LRO Design Status - Massachusetts Institute of Technology · LRO Design Status Craig Tooley - GSFC/431 LRO Project Manager November 14, 2005. 2 LRO REDESIGN TO UTILIZE SRM/MONO-PROPELLANT

1LRO REDESIGN TO UTILIZE SRM/MONO-PROPELLANT PROPULSION SYSTEM

LRO Design Status

Craig Tooley - GSFC/431LRO Project Manager

November 14, 2005


2008 Lunar Reconnaissance Orbiter (LRO)First Step in the Robotic Lunar Exploration Program

Robotic Lunar Exploration ProgramRobotic Lunar Exploration Program

LRO Objectives

• Characterization of the lunar radiation environment, biological impacts, and potential mitigation. Key aspects of this objective include determining the global radiation environment, investigating the capabilities of potential shielding materials, and validating deep space radiation prototype hardware and software.

• Develop a high resolution global, three dimensional geodetic grid of the Moon and provide the topography necessary for selecting future landing sites.

• Assess in detail the resources and environments of the Moon’s polar regions.

• High spatial resolution assessment of the Moon’s surface addressing elemental composition, mineralogy, and Regolith characteristics


LRO Payload Suite

Tectonic, impact and volcanic processes,

resource evaluation, and crustal evolution

Surface Landing hazards and some resource identification

1000’s of 50cm/pixel images (125km2), and

entire Moon at 100m in UV, Visible

LROC

Radar imaging, possibly including bi-static

imaging with Chandrayaan-1

~50 m scale polar topography at < 1 m vertical, roughness

Maps of hydrogen in upper 1 m of Moon at

5km scales

Maps of frosts in permanently shadowed

areas, etc.

300m scale maps of Temperature, surface ice,

rocks

Tissue equivalent response to radiation

Measurement

Source, history, deposition of polar

volatiles

Demonstrate new lightweight communication and

navigation technology, Locate potential water-ice

Mini-RF(technology demonstration)

Geodetic topography for geological evolution

Safe landing sites and surface navigation

LOLA

Source, history, migration and

deposition of polar volatiles

Locate potential water-ice in lunar soil

LEND

Locate potential water-ice on the surface

LAMP

Determines conditions for systems operability and

water-ice location

Diviner

Radiation boundary conditions for

biological response

Safe, high performance, lighter weight space vehicles

CRaTER

ScienceBenefit

ExplorationBenefit

INSTRUMENT

Tectonic, impact and volcanic processes,

resource evaluation, and crustal evolution

Surface Landing hazards and some resource identification

1000’s of 50cm/pixel images (125km2), and

entire Moon at 100m in UV, Visible

LROC

Radar imaging, possibly including bi-static

imaging with Chandrayaan-1

~50 m scale polar topography at < 1 m vertical, roughness

Maps of hydrogen in upper 1 m of Moon at

5km scales

Maps of frosts in permanently shadowed

areas, etc.

300m scale maps of Temperature, surface ice,

rocks

Tissue equivalent response to radiation

Measurement

Source, history, deposition of polar

volatiles

Demonstrate new lightweight communication and

navigation technology, Locate potential water-ice

Mini-RF(technology demonstration)

Geodetic topography for geological evolution

Safe landing sites and surface navigation

LOLA

Source, history, migration and

deposition of polar volatiles

Locate potential water-ice in lunar soil

LEND

Locate potential water-ice on the surface

LAMP

Determines conditions for systems operability and

water-ice location

Diviner

Radiation boundary conditions for

biological response

Safe, high performance, lighter weight space vehicles

CRaTER

ScienceBenefit

ExplorationBenefit

INSTRUMENT

Cosmic Ray Telescope for the Effects of Radiation

Lyman-Alpha Mapping Project

Lunar Exploration Neutron Detector

Lunar Orbiter Laser Altimeter

Lunar Recon Orbiter Camera

IR Thermal Imager


• Launch in late 2008 on a Delta II rocket into a direct insertion trajectory to the moon.

• On-board propulsion system used to capture at the moon, insert into and maintain 50 km altitude circular polar reconnaissance orbit.

• 1 year mission• Orbiter is a 3-axis stabilized,

nadir pointed spacecraft designed to operate continuously during the primary mission.

LRO Mission OverviewFlight Plan – Direct using 3-Stage ELV

Solar Rotating Coordinates

Earth

Moon at encounter

Cis-lunar transfer5.1978 day transferLaunch C3 –2.07 km2/s2

1-dayLunar Orbit

Sun direction

Nominal Cis-lunar Trajectory

Cis-Lunar Transfer

12-hour orbit

6-hour orbit

100 and 50kmmission orbits

Insertion and CircularizationImpulsive ΔVs (m/s)

1 – 344.242 – 113.063 – 383.914 – 11.455 – 12.18


LRO Mission – Current Status• LRO in combined Phase A/B which was begun in earnest February 2005 after instrument selection and

receipt of funding.– Spacecraft and Ground System being developed in-house at GSFC– Instruments selected via competitive AO, targeted high heritage instruments– Investigation Data Products delivered to the Planetary Data System (PDS)

• Successfully completed Mission System Requirements Review (SRR) 8/18/05 and Instrument Preliminary Design Reviews (PDR) 10/05/05.

• 6 of the 9 spacecraft subsystem Peer Design Reviews completed– Thermal, C&DH, Power, Software, Communication, GN&C

• Major Procurements & Hardware Development Proceeding– Spacecraft Computer contract in place and past PDR– Data Recorder acquisition on going– Negotiation and award of instrument Phase C/D/E contracts to be complete by end of this December.– Power system electronics designed and being breadboarded– Battery RFP scheduled for release 11/17/05– All mono-propellant system components except tank on order– Mechanism components, ACS sensors, SRM, and communication component RFPs to be released Dec. 2005– Contract in-place for White Sands Ground Station

• Development paused to address increased risk posed by Nutation Time Constant issue– PDR and remaining 3 Subsystem Peer Reviews rescheduled later as part of chosen solution

• Mission budget and schedule baseline to be established through Confirmation Process following PDR.

• Project is adequately staffed and making excellent progress.


LRO Development Start-Up Chronology

6/16 – 10/05IPDRs

2004 2005 2006 2007 2008

6/18AO Release

11/15AO Select

4/15IPDRs

4/28IAR

5/15PDR

10/15Confirm

2/15ICDRs

4/15CDR

10/15Instrument Delivery to

I&T

10/15LRD


I&T

10/15LRD


I&T

10/31LRD

12/23AO Select

6/18AO Release

6/18AO Release

2/1Funding Available

6/16 – 10/05IPDRs

8/18SRR

11/15PDR

11/16IBR

12/27NAR

3/1 – 5/1ICDRs

9/7CDR

11/1CDR

2/1PDR

3/1 – 5/1ICDRs

IBR NAR

8/18SRR

12/23AO Select

2/1Funding Available

AO HQ:SMD RLEP:GSFC

4/15Mini-RF Added to LRO Payload

10/1Start RLEP $

12/03PER

12/03PER

12/03PER

7/15MRR

7/15MRR

7/15MRR

Post SRR HQ:ESMD RLEP:ARC

Post NTC IssueHQ:ESMD RLEP:ARC

• Project initially planned with SMD as in-house SMEX/Skunk-works type effort with early very aggressive milestones established to ensure high heritage instruments and rapid start of development. Programmatic changes, the consequences of intentionally planning very aggressive target dates, and the recent NTC technical issue have all contributed to the flux in LRO’s near term milestone schedule

• The Project believes the overall rapid development plan continues to be sound and the 2008 launch requirement is still achievable.

3/4/04ORDT

3/4/04ORDT

3/4/04ORDT


LRO Orbiter Design Development

SOLAR ARRAY

HGA

PROPULSION MODULE

AVIONICS MODULE

INSTRUMENT MODULE

Diviner

LEND

LAMP

AO Concept Present Redesign(PDR Baseline)

SRR Baseline


LRO Nutation Time Constant Issue and Redesign Trade Study


LRO Nutation Time Constant (NTC) IssueTechnical Background

• NTC requirement originates from stability requirements of Delta II spin stabilized third stage. NTC is a function of propellant mass, tank configuration, and overall stack mass properties.

– Requirement: NTC >150 sec at SRM ignition & >50 sec at burn-out.– The NTC characterizes the rate of growth of the nutation angle (coning) as a

prolate spinner dissipates energy and evolves toward a stable (flat spin) state.

– The Delta II third stage employs a simple active nutation control system to correct for this during third stage flight.

• Initial technical meetings with KSC Launch Vehicle group at the beginning of the Project identified the need to address tank baffle design for NTC control early.– The problem is largely empirical and is always mission unique. – LRO’s large propellant load and mass fraction at the edge of Delta II

experience base.• Identified and tracked as a significant risk early in the Project.


SOLAR ARRAYPROPULSION

MODULE

AVIONICS MODULE

INSTRUMENT MODULE

Mini-RF

LOLA

LAMP

LROC

CRaTER

SOLAR ARRAY

HGA

PROPULSION MODULE

AVIONICS MODULE

INSTRUMENT MODULE

Diviner

LEND

LAMP

Dry: 603 kg

Fuel: 714 kg

575 Gb/dayMeasurement Data Volume

745 WPower

1317 kgMass

LRO Characteristics

LRO SRR BaselineLarge Mono-Propellant Design Baselined for Simplicity, Ease of

Operation, and Lower Cost

Delta II Third Stage

LRO


• LRO Propulsion consulted with and then contracted with the experts (Hubert Astronautics, John Harrison, PSI) recognized by KSC and Boeing to support LRO propulsion system design.

– Both semi-empirical and analytical methodologies applied to LRO tank options as design progressed and then to the selected baseline (single cylindrical) tank design. Specification and contract for early drop-tower testing also begun.

– Initial assessment was that with adequate analysis and test verification a viable tank baffle design would be straightforwardly established.

– Subsequent analysis raised concerns that an acceptable tank/baffling design might not be achievable.

• Further investigation included consultations with MSFC, NRL propulsion group, Boeing, and KSC Launch Vehicle Group as well as continued work by GSFC Propulsion Group.

LRO Nutation Time Constant (NTC) IssueEvolution of the Risk


• Expert consultations and further study led the Project to conclude that the risk of an unacceptable schedule and cost impact due to the failure to achieve an acceptable NTC with the baseline large mono-propellant design was too high to be accepted if alternative designs were viable.

• Furthermore, even if an acceptable NTC tank baffle design is eventually found, the risk of a protracted, expensive effort to achieve this is also judged unacceptable if a viable, more deterministic design solution can be identified

LRO Nutation Time Constant (NTC) IssueDecision to Seek Alternative Design


NTC Issue Redesign Trade Study Approach

• Ground Rules for Identifying Viable Options:– Risk to schedule must be judged significantly lower than continuing forward with NTC issue.– Adequate mass margin must be preserved.– Impacts to Instrument requirements and interfaces must be minimal.

• A variety of design options were examined including:– 3 different mono-propellant tank configurations– 3 different bi-propellant configurations (single mode, dual mode, and single mode on 2-stage Delta)– 3 different “mixed” options (solid/mono-prop, EELV launch, inertial 3rd stage on Delta)– The trade study leveraged propulsion work done earlier in the Project prior to the selection of the

large mono-propellant system.

• These converged to 4 Redesign Options which the trade study further compared:

– 1)Baseline large mono-propellant design– 2)Single Mode Bi-propellant– 3)SRM + small mono-propellant system design– 4)Non-spinning launch vehicle (EELV)


• Large schedule and cost impact if a tank solution is not found • Will not know whether there is a problem until tests and analysis are completed 4 to 6 months from now• Mass margin requires UltraFlex lightweight solar array which is much lighter but is less mature technology.

Disadvantages

• Substantial mechanical, GN&C and thermal analysis had already been done on this design (within two weeks of PDR-level peer review)• Simplest propulsion solution—one system handles all maneuvers• Simplest Mission profile

Advantages

24%LRO Mass Margin711 kg.Fluid MassDelta 2925H-9.5Launch Vehicle

Design Option: SRR Baseline Large Mono-Propellant (One 35”x57” cylindrical tank)

NTC Issue Redesign Trade Study


• NTC issue still a significant risk• Complex propulsion system

– High schedule risk • Complication of LRO providing structural/thermal system as GFE.

– Extra $7-9 million• Need to redesign the structure and thermal system, reanalyze ACS• Potential contamination issue throughout the mission

Disadvantages

• Increases Mass Margin• Modestly reduced fluid mass arranged in a manner similar to a ultimately successful design (Messenger)

Advantages


Design Option: Single Mode Bi-Propellant System (Two 32” Spherical Tanks)



• Impacts operations concept during transit to the moon–Must keep solar arrays stowed (use traditional array design facing out)–Need heater power to keep STAR motor at a uniform temperature

• Need a spin-up/spin-down system for STAR-27 burn• Need nutation control during STAR-27 burn• Need to redesign the structure and thermal system, reanalyze ACS• $3 to 4 million for SRM and associated new hardware• Thermal and contamination concerns from solid motor

Disadvantages

• Drastically reduces fluid volume and therefore NTC issue• Adequate mass margin• Mono-prop system can be built in-house, most components already in-hand, tank procurement comfortable.• Short, robust LOI burn• Simple, smaller mono-prop system, an advantage during I&T.• Flight SRM not required until LRO arrives launch site.

Advantages


Design Option: Star-27 SRM + Small Mono-Propellant (One 32” spherical tank)



• Cost impact not known• Contrary to Formulation Authorization Document and Level 1 Req.• Need to redesign the structure and thermal systemDisadvantages

• 3-axis stabilized launch vehicle, eliminates NTC concern & spin balance• Retains simplicity of original baseline• Reduce schedule risk (would allow use of on-hand TDRS tanks, traditional solar array)• Modest mass margin gain (tank limited)• Increased fairing volume increases design flexibility• Avoids potential issues with using last Delta II

Advantages

28% (Launch Vehicle has additional 1043 kg excess capability)LRO Mass Margin870 kg.Fluid MassDelta 4020Launch Vehicle

Design Option: EELV Launch w/ Mono-Propellant (Two 40”oblate spheroid TDRS tanks)



• LRO Project selected SRM+small mono-propellant system as best choice for redesign.– Drastically reduces NTC risk– Small, simple mono-prop system simplifies Orbiter development– SRM and small mono-prop system have low acquisition risks– Development challenges of SRM Lunar Orbit Insertion Stage (LOIS)

well understood– Simple I&T since flight SRM is mated to Orbiter at KSC and Orbiter is

significantly smaller during most of I&T• EELV determined to be best technical/risk solution but

dismissed as outside the Project’s “trade-space”.– To exploit advantages LRO must be baselined on an EELV before

the redesign is completed. The schedule will not allow another redesign cycle later in the program.

NTC Issue Redesign Trade StudySelected Approach


Implementation of SRM + Small Mono-Propellant Redesign of LRO


Programmatics of LRO Redesign

• Schedule– Development schedule must be adjusted to allow mechanical and thermal

and GN&C subsystems to recover from design changes before PDR.– Sufficient time for Lunar Orbit Insertion Stage (LOIS) subsystem preliminary

design must also be provided.– Additional experienced staff brought on-board, subsystems can and are

committed to recovering in time for PDR 2/6/06 & CDR 11/1/06. – Other subsystems and instruments minimally affected– If necessary LRO will plan for ∆PDRs if specific aspects lack sufficient

maturity– Technical Peer reviews prior to PDR will include a LOIS Peer review

• Cost– Impact of adding LOIS estimated to be ~$4.0 - 6.0M

• $3-4M for hardware plus $1-2M additional staffing


LRO Redesign Technical Tasks

• SRM powered flight phase is a new mission phase with unique requirements and constraints.– Trajectory must be adjusted to use single SRM maneuver for LOI– Orbiter must be power positive and establish communications during cruise

without deployments.– Orbiter thermal design must account for SRM plume and soak-back heating– LOIS subsystem and/or GN&C system must provide spin-up/down and

nutation control during SRM powered flight.– Orbiter must jettison system for SRM after burn– Contamination from SRM must be understood and mitigated if necessary– SRM represents addition range safety hazards that must be controlled – Reconfigured thermal and mechanical systems must meet all mission

requirements and accommodate new LOIS subsystem


Redesign of LRO w/SRM

Star-27 w/ Thermal Shield

Star-27 Separation Plane

Mini-RF

S/A Deployed Area 10.1m2

Thermal Radiator

High Gain Antenna System

LRO Spin/De-Spin Motors (4)

LOLA

LROC

Delta II Fairing

CRaTER

LEND

Diviner

Service Module (Avionics & Propulsion)

LAMP


LRO System DiagramPreliminary w/addition of SRM

PDE

Propulsion

LROC

Mini-RF

LOLA

LEND

LAMP

Diviner

CRaTER C&DH

DIB

HK/IO

Thermal

LVPC

SBC

PSE

PMC

SAM

OM - A

OM - B

OM - C

OM - D

Solar Array

Gimbal Controller

S-Xpndr

KaModulator

IMU

Spac

eWire

Net

wor

k

LAMP Science Data (LVDS, 1Mbps)

MIL-STD-1553 Network

HGA Gimbals

Omnis

HGA

CSS (8)

IRW (1)

HGA Deploy Actuation

Cmds (4kbps) & Low Rate Tlm (2Mbps max)

Hi-Rate Tlm (I&Q, 115 Mbps Max)

Thermistors

Unsw +28V

Uns

w. +

28V

Pow

er B

us

cPCI

NAC1 WACNAC2LROC Sw +28V

LOLA Sw +28V

LEND Sw +28V

LAMP Sw +28V (RED)

DLRE Sw +28V

CRaTER Sw +28V

DDA

Mini RF Electronics(Sw +28V)

Mini RF TX(Sw +28V)

IRW (3)

IRW (4)

IRW (2)

LAMP Cmd/Tlm Data (RS-422 UART @ 38.4kbps)

DLREt Survival Heaters (Sw +28V)

Thermistors (4)

Thermistors (2)

Thermistors (?)

32Mbps Tx/4Mbps Rx

125M

bps

Tx/1

0Mbp

s R

x (S

cien

ce K

a D

L)

40Mbps Tx/4Mbps Rx

4Mbp

s Tx

/2M

bps

Rx

(HK

S-B

and

DL)

Closed Loop Heaters (5)Sw +28V

(Ops Heaters)

+28V

Bac

kpla

ne

Unsw +28V(3 services)

Sw +28V(12 services)

Li-Ion Battery80 AHr

Single point

Ground

HGA Gimbal

Controller

Star

Tracker

(1)

StarTracker (2)

Uns

w +

28V

(Rec

eive

r)

Sw +

28V

(Tra

nsm

itter

)

Sw +28V(Ka Transmitter)

Sw +28V

Sw +28V

Sw +28V

Sw +28V

Sw +

28V

Sw +28V

Sw +28V

Solar Array

24 Cells per string10 Strings per Segment

14 Segments

SA Gimbal

SA Gimbal

Sw +28V

Instrument Survival Heater Bus (Unsw +28V)

Thermistors (2)

Thermistors (?)

DLRE Op Heaters (Sw +28V)

LAMP Sw +28V (PRI)


Instrument Ops Heater Bus (Sw +28V)


Instrument Ops Heater Bus (Sw +28V)


LROC Decontamination Heater (Sw +28V)


Omni/HGA Select (28V pulses)

Cmd/Tlm (RS-422 [email protected])

Ka Comm

S-Band Comm

Sw +28V

USO9500

20 MHz Clock

USO9600

20 MHzClock

1 PPS

Sw +28V

S/ADeploy

S/ADeploy

HGA Gimbals

HGADeploy

HGADeploy

S Comm Card Reset(sbc cmd 0)

PSE PMC Reset (sbc cmd 2)

PSE OM “Jumper ON” (sbc cmd 3)

PSE SAM “All Segments ON” (sbc cmd 4)

C&D

H L

VPC

Pow

er C

ycle

(h/w

cm

d 1)

C&

DH

LVP

C P

or (h

/w c

md

2)

PSE

OM

“All

OFF

” (h/

w c

md

3)

PSE

SAM

“Nor

mal

” (h/

w c

md

4)

C&DH LVPC Power Cycle (h/w cmd 1)

C&DH LVPC POR (h/w cmd 2)

PSE OM “All OFF” (h/w cmd 3)

PSE SAM “Normal” (h/w cmd 4)

Prop/Dep-A

Prop/Dep-B

Prop/Dep-C

Prop/Dep-D

Sw +28V

SA Deploy Actuation

x14

11w/ OVbckp

1

2







Bat

tery

ON

/OFF

Lin

e

DPC/GSE Power

N2H4

Propellant

Tank

With PMD

R

R

R

R

P

P

MIL-STD-1553 Network

KaTWTAModulator Cmd/Tlm (RS 422)

S/A Hinge

HGA Hinge

TWTAEPC

HighVoltage

Sw +28V

+15V

4

4

ELV Sep Switch

Sw +28V

Sw +28V

Sw +28V

ELV Sep Switch

ELV Sep Switch

ELV Sep Switch

Sw +28V(S/C Ops Heater Bus)

Sw +28V(S/C Survival Htr Bus)

ATA Interface

50M

bps

Tx4M

b ps

Rx

Helium Pressure Tank

SpinRocket

SpinRocket

SpinRocket

SpinRocket

SRM Separation

Safe & Arm

Spin Up

Spin Down

Pyrotechnic Ordnance

Train

SRM Fire

1A

3A

2A

1B

4A

2B

3B

4B

UmbilicalConnector

Motor Control

Position Status

To S/C Tlm

Lunar Orbit Insertion Stage

NCT

N2H4

Propellant

Tank

With Diaphragm

P

P

Fill/DrainValvePanel

Nutation Control Electronics

STAR 27H

Safe/ArmS/C Skin

Connector

Lunar Orbit Insertion Stage Jettison

Thruster/Latch valve Cmds/Tlm



Nutation Control Thruster Cmds

Sw +28V

8

4

4

2

Changes from SRR Baseline

New LOIS Subsystem

Rigid Array

Smaller mono-prop Subsystem


22 suns

9 suns 16 suns

6 suns

1 sun

No Plume Shield Effect of Plume Shielding

Motor soak back will be negligible due to 4 bolt mounting configuration and MLI blanketing on case and nozzle

8 suns

1 sun = 1372 W/m²

Plume Heating (Convective & IR)Instantaneous Peak Heating Predictions

Continuous lightweight shield, alternative designs use local shields


LRO SRM Powered Flight

• Preliminary powered flight simulations show system can be controlled with acceptable flight performance using a reasonably sized Nutation Control System (NCS) (~5 nm).

• NCS design to be derived from Triana system that was built and tested.


31 flights (100% success)Original Qualified in 1975H designation is upgraded version

Size: 48” long x 27.3” diameterIsp: 294 secMax Propellant Load: 339 kgMax Thrust: 24500 N

0

1000

2000

3000

4000

5000

6000

0 5 10 15 20 25 30 35 40 45 50

TIME (sec)

VAC

UU

M T

HR

UST

(lbf

STAR-27H Solid Rocket Motor


• ATK is upgrading the Star-27 using flight proven design features from other STAR motors.

– Increased performance and replacement of obsolete materials• Upgrade to Head-End Web Grain Design

– Flow successfully on over 470 motors• Upgrade to HTPB type propellant

– Higher Performance, – 280 STAR motor flights with this formulation– Demonstrated up to 15 months exposure to space (Magellan)

• Upgrade Nozzle to Star 30C design– No significant modifications needed for Star 27 case– Provides higher thermal margins and increased Isp

– Upgrades being made for another 2008 mission (IBEX, launch mid-2008)• Upgraded motor will undergo static test fire

– Upgrade test and certification program schedule within LRO schedule

• LRO believes this is a modest risk but plans to engage ATK to explore accelerating the upgrade qualification program. Initial discussions promising.

– ATK report only 3 failures out of 2600 Star motor flights. The three failures were all traced to materials no longer used on STAR motors. The current Star motor configurations have 100% success rate with more than 83 flights

STAR-27H Solid Rocket MotorUpgrade


Summary

• The LRO Project has concluded that modifying the LRO design to use a SRM for Lunar Orbit Insertion and a small mono-propellant system for orbital maneuvering is the best viable solution for 2008 launch of within the original SMD direction to “be compatible with a Delta II class launch vehicle”.

• LRO has begun to proceed with redesign.


Back-Ups


Propulsion Options Examined• Baseline large mono-prop (35" cylindrical tank) • Single-mode bi-propellant

– MMH, NTO– Tanks: 1-32" Fuel, 1-32" Ox spheres on center – One 100 lbf+ Eight 5lbf thrusters– 221 kg fuel, 364 kg ox

• STAR-27 with mono-prop regulated (1-32" PMD Tank)• EELV, mono-prop in 2 TDRSS tanks

• Dual-mode bi-propellant– N2H4, NTO– Tanks: 3-Fuel, 1-Ox : 24.6"x 29"– One 100 lbf + Eight 5 lbf thrusters– 443 kg fuel, 172 kg ox– Rejected: similar complexity to single mode, but margin is worse

• Bi-prop on a 2-stage Delta – Rejected: only 17% margin, radiation problems with phasing loops

• Mono-prop 42" tank with 2-19" – Rejected: only 17% margin

• Mono-prop single 42" tank without extended mission DeltaV– Rejected: does not meet requirements, not clear that 42” tank is any better than current tank

• Bi-prop on a 2-stage Delta – Rejected: only 17% margin, radiation problems with phasing loops

• Inertial 3rd stage– Rejected: not enough margin, too much complexity (building a launch vehicle)


Redesign Options - Mass

Option BaselineSingle-mode

Bi-propSTAR-27 with

mono-prop

EELV, mono-prop in 2 TDRSS

tanks

Launch VehicleDelta 2925H-9.5,

3-stageDelta 2925H-9.5,

3-stageDelta 2925H-9.5,

3-stageDelta 4020

EELVLaunch Capability (kg) 1480 1480 1480 2855Launch Mass for this config (kg) 1476 1480 1480 1812Excess Launch Capacity (kg) 4 0 0 1043Fluid Mass (kg) 711 586 285 873Propellant Mass (kg) 711 586 626 873Ejected Mass (kg) N/A N/A 50 N/ADry Mass Capability (kg) 765 894 804 939Propulsion System Mass (kg) 96 112 68 130Baseline CBE for Non-Prop (kg) 523 523 523 523Non-Prop Impacts (kg) 0 45 64 80Total Dry Mass (kg) 618 680 654 732Dry Margin (kg) 147 214 150 207Margin (% of total Dry Mass) 24% 32% 23% 28%


LRO Mission Timeline

Nom Mission Ext Mission Disposal Pre-launch Launch Cruise LOI Commission

MCC Track

Point for LOI

Sun Point Omni Comm Track Point for MCC


Sun Point

Cmd Deploy

De-spin Separate


Trim Orbit Unload Mom Inertial Pnt HGA Comm Lunar Pnt Checkout/Cal

Sun Point

Track

Unload Mom

Unload Mom


3rd Stage De-spin Separate 1st Stage 2nd Stage Coast Reorient Spin-up

Burn SRM Spin-up Unload Mom

Take Data


Trim Orbit Unload Mom Inertial PntSend Data Lunar Pnt Periodic Cal Track


Mission Timeline - Changes for SRMLunar Cruise


MCC Sun Point Track Rate Null Sun Point Omni Comm Track Point for MCC

Component Allocation CBE MarginPSE 67.1 45S-Comm 48.3 36RW 72 64C&DH 100 85Gyro 36 32LEND 13 13CRaTER 9 5SubTotal H/W 345.4 280 23%

Propulsion Heaters 90.96 75.80Total Essential Heaters CBE 90 75.00Total SC Op Heaters CBE 0 0.00Total Gimbal Op Heaters CBE 0 0.00Total Instr Op Heaters CBE 0 0.00Total Instr Surv Heaters CBE 86.52 72.10Total Diviner Op Heaters CBE 0 0.00Total Diviner Surv Heaters CBE 15.6 13.00Total Deployment Heaters CBE 57.6 48.00SubTotal Thermal 340.68 283.90 20%

Total 686.08 563.90 22%

Max Allocation 700 563.90 24%

- Rate null is likely propulsive (unchanged)

- Solar array remains stowed (1/3 power) ->

- High gain remains stowed (unchanged)

- Comm is Omni/S-band only (unchanged)

- 3-axis, reaction wheel control (unchanged)

POWER (Watts)


Mission Timeline - Changes for SRMLunar Orbit Insertion (LOI)


- Proper attitude attained via 3-axis, reaction wheel control

- All momentum removed from system prior to open-loop spin-up

- Spin-up done using 2 small Star-3 solid rocket motors (SRM’s)

- Main Star-27 SRM completes its burn in 47 seconds (2.4 g max)

- De-spin delayed for approximately 2 minutes while SRM chuffs

- De-spin done with 2 small, appropriately canted, Star-3 SRM’s

- Separation system consists of 4 explosive bolts and springs

- After separation, LRO back under 3-axis, reaction wheel control

- Final maneuver removes residual momentum from system

- Final maneuver serves as collision avoidance maneuver

- Entire sequence nominally within view of LRO ground station

Point for LOI Sun Point De-spin Separate Unload H Burn SRM Spin-up Unload H

Delta V Budget

6251326Totals

14--Residuals

16--Momentum Unloading

820Margin

53125Extended Mission

76180Mission Maintenance

63148Additional Orbit Trim

5--Nutation Control

340 (solid)778Solid Rocket Burn

5075Mid-Course Correction

Fuel Mass (kg)

Delta V (m/s)

Event

Delta V Budget

6251326Totals

14--Residuals

16--Momentum Unloading

820Margin

53125Extended Mission

76180Mission Maintenance

63148Additional Orbit Trim

5--Nutation Control

340 (solid)778Solid Rocket Burn

5075Mid-Course Correction

Fuel Mass (kg)

Delta V (m/s)

Event


Ground System Architecture Overview

Documents

LRO Design Status - Massachusetts Institute of Technology · LRO Design Status Craig Tooley - GSFC/431 LRO Project Manager November 14, 2005. 2 LRO REDESIGN TO UTILIZE SRM/MONO-PROPELLANT