# Presentation: 645118

## Space Systems:Spacecraft and Launch Vehicle

**Henry Heetderks**
**Space Sciences Laboratory, UCB**
**July 2002**

## The  Observatory consists of a custom scientific package integrated with a system which performs standard Spacecraft functions
The Science Instrumentation is of necessity a one of a kind development
Since the the instrument is already a development item,  any unusual or non-standard functions which may require R&D will be performed on the instrument side
Spacecraft functions will be performed with off-the -shelf “commodity” components which can be procured from multiple vendors

Additionally, to strengthen the commoditization of the Spacecraft,  interfaces between S/C components will use industry standard protocols chosen for simplicity and interoperability,  and will be maintained by the SNAP Project System Engineering group
This will help to control vendor costs and facilitate the development of workarounds in the event of delivery difficulties

The observatory mechanical layout is designed for ease of assembly to maximize flexibility during Integration and Testing

This will facilitate the purchase of a “standard” S/C bus thru the NASA RSDO

- The Science Instrumentation is of necessity a one of a kind development
- Since the the instrument is already a development item,  any unusual or non-standard functions which may require R&D will be performed on the instrument side
- Spacecraft functions will be performed with off-the -shelf “commodity” components which can be procured from multiple vendors
- Additionally, to strengthen the commoditization of the Spacecraft,  interfaces between S/C components will use industry standard protocols chosen for simplicity and interoperability,  and will be maintained by the SNAP Project System Engineering group
- This will help to control vendor costs and facilitate the development of workarounds in the event of delivery difficulties
- The observatory mechanical layout is designed for ease of assembly to maximize flexibility during Integration and Testing
- This will facilitate the purchase of a “standard” S/C bus thru the NASA RSDO

**Spacecraft Bus Development Philosophy**

## Attitude Control System (ACS)
3-axis Stabilized Vehicle
Need  <.03  Arc-sec,  1 Sigma Stability [requirement]
Accuracy is Achieved by using Data from Focal Plane Guider
Need Independent Fail-Safe for Sun Avoidance
Drives Requirements on Structure Stiffness and Fuel-Slosh

Propulsion System
Dumps Momentum Wheels
Performs Orbit fine adjust and station keeping
Does End-of-Life Disposal

Command and Data Handling
Provide CCSDS 1 KHz S-Band Communications to 25 Re  
Provide CCSDS 300 MHz Ka-Band Telemetry to 60000 Km Altitude
Autonomous Operations for Two Orbits without Ground Intervention

**3-axis Stabilized Vehicle**
**Need  <.03  Arc-sec,  1 Sigma Stability [requirement]**
**Accuracy is Achieved by using Data from Focal Plane Guider**
**Need Independent Fail-Safe for Sun Avoidance**
**Drives Requirements on Structure Stiffness and Fuel-Slosh**
**Propulsion System**
**Dumps Momentum Wheels**
**Performs Orbit fine adjust and station keeping**
**Does End-of-Life Disposal**
**Command and Data Handling**
**Provide CCSDS 1 KHz S-Band Communications to 25 Re  **
**Provide CCSDS 300 MHz Ka-Band Telemetry to 60000 Km Altitude**
**Autonomous Operations for Two Orbits without Ground Intervention**

**Spacecraft Bus Key Requirements**

## Spacecraft Bus Key Requirements

**Structure**
**Provide connection between the Instrument (~1000 Kg) and the Launch Vehicle**
**Provide mounting of non-Optics Bench Components**
**Must Meet Requirement on Minimum Stiffness from ACS**
**Power System****	**
**Provide  275 W max, 130 W avg, 30 W standby to Instrument**
**Design for Mission Lifetime of 4 years**
**Quality level: Fully redundant**
**Provide End-of-life Disposal as Required (NASA NPD 8710.3)**

## Spacecraft Bus Key Requirements

**Electronics Needs Radiation Tolerant Design **
** ****Mission Life of 4 years**
** ****Margin of at least 50%**
**Can be Achieved with Shielding and Inherent Parts Tolerance**
**E.g:  ~100 Krad parts with 3.5mm total shielding**

## 3-axis Stabilized, 4 Reaction Wheels, IRUs, Sun Sensors,  Coarse Star Trackers with Guider in OTA,  no Torquer Bars
Sun side w/ rigid body mounted solar arrays & anti-sun side w/ radiators 
Standard Hydrazine propulsion system, 22 N thrusters, ~130 kg total propellant required:  30 kg for attitude and station keeping, 100 kg for disposal
5.6 Tbits SSR, storage for science data. (Peak data readout rate ~55 Mbps; 300Mbps downlink rate). 
Ka band downlink @ 300 Mbps to Berkeley and Perth 
Gimbaled .5m Ka  band HGA (X3); S-band omnis; S-Band T&C thru omni antennas

**Sun side w/ rigid body mounted solar arrays & anti-sun side w/ radiators **
**Standard Hydrazine propulsion system, 22 N thrusters, ~130 kg total propellant required:  30 kg for attitude and station keeping, 100 kg for disposal**
**5.6 Tbits SSR, storage for science data. (Peak data readout rate ~55 Mbps; 300Mbps downlink rate). **
**Ka band downlink @ 300 Mbps to Berkeley and Perth **
**Gimbaled .5m Ka  band HGA (X3); S-band omnis; S-Band T&C thru omni antennas**

**Spacecraft Bus Baseline Concept**

## Spacecraft Electrical Block Diagram

## SNAP Observatory in Delta IV Fairing

**SNAP Observatory in Delta IV Fairing**

**Observatory Model made by**
**Robin Lafever, LBNL Engineering**

## SNAP Observatory

**SNAP Observatory**

- Main Baffle Assembly

- Door Assembly

- Solar Array, ‘Sun side’

- Secondary Mirror
- Hexapod
- Bonnet

- Secondary Metering structure

- Primary Mirror

- Optical Bench

- Instrument Metering Structure

- Tertiary Mirror

- Fold-Flat Mirror

- Spacecraft

- Shutter

- Instrument Bay

- Instrument Radiator

- Solar Array, ‘Dark side’

- Hi Gain Antenna

- Solid-state recorders

- ACS
- CD & H
- Comm
- Power
- Data

- CCD detectorsNIR detectorsSpectrographFocal Plane guiders
- Cryo/Particle shield

## SNAP Observatory Cut-Away

**SNAP Observatory Cut-Away**

**Stray Light/Baffeling**:
- John Bercovitz
- Susanna Deustua
- Mike Lampton

**Shutter Assembly:**
- Chris Bebek
- John Bercovitz
- Mike Lampton
- Michael Levi
- Robin Lafever

**Spacecraft:**
- Bobby BesunerDave Pankow
- Robin Lafever
- Michael Levi

**Spectrograph:**
- Robin Lafever
- Eric Prieto

**Instrument Electronics****Packaging**
- Natalie Roe
- Henrik Von Der Lippe
- Robin Lafever

**Focal Plane Assembly:**
- Chris Bebek
- John Bercovitz
- Robin Lafever

**Optical /Thermal Mounts:**
- Mike Lampton
- Robin Lafever
- Dave Pankow

**Baffle Structure/ Metering structure:**
- Dave Pankow
- Bobby Besuner

## SNAP Observatory

**Mounting of HGA’s shown in Spacecraft Structure**

**Front Covers Open**

## Snap Observatory

**Snap Observatory**

**Shows Launch Vehicle Attach Ring and Propulsion System Components**

## Observatory Exploded View showing S/C

## Spacecraft Bus Subsystems Status

**Attitude Control System**
**Consists of a stand-alone system with a single interface to the C&DH**
**A candidate set of components were identified by GSFC IMDC:**

**Since the ACS requirements are at the high end of routine practice,  we plan to perform additional development and modeling work during the R&D phase**
**This will be described by M. Lampton in a later section**

## Power System
Solar Array.   We are baselining 2.5 M2 solar array area using Triple Junction Gallium Arsinide (TJGaAs) solar cells to make maximum use of the array area and mass.
Battery.   A number of candidate batteries are available, requires a 40 Ah unit.  
Charge Controller and Interface.  A dedicated stand-alone unit will be used.
Power System Requirements will be refined during the R&D phase.  Since these are commodity items, no additional development will be done.

Telemetry System
The current baseline is to use S-band for commands and low rate engineering, and to use a Ka band transmitter with a steered antenna for the 300 Mhz Science Data
Since Ka band infrastructure is currently in development,  the maturity of this technology will be closely monitored during the R&D phase.
The T/M system is described in greater detail in a following section

**Solar Array.   We are baselining 2.5 M****2**** solar array area using Triple Junction Gallium Arsinide (TJGaAs) solar cells to make maximum use of the array area and mass.**
**Battery.   A number of candidate batteries are available, requires a 40 Ah unit.  **
**Charge Controller and Interface.  A dedicated stand-alone unit will be used.**
**Power System Requirements will be refined during the R&D phase.  Since these are commodity items, no additional development will be done.**
**Telemetry System**
**The current baseline is to use S-band for commands and low rate engineering, and to use a Ka band transmitter with a steered antenna for the 300 Mhz Science Data**
**Since Ka band infrastructure is currently in development,  the maturity of this technology will be closely monitored during the R&D phase.**
**The T/M system is described in greater detail in a following section**

**Spacecraft Bus Subsystems Status**

## Command and Data Handling Unit
The Observatory Architecture is designed to allow this function to be performed by one of a number of generic units
CCSDS used for both up link and down link
SSR is included in the Science Instrument -- simplifies C&DH procurment
Detailed requirements will be developed in the R&D phase to allow this system to be accurately costed by the RSDO

Propulsion System
This system consists of a set of three diaphragm tanks and eight thrusters.  It is described in more detail in a later section
Refined requirements will be developed in the R&D phase
We consider this to be a generic item which does not require development in the R&D phase

**The Observatory Architecture is designed to allow this function to be performed by one of a number of generic units**
**CCSDS used for both up link and down link**
**SSR is included in the Science Instrument -- simplifies C&DH procurment**
**Detailed requirements will be developed in the R&D phase to allow this system to be accurately costed by the RSDO**
**Propulsion System**
**This system consists of a set of three diaphragm tanks and eight thrusters.  It is described in more detail in a later section**
**Refined requirements will be developed in the R&D phase**
**We consider this to be a generic item which does not require development in the R&D phase**

**Spacecraft Bus Subsystems Status**

## Integration and Test Plan Development in R&D

**Spacecraft Systems Level Tests**
**Initial testing of the Spacecraft will be done at the subsystem level**
**Power -- Use spare battery and Solar Array simulator**
**ACS -- Use a combination of actual hardware with simulated sensor inputs**
**Testing will be based on continued use of the ASC model system developed during the R&D phase**
**C&DH and T/M -- Functional test development is straight forward**
**Propulsion -- Testing by the Vendor at the subsystem level only**
**A Spacecraft level Functional Test will be developed from the individual subsystem tests**
**Instrument functions will be added to the Spacecraft Functional Test, to generate an Observatory level Functional Test**
**Stimulation units built into the sensors will enable a complete end-to-end test of the electronics**
**Procedures to verify CCD dark current will be addressed in the R&D phase**
**This test will be used to verify operation after environmental testing, and during the EMI/EMC test**

## Integration and Test Plan Development in R&D

**Observatory end-to-end Optics Test**
**This test uses pin hole emitters built onto the focal plane along with a high quality optical flat to do a full end-to-end test over the entire surface of each of the mirrors**
**The full Observatory Optical Systems and Electronics will be exercised in this test including all relevant Spacecraft Systems**
**Performance of this test will require a dedicated facility,  for which the requirements will be developed during the R&D phase.**
**A complete set of preliminary plans for the Integration and Test activities at each level of assembly will be developed during the R&D phase to allow an accurate cost of this activity to be developed**

## Reflex Mirror Test Set-up

**Optical components**

## Reflex Mirror Test Set-up

**Satellite end-to-end test**

##  Delta 2920H-10L
Lift Capacity and Fairing Volume Marginal

 Delta 3940-11
Lift Capacity and Fairing Volume Adequate 
GSFC “actual” Cost Estimate is $80M to $84M
NASA & Boeing Describe this as a TRANSITION VEHICLE 

Atlas IIIB (or larger)
Lift Capacity Comfortable, Fairing Volume Tight but Adequate

Zenit 3SL w/ Block DM-SL restartable upper stage (SEA LAUNCH)
Liftoff Capability Comfortable, Fairing Volume Comfortable
 Not a US Owned (majority holder) Company

Delta 4240-12 (BASELINED)
Liftoff Capability of 2020kg Direct to Selected Orbit
Fairing Volume Adequate 

- Lift Capacity and Fairing Volume ***Marginal***
-  **Delta 3940-11**
- Lift Capacity and Fairing Volume Adequate
- GSFC “actual” Cost Estimate is $80M to $84M
- NASA & Boeing Describe this as a ***TRANSITION VEHICLE***
**Atlas IIIB **(or larger)
- Lift Capacity Comfortable, Fairing Volume Tight but Adequate
**Zenit 3SL w/ Block DM-SL restartable upper stage (SEA LAUNCH)**
- Liftoff Capability Comfortable, Fairing Volume Comfortable
-  ***Not a US Owned*** (majority holder) Company
**Delta 4240-12 *****(BASELINED)***
- Liftoff Capability of 2020kg Direct to Selected Orbit
- Fairing Volume Adequate

**Candidate Launch Vehicles**

## NASA GSFC Estimates of Delta IV Mass to Orbit

- Orbit  (29 deg inclination)       Vehicle                 Observatory Mass
- 7000 km x 155830 km	Delta 4040-12	    1350 kg
-   				Delta 4240-12	    2330 kg
- 				Delta 4450-14	    2600 kg
- 8000 km x 154830 km	Delta 4040-12	    1250 kg
-   				Delta 4240-12	    2210 kg
- 				Delta 4450-14	    2450 kg
- 10000 km x 152830 km	Delta 4040-12	    1100 kg
-   				Delta 4240-12	    2020 kg     (SNAP Baseline)
- 				Delta 4450-14	    2230 kg
- 				Delta 4040-12/SC	    2700 kg     (Note 1)
- Note 1:  For the last case, spacecraft hydrazine propulsion is used to raise the perigee from  vehicle parking orbit altitude to 10000 km.  The mass of 2700 kg is the final spacecraft mass after an expenditure of 311 kg of hydrazine with a specific impulse of 220 seconds.

## Results of NASA Delta IV Dispersion Analysis

- Because the final SNAP orbits are high altitude, the mass of spacecraft   hydrazine required to correct 3 vehicle dispersions is quite small.  Hydrazine masses are shown in the following table for initial spacecraft masses specified on the previous slide as the vehicle capabilities.
- Perigee, km     Dispersion	  Dispersion, km	   Hydrazine,kg     Hydrazine, kg
-                                                                            Delta 4240	Delta 4450
-     7000	         3 Perigee	           54	            1.2	        1.4
-     7000	         3 Apogee	       2220	            4.2	        4.7
-   10000	         3 Perigee	           73	            1.3	        1.4
-   10000	         3 Apogee	       2190	            4.0	        4.4

## Spacecraft Systems Summary

**Bus Requirements Are Well Understood**
**Concept Is Well Underway**
**We understand the Trade Space**
**We have a Working Concept with Few Tall Poles**
**Detailed Interface Definitions will be Generated during R&D**
**Details Will Flush Out Latent Problems **
**Will Help Firm Up Plans**
**We have identified several Launch Vehicles which will deliver SNAP to our chosen orbit**
**Development of the baselined Delta IV will be carefully tracked during the R&D phase **
**Integration and Test**
**Early Planning of S/C, Instrument and Telescope I&T Essential to Development of Instrument Design**
**I&T Schedule and Facilities Planning Will Enable Valid Cost Estimates**