall be capable of surviving indefinitely when their temperatures are within the survival limits shown in Table 5-7 without damage or permanent performance degradation.

b. The pressure transducers shall meet all performance requirements when powered “ON” anywhere within the Operational and Protoflight/Qualification limits shown in Table 5-7.

c. The pressure transducers shall demonstrate turn on at the Minimum and Maximum Protoflight/Qualification (in spec) limits shown in Table 5-7.

Table 5-7 Temperature Limits at Box Mounting Interface, Unwetted

Condition Minimum Temperature (ºC) Maximum Temperature (ºC) Operational +10 +50 Acceptance Test +5 +55 Protoflight/Qualification Test 0 +60 Survival (Unpowered) -5 +65

Note: the hardware must meet all performance specifications within the qualification temperatures. Operation is not required at survival temperature.

Table 5-8 Propellant-Wetted In-Flight Software Temperature Limits

Software Temperature Limit Minimum Temperature (ºC) Maximum Temperature (ºC) Green +10 +50 Yellow +6 +60 Red +5 +65

Note: Table 5-8 used for RST Software purposes only. Verification at these temperatures are not required, unless the temperature coincides with acceptance/qual temperatures.

5.8 Charged Particle Radiation Requirements

Components containing electronic parts will be exposed to a natural space radiation environment that consists of: (1) trapped particles which include electrons, protons, and heavier ions; (2) particles from solar events (coronal mass ejections and flares); and (3) galactic cosmic ray particles.

5.8.1 Definitions

Total Ionizing Dose (TID) - the mean energy deposited by ionizing radiation in a device region divided by the mass of the region. This is often given in units of rad(Si), where 1 rad(Si) = 100 erg deposited per gram of silicon.

Radiation Design Margin (RDM) - the ratio of the derated hardness capability of the part to the estimated dose at the part location.

Enhanced Low Dose Rate Sensitivity (ELDRS) - used to refer to a part that shows enhanced radiation-induced damage at dose rates below 50 rad(Si)/s. The enhancement is the result of true dose rate effects.

Displacement Damage Dose (DDD) - the mean energy deposited by ionizing radiation in a device region that goes into atomic displacements divided by the mass of the region. There is no official unit for DDD. One such unit is MeV/g

Non-Ionizing Energy Loss (NIEL) - a measure of the rate of energy loss due to atomic displacements as a particle traverses a material.

Single Event Effect (SEE) - any measurable effect to a circuit due to an ion strike. This includes, but is not limited to, single event upsets (SEUs), single event transients (SETs), single hard errors (SHEs), single event latchups (SELs), single event functional interrupts (SEFIs), single event burnouts (SEBs), single event gate ruptures (SEGRs), and single event dielectric ruptures (SEDRs).

Single Event Upset (SEU) - a change of state or transient induced by an energetic particle such as a cosmic ray or proton in a device. This may occur in digital or analog, circuits and may have effects in surrounding interface circuitry (a subset known as SETs). These are “soft” errors in that a reset or rewriting of the device will usually return the device to normal behavior thereafter.

The general goal for non-destructive events such as SEUs or SETs is not to avoid them completely, but to manage their impact through robust circuit design, automatic correction, and/or operational activities based on knowledge from ground radiation tests and circuit/system analysis.

Single Hard Error (SHE) - a SEU that causes a permanent change to the operation of a device.

An example is a stuck bit in a memory device.

Multiple Bit Upset (MBU) - an event induced by a single energetic particle such as a cosmic ray or proton that causes multiple upsets or transients during its path through a device or system in a single logical structure (ex., 2 bits affected in a single 16-bit word).

Single Event Functional Interrupt (SEFI) - a condition that causes loss of device functionality due to a single event in a control portion of a device. It generally requires a device reset or a re-initialization to resume normal device operations, but, for some devices, a power cycle is necessary to resume normal device operations. The general goal for non-destructive events such as SEFI is to avoid them, however, managing their impact through robust circuit design, automatic correction, and/or operational activities may be considered.

Single Event Latchup (SEL) - a condition that may cause device failure due to a single event induced high current state. A SEL may or may not cause permanent device damage, but requires power cycling of the device to resume normal device operations. In addition, susceptible devices have the concern for latent damage (device does not fail from the immediate single particle event, but reliability is degraded and premature failure may occur).

Single Event Burnout (SEB) - a condition that can cause device destruction due to a high current state in a power transistor.

Single Event Gate Rupture (SEGR) - a single ion induced condition in power MOSFETs that may result in the formation of a conducting path in the gate oxide.

Linear Energy Transfer (LET) - a measure of the energy deposited per unit length as an energetic particle travels through a material. The common LET unit is MeV*cm2/milligram (mg) of material.

Threshold LET (LETth) - the maximum LET at which no SEE is observed at a particle fluence of 107 ions/cm2.

5.8.2 Total Ionizing Dose

5.8.2.1 Minimum TID Tolerance for EEE Parts and Materials

All EEE parts shall tolerate a minimum Total Ionizing Dose (TID) of 29.0 krad (Si).

This value is based on radiation analysis data in the RST radiation analysis using the Total Ionizing Dose 95% confidence level curve with an equivalent of 2.54 mm (100 mils) of Aluminum shielding and factoring for extended mission considerations.

5.8.3 Displacement Damage Dose

The displacement damage dose or non-ionizing dose is the mean energy deposited in a material that goes into atomic displacements divided by the mass of the material. It is analogous to the ionizing dose except that the energy considered produces displacements in a semiconductor lattice.

The silicon and gallium arsenide EEE parts susceptible to displacement damage degradation shall be able to tolerate a minimum DDD of 7.58 x 108 MeV/g (or a 10 MeV equivalent proton fluence of 9.60×1010 cm-2 ) and 6.26 x 108 MeV/g (or a 10 MeV equivalent proton fluence of 9.50×1010 cm-2 ), respectively.

This meets the DDD exposure for a 10.5-year mission with orbit about the Sun-Earth second Lagrange point, L2, at the 95% confidence level and nominal 100 mils of aluminum shielding.

5.8.4 Single Event Effects

a. The pressure transducers shall be designed to avoid or tolerate errors due to non-destructive Single Event Effects (SEE).

b. All EEE parts shall have LET thresholds for SEL greater than 75 (MeV·cm2)/mg for the destructive events. Parts with LET thresholds between 37 MeV-cm2/mg and 75 MeV-cm2/mg could be used if the evaluation of the probability and outcome of the destructive event shows that there would be no mission impact. Reference the RST Radiation Specification (RST-REF-06763) for details on the environments.

5.8.5 Charging Environment

The pressure transducers shall be designed to withstand the degradation of surface materials and associated surface charging effects due to the radiation environment for the RST mission orbit.

6 CLEANLINESS

The requirements in this section ensure the cleanliness of the pressure transducers at delivery, so as not to adversely affect their 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 their design, materials of construction, or operation.

6.1 Surface Contamination

6.1.1 Surface Contamination Levels at Delivery

6.1.1.1 External Cleanliness

The external surfaces of the pressure transducer shall meet IEST-STD-CC1246E Particulate Cleanliness Level 300 and shall also be visually inspected and meet Level VC-0.5-1000 + UV, or equivalent, when inspected with both UV and white light in a darkened room

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 pressure transducer contractor shall not employ any of the following particle generating materials or processes into the pressure transducer 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 pressure transducer materials shall have a total mass loss (TML) less than 1.00% and a collected volatile condensable mass (CVCM) less than 0.10% at 125C under vacuum for twenty-four hours, 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 EEE Parts

The pressure transducer contractor’s Quality Assurance system for EEE parts will be in accordance with the requirements in the SOW, RST-PROP-SOW-0017.

7.1.2 Materials

The pressure transducers will be comprised of materials and processes in accordance with the requirements in the SOW, RST-PROP-SOW-0017.

7.1.2.1 Material Conductivity

All parts should be passivated and mounting surfaces on the pressure transducers shall be conductive as defined in Section 3.4.

7.2 Electrical

7.2.1 Test Sensors

a. Test sensors shall be designed for flight.

b. Unless specified to be removed before flight, test sensors shall not be removed prior to flight.

7.2.2 Interface Requirements

7.2.2.1 Connector Selection

Connector Specifications

a. Connectors / Receptacles shall be selected from EEE-INST-002, Section C2 and approved by the COR. The Receptacles mounted on the pressure transducers shall be Hermetic Receptacles. The mating connectors shall be supplied with the pressure transducers.

b. Environmental seals shall not be used in connectors especially if made from silicone without the explicit approval by the NASA/GSFC COR.

External Box Power Connectors The use of micro-miniature D connectors shall be avoided for power interfaces.

Contact Derating The current carrying capacity of the contacts shall be derated for continuous operation at the required current levels in a vacuum, as defined in Section C2 of EEE-INST-002.

Redundant Contact Derating When redundant contacts are used for a single power source, each contact shall meet the required derating criteria.

7.2.2.2 Signal Segregation

a. Wherever possible, different classes of signals (power, digital, analog, etc.) shall be separated by using separate connectors.

b. If separate connectors are not feasible, classes of signals within a common connector shall be isolated from one another. Connector pin assignments should be such that sensitive circuits are separated from potential interference sources.

7.2.2.3 Test and Flight Signal Isolation

Test signals and flight signals shall not be located in the same connector.

7.2.2.4 Test Interfaces

Component test signals that require access during observatory-level testing will be handled as follows:

Facility-Induced Noise All test signals should be protected or isolated from facility-induced noise.

Facility-Induced ESD GSE Malfunction All test signals shall be protected or isolated from facility-induced ESD GSE malfunction.

Facility-Induced GSE Malfunction All test signals shall be protected or isolated from facility-induced GSE malfunction.

7.2.3 Dielectric Structures

All dielectric structures shall have a charge bleed path to the spacecraft interface, designed to route the discharge into the spacecraft structure in a controlled fashion.

Dielectric structures with a bulk resistivity of greater than 1012 ohm-cm shall not be used.

7.3 Electromagnetic Compatibility

The design will be such that they are electromagnetically compatible with each other and with all the external environments in which they are expected to operate. Specifically, the components will not: (a) generate electromagnetic interference that could adversely affect the operation and safety of each other, the launch vehicle, and other equipment present during ground operations;

and (b) be susceptible to emissions that could adversely affect their operation and safety. These emissions may originate from each other, from the launch vehicle, or from other equipment present during ground operations.

The mode of transmission of electromagnetic interference may be conduction through wires or radiation through space. Electromagnetic compatibility requirements are designed to limit both conducted and radiated emissions and to minimize susceptibility to both types of emissions.

These requirements can be grouped into four types:

(1) Conducted emission requirements are designed to prevent excessive noise from being induced on the spacecraft power bus by components, subsystems, and instruments that are connected to the bus.

(2) Conducted susceptibility requirements are designed to insure that any component, subsystem, or instrument connected to the spacecraft power bus is not unduly susceptible to noise expected to be present on the power bus.

(3) Radiated emission requirements are designed to prevent excessive noise from being radiated by the observatory and its components.

(4) Radiated susceptibility requirements are designed to ensure that spacecraft components, subsystems, or instruments are not unduly susceptible to radiated noise emanating from the observatory and other sources.

The following table summarizes the emission and susceptibility requirements.

Table 7-1 Emission and Susceptibility Requirements

Test Types Applicability Test Reference

(MIL-STD-

461F or 461C)

GEVS

Paragraph Reference

Conducted Emissions

(CE)

Components with Unregulated Power Input

CE101 2.5.2.1.1

CE102/CE03 2.5.2.1.1

Common Mode

Test 2.5.2.1.2

Conducted Susceptibility (CS)

Components with Unregulated Power Input

CS101 &

CS02/CS114

2.5.2.2.1 & 2.5.2.2.4.2

Radiated Emissions

(RE)

Applicable to all components

RE102 2.5.2.3.2

Radiated Susceptibility (RS)

Applicable to all components

RS103 2.5.2.4.2

RF

Component/Antenna Test

RF Transmitters & Receivers only

CE106 2.5.2.1.4

CS103 2.5.2.2.2

CS104 2.5.2.2.2

7.3.1 Conducted Emissions

Conducted emission requirements are designed to prevent excessive noise from being induced on the spacecraft power bus by components that are connected to the bus.

7.3.1.1 Applicability of Conducted Emissions

The conducted emissions requirements are applicable for the primary power (+28V bus) and return leads that supply power to the component. The requirements do not apply to secondary power wires or signal wires.

7.3.1.2 CE101 – Differential Mode Current Emission Limits (Low Frequency) The pressure transducers shall not produce reflected ripple greater than the CE101 limits from 30 Hz to 150 KHz shown in Figure 7-1, when power leads are tested to a modified MIL-STD-461F CE101 (extended frequency range) test method.

7.3.1.3 CE102/03 – Differential Mode Current Emission Limits (Low Frequency) The pressure transducers shall not produce reflected ripple greater than the limits from 150 KHz to 50 MHz shown in Figure 7-1, when power leads are tested to the MIL-STD-461 CE102/CE03 test method.

Figure 7-1 CE101/CE03 Differential Mode Current Emission Limits

7.3.1.4 Common Mode Current Emissions Limits

Common Mode Current Emissions (CMCE) from the pressure transducer power and return leads shall not exceed the limits from 30 Hz to 200 MHz in Figure 7-2, referenced to chassis, when tested per GSFC-STD-7000A, Section 2.5.2.1.2.

10 100 1000 10000 100000 1000000 10000000 100000000 dB uA

Frequency Hz

< 500 Hz < 5 kHz < 50 kHz

CE101 30 Hz to 150 KHz CE03 150 kHz to 50 MHz

Measurement Bandwidth < 5 Hz

Component and Instrument Test Levels (DM)

(110 dBuA, 30 Hz) (110 dBuA, 350 Hz)

(62 dBuA, 150 kHz)

(26 dBuA, 2 MHz)

(26 dBuA, 50 MHz)

Figur…