F.  MISSION IMPLEMENTATION

F.1.Mission Overview

The HESSI Mission involves a single instrument consisting of an Imaging System, a Spectrometer, and Instrument Electronics, mounted in a simple Sun-pointed, spin-stabilized spacecraft. The instrument and science objectives are summarized in foldout Figure D-11 and the spacecraft in foldout Figure F-1 (end of this section).

HESSI will be launched in mid-2000 on a small-fairing SELVS II vehicle into a 600-km circular, 38° inclination orbit, in a standard configuration with the spacecraft ON and the instrument OFF (detectors warm). Following injection into orbit, the spacecraft will reorient towards the Sun, deploy its solar arrays, and spin up to 15 RPM.

The spacecraft will then transition from course to fine Sun sensor pointing, turn on the SAS and RAS, and position the Sun within the SAS field of view. After confirming SAS functionality, the spacecraft will be directed to use SAS signals, providing a fully-automated closed-loop pointing system referenced to the Imaging Systems’ boresight.

Once the operational attitude, orientation and spin rates have been achieved, the instrument electronics and cryocooler will be powered. Using a preprogrammed thermal profile, the cryocooler will cool the spectrometer to operational temperatures within a few days. This will be followed by a brief detector checkout period in which high voltages are turned on before normal operations begin.

In normal operation the GeDs are kept <75 K, GeD high voltage is on, and observations are taken continuously. Because of the large thermal mass of the GeDs the cryocooler can be cycled over time scales of hours if needed. Energy and arrival time of every photon detected, together with instrument SAS and RAS attitude data, are stored in the spacecraft’s 2 Gbyte mass memory and then telemetered. Ground data systems will convert these data into X-ray and gamma-ray images and spectra.

A LEO-T ground station at UCB is planned for all command and data reception. A Mission/Science Operations Center at UCB will operate the spacecraft and instrument, write the data onto CD-ROMs, and distribute them to the Solar Data Analysis Center (SDAC) at GSFC and the High Energy Data Center (HEDC) in Zurich. The SDAC will archive and distribute both data and analysis software to outside users in the U.S., and provide context observations from other spacecraft and ground instruments. The HEDC will perform the same functions in Europe. A program of ground observations is supported directly by HESSI to provide critical context data.

The Mission Elements are summarized in Table F-1 and discussed below. The spacecraft is described in detail in section F.2 and ground systems in section F.3, and potential risk areas and mitigation in section F.4.

Product Assurance

Product Assurance and Configuration Management will be based on those used successfully on SMEX FAST, SOHO XDL, Wind 3D Plasma, Polar EFI, etc. An overall CM Plan will be developed which incorporates instrument, spacecraft, and ground segment documentation.

Table F-1 Mission Elements

Table F-2. Instrument Requirements on Spacecraft

Launch Date

The required HESSI launch date and useful mission lifetime are determined by the timing of the next solar activity maximum. Predictions based on the last two solar cycles indicate that several thousand hard X-ray flares and of order a hundred gamma-ray flares will be detected in a two- or three-year HESSI mission starting in mid-2000 (Figure D-10). Even a launch as late as the end of 2001 or an unusually early solar maximum would not reduce the predicted number of hard X-ray flares below 1500. Thus, the first of the two SMEX missions being selected through this AO, with a launch in mid-2000, would be ideal for HESSI.

Orbit

We propose to launch HESSI into a simple 600 km circular orbit, chosen for >3-year orbit lifetime, inclined at 38° for communication and operation simplicity. While this orbit is higher in radiation than an equatorial orbit, tests of the GeD detectors show that radiation test damage is not a problem for a 2-3 year mission if the GeD are kept at temperatures below 75 K, easily achieved with HESSI cryocooler and cryostat design.

Instrument Accommodation

The HESSI instrument places modest requirements on the design and operation of the spacecraft (Table F-2) so a simple, spin-stabilized spacecraft with extensive use of currently available, space-qualified components and subsystems is adequate.

Ground Based Program

Ground-based observatories are a unique source of the optical and radio data (Table F-3) that are crucial to the successful interpretation of the HESSI data. We have allocated HESSI funds for the necessary upgrading of hardware at these observatories to provide measurements with adequate temporal resolution, dynamic range, and data handling capabilities for comparison with HESSI’s observations. For two facilities that lack long-term support, we will provide the minimal support during the mission to ensure the operation of these critical and unique instruments.

Table F-3. HESSI US Ground-based Program

F.2. HESSI Spacecraft

Our guiding philosophy in the design of the HESSI spacecraft is to drive toward firm requirements definition, conservative design margins, utilization of standardized and proven electronic interfaces, minimum parts count, proven designs, operational simplicity, and maximum use of existing plans, procedures, and processes. The HESSI spacecraft, shown in Figure F-1, is consistent with this philosophy and accommodates all of the HESSI instrument requirements using space-qualified components and approaches. Table F-4 is a summary of the mass and power characteristics of the spacecraft and illustrates our conservative margin approach; we have designed-in allowances for 15% growth in spacecraft bus mass, 20% in instrument mass, and 20% in power. In addition to these growth allowances, our design incorporates 47% launch vehicle mass contingency.

Table F-4. Mass and Power Characteristics

Structure & Mechanisms

The primary structure is composed of a single aluminum honeycomb panel and a thrust tube to carry loads from the launch vehicle adapter ring. The open structural design permits thermal radiator area for heat rejection and enables access to spacecraft and instrument components throughout all phases of integration and test to reduce cost and schedule. All spacecraft components are attached to the mid-deck panel allowing the spectrometer to have an unobstructed radial field-of-view. The spectrometer and the imager assembly are structurally independent from one another, allowing separate bolt-on installation - the imager assembly is installed from above the mid-deck and the spectrometer is installed from below.

The mechanism subsystem consists of the mechanisms to support the deployment of four identical solar array wings. A single shaped memory alloy (SMA) actuated release device preloads the wing against the cup and cone snubbers in the stowed configuration. Advantages of the SMA include lower cost, lower mass, higher reliability, negligible shock loads, and the capability of being operated through repeated test cycles allowing protoflight qualification of the actual flight units. HESSI contains no pyrotechnic devices, thereby minimizing safety and range interface concerns.

Spacecraft dynamic balance is critically important to the operation of the instrument, and spacecraft components have been located with considerable attention paid to inertia properties. The result is an efficient, symmetrical, well-balanced spacecraft design that is ideally suited for this spin-stabilized configuration. In addition to the design efforts, the spacecraft will be spin-balanced following final system test, and linear mass drivers will be used on-orbit to do the fine adjustments that may be necessary following deployment of the solar array wings.

Electrical Power

The electrical power subsystem consists of the four silicon-cell solar array wings, a single Common Pressure Vessel (CPV) NiH₂ battery and all spacecraft cabling. Battery charging, power conversion and distribution are performed within the integrated electronics module (IEM) of the Command & Data Handling subsystem. Each solar array wing consists of Al facesheet/Al honeycomb substrate with 2 ohm-cm, front surface passivated silicon cells and produces over 81 Watts at end of life for a total spacecraft power of 325 W. The battery stores power in ten 12 amp-hour CPV cells connected in series to produce 28 ±4V. The IEM hosts the power input/output (PIO) board and the charge control board (CCB). The PIO, developed by Spectrum for Lunar Prospector, provides VME-controlled, 28 ±4 V switched power outputs. The CCB, also developed for Lunar Prospector, uses pulse-width modulated FET switches to control the direct flow of array current to the battery. The cabling approach is based upon Spectrum’s MightySat and New Millennium Program (NMP) DS-1 designs.

Command & Data Handling

The Command and Data Handling subsystem is built around the VME-based integrated electronics module (IEM) which contains the 1750A CPU, communications interface board (CIB), payload and attitude control interface (PACI), charge control board (CCB), power input/output (PIO), 2 Gbyte solid state memory, and 5 spare card slots. The 1750A has been developed by Southwest Research Institute (SWRI) and Spectrum Astro in support of MSTI, MightySat and NMP DS-1 and has been previously flown on MSTI-1, MSTI-2, and MSTI-3.The CIB, PACI, CCB, and the PIO were developed by Spectrum in support of Lunar Prospector, Mars-98 Orbiter and Lander, and NMP DS-1. The solid state memory is supplied by SEAKR based on space-qualified hardware produced for NASA’s SSTI program.

Attitude Control System (ACS)

The Attitude Control System consists of four coarse Sun sensors, a fine Sun sensor, three torque rods, a magnetometer, two linear mass drivers, and a passive nutation damper. The spacecraft will separate from the launch vehicle in 3-axis mode, deploy the four body-fixed solar array panels, acquire the Sun using four coarse Sun sensors with 4p steradian coverage, spin-up to 15 RPM and perform Sun acquisition to 0.2° using the fine Sun sensor. This Sun-pointing attitude will be maintained throughout the mission using Sun attitude data from the spacecraft fine Sun sensor and the instrument SAS. Spacecraft attitude will be continuously changed to follow the Sun using internally redundant torque rods in combination with magnetic field data from the 3-axis magnetometer. The passive nutation damper is used to damp nutation which can also be actively controlled using the torque rods. The linear mass drivers allow fine control of the spacecraft moments and cross products of inertia in the final deployed configuration ensuring stable, smooth, spin during the entire mission.

Both the instrument and the spacecraft have been designed to operate autonomously for weeks at a time. Following the initial attitude acquisition, even if the Attitude Control subsystem should fail "off", the spacecraft spin axis will remain fixed in inertial space.

Telecommunications

The Telecommunications subsystem consists of a 5-Watt STDN S-band transponder, RF assembly, and two omni antennas and ensures full downlink capability regardless of spacecraft attitude. The downlink rate to the UC Berkeley ground facility is 3.5 Mbits/second with 2.6 dB link margin, using a ground antenna diameter of 5 meters as shown in Table F-5. The data is BPSK coded using CCSDS recommended r=1/2, k=7 concatenation with RS(233,255). This approach results in an average of 11.2 Gbits of data downlinked per day.

Table F-5. Link Margin

Thermal Control

HESSI uses a proven, cold-biased design with flight proven technologies to provide an inherently reliable thermal control architecture. All thermal control components are standard, off-the-shelf hardware. The instrument is thermally isolated from the spacecraft to ensure that spacecraft thermal properties do not affect science operations.

F.3. Ground Systems

The HESSI mission operations scenario (see Figure F-2) is simple and efficient. Both the spacecraft and the instrument are fully autonomous during normal operations. No commands need be sent to the spacecraft for days or even weeks at a time, other than those required to activate the transmitter and dump data during each ground-station pass. Thus, once all spacecraft and instrument functions have been activated and verified after launch, the mission operations scenario consists merely of reading out science data from the onboard memory each orbit and verifying the health and safety of the mission.

We propose to install a single LEO-T ground station at Berkeley for commanding and data reception. LEO-T is a small autonomous station for tracking low Earth orbiting spacecraft. It includes a 5-meter dish, RF electronics, and a CCSDS-compatible front-end processor. It is identical to the station being installed in Puerto Rico for the FUSE mission. HESSI will make 6 passes per day of about 7 to 10 minute duration over the UCB station. A 2.6 dB link margin from HESSI to LEO-T can be maintained at 3.5 Mbps transmit rate, giving a daily downlink data volume of 11 Gbits (Table F-5). The FUSE ground station can be used as a back-up for HESSI, and the Berkeley ground station can be made available to NASA for other missions.

HESSI will use the Integrated Test and Operations System (ITOS) to satisfy both Integration and Test (I&T) and Mission Operations Center (MOC) requirements (Fig. F-2). ITOS was developed at GSFC as an I&T system for SMEX missions, and was used by the FAST spacecraft for I&T. Later ITOS evolved into a system which satisfies both I&T and MOC needs, and is being used in both capacities by the TRACE mission. ITOS can perform health and safety monitoring, commanding, ground station control, mission planning, and orbit and attitude determination. Using the same system for I&T and MOC greatly reduces costs and schedule risks. Furthermore, UCB personnel are already familiar with ITOS from their FAST experience. The Science Operations Center will be nearly identical to the system developed for FAST at Berkeley. This autonomous system performs instrument monitoring, selected data reduction (with data plots automatically posted to a page on the WWW), and data distribution in the form of CD-ROMs (automatically labeled). The entire system operates with no human intervention. Data will be transferred by CD-ROM to the SDAC data archive at GSFC and the HEDC archive at Zurich from which it can be accessed generally.

Co-locating the ground station, Mission Operations Center, and Science Operations Center at Berkeley eliminates ground communications costs and reduces personnel requirements during operations. The whole system can be operated by two people on a one shift per day basis (plus student help and occasional consulting by subsystem experts). Automated spacecraft and instrument health monitoring at the MOC (an off-site operator is notified in case of a problem) is desirable to minimize possible data loss, but is not required. Mission planning and real-time support is required only during launch and initial orbit acquisition.

F. 4. Potential Risk Areas

The HESSI team has worked diligently to remove schedule, cost, and technical risks from the program and will work with the SMEX office to assess and mitigate potential risk elements. Because essentially all the required development work for the HESSI instrument has already been done, and the spacecraft consists almost entirely of already qualified subsystems, the risks involved for HESSI are minimal, much lower than for the typical SMEX. Below, we discuss specific possible risk areas that have been identified, and possible mitigation.

Schedule

The schedule is tight but we have done all the development work and made the arrangements required to make it clearly achievable. There is 14 to 16 weeks of distrbuted slack for the instrument and spacecraft, respectively, for a July 1, 2000, launch. There is an additional 3 months of contingency to meet the launch requirement of September 2000 for the first SMEX chosen under this AO. Delays will decrease the number of flares detected (Figure D-10) but even a launch at the end of 2001 would still be scientifically acceptable.

Cryocooler

The Sunpower cryocooler is new technology. It has been fully flight-qualified and tested with the GeDs so we are very confident that it will be successful, but it has not flown previously. If significant problems arise, we can go to another flight-qualified cooler (3 others have been tested by UCB and proven compatible with GeDs).

PSI Fabrication of the Imaging System

PSI (Paul Scherrer Institute) has extensive experience with fabrication of space flight hardware and has always delivered successfully. However, we have costed the effort required to duplicate PSI’s tasks (given in the budget as PSI contribution) in the U.S., and are confident that they can be done by UCB or a subcontractor for that cost.

Grid Fabrication

Grids #3-9 are being fabricated by Co-I van Beek. The technology is fully developed, but if problems arise these grids can be replaced by the JPL gold foil grids with loss of imaging at high energies (above ~40 keV). This would be an acceptable science fallback. The gold foil grids can also be fabricated by outside vendors as well as JPL.

GeD fabrication

In the very unlikely event that problems arise at ORTEC, the GeDs can be fabricated at UCB with some delay in the schedule (see above), but no other impact.