Attachment B WFIRST-COMM-SPEC-0090-.pdf

Effective Date: February 6, 2020

Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

National Aeronautics and Space Administration

Goddard Space Flight Center Greenbelt, Maryland

WFIRST-COMM-SPEC-0090, Revision -

Wide Field Infrared Survey Telescope (WFIRST), Code 448

WFIRST High Gain Antenna

Specification

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GSFC WFIRST CMO

February 10, 2020

Released

WFIRST High Gain Antenna Specification WFIRST-COMM-SPEC-0090, Revision -

Effective Date: February 6. 2020 ii Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

WFIRST High Gain Antenna

Specification Review/Signature/Approval Page

Prepared by:

Rene Gosselin

Approved by:

Jason Hylan iii Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

Preface

This document is a Wide Field Infrared Survey Telescope (WFIRST)

Configuration Management (CM)-controlled document. Changes to this document require prior approval of the applicable Configuration Control Board (CCB) Chairperson or designee.

Proposed changes shall be submitted to the WFIRST 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:

WFIRST Configuration Management Office

Mail Stop 448

Goddard Space Flight Center

Greenbelt, Maryland 20771 iv Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

Change History Log

Revision Effective Date Description of Changes

(Reference the CCR & CCB/ERB Approval Date)

Revision - February 6, 2020 Initial Release of Document per WFIRST-CCR-0085 v Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

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 HGA Description

2.2 Ground Support Equipment Description

3 FUNCTIONAL/PERFORMANCE REQUIREMENTS

3.1 HGA Flight Unit Functional/Performance Requirements

3.1.1 HGA Coordinate System

3.1.2 HGA Subassemblies

3.1.2.1 Primary HGA Reflector Subassembly

3.1.2.2 Ka-Band Secondary Frequency-Selective Surface (FSS) Subassembly

Embedded FSS Surface Material

3.1.2.3 Ka-Band Feed Horn Subassembly

3.1.2.4 S-Band Feed Subassembly

3.1.2.5 S-Band Cable

3.1.2.6 Support Struts

3.1.2.7 Mounting Plate

3.1.3 HGA Assembly Performance

3.1.3.1 Surface and Alignment Accuracies

3.1.3.2 As-Built Alignment and Pointing Errors

3.1.3.3 On-Orbit Deformations

3.2 HGA Assembly Resource Allocations

3.2.1 HGA Assembly Mass Allocation

3.3 Power

3.4 Electrical Grounding

3.4.1 Ground Strap, Non-Conductive Surface

3.5 Signal and Data Interfaces

3.5.1 RF Signals

3.5.1.1 RF Signal Interface

3.5.1.2 Multipaction

3.6 Operating Modes

3.7 Command and Data Services

3.8 Flight Software

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

vi Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

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 Fields of View

4.6 Alignment

4.6.1 Alignment Location/Orientation Accuracy

4.6.2 Alignment Knowledge Accuracy

4.6.3 Alignment Stability

4.6.4 Alignment Method

4.6.5 Alignment Measurement Tooling/Features

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 Acoustics

5.6 Pressure

5.6.1 Operating Pressure Range

5.6.2 Maximum Depressurization Rate

5.6.3 Launch Vehicle Environmental Control System Impingement Velocity

5.7 On-Orbit Dynamic Environment

5.8 Ground Environments

5.9 Thermal Requirements

5.9.1 Flight Interface Design Temperature Limits

5.9.2 Allocation of Spacecraft Monitored Temperature Sensors

5.10 Charged Particle Radiation Requirements

5.10.1 Definitions

5.10.2 Total Ionizing Dose

5.10.2.1 Minimum TID Tolerance for EEE Parts and Materials

5.10.3 Displacement Damage Dose

5.10.4 Single Event Effects

5.10.5 Charging Environment

5.11 Plume Impingement

6 CLEANLINESS

6.1 Surface Contamination

6.1.1 Surface Contamination Levels at Delivery

6.1.1.1 Particulate Contamination

6.1.1.2 Molecular Contamination – Exposed Surfaces

6.1.1.3 Molecular Contamination – Covered Surfaces

6.1.2 Surface Contamination Generation

6.1.2.1 Particulate Generation

vii Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

6.1.2.2 Molecular Generation

Material Selection Material Selection - Silicones

Assembly Outgassing

6.1.2.3 Intentional and Unintentional Vents

6.1.2.4 Intentional and Unintentional Vents – Filters

6.1.2.5 Intentional Vents – Pressure Buildup Prevention

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

6.2 Electrostatic Cleanliness

6.2.1 Conductive Surface Ground Path

6.2.2 Conductive Surface Resistivity

6.2.3 Closeout of Gaps and Apertures

6.2.3.1 Conductive Tape Surface Resistivity

6.2.3.2 Conductive Tape, Grounding

6.2.4 Exposed Harness Specific Requirements

6.2.5 Thermal Blankets

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

7.1.3 Software Assurance

7.2 Electrical

7.3 Safety

7.4 Electromagnetic Compatibility

7.5 Identification and Marking

7.6 Workmanship

7.6.1 Workmanship Standards

7.6.2 Connector

7.6.2.1 GSE Cable Connectors

Accessibility

7.6.2.2 Test Connectors

7.6.2.3 Connector Identification

7.6.2.4 Protection of Unused Test Connectors

7.7 Reliability and Mission Lifetime

7.7.1 Mission Life

7.7.2 Operating Time

7.7.2.1 Total Time

7.7.2.2 Failure Free Time

7.8 Ground Handling

7.8.1 Ground Support Equipment (GSE) Design

7.8.2 Lifting Hardware

7.8.3 Dollies, Stands, Shipping Containers

viii Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

7.8.4 Lift Point Capability

7.8.5 Lift Point Locations

7.8.6 Lift Point Proof Testing

7.8.7 GSE Cleanliness

7.8.8 GSE Bakeout

7.8.9 Test Harness

8 MECHANICAL DESIGN REQUIREMENTS

8.1 Mechanical Factors of Safety

8.2 Metallic Reliability

8.3 Fracture Control Requirements

8.4 Metallic Pressurized Lines and Fittings Burst Factor

8.5 Electronics Box Minimum Vent Area

8.6 Materials

8.6.1 Dissimilar Metals

8.6.2 Material Allowables and Stiffness

8.6.3 CTE Data

8.6.4 Stress Corrosion Cracking

8.7 Joints

8.7.1 Joint Edge Distance

8.7.2 Bonded Joint Allowable

8.7.3 Reserved

8.7.4 Fastened Joints

8.7.4.1 Fastener Locking

8.7.4.2 Fastened Joint Margin of Safety

8.7.4.3 Installation Torque Documentation

8.7.4.4 Seal Analysis

8.7.4.5 Critical Fasteners

8.8 Mechanism Design

8.9 Mechanism Verification Requirements

9 LOGISTICS

9.1 I&T Deliverables

9.2 Ground Support Equipment

9.3 Transportation Equipment

10 VERIFICATION REQUIREMENTS

10.1 Verification Methods

10.1.1 Inspection

10.1.2 Analysis

10.1.3 Test

10.2 Inspection Requirements

10.2.1 Visual Inspection

10.2.2 Physical Measurement

10.2.3 Documentation Search

10.3 Analysis Requirements

10.4 Test Requirements

10.4.1 Definitions

ix Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

10.4.2 Test Factors

10.4.3 Thermal Cycle Prior to Structural Testing

10.4.4 Powered-on Conditions for Test

10.4.5 Proof Testing Factors

10.4.6 Test Tolerances

10.4.7 Test Restrictions

10.4.7.1 Failure During Tests

10.4.7.2 Modification of Hardware

10.4.7.3 External Adjustment

10.4.7.4 Re-Test Requirements

10.5 Required Tests

10.5.1 Performance Tests

10.5.1.1 Antenna Performance Tests

10.5.1.2 Antenna VSWR Performance Test

10.5.1.3 Antenna Pattern Performance Test

10.5.2 Mass Properties Measurement

10.5.3 Static Loads/Strength Test

10.5.3.1 Sine Burst

10.5.3.2 Static Pull

10.5.4 Frequency Signature Survey

10.5.5 Modal Survey

10.5.6 Sine Vibration

10.5.7 Random Vibration

10.5.8 Acoustic Test

10.5.9 Proof Pressure Test

10.5.10 Shock

10.5.11 Thermal Vacuum Bake-out

10.5.12 Thermal Vacuum Test

10.5.12.1 Thermal Vacuum Test Parameters

10.5.12.2 Thermal Vacuum Test Profile

10.5.12.3 Thermal Vacuum Test Verification

10.5.13 Thermal Balance Requirement

10.5.14 Thermal Cycling Testing - Ambient

10.5.15 Harness Tests

10.5.16 EMI/EMC Tests

APPENDIX A ABBREVIATIONS AND ACRONYMS

x Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

List of Figures

Figure 2-1 HGA Dual-Band Operation Figure 3-1 HGA Subassembly Identification

Figure 3-2 HGA STEP File to define Reflector Surfaces Figure 3-3 Secondary FSS Subassembly Over S-band Feed Figure 3-4 FSS RF Layer Construction Figure 3-5 Ka-band Feed Horn Subassembly Figure 3-6 S-band Feed Antenna with Example Mounting

Figure 10-1 Thermal Vacuum Profile

List of Tables

Table 5-1 HGA Design Limit Loads Table 5-2 HGA Sine Vibration Environment Table 5-3 Random Vibration Limit Levels for Components Weighing Less Than 22.7 kg

Table 5-4 ASD Limit Load Reduction Table 5-5 Minimum Workmanship ASD Level for Components Weighing Less than 45.4 kg. . 16 Table 5-6 Minimum Workmanship Reduction Table 5-7 Internal and External Shock Source Limit Levels Table 5-8 Acoustic Limit Sound Pressure Level (SPL)

Table 5-9 Depressurization Profile Table 5-10 Temperature Limits

Table 8-1 Design Factors of Safety Table 8-2 Minimum Edge Distance Table 10-1 Test Factors and Durations

Table 10-2 Proof Test Factors Table 10-3 Test Tolerances

Table 10-4 Thermal Vacuum Test Parameters

Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

1 INTRODUCTION

1.1 Purpose

The Wide Field InfraRed Survey Telescope (WFIRST) 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 specification document defines the requirements for the WFIRST High Gain Antenna (HGA).

1.2 Scope

This specification describes the electrical, mechanical, environmental, and verification testing requirements for a space-qualified HGA for the NASA Goddard Space Flight Center (GSFC)

WFIRST Mission.

1.3 Related Documentation

The latest versions of all documents below should be used. WFIRST documents can be obtained from the CM tool.

1.3.1 Applicable Documents [and Forms]

The following documents are referenced within this document and are directly applicable or contain policies or other directive matters that are binding for the contents of this document. In the event of conflict between an Applicable Document and the content of this document, the

WFIRST Project Configuration Change Board (CCB) has the final authority for conflict resolution.

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

EEE-INST-002 Instructions for EEE Parts Selection, Screening, Qualification, and Derating

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

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-5020 Requirements for Threaded Fastening Systems in

Spaceflight Hardware

NASA-STD-8719.24 NASA Expendable Launch Vehicle Payload Safety

Requirements file:///C:/Users/ddusterw/AppData/Local/Temp/1/from

Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

GSFC-STD-7000A General Environmental Verification Standard (GEVS)

541-WI-5330.1.41 Fastener Locking Using Arathane 5753

WFIRST-SYS-PLAN-0051 WFIRST Contamination Control Plan

WFIRST-COMM-LIST-0061 WFIRST High Gain Antenna Deliverable Items List and

Schedule (DILS)

WFIRST-COMM-SOW-0026 WFIRST High Gain Antenna Statement of Work

WFIRST-COMM-ANYS-0153 High Gain Antenna (HGA) Reference STEP file

2221336 ANTENNA MICD, 1.7 METER DIAMETER, HGAS

2237031 FSS MEMBRANCE, MICD, HGA WFIRST

2237040 KA-BAND FEED ASSEMBLY, MICD, HGA WFIRST

2237050 S-BAND FEED ASSEMBLY, MICD, HGA WFIRST

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

None

Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

2 CONTRACT DESCRIPTION

2.1 HGA Description

The WFIRST HGA assembly is a passive microwave antenna that is a component of the WFIRST spacecraft deployable High Gain Antenna System (HGAS), capable of providing S-band and Ka-band communications as a connected part of the RF Communications Subsystem. After spacecraft launch and deployment of the HGAS boom, a gimbal mechanism located at the end of the boom will point the HGA to enable the transmission and reception of free-space RF radiation energy between the spacecraft and one of NASA’s communications assets on or near the earth.

The HGA is designed to operate by transmitting and receiving at both S-band and Ka-band frequencies using separate feeds and radiation paths for each of the two bands. The overall dual-band RF operation is depicted below in Figure 2-1. For both frequency bands, the main far-field radiation beam of the antenna is collimated by the energy distribution on the solid primary reflector surface. The antenna operates reciprocally between transmit and receive modes, although on-orbit the HGA is only used at Ka-band for transmission. A significant challenge in the fabrication of this antenna is achieving the required mechanical alignment of the feeds and the specified surface properties of the reflectors over the orbital temperature range, since all are critical to achieve the overall antenna performance.

Figure 2-1 HGA Dual-Band Operation

For the case of Ka-band frequencies, a conical Ka-band feed horn is mounted at the base of the primary reflector surface. Energy fed into the Ka-band horn radiates to the convex side of a curved

Frequency Selective Surface (FSS) subassembly that is designed to reflect only Ka-band radiation, and thereby at that frequency operate as a secondary reflector to illuminate the primary reflector surface. Both the FSS subassembly and primary reflector surfaces are shaped for optimum gain performance. The design and location of the FSS subassembly is critical to properly reflect Ka-band radiation and ensure the solid primary reflector surface is correctly illuminated to generate the far-field Ka-band antenna beam. The RF design of the FSS is considered as Government

Furnished Equipment (GFE).

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For the case of S-band frequencies, an S-band feed assembly is mounted inside the concave side of the FSS subassembly, and its phase center is located at the focal point of the solid primary reflector surface. This concave surface consists of a dielectric layer with the FSS embedded on the concave side. The FSS is designed to be highly RF transparent to the S-band frequency radiation propagating to and from the S-band feed above it, and highly RF reflective to Ka-band radiation.

Radiation transmitted from the S-band feed passes directly through the FSS subassembly with very little attenuation and is reflected by the solid primary reflector surface, which collimates the energy into the far-field S-band antenna beam.

2.2 Ground Support Equipment Description

The following items shall be delivered as GSE:

Handling Fixture for the HGA assembly – Shall enable easy removal from the storage container and lifting by hand for integration and testing purposes.

Storage Container / Transit case – Shall have polymer surfaces that can be wiped down for cleanroom usage. Use of internal foam is allowable, but only if enclosed with Mylar film, or other approved barrier material. The container shall have casters, provide water-resistance, and feature lift-points.

Requirements for all GSE are further detailed in Section 9.2.

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3 FUNCTIONAL/PERFORMANCE REQUIREMENTS

This section defines the functional and performance requirements for the HGA as defined in

Section 3.1. All of the written requirements in this document apply over all environmental conditions, such that the HGA assembly survives the launch environment and performs as specified while in orbit until the end of spacecraft life (EOL).

3.1 HGA Flight Unit Functional/Performance Requirements

In the requirements below, the functioning surface of the reflectors refers to the side that is illuminated by the RF energy. For the primary reflector, this refers to the face pointing towards the

S-Band Feed. For the FSS subassembly, the concave surface facing the primary reflector and the

Ka-band feed horn operates as a secondary reflector for Ka-band frequencies.

3.1.1 HGA Coordinate System

The origin of the HGA coordinate system coincides with the center of the aperture of the Ka-band feed horn, at room temperature. The positive Z-axis extends from the origin in a direction normal to the mounting plate, pointing toward the secondary. The X-axis and the Y-axis are in the plane of the feed horn aperture with orientation defined in the contractor-provided Mechanical Interface

Control Document (MICD).

3.1.2 HGA Subassemblies

The HGA assembly consists of: (1) A solid primary antenna reflector; (2) A curved FSS subassembly incorporating a conformal FSS layer notionally supported by three struts that are also designed to support the S-band feed assembly; (3) A Ka-band feed horn with S-band cross-pol cup

(supplied as GFE); (4) A low profile S-band feed (supplied as GFE) housed inside the convex side of the curved FSS subassembly and positioned to radiate towards the focal point of the primary reflector; (5) the S-band feed cable; (6) Support struts suspending the S-Band Feed and FSS subassembly; and (7) the Feed Support cone. This is illustrated in the figure below.

Figure 3-1 HGA Subassembly Identification

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3.1.2.1 Primary HGA Reflector Subassembly

The nominal diameter of the primary reflector shall be 1.7 meters (m) and shall be rotationally symmetric. The precise shape of the primary reflector shall meet the provided RF-active surfaces as defined in the High Gain Antenna (HGA) Reference STEP file, WFIRST-COMM-ANYS-0153.

The primary reflector as viewed from the front shall accommodate the installation and removal of the entire HGA assembly. Final design implementation shall obtain NASA/GFSC COR approval.

Figure 3-2 HGA STEP File to define Reflector Surfaces

The surface of the primary reflector shall be less than 6% specular in the primary solar wavelengths to prevent the focusing of solar flux and thus accumulating heat, on the secondary FSS subassembly. If reflector surfaces are implemented using composite fibers, the toe size, weave and orientation shall be globally and locally net symmetric so that the reflector does not show a directional preference for electric polarization. Reflector materials chosen shall constrain the RF reflection loss to 0.1 dB, compared to a perfect electric conductor, and the selection shall obtain

NASA/GSFC COR approval.

3.1.2.2 Ka-Band Secondary Frequency-Selective Surface (FSS) Subassembly

The nominal diameter of the secondary FSS subassembly shall be 0.214 meters (m) and shall be rotationally symmetric. The precise shape of the curved FSS subassembly shall meet the provided

RF-active surfaces as defined in the High Gain Antenna (HGA) Reference STEP file, WFIRST-

COMM-ANYS-0153. The FSS covers the S-band feed as shown in the figure below.

Figure 3-3 Secondary FSS Subassembly Over S-band Feed

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Embedded FSS Surface Material

The FSS subassembly shall incorporate a precision-designed Frequency Selective Surface (FSS) membrane (provided as GFE) that enables the passage of the S-band frequencies with very little attenuation while reflecting the Ka-band frequencies. The main physical characteristics of the FSS subassembly are shown in the figure below and FSS MEMBRANCE, MICD, HGA WFIRST drawing 2237031.

During manufacture of the FSS subassembly, the FSS membrane material shall be conformed to the contours of the curved secondary dielectric material surface, as depicted in figure below. The contractor may choose an alternative method, to be approved by the NASA/GSFC COR, which ensures that the conformal mapping scheme does not lead to deviations of the x and/or y values of more than 5%.

Figure 3-4 FSS RF Layer Construction

The contractor shall propose a dielectric material with (εr=3.2 ±0.1 and tanδ<0.004, preferably

Kapton) subject to NASA/GSFC COR approval. When the assembly is bonded, the non-metallized

Kapton surface remains exposed, as the outer-most surface.

3.1.2.3 Ka-Band Feed Horn Subassembly

The Ka-band feed horn with polarizer (supplied as GFE), as defined in the KA-BAND FEED

ASSEMBLY, MICD, HGA WFIRST drawing 2237040 shall be mounted as part of the HGA, centered about the Z-axis and with phase center at the HGA coordinate system origin at room temperature, as defined in Section 3.1.1, and such that all performance requirements shall be satisfied over all conditions.

Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

Figure 3-5 Ka-band Feed Horn Subassembly

3.1.2.4 S-Band Feed Subassembly

The S-band feed assembly (supplied as GFE), as defined in the S-BAND FEED ASSEMBLY, MICD, HGA WFIRST drawing 2237050 shall be mounted inside the concave side of the FSS subassembly, as shown in Figure 3-6 S-band Feed Antenna with Example Mounting below. The

FSS subassembly is transparent for S-Band radiation but is a reflective surface for Ka-Band frequencies, therefore it serves as a secondary reflector only at Ka-Band frequencies.

The S-band feed shall be incorporated into the HGA such that all performance requirements are satisfied over all conditions.

Figure 3-6 S-band Feed Antenna with Example Mounting

3.1.2.5 S-Band Cable

The S-Band Coaxial feed cable will be provided as GFE. The length and other properties shall be approximately 1 meter, subject to the engineering design trade-offs encountered during development of the strut assembly used to support the S-band feed.

The S-Band coaxial cable shall be routed through and secured properly to one of the struts and continued on the back side of the main reflector to an accessible location on the back plate where it will be connected to a double-sided bulkhead TNC connector compliant with EEE-INST-002

Level 2 requirement. The mechanical routing and attachment of the two TNC connectors, at the

S-band feed assembly end and the bulkhead end, is to be proposed by the vendor, and is subject to

NASA/GSFC COR approval.

The feed cable shall be routed along one strut and penetrate the primary reflector.

Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

3.1.2.6 Support Struts

Three struts shall be used to support the secondary FSS subassembly above the primary reflector and shall initially extend outward to not violate the stay-out region as defined in the High Gain

Antenna (HGA) Reference STEP file, WFIRST-COMM-ANYS-0153.

The total blockage by the physical structure of the struts shall not exceed 4% of the total projected area of the primary reflector that is not blocked by the FSS subassembly. To minimize diffraction a wedge shape shall face the primary reflector.

The structural interface to the primary reflector shall remain within the envelope requirements, noting the three provisions for these attachments.

Struts shall support the mass (approximately 250 grams) and moment of the S-band feed antenna, in addition to a hat coupler of 1.0 Kg mass.

Alternate designs toward improvement of the reflector assembly’s mechanical performance may be considered; however, their implementation shall require NASA/GSFC COR approval pending an RF performance analysis impact.

3.1.2.7 Mounting Plate

The HGA assembly shall include an aluminum “HGA Mounting Plate” with an aluminum or composite Feed Support for interfacing the HGA to the spacecraft two-axis gimbal.

The HGA Mounting Plate shall support the following:

The Ka-band feed horn

An S-band TNC pass-through female-female-bulkhead connector

Accessibility to the S-band TNC connector and the Ka-band feed horn flange

A thermal isolation system, isolating the Ka-band horn from the HGA reflectors and isolating the HGA reflectors from the Mounting Plate

Provisions for grounding attachment

The HGA mounting plate shall make provision for the HGA to be removed from the front

“radiating side” with fasteners strategically located to minimize RF effects – such as aligned with the projections of the struts and secondary reflector. Removal of the reflector assembly may leave behind the Ka-band feed assembly to eliminate waveguide disconnections.

The Feed Support shall provide, from the front radiating side of the HGA, physical access to the side port of the polarizer, which is part of the Ka-band Feed Horn Subassembly for GSE.

3.1.3 HGA Assembly Performance

3.1.3.1 Surface and Alignment Accuracies

All calculations shall be made in the HGA coordinate system as defined in Sections 3.2.1.

All Root Mean Squared (RMS) calculations shall be weighted equally throughout the reflector surface on an area basis and use enough data points to ensure that any error stays within this specification limit.

Use or disclosure of data contained on this page is subject to the restriction(s) on the title page of this document.

The alignment of each reflector shall be defined through a best-fit routine to the nominal surface, accounting for the specified surface errors.

3.1.3.2 As-Built Alignment and Pointing Errors

The as-built RMS error of the reflector shapes, as compared to the ideal shape, at ambient temperature shall not exceed 0.50 mm for the primary and 0.20 mm for the secondary.

The as-built lateral (X/Y-axis) alignment error of each reflector, as compared to the ideal location, shall not exceed 0.40 mm for the primary and 0.30 mm for the secondary. The as-built alignment error along the z-axis, as compared to the ideal location, may deviate up to 0.5 mm.

The intent of these error allocations is to constrain the as-built RF pointing error. The as-built

Pointing error shall not exceed 0.15 degrees. However, the contractor may instead elect to meet this requirement in terms of the specified angular pointing error. The government will use commercially available physical optics software and/or hardware to validate this specification.

3.1.3.3 On-Orbit Deformations

The on-orbit rms error of the reflector shapes, as compared to the as-built shape, shall not exceed

0.075 mm for the primary FSS subassembly and 0.030 mm for the secondary.

The on-orbit lateral (x/y-axis) alignment error of each reflector, as compared to the as-built location, shall not exceed 0.035 mm for the primary and 0.025 mm for the secondary.

The on-orbit alignment error along the z-axis, as compared to the as-built location, shall not exceed

0.5 mm.

The intent of these error allocations is to constrain the on-orbit RF pointing error (after the as-built error has been removed through on-orbit calibration). The on-orbit RF pointing error shall not exceed 0.02 degrees. However, the contractor may instead elect to meet this requirement in terms of the specified angular pointing error. The government will use commercially available physical optics software to validate this specification if so elects.

3.2 HGA Assembly Resource Allocations

3.2.1 HGA Assembly Mass Allocation

Requirements in this section are for the overall assembly. Section 4.2.1 contains more requirements on mass properties.

The HGA Assembly shall have a mass of less than or equal to 16 kg.

The HGA shall be nominally rotationally symmetric, such that the center of gravity (CG) lies along the mechanical boresight (Z) axis within a 5 mm radius, and at a maximum of 200 mm from the

HGA mounting interface plane.

Maximum principal moments of inertia values, as calculated about the HGA CG, shall not exceed the following:

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Ixx, Iyy = 3.0 kg m2

Izz = 5.5 kg m2

3.3 Power

Reserved

3.4 Electrical Grounding

3.4.1 Ground Strap, Non-Conductive Surface

The HGA shall contain provisions for a grounding strap to be attached from the component chassis for connection to the spacecraft conductive structure, including for electrostatic cleanliness as specified in Section 6.2.

3.5 Signal and Data Interfaces

3.5.1 RF Signals

3.5.1.1 RF Signal Interface

Accommodations shall be made for the WR-34 waveguide interface from the Ka-band feed horn assembly, and for the S-band RF cable interface.

3.5.1.2 Multipaction

The HGA shall be designed to preclude damage or measurable degradation in performance due to multipaction while operating in a vacuum environment. This requirement is not levied on the vendor for any GFE provided hardware.

3.6 Operating Modes

Reserved

3.7 Command and Data Services

Reserved

3.8 Flight Software

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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 contractor-provided Mechanical Interface Control

Drawing (MICD), which will be further 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

Reserved

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 HGA shall be determined to within ±5 mm relative to the boresight

(axial) and radial directions.

4.2.4 Determination of Moments and Products of Inertia

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

4.3 Physical Envelope

The HGA outside dimensions, including mounting flanges and connectors, shall not exceed the volume envelope as defined in ANTENNA MICD, 1.7 METER DIAMETER, HGAS drawing

2221336.

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4.4 Mounting

a. Mounting interface flatness, and co-planarity requirements for the component side of the interface (including brackets, if any, and shims) shall be as defined in the MICD.

4.5 Fields of View

The secondary reflector shall have an unobstructed view of the central portion of the primary reflector as defined by the “stay-out” regions defined in the High Gain Antenna (HGA) Reference

STEP file, WFIRST-COMM-ANYS-0153, as applied to the struts and any other obstructions.

4.6 Alignment

4.6.1 Alignment Location/Orientation Accuracy

Alignment location accuracy shall be suitable to ensure meeting the accuracy requirements of

Section 3.1.3 with margin.

4.6.2 Alignment Knowledge Accuracy

Alignment knowledge accuracy shall be suitable to ensure meeting the accuracy requirements of

Section 3.1.3 with margin.

4.6.3 Alignment Stability

The assembly shall be rigidly bonded to ensure stability, and over temperature as confirmed during thermal verification.

4.6.4 Alignment Method

The subassembly parts of the HGA shall be measured and aligned as a complete assembly using a laser tracker to quantify the alignment, or with another NASA/GSFC COR approved method.

4.6.5 Alignment Measurement Tooling/Features

Three or more tooling/features shall remain as part of the HGA enabling a laser tracker to re-establish the coordinate system used during surface accuracy configuration and photogrammetry.

They shall be located to minimize negative effects to the RF performance.

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5 ENVIRONMENTAL REQUIREMENTS

Environmental design requirements are specified in this section.

The HGA shall meet its performance requirements in section 3.0 of this document 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 HGA 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 HGA Design Limit Loads

Component Mass (Kg) Limit Load (g)

5 30.1

10 26.5

20 22.4

5.2 Frequency Requirement

5.2.1 Stowed Fundamental Launch Frequencies

The HGA shall have a fundamental frequency greater than 75 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 and discussed in Section 10.5.5.

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

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5.3 Vibration

5.3.1 Sinusoidal Vibration

The HGA 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 Table 10-1.

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 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 until higher levels of assembly upon approval by the

NASA/GSFC COR.

Table 5-2 HGA Sine Vibration Environment

Frequency (Hz) Limit Level

5 to 100 10.0 g

5.3.2 Random Vibration

The HGA shall withstand the random vibration environment in Table 5-3 or Table 5-5, whichever is applicable.

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. For components weighing more than 22.7 kg, relief can be provided in accordance with

Table 5-4. Levels defined are acceptance levels. Analysis is conducted against acceptance with the safety factors defined in Table 8-1. Testing factors are defined in Table 10-1. Table 5-5 defines minimum workmanship for components weighing less than 45.4 kg. For components weighing

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more than 45.4 kg, relief can be provided in accordance with Table 5-6. 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 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 59kg. Above

59 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 kg, levels are maintained at the 182 kg level.

Table 5-5 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

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Table 5-6 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 182kg, the workmanship spectrum is 0.01g2/Hz from 20 to 2000 Hz.

5.4 Shock

The HGA shall be designed to meet its performance requirements after being subjected to the shock environment in Table 5-7, applied at the HGA interface to the WFIRST spacecraft structure.

Levels defined are acceptance levels. Testing factors are defined in 10.4.2.

Table 5-7 Internal and External Shock Source Limit Levels

Component Frequency (Hz) Limit Level

High Gain Antenna

100 100 g

100 - 2000 6.5 dB/oct

2000 - 10000 2500 g

Note: A shock susceptibility and attenuation assessment is to be performed on the HGA. 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 HGA.

If the environment is low risk, then the shock test may be deferred to the Observatory level with concurrence from the NASA/GSFC COR.

5.5 Acoustics

The HGA shall be designed to withstand, without any damage or degradation of performance, the acoustic loads shown in Table 5-8.

Levels defined are acceptance levels. Analysis is conducted against acceptance with the safety factors defined in Table 8-1. Testing factors are defined in Table 10-1.

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Table 5-8 Acoustic Limit Sound Pressure Level (SPL)

One-Third Center Frequency (Hz)

Limit Levels (dB)

31.5 123.5

40 127.5

50 130.0

63 131.5

80 132.5

100 133.0

125 133.0

160 133.0

200 133.0

250 133.0

315 133.0

400 131.0

500 129.0

630 126.5

800 124.5

1000 122.5

1250 120.7

1600 118.3

2000 116.5

2500 115.0

3150 113.0

4000 111.5

5000 109.5

6300 107.5

8000 106.0

10000 104.0

OASPL 143.1

5.6 Pressure

5.6.1 Operating Pressure Range

The HGA 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).

5.6.2 Maximum Depressurization Rate

The HGA shall be designed to meet all performance requirements after exposure to a depressurization profile shown in Table 5-9 during launch and ascent.

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Table 5-9 Depressurization Profile

Transonic Late Early sec psi sec psi sec psi

0.0 14.70 0.0 14.70 0.0 14.70

10.0 14.20 10.0 14.40 10.0 14.00

15.0 13.70 15.0 14.00 15.0 13.40

20.0 12.95 20.0 13.35 20.0 12.55

25.0 11.95 25.0 12.45 25.0 11.45

41.2 7.90 44.8 7.90 28.0 10.64

43.2 7.70 46.8 7.80 30.2 9.75

45.2 6.24 48.8 6.50 32.2 9.00

46.6 5.40 50.4 6.00 33.5 8.50

48.7 4.30 52.1 5.50 35.0 8.00

50.9 3.30 53.8 5.00 36.4 7.50

54.0 2.40 55.3 4.60 37.9 7.00

58.0 1.60 56.8 4.20 39.4 6.50

65.0 1.00 58.4 3.80 41.0 6.00

70.0 0.80 60.0 3.40 42.7 5.50

80.0 0.50 61.4 3.10 44.4 5.00

110.0 0.30 62.8 2.80 46.3 4.50

150.0 5.00 64.2 2.50 48.0 4.05

65.5 2.25 50.0 3.75

66.9 2.00 52.0 2.88

68.0 1.80 57.0 2.00

68.8 1.68 65.0 1.30

75.0 1.10 80.0 0.60

85.0 0.60 110.0 0.30

110.0 0.30 150.0 0.05

150.0 0.05

5.6.3 Launch Vehicle Environmental Control System Impingement Velocity

5.7 On-Orbit Dynamic Environment

Reserved

5.8 Ground Environments

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

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5.9 Thermal Requirements

5.9.1 Flight Interface Design Temperature Limits

a. The HGA shall be capable of surviving indefinitely when its temperatures are within the survival limits shown in Table 5-10 without damage or permanent performance degradation.

b. The HGA shall meet all performance requirements anywhere within the Operational and

Protoflight/Qualification limits shown in Table 5-10.

Table 5-10 Temperature Limits

Minimum Temperature (ºC) Maximum Temperature (ºC)

Operational (In Spec) 0 +128

Protoflight/Qualification

(In Spec)

-10 +138

Survival (Unpowered) -138 +143

Thermal gradients of up to 70 ºC within these limits across the primary reflector opposing edges shall be considered as an operational condition.

5.9.2 Allocation of Spacecraft Monitored Temperature Sensors

Reserved

5.10 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.

5.10.1 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.

5.10.2 Total Ionizing Dose

5.10.2.1 Minimum TID Tolerance for EEE Parts and Materials

The HGA shall tolerate a minimum Total Ionizing Dose (TID) of 29 krad (Si) over 10.5 years for a 100 mil aluminum shell.

5.10.3 Displacement Damage Dose

Reserved

5.10.4 Single Event Effects

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5.10.5 Charging Environment

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

5.11 Plume Impingement

The HGA assembly shall be designed to withstand any degradation of surface materials and performance due to periodic impingement of plumes from the spacecraft thrusters. Plume species may include any and all of N2, H2, and NH3.

Anticipated heating from plume impingement is included in the operational temperature limits stated in Section 5.9.1.

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6 CLEANLINESS

The requirements in this section ensure the cleanliness of the HGA 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 Particulate Contamination

The HGA shall meet IEST-STD-CC1246E VC-0.5-1000 + UV, or equivalent, when inspected with both UV and white light in a darkened room.

6.1.1.2 Molecular Contamination – Exposed Surfaces

If the HGA has surfaces exposed to the space environment, all such surfaces shall meet a molecular surface cleanliness level of IEST-STD-CC1246E R3.3E-1 on all external and critical surfaces, when tested in accordance with IEST-STD-CC1246E VC-0.5-1000 + UV, or equivalent.

6.1.1.3 Molecular Contamination – Covered Surfaces

a. All HGA surfaces not exposed to the space environment shall meet IEST-STD-CC1246E

VC-0.5-1000 + UV, or equivalent, when inspected in a darkened room with white and UV light.

b. The surfaces shall be free of molecular contamination – for example films, spots, or other.

6.1.2 Surface Contamination Generation

6.1.2.1 Particulate Generation

The HGA contractor shall not employ any of the following particle generating materials or processes into the HGA 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.).

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Foams, highly textured materials.

Trapped debris in holes.

6.1.2.2 Molecular Generation

Material Selection

The HGA 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.