# Presentation: 848264

## SuperNova / Acceleration Probe (SNAP)

**June 28, 2001**

**SuperNova / Acceleration Probe (SNAP)**

**System Overview**

## Driving Requirements and Assumptions
Options
Selected Configuration and Rationale
Technologies Required
Mass and Power Summary
Requirements Verification
Additional Trades
Risk Assessment
Issues and Concerns

**Options**
**Selected Configuration and Rationale**
**Technologies Required**
**Mass and Power Summary**
**Requirements Verification**
**Additional Trades**
**Risk Assessment**
**Issues and Concerns**

**Outline**

## Overview

**Mission objective:**
- Determine the magnitude-redshift relationship for supernovae of type 1A over redshift range 0.3<Z<1.8
- Determine the distribution of gravitational potentials along cosmologically significant lines of sight
- Determine the magnitude-redshift relationship for other supernova types
**Additional constraints, challenges, and measurements**
- Orbit w/ adequate thermal environment, good observing efficiency
- Pointing stability and tracking / Observatory stiffness
**Primary purpose of this study:**
- Establish/validate baseline mission configuration
**Length of study:  **
- 4 days

## Driving Requirements & Assumptions

**Orbit:  **
- Best candidates to date are 19 Re x 57 Re w/ lunar assist or 38 Re circular achieved w/ lunar assist
**Launch year: 2008 **
**Lifetime: 2 year required, 5 years goal  **
**Quality level: Selective redundancy**
**End-of-life disposal: Not required for orbits > GEO**
**Instrument support:**
- Mass: 700 kg
- Power:  135 W max, 85 W avg, 22W stdby
- Average Instrument Data Rate: 40 Mbps continuous

## Orbit Options

**LEO orbits:**
- PROs: Simplify launch vehicle and propulsion requirements, easy RF comm
- CONs: Enormous loss of observing efficiency, huge thermal problems for Payload, Bus, and Radiators
- Disposal required
**HEO orbits:**
- PROs: Easy observing, good thermal environment
- CON: Radiation environment problems
**SHEO orbits w/ lunar assist:**
- PROs: Easy observing, good thermal environment
- CONs: Requires “exotic” launch vehicle, difficult downlink and ground station situation that limits inclinations, orbits. Difficult orbit calculations.
**Sun-synchronous drift-away orbits:**
- PROs: Perfect observing w/o eclipses, best thermal environment
- CONs: RF communication difficult, especially at EOL.  Possible mass constraint.
- Final orbit selection after detailed analyses. For some orbits, minor errors can easily propagate into fatal dispersion

## Launch Vehicle Options

** ****Delta 2920H-10L**
- Liftoff capability marginal but fairing is way too small
- GSFC “actual” cost estimate is $67M to $72M
- Maiden flight w/ SIRTF in 2002
** ****Delta 3940-11**
- Liftoff capability adequate
- GSFC “actual” cost estimate is $80M to $84M
- Liftoff capability adequate, fairing volume tight but adequate
** ****Atlas IIIB**
- Liftoff capability comfortable, fairing volume tight but adequate
**Zenit 3SL w/ Block DM-SL restartable upper stage (Sea Launch)**
- Liftoff capability comfortable, fairing volume comfortable
- For details see “SNAP_Launch_Vehicles.xls”

## RSDO Bus Options

**Ball Aerospace BCP 2000, Bus dry mass = 608 kg****	****	**
- Payload Power (OAV) (EOL) / Mass Limit: 730 W / 380 kg
**Spectrum Astro - SA 200HP, Bus dry mass = 354 kg**
- Payload Power (OAV) (EOL) / Mass Limit: 650 W / 666 kg
**Orbital StarBus, Bus dry mass = 566 kg ****	**
- Payload Power (OAV) (EOL) / Mass Limit: 550 W / 200 kg
**Lockheed Martin - LM 900, Bus dry mass = 492 kg**
- Payload Power (OAV) (EOL) / Mass Limit: 344 W / 470 kg
**Orbital - Midstar, Bus dry mass = 580 kg**
- Payload Power (OAV) (EOL) / Mass Limit: 327 W / 780 kg
- For details see “SNAP_ Candidate_RSDO_Busses.xls”
- All above busses are designed for multi-year missions w/ redundant components.
- Mission Unique spacecraft structure and several significant subsystem upgrades are required.

## Earth Shield for Detector Radiator
Use in LEO to eliminate thermal coupling between Earth and Radiator
May have to be actuated similar to a solar array drive
Natural frequency must be > 1Hz (preferably >> 1Hz)
Rigid cylinder shape could be considered
PROs:
Would allow passive cooling of Radiator to 130K even at LEO
CONs:
Extra mechanism, increased risk
Complicates I&T
DISPOSITION:
Option dismissed, as LEO option was dropped due to several other problems

**Earth Shield for Detector Radiator**
- Use in LEO to eliminate thermal coupling between Earth and Radiator
- May have to be actuated similar to a solar array drive
- Natural frequency must be > 1Hz (preferably >> 1Hz)
- Rigid cylinder shape could be considered
- PROs:
- Would allow passive cooling of Radiator to 130K even at LEO
- CONs:
- Extra mechanism, increased risk
- Complicates I&T
- DISPOSITION:
- Option dismissed, as LEO option was dropped due to several other problems

## Other Options Considered

**Slew rate vs. jitter**
- Select small reaction wheels for low jitter or “Heavy Duty” (higher jitter) reaction wheels for fast slew
- DISPOSITION:
- Analysis shows that the actual “driver” is solar wind, requiring bigger wheels
- 30 Nms/.05Nm wheels selected
- Use isolation mount

## Baseline Configuration & Rationale

**3-axis stabilized, 4 Reaction wheels, IRUs, no torquer bar**
**Sun side w/ rigid body mounted solar arrays & anti-sun side w/ radiators **
**Standard Hydrazine propulsion system, 1 lbs.thrusters, ~150 kg total propellant required @ 100 m/s for apogee lowering, corrections and ACS.**
**74 Gbits SSR, storage only for spectroscopy data. (Avg. data rate ~52 Mbps; lossless compression plus overhead). **
**Continuous Ka band downlink @ 55 Mbps to 3 Northern Latitude ground stations (Berkeley, France, Japan). **
**3 gimbaled .7m Ka band HGAs; S-band omnis; S-Band T&C thru omni antennas**

##  Preliminary Bus Subsystems Mass [kg]

**	**
**Bus Structure****	****	****	****	****143.0****	**
**Payload Mount****	****	****	****	****9.0****	**
**Antenna support****	****	****	****30.0****	**
**ACS****	****	****	****	****	****50.0****	**
**C&DH****	****	****	****	****	****12.0****	**
**Power Electronics****	****	****	****15.0****	**
**Battery****	****	****	****	****81.2****	**
**Solar Arrays****	****	****	****	****10.5****	**
**Thermal Hardware****	****	****	****67.0****	**
**RF Communications****	****	****	****53.0****	**
**Bus Harness****	****	****	****	****8.0****	**
**Separation System, spacecraft side****	****8.0**
**	**
_**Bus Subsystems Total**__**	**__**	**__**	**__**486.7**__**	**_
_**	**_

##  Preliminary SNAP Obesrvatory Mass [kg]

**	**
**Payload Total****	****	****	****	****700**
**Bus Subsystems Total****	****	****	****490****	**
**Propulsion Total (estimate)****	****	****170**
_**SNAP Observatory**_** **_**Total**__**	**__**	**__**1360**_

##  Preliminary Cost Summary [$M]

**Cost information**
- Details in spreadsheet
- Subsystem cost estimates included
- Other costs are rough estimates

_**SNAP MISSION COST**_** ****	****  **_**	**__**w/ Contingency**_
- SCIENCE
- SPACECRAFT BUS
- MISSION INTEGRATION
- OPERATIONS
- LAUNCH VEHICLE
_**TOTAL (excl. data analysis)**__**	**__**	**__**$307.8M**_**	****	**
**	****	****	**

## Requirements Verification

**Standard functional and environmental verification per GEVS**
- Tight contamination control required
**Main challenges are in Instrument verification **
- Ideally, observatory level thermal vacuum / thermal balance test is combined with and end-to-end image quality verification
- May use double-pass test

## Additional Trades to Consider

**Revisit possibility of using LEO**
- Lower mission cost
- More “realistic” launch / launch vehicle configuration
- May eliminate propulsion
- Simplifies RF Comm
- Must assess disposal issues

## Risk Assessment, Technologies

**Risk on spacecraft bus is generally low, with well-understood technologies and readily available components.**
**Designed subsystem for a 3 year mission, solar array sized to 5 years**
**Higher risk on instrument, especially on the enormous CCD cluster**
**No significant technology development required for bus**

## Mass, Power, and Cost information
“SNAP_Mass&Cost_Summary.xls”
Useful Web sites
Access to Space at http://accesstospace.gsfc.nasa.gov/ provides launch vehicle performance information and other useful design data.
Rapid Spacecraft Development Office at http://rsdo.gsfc.nasa.gov/ provides spacecraft bus studies and procurement services.

- “SNAP_Mass&Cost_Summary.xls”
**Useful Web sites**
- Access to Space at http://accesstospace.gsfc.nasa.gov/ provides launch vehicle performance information and other useful design data.
- Rapid Spacecraft Development Office at http://rsdo.gsfc.nasa.gov/ provides spacecraft bus studies and procurement services.

**Overview, Supporting Data**