GOODRICH CORPORATION Electro-Optical Systems 100 Wooster Heights Road Danbury, CT 06810-7589 Telephone (203) 797-5000

Document No.: P-4555 Rev A

White Paper Proposal for Wide Field Camera 3 Re-host Mission

31-March-2005

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This proposal or quotation includes data that shall not be used or disclosed outside the Contractor/theGovernment and shall not be duplicated, used or disclosed - in whole or in part - for any purpose other than toevaluate this proposal or quotation. If, however, a contract is awarded to this offeror or quoter as a result of - orin connection with - this submission of this data, the Contractor/the Government shall have the right to duplicate,use or disclose the data to the extent provided in the resulting contract. This restriction does not limit theContractor’s/the Government’s right to use information contained in this data if it is obtained from anothersource without restriction. The data subject to this restriction is contained in all sheets.

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Preface This document is a collaborative effort of Arizona State University Department of Physics and Astronomy, Goodrich Electro Optical Systems, General Dynamics C4 Systems - Spectrum Astro Space Systems. Points of contact are:

Jon Morse, PhD Assocoiate Professor Arizona State University Department of Physics and Astronomy Tempe, AZ 85287 480.965.2552 Jon.Morse@asu.edu Linda Abramowicz-Reed Advanced Development Group Goodrich Electro-Optical Systems 100 Wooster Heights Road Danbury, CT 06810 203.797.5764 linda.abramowicz-reed@goodrich.com John Dyster Director, Commercial and International Programs General Dynamics Corporation C4 Systems Spectrum Astro Space Systems 1440 N. Fiesta BLVD Gilbert, AZ 85233 480.892.8200 John.Dyster@specastro.com

Acknowledgements Paul Scowen (ASU) made important contributions to the text. Some of the science programs discussed are based on programs conceived by members of the Orion MIDEX and HORUS Origins Probe science teams.

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Table of Contents

1. OVERVIEW.............................................................................................................. 1

2. SCIENCE PROGRAM ............................................................................................. 1

3. OPTICAL BASELINE AND TRADE SPACE........................................................... 3

4. RE-HOST HARDWARE........................................................................................... 4

4.1 SCIENCE PAYLOAD ..................................................................................................................................5

4.2 SPACECRAFT..............................................................................................................................................6

5. COST AND SCHEDULE........................................................................................ 10

6. SUMMMARY.......................................................................................................... 11

Figures Figure 1. Mars at 0.1 Arcsecond Resolution................................................................................... 2

Figure 2. Optical Layout of WFC3................................................................................................. 4

Figure 3. Science Payload............................................................................................................... 5

Figure 4. WFC3 Mounting Concept ............................................................................................... 6

Figure 5. FGSs Mounted to Optical Bench..................................................................................... 6

Figure 6. WFC3 Re-host Modular Design...................................................................................... 7

Figure 7. Spacecraft Concept Summary ......................................................................................... 9

Figure 8. WFC3 Re-host Mass Summary and Launch Vehicle Configuration ............................ 10

Figure 9. WFC3 Re-host Schedule ............................................................................................... 12

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White Paper Proposal for WFC3 Re-host Mission

1. OVERVIEW

If the next servicing mission to the Hubble Space Telescope (HST) does not occur, the Wide Field Camera 3 (WFC3) could be made available by NASA for use aboard a free-flying mission. We describe the science program and technical implementation for a 1.2-meter space telescope mission to re-host WFC3 in low-Earth orbit (LEO) for a total mission construction cost of ~$300M, including reserves and launch vehicle. Scaling the HST optical design to a 1.2-m aperture yields a pixel scale and field of view for the WFC3 ultraviolet-visible camera (UVIS channel) of 0.08”/pixel and 5.4’×5.4’, respectively, and for the near-infrared camera (IR channel) of ~0.26”/pixel and 4.2’×4.6’. Because of the reduced aperture compared to Hubble, not all of the original WFC3 science programs are optimized. However, in this cost-driven approach, the mission will still address several of NASA’s highest scientific priorities. The science program will utilize the unique capabilities of WFC3 – particularly its broad wavelength coverage and large filter set – to study the origin of planetary systems, temporal behavior of Solar System objects (planets, moons, and trans-Neptunian objects), and characteristics of extrasolar planets via transits and microlensing. With its larger field of view and higher detector quantum efficiencies at most wavelengths, our WFC3 Re-host mission will be just as efficient as its predecessor instrument, WFPC2, has been on HST for conducting large survey programs.

The mission development time is ~50 months; launch could occur as early as 2010 if Phase A commenced in FY06. We conclude that WFC3 could be re-hosted on a ~1.2m telescope within a Discovery mission budget, but could not meet the required mission performance within a ~$250M (MIDEX) cost envelope that includes >25% reserves and launch vehicle, primarily due to WFC3’s large mass and the spacecraft’s exquisite pointing requirements.

2. SCIENCE PROGRAM

WFC3 would advance HST’s discovery capability through its combination of large wavelength coverage, field of view, and high sensitivity. The primary science programs for which it was designed include characterizing very distant galaxies; observing distant Type Ia supernovae to probe the nature of the mysterious dark energy that pervades the Universe; studying star formation in the Milky Way, nearby galaxies, and distant primordial objects; determining the baryonic mass by detecting brown dwarf stars at the limit of hydrogen-burning and extrasolar Jupiter-like planets; and studying planets, moons, and other debris in the Solar System. A document by M. Stiavelli, and R. W. O'Connell (editors), titled: “Hubble Space Telescope Wide Field Camera 3, Capabilities and Scientific Program,” outlines the original scientific programs and capabilities of WFC3. It can be found at http://wfc3.gsfc.nasa.gov/ white_paper.html.

With a 1.2-m aperture, but using the same f/ratio, the WFC3 imaging modes would have ~50% coarser pixel sampling than on HST (~0.078 arcsec/pixel in the UVIS channel, and ~0.26 arcsec/pixel in the IR channel). Even with this aperture reduction, the pixel size still enables imaging capabilities that far exceed ground-based instruments over its wavelength range in terms

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of its spatial resolution, field of view, and sensitivity. The diffraction limit λ/D is ~0.1 arcsec for a 1.2-m telescope at visible wavelengths. We know from WFPC2 data that this spatial resolution samples important physical scales (Figure 1) and cannot be matched from the ground over a ~5 arcmin field. WFC3 also observes at UV wavelengths that are not accessible from the ground. The WFC3 Re-host mission would have ~30-50 times higher spatial resolution than GALEX and the Spitzer Space Telescope, following up pioneering observations from these missions with panchromatic imaging at wavelengths from the near-UV to the near-IR. Its field of view is ~3 times larger than HST-WFPC2. When combined with its higher detector quantum efficiencies at most wavelengths, our WFC3 Re-host mission will survey just as efficiently as HST-WFPC2.

The principal science themes of WFC3 on the HST include:

1. Galaxy assembly in the high-redshift (4 < z < 10) Universe;

2. Nature of the mysterious dark energy that accelerates the expansion rate of the Universe;

3. Global processes of star formation in nearby galaxies;

4. Old stellar populations and very low-mass stars;

5. Environments of star formation in the Milky Way; and

6. Properties of planets, moon, and relic remnants of the early Solar System.

Our proposed WFC3 Re-host mission would not have the collecting area available to retain a large impact on the first two themes, though these themes are central to the science programs of the James Webb Space Telescope (JWST) and the Joint Dark Energy Mission (JDEM). The re-host mission could, however, uniquely and profoundly impact themes 3 – 6. Adding studies of extrasolar planets via transits and microlensing to the list and aligning the WFC3 Re-host mission with key themes in NASA’s Vision for Space Exploration, our reconstituted science program for the WFC3 Re-host mission includes:

1. Understand how planetary and stellar systems form and evolve;

2. Characterize the properties and distribution of extrasolar planets;

3. Learn how the Sun’s family of planets and minor bodies originated;

4. Understand the initial stages of planet and satellite formation; and

5. Study the processes that determined the original characteristics of the bodies in the Solar System.

Figure 1. Mars at 0.1” resolution

• With its larger field of view and higher detector quantum efficiencies at most wavelengths, our WFC3 Re-host mission is just as efficient as HST-WFPC2 for conducting large surveys.

• The 0.1” resolution offered by our mission provides Hubble-class access to nearby galaxies; stellar and planetary system formation in the Milky Way; and the planets, moons, comets, asteroids and distant populations in the Solar System

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3. OPTICAL BASELINE AND TRADE SPACE

There are several choices for optical configurations when re-hosting WFC3 on alternate telescopes and spacecraft. This section summarizes the ground rules, choices, and trades that led to our baseline optical architecture. This architecture becomes the foundation for the proposed mission discussed herein and aids in demonstrating the technical feasibility and cost credibility of the WFC3 Re-host program.

The ground rules for the program include the following:

1. Re-host WFC3 on a new 1.2-meter telescope.

- Minimum aperture diameter that delivers ground-breaking science not possible from the ground or any other planned mission.

- The low payload mass of a 1.2-m system allows for a low-risk, cost-effective mission.

2. Preserve the f/31 final f/number for WFC3.

- Minimize modifications to WFC3 as compared to the HST configuration.

3. Maintain the highest possible throughput consistent with the wide wavelength response of the instrument.

- Allows faint objects to be detected with the combined telescope and instrument, especially important at UV wavelengths.

4. Use a classical Ritchey-Chretien (R-C) optical telescope assembly (OTA) design.

- Improves fine guidance sensor performance and dramatically simplifies OTA integration and test.

The following table summarizes the payload configuration that results from the ground rules.

Ground Rule Effect on Baseline Design

(1) Re-host WFC3 on 1.2 m telescope Lengthen the OTA to compensate for increase in field curvature. (Easiest solution added to baseline.)

(2) Preserve the f/31 final f/number of WFC3 Telescope is f/24 Ritchey-Chretien type

(3) Maintain highest possible throughput Remove WFC3 fold mirror and mount instrument axially.

(4) Classical R-C telescope design Replace pupil mirrors in WFC3. (Current mirrors correct for HST spherical aberration.)

In order to simplify the fine guidance sensor design and maximize performance, the 1.2-m telescope used for re-hosting the WFC3 will not emulate the HST spherical aberration. This approach requires us to exchange at least one optic in the WFC3. However, it allows us to reduce risk in the optical alignment tolerances, and maximize the instrument performance.

We now discuss the WFC3 optical configuration and planned modifications, focusing on the UVIS channel of WFC3. (The discussions that follow also apply to the IR channel, though the

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details such as image scale and total field coverage differ.) The WFC3 was designed to optimally match the optical characteristics of the Hubble Space Telescope. Once mounted in the HST Optical Telescope Assembly (OTA), it would provide a field of view (FOV) in the UVIS channel of 160 arcseconds per side, with an image scale of 0.039 arcseconds per pixel.

The basic design involves an on-axis pick-off mirror that directs the incident light from the OTA onto a relay mirror residing near the rear of the instrument structure. This relay mirror creates an image of the OTA exit pupil near the front of the instrument structure. The relay mirror is tilted in order to offset the pupil image from the incoming light. A second mirror resides at the pupil image, as shown in Figure 2. This aspheric mirror is designed to compensate for the prescription of the HST OTA Primary Mirror (PM). The curvature and separation of the two mirrors in the instrument, combined with the f/24 incident beams, result in an f/31 optical speed at the camera focal plane. The relay mirror and the mirror containing the compensating asphericity will be readily changed-out.

These modifications to the WFC3 are low risk. Three other radial bay instruments, the HST Fine Guidance Sensors (FGS), were successfully refurbished with mechanisms, optics, restored latches, and electronics. The alignment tolerances for these refurbished FGSs were much more stringent than required for WFC3. All three FGS restoration programs met performance, cost, and schedule requirements.

4. RE-HOST HARDWARE

The major components of the system are the science payload and the spacecraft. The science payload includes the R-C telescope, two Fine Guidance Sensors (FGS) and the modified WFC3. The spacecraft portion includes the following systems: Attitude Determination and Control System, electrical power and management, command and data handling, telecommunications, and structures. The launch vehicle is a Delta II 7420-10, 3-m fairing.

Obtained from: /wfc3.gsfc.nasa.gov/technology/schematics.php#

Figure 2. Optical layout of WFC3.

Compensating Aspheres

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4.1 SCIENCE PAYLOAD • The OTA Light Shield Assembly: The assembly contains the sun shield, aperture door, the main baffle and forward shell. The main baffle and the forward shell are mounted to the optical bench. The assembly assists in providing stray light suppression and contamination control. The main baffle prevents direct illumination of the primary mirror from objects greater than 26 degrees from the telescope boresight. The aperture door is mechanized for the required cycles of the mission.

Primary Mirror Assembly (PMA): The PMA consists of the 1.2-m Primary Mirror and its mounts to the Optical Bench. The Primary Mirror mounts to the Optical Bench with a three bipod-flexure system that eliminates stress. The flexures limit the parasitic moments that would otherwise degrade mirror figure. The bipod’s design is optimized to minimize gravity vector distortions and tailor the natural frequency of the mirror on its mount.

Metering Structure Assembly (MSA): This lightweight, stiff, dimensionally stable metering truss is a graphite composite truss/ring structure that supports the secondary mirror subassembly. The truss will maintain the required dimensional control between the PM and secondary mirror. We expect that the truss will be thermally passive. The lightweight, durable, and predictable properties of the graphite cyanate-ester composite structure makes it the material of choice. We have extensive space experience with graphite cyanate-ester composite structures combined with bonded metal fittings for interfacing to mirrors and system mounts. This material closely matches the CTE of mirror substrates and provides stable, rigid performance. It is a standard for stable space structures

The Secondary Mirror Assembly (SMA): The secondary mirror subassembly contains the secondary mirror and the actuators that align the secondary mirror to the primary both in-orbit and on the ground. A hexapod mechanism provides the required 6 degrees of freedom similar to those flown on other missions, including HST.

The Re-host Mission Optical Bench: The optical bench (or main support system) is the load bearing structure for the system and includes the support truss for the Wide Field Camera 3. Mounted to the bench are the PMA, the Light Shield Assembly, the Central Baffle, the Fine Guidance Systems, and WFC3 (Figure 4). The characteristics of this structure as well as the WFC3 supporting truss provide the strength and stiffness required for surviving launch loads and for on-orbit stability.

• Instrument Section The instrument section includes the modified WFC3 and the two FGSs.

Wide Field Camera 3: The WFC3 instrument hosts two channels, a UVIS channel with two 2k×4k pixel E2V Technologies CCDs and an IR channel with a 1k×1k Mercury-Cadmium-

Figure 3. Science payload.

Optical BenchPrimary Mirror

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Telluride detector built by Rockwell Scientific. Together the two channels span 200 to 1700 nanometers of the electromagnetic spectrum. In addition to the detectors, each channel contains a filter wheel mechanism for imaging at specific wavelengths.

Optical element modifications to WFC3 will be accomplished per the description provided in Section 3 and do not include alterations to the detectors or filter wheels. In addition, the radiator will be tailored to the new spacecraft configuration while maintaining its functionality.

The WFC3 main attachment points to the optical bench are displayed in Figure 4. The current WFC3 is fitted with a set of latches that were designed for the HST and for storage inside its instrument carrier in the Shuttle cargo bay. Our baseline is to re-use part or all of the mounting latches for the WFC3 as part of the support structure. The latches were designed to withstand Shuttle launch and landing loads. If modifying the latch system should be required, we have experience re-using latches for HST Servicing Missions and with mounting fixtures for a variety of space systems. Fine Guidance Sensors: FGSs are required to provide pointing error signals from the science payload to the spacecraft Attitude Determination and Control System (ADCS) to an accuracy of a few milliarseconds. Pick-off mirrors, mounted on the optical bench with a flexure system, divert the beam into the guiders. The enclosures for fine guidance assemblies will be mounted on a removable graphite composite sub-frame. We use a technique similar to our robust Star Tracker center of mass (COM) algorithms to compute the target centroid and error signal. We plan two FGSs for the Re-host Mission for redundancy and to increase guide star availability.

The units will likely be CCD or CMOS-based systems. The FGS detector plane is sized so that there is a high probability of finding candidate stars in a guidance unit when viewing typical star forming regions. We will also implement moving source tracking for observing Solar System targets.

4.2 SPACECRAFT In the spacecraft segment we concentrate on the integrated observatory that results from the optical baseline discussed in Section 3 and environmental aspects of the mission.

• Observatory Configuration In this section we summarize the spacecraft modular design as well as environmental and other technical considerations.

Configuration: WFC3 Re-host, depicted in Figure 6, is a modular design that integrates the WFC3 instrument along the primary telescope into an instrument/optic module. The module is

Figure 4. WFC3 mounting concept.

Figure 5. FGSs mounted to optical bench.

Pick-off Mirror

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assembled, tested, and calibrated independently, as is the light shield and aperture door assembly. The instrument/optic module and light shield are then integrated to the spacecraft bus to form the complete observatory.

Figure 6. The WFC3 Re-host mission features a modular design.

Communications, Orbit, and Launch Vehicle: The WFC3 Re-host observatory mission concept utilizes the same ground communications architecture as HST. Although several circular and elliptic orbits for WFC3 Re-host are possible while still utilizing the Tracking and Data Relay Satellite System (TDRSS), the current HST orbit (~600 km circular orbit at 28.5° inclination) provides a reasonable baseline for calculating launch performance and preliminary assessments of pointing performance. The observatory is injected into its operational orbit by a Delta II 7420-10 launch vehicle.

Pointing Performance: The WFC3 Re-host pointing goal is stringent but relaxed by a factor of two compared to HST. Like HST, WFC3 Re-host requires a fine guidance sensor signal from the instrument optics module FPA to the spacecraft bus ADCS.

Jitter: To maintain the pointing capability, WFC3 re-host spacecraft bus components with moving parts, such as the solar arrays and reaction wheels, are mechanically isolated. Since the solar arrays are stationary during an observation, pointing and jitter concerns are limited to settling time and thermal snap response. WFC3 Re-host uses the same type of dampers used by HST to settle out solar array disturbances. To avoid jitter and rigid body mode wobble, the Reaction Wheels are mechanically isolated from the spacecraft bus and (speed) biased well above the corner frequency of the isolators (similar to HST). The reaction wheels are analog torque commanded, eliminating any torque quantization issues.

Orbit Environments: Disturbance torques on WFC3 Re-host from gravity gradient, magnetic, and aerodynamic torques are well understood. With the configuration and mass properties of the observatory reasonably defined, the gravity gradient torque is predicted to high accuracy and a feed-forward control signal produces a precise countervailing torque from the reaction wheels. A high closed loop gain at low frequencies removes residual momentum. Magnetic torques used to unload momentum from the wheels are produced from oversized torque rods and fired in a linear fashion. A feed-forward torque is used to remove residual momentum. A feed-forward to

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the reaction wheels, based on the last direction the torque rods were fired, is used to cancel out residual magnetic moments that may cause torques that could disturb the observatory.

Contamination Control: Control of both molecular and particulate contamination throughout the manufacture and integration of WFC3 Re-host components and assemblies is critical to on-orbit performance. This is especially true since the telescope and instrument must perform at ultraviolet wavelengths and the instrument/optics module is enclosed by the spacecraft bus structure. Molecular and particulate contamination analyses are performed to assess the contamination effects on critical surfaces, to identify requirement incompatibilities, and to develop an effective contamination control program. Total source outgassing for both the instrument and spacecraft bus is analyzed using Molecular Flux (MOLEFLUX) modeling. Contamination budgets are derived from science throughput requirements, observation wavelengths, analysis of observatory geometry, and temperatures of sensitive surfaces.

Thermal Environment: While WFC3 Re-host instrument/optics module is contained inside the spacecraft bus, it is thermally decoupled from the spacecraft and provides for its own thermal control. The instrument/optics module is attached to the spacecraft bus using a low conductivity flexure mounting scheme. The flexures isolate the instrument/optics module from spacecraft bus mounting and thermal distortions. Other electronics boxes (external to the instrument modules) necessary to support WFC3 Re-host instruments are mounted directly to the spacecraft structure, which provides for their thermal dissipation.

• Spacecraft Bus Description A pictorial summary chart of the spacecraft is provided in Figure 7. The WFC3 Re-host spacecraft bus is a redundant zero-momentum biased, three axis stabilized spacecraft adapted from the GSFC Rapid Spacecraft Development Office (RSDO) catalog. The ADCS uses finely balanced reaction wheels arranged in a configuration providing the required momentum envelope. Excess momentum is off-loaded with linearly actuated oversized magnetic torque rods to minimize the magnitude and time duration of jitter disturbances. Attitude knowledge is provided by three Star Trackers with the fine guidance sensors providing updates for fine pointing during observations. The spacecraft has two single axis articulated solar arrays providing power. Energy storage is provided by an internally redundant (spare cells) Nickel-Hydrogen battery. The spacecraft Command and Data Handling (C&DH) subsystem includes a solid-state recorder for processing of science and spacecraft data. The C&DH is a block redundant compact Peripheral Component Interface (cPCI) system with a RAD 750 processor. Our baseline telecommunications subsystem, originally developed for the Gamma Ray Large Area Space Telescope (GLAST) program, uses redundant multi-mode S-band transceivers (MMTs) for uplink and downlink of commands, telemetry and (as a back-up) instrument data. The MMT is capable of a real time link through TDRSS S-band Single Access (SA) and Multiple Access (MA) downlink for safe mode alerts as well as the standard ground network communications. While HST uses a 1 megabit per second S-band science data downlink through TDRSS SA Antenna (SAA) services, WFC3 uses redundant X-band transmitters operating at 1.3 Mbps downlink for Stored Science data and spacecraft state of health data.

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Figure 7. Spacecraft concept summary.

Our WFC3 S/C Concept Leverages Our DS1, OrbView 5 and Coriolis Experience with Optical Payload Missions.

Mission• Support Payload Requirements &

Provide Instrument Data to Ground• 3-Year Lifetime w/ 5-Year Goal

(>0.9 Bus Reliability at 3 years)• 600 km circular orbit at 28.5

degree inclinationAttitude Control• Zero Momentum 3-Axis Control • Control with Wheels & Magnetic

Momentum Management • Attitude Determination Using Star

Tracker & Precision Gyro• Requires Fine Guidance Sensor

input to complete final pointing• Autogenerated Flight Software• Vibration isolation to reduce jitter

Launch Vehicle• Delta II 7420-10• 3 m Fairing w/ 6915 PAF• 2800 kg Capability to 600 km orbit

Structure & Thermal• Aluminum Honeycomb & Frame• Cold-biased Thermal with MLI,

Heaters, & Passive RadiatorsCommand & Data Handling• cPCI Architecture• R750 CPU Operating at 80 MIPs• Standard Interfaces (1553, RS-422)• Custom High Speed Interface @

~ 18 Mbps (LVDS)• C++ Flight Software• Solid State Recorder – SEAKR (Up

to 512 Gb Storage Capability)Telecommunications• Science & Bus Data Downlink at 1.3

Mbps Using X-Band• Multiple Omni Antennae• 2 kbps Command Uplink in nominal

pointing (S-band)Propulsion• No Propulsion System On-board

Electrical Power• Deployed, Steerable 11 m2

GaAs MJ Solar Array• 2,340 W EOL Array Output• One 78 Ahr NiH2 Battery • 28 V Unregulated Power Bus

Our WFC3 S/C Concept Leverages Our DS1, OrbView 5 and Coriolis Experience with Optical Payload Missions.

Mission• Support Payload Requirements &

Provide Instrument Data to Ground• 3-Year Lifetime w/ 5-Year Goal

(>0.9 Bus Reliability at 3 years)• 600 km circular orbit at 28.5

degree inclinationAttitude Control• Zero Momentum 3-Axis Control • Control with Wheels & Magnetic

Momentum Management • Attitude Determination Using Star

Tracker & Precision Gyro• Requires Fine Guidance Sensor

input to complete final pointing• Autogenerated Flight Software• Vibration isolation to reduce jitter

Launch Vehicle• Delta II 7420-10• 3 m Fairing w/ 6915 PAF• 2800 kg Capability to 600 km orbit

Structure & Thermal• Aluminum Honeycomb & Frame• Cold-biased Thermal with MLI,

Heaters, & Passive RadiatorsCommand & Data Handling• cPCI Architecture• R750 CPU Operating at 80 MIPs• Standard Interfaces (1553, RS-422)• Custom High Speed Interface @

~ 18 Mbps (LVDS)• C++ Flight Software• Solid State Recorder – SEAKR (Up

to 512 Gb Storage Capability)Telecommunications• Science & Bus Data Downlink at 1.3

Mbps Using X-Band• Multiple Omni Antennae• 2 kbps Command Uplink in nominal

pointing (S-band)Propulsion• No Propulsion System On-board

Electrical Power• Deployed, Steerable 11 m2

GaAs MJ Solar Array• 2,340 W EOL Array Output• One 78 Ahr NiH2 Battery • 28 V Unregulated Power Bus

Spectrum Astro

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The Spacecraft has adequate margin on a Delta II 7420-10 Two Stage Launch Vehicle With Standard Length Fairing.

Adequate Clearance with Static Envelope assures

dynamic clearance

Delta II 3m Diameter Fairing in 2-Stage Configuration

6915 Payload Attach Fitting

The Spacecraft has adequate margin on a Delta II 7420-10 Two Stage Launch Vehicle With Standard Length Fairing.

Adequate Clearance with Static Envelope assures

dynamic clearance

Delta II 3m Diameter Fairing in 2-Stage Configuration

6915 Payload Attach Fitting

• Margins and Launch Vehicle Summary The spacecraft mass and power were computed for the WFC3 Re-host mission. The chart displayed in Figure 8 demonstrates that we have ample mass margin for a Delta II 7420-10 launch vehicle to achieve an orbital altitude of 600 km. In addition, bulk dimensions of the spacecraft have adequate margin for the proposed (two stage) launch vehicle as displayed in Figure 8.

Our WFC3 S/C Concept Provides an Adequate Mass Margin to Confidently Proceed Into Next Design Phase.

MassCruise Power

Pass Power

Ops Power

SUBSYSTEM kg % (W) (W) (W)Structure 224.5 13% 0.0 0.0 0.0

Attitude Control 66.6 4% 77.0 90.4 90.4Electrical Power 148.6 9% 100.8 100.8 100.8

Propulsion 0.0 0% 0.0 0.0 0.0Command & Data Handling 59.6 4% 54.6 54.6 54.6

Thermal 19.6 1% 29.2 6.8 29.2TT&C 34.3 2% 118.8 111.8 66.3

Cabling 80.9 5% 75.0 75.0 75.0Uncertainty 55.0 3%Spacecraft 689.1 41% 455.4 439.4 416.3

Payload 791.8 47% 446.0 446.0 446.0Payload Uncertainty 198.9 12%

Dry Satellite 1679.8 100% 901.4 885.4 862.3Propellent 0.0 0%

Total 1679.8 100%Throw Weight 2800

Margin 1120.2 67%

Our WFC3 S/C Concept Provides an Adequate Mass Margin to Confidently Proceed Into Next Design Phase.

MassCruise Power

Pass Power

Ops Power

SUBSYSTEM kg % (W) (W) (W)Structure 224.5 13% 0.0 0.0 0.0

Attitude Control 66.6 4% 77.0 90.4 90.4Electrical Power 148.6 9% 100.8 100.8 100.8

Propulsion 0.0 0% 0.0 0.0 0.0Command & Data Handling 59.6 4% 54.6 54.6 54.6

Thermal 19.6 1% 29.2 6.8 29.2TT&C 34.3 2% 118.8 111.8 66.3

Cabling 80.9 5% 75.0 75.0 75.0Uncertainty 55.0 3%Spacecraft 689.1 41% 455.4 439.4 416.3

Payload 791.8 47% 446.0 446.0 446.0Payload Uncertainty 198.9 12%

Dry Satellite 1679.8 100% 901.4 885.4 862.3Propellent 0.0 0%

Total 1679.8 100%Throw Weight 2800

Margin 1120.2 67%

Figure 8. WFC3 Re-host mass summary and Launch Vehicle configuration.

5. COST AND SCHEDULE

The WFC3 Re-host program schedule is attached in Figure 9 and demonstrates that the time required to design, manufacture, integrate, and test the observatory falls within the Discovery Program estimate for program start to launch date. The time spans are based on historical performance on similar efforts. The schedule includes approximately 6-months of margin on the critical path.

The real-year (RY) ROM cost estimate presented herein assumes a cost reimbursable contract and includes the science payload (the OTA, modified WFC3 and the FGS), the Spacecraft Bus Assembly, observatory integration and test, and a 30-day on-orbit checkout. These ROMS are based on actual performance costs and estimates performed recently for similar hardware.

Item Cost (RY$M) Phase A 1 Phase B 15 Phase C/D 164 30% Reserves 54 LV 84 Total 318

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The phases in the previous table are defined as follows: Phase A is the concept study; Phase B includes detailed design work and ends with the Preliminary Design Review (PDR); Phase C/D covers construction from the completion of the PDR through launch and mission check-out.

6. SUMMMARY

The WFC3 Re-host mission offers a practical and cost-effective option to address several of NASA’s highest scientific priorities should the next servicing mission to HST not occur. Its larger field of view and high detector quantum efficiencies ensure a frontier science program at modest cost. This mission leverages prior NASA investments in WFC3 design and construction, and will build on HST’s rich legacy of scientific discovery.

Our team approach offers several benefits, including:

• Low-risk use of existing HST legacy hardware and software capabilities. - WFC3 - HST latches - Ground system scheduling, commands, and command groups

• Synergy with other spacecraft and payload designs, hardware, and capabilities. - 1.2-meter OTA system (scaled from 1.0-meter hardware produced) - Similar spacecraft experience with DS1, OrbView5, and Coriolis - Experience with HST latch design, manufacture, and retrofitting - Refurbished three other HST radial bay units (FGSs)

• Conservative low-cost mission. - Dedicated instrument (not time sharing with additional instruments) - Adequate mass margin for LEO and a two stage Delta II rocket - Ample cost margin - Executable schedule with margin

P-4555Rev AUse or disclosure of this data is subject to the notice or restrictions on the title page of this document.

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Figure 9. WFC3 Re-host Schedule

ID Task Name1 WFC-3 Re- Host Mission2 WFC-3 Mods3 Phase A study 4 Phase B Kick off5 SRR6 Preliminary Design 7 PDR8 STE , GFE prep for WFC-3 Delivery9 Detailed Design 10 CDR11 Delivery of WFC-3 for De-integration12 De-integration of WFC313 Re-integration and Periodic Testing14 Environmental and Verification Tests15 Prep and Delivery to S/C Integ Site16 Margin17 Fine Guidance Sensors18 Phase A study 19 Phase B Kick off20 SRR21 Preliminary Design 22 PDR23 Long Lead Item Procurement24 Detailed Design 25 CDR26 Mech/Elec Parts Procurement and Fab27 Board-level chekout28 System Assembly and Test29 Environmental and Verification Tests30 Prep for Payload Integration31 Margin32 OTA and Optical Bench33 Phase A study 34 Phase B Kick off35 SRR36 Preliminary Design 37 PDR38 Detailed Design39 CDR40 Structures and Parts Procurement41 Small Optics Procurement42 Mech/Elec Parts Procure and Fab43 Subassemblies44 PM Blank Procurement45 Primary Mirror Fabrication 46 Telescope Integration and Alignment47 Test48 Payload I&T49 Payload Integration50 Payload Verification51 Prep and Delivery to S/C Integ Site52 Margin53 General Dynamics Spacecraft54 Phase A Kick-off55 Phase A56 Down select57 Phase B Kick off58 SRR59 Software Development60 Preliminary Design61 PDR62 Prepare For Conf Review63 Confirmation Review64 Detailed Design65 CDR66 Bus Manufacture I&T95 Margin96 Payload Delivery97 Payload Delivery98 Observatory Integration99 Observatory Test100 PSR101 Margin102 Pack and Ship103 Launch Site Activities104 LRR105 Margin106 Launch107 LEOP and IOT

8/211/16

12/153/233/26

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11/261/7

3/19

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1/164/174/18

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8/2010/1

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