Robotic Lunar Lander Concept

International Space Development Conference

Reginald Alexander

Greg Chavers

Tom Percy

May 26, 2018

Background Information

• Performed a short study of cryogenic lunar lander concepts in Summer of 2017
• Multi-center Lander Tech Office 2-phase effort to define a trade space and develop a concept to land cargo on the moon using cryo propellants
• Purpose: Investigate viability of cryo propellant lander within the constraints of existing launch vehicle capabilities

​

• Findings:
• For 500 kg payload, lander wet mass exceeded Atlas V 551 capability
• Cryogenic propellants trade better as landed payload grows
• Cryogenic propulsion systems can enable more ambitious missions if more capable launch vehicles are available

After the Study

• Team identified several areas for improvement
• electric-Pump fed methane thrusters may save mass over a pressure-fed system and enables improved engine performance
• Landing legs may enable reuse and provide more stable landing platform
• Payload access to the surface is challenging as landers grow in physical size due to increased propellant loading and lower density propellants
• Structural optimization and reconfiguration of concept can reduce overall lander mass

​

• Team took on a new perspective on launch vehicle performance
• Newly emerging launch vehicles promise increased payload capacity
• Fitting the methane lander in existing launch vehicles is challenging
• Leveraging new launch vehicles allows for an increase over the previous 500 kg landed mass target

The team determined that next lander concept study would leverage work completed in September, 2017 with focused improvements and an eye towards emerging launch vehicles and large landed payloads

Objective Statement and Approach

• Study Objective: Update concept based on previous findings and design a lunar robotic lander concept that could support the demonstration of active cryo-fluid management technologies for NASA and serve as a workhorse lunar surface cargo delivery vehicle
• The lander should support the following:
• Short term goal: Demonstration of long-duration (longer than standard lunar mission) active cryogenic fluid management technologies
• Long term goal: Landing 1000 kg of cargo on the lunar surface using LOX/CH4 propellant with a lander concept that is operationally and economically appealing to a private landing services provider
• Modular cryo system that the end user can modify as needed (i.e. removing long-duration CFM components)
• Mission portfolio approach
• Identify of portfolio of missions that the lander should be capable of executing to varying levels of performance
• Select 1 mission to set the baseline design
• Determine what performance the lander can achieve in the other missions

​

Key Concept Ideas

• Workhorse Lander: Flexibility to support a range of lunar landing missions while filling a gap in payload delivery capability

​

• Demonstration of Technology: NASA uses the lander design to demonstrate feed-forward technologies in propulsion and cryogenic fluid management

​

• Forward-Leaning in Specific Areas: Lander concept relies on methane propulsion and associated CFM technologies, applying commercial and government technology development programs already underway, while employing high-TRL components in other areas to maintain affordability

​

• Applications for the Future: Applying advances in cryo propulsion, the lander lays the groundwork for more ambitious endeavors in the future, including human exploration beyond Low Earth Orbit.

​

Mission Modes: ΔV Map

Launch

TLI

Moon

Surface

Earth

3200 m/s

2500

900

425

730

1800

730

DSG

DSG

Lunar Orbit

Lunar Orbit

LV/US

US/L

US/L

US/L

US/L

L

L

Potential Elements to Perform Maneuvers

LV = Launch Vehicle

US = Upper Stage

L = Lander

O = Other

3000

L/O

1800

L

Multiple Sites

Potential Missions

Polar Only

Global Access

Landing Profile

= Loiter time (up to 14 days) required

= Active CFM required

Surface Mission Profile

Crater Exploration

= Restart required

= Additional DV margin required

Surface Hopping

Return to Orbit

Reusable Lander

= Return DV required (By ISRU or in-space prop transfer), at least 1900-2500 m/s

= Lunar Surface Day / Night survival considerations

Baseline Mission Description

• Baseline mission was selected to serve as the sizing case for the lander concept

​

• Mission Profile:
• Deliver 1000 kg of payload to the lunar surface
• Layover in near-rectilinear halo orbit (NRHO) for potential stay at the Deep Space Gateway facility
• Transfer from NRHO to low lunar orbit (LLO) for phasing and precision landing navigation
• Global lunar surface access can be achieved through a loiter period in LLO of up to 14 days

Mission Modes: ΔV Map

Launch

TLI

Moon

Surface

Earth

3200 m/s

2500

900

730

DSG

Lunar Orbit

LV/US

US/L

US/L

L

Potential Elements to Perform Maneuvers

LV = Launch Vehicle

US = Upper Stage

L = Lander

O = Other

3000

L/O

1800

L

Multiple Sites

Landing

20 m/s (DOI)

1640 m/s (Braking)

220 m/s (Approach)

50 m/s (Vertical Drop)

NRO / DSG

TCM’s

30 m/s

Notes:

* All DVs except for landing are ideal/impulsive.

* Guestimate (placeholder) NRO loiter of 10.9 days

1930 m/s

L

Landing ~ 65 min

TLI + 30.6 days

LLO

(14 day loiter)

178 m/s

US/L

TLI + 4.1 days

Lunar

Flyby

NRO Arrival

250.5 m/s

US/L

TLI + 5.1 days

NRO Departure

250.5 m/s

US/L

TLI + 16 days

LOI

648.4 m/s

US/L

TLI + 16.5 days

Segment

TCM’s

10 m/s

Lander-Cargo

1 m

5.5 m

6.5 m

3.5 m

Cargo

2 x 3 x 1.5 m

Lander-Cargo

Concept Analysis

• CFM
• Baselined active cryo storage for longer-duration missions
• Removable parts for short-duration missions
• 2 cryocoolers required; 0.650 kW power req.

​

• Propulsion
• 8 x ePump-driven 1,400 lbf Lox/LCH4 main engines
• 16 x press-fed 30 lbf Lox/LCH4 RCS thrusters
• 67 kW required operational power to run ePumps

​

• Power:
• Single ultraFlex solar array for steady-state operations
• Batteries for propulsion system are significant challenge due to rapid discharge requirement to support electric pump operations
• Flight heritage battery solution heavy given discharge requirements

​

​

Concept Analysis

• Structures
• Full FEA performed for Earth Launch / Ascent, Propulsive Lunar Descent, and Lunar Landing
• Aluminum primary frame structure
• Composite tank support struts to minimize thermal conductivity

​

• Avionics
• 1-fault tolerant critical systems w/ component redundancy
• X-band comm to DSN
• Autonomous landing & hazard avoidance system based on LaRC/JPL work underway for lander project office
• Automated Rendezvous & Docking bolt-on avionics kit identified for return-to-orbit missions

Baseline Lander MEL

Payload = 1000 kg

​

Total Launch Mass = 15387.2 kg

CFM Demo Mission Description

• A potential first mission for the lander concept is a technology demonstration mission
• Demonstrate general mission operations
• Demonstrate Lox/LCH4 landing propulsion
• Demonstrate long-duration cryo-fluid management

​

• Mission Profile:
• Lander payload is replaced with CFM demonstration payload for use prior to lunar landing
• Follow same general mission profile as baseline lander mission
• Extend stay in both NRHO and LLO to achieve various CFM technology demonstration goals
• Lunar landing at the end of the mission demonstrates landing propulsion

​

Top Level CFM Demo Mission Requirements

• Must fit within the lander design for the operational reference mission
• Propellant loads limited to lander design tank volumes
• Must leverage CFM technologies already built into the operational lander design to the greatest extent possible
• Add CFM Demonstration payload to supplement demonstration goals
• Must end with a lunar landing demonstration
• Nominal mission duration and operations are set however, if off nominal performance is revealed, the in-space portion of the mission will be cut short to ensure enough propellant is available to land on the moon
• i.e. Demonstrate CFM for X days OR until propellant load = Y kg, whichever limit is reached first, then immediately initiate landing sequence

Cryogenic Fluid Management Across Multiple Propulsion Pieces

Red numbers indicate technologies that need to fly to reach TRL 6.

Does not capture effects of scale.

Fluid specific technologies may be shown in multiple locations.

10

4

10

2

14

Demonstrated on Lander

Demonstrated by adding a receiver tank on the payload

Deep Space Transport

(LOX/LCH₄)

CFM Tech: Lander vs Demo Payload

• Lander-Only Demo
• Captures ~80% of technologies to be demonstrated
• Reduces complexity and cost
• Requires addition of second set of avionics for instrumentation and data transmission

• Lander w/ Payload Demo
• Captures 100% of technologies to be demonstrated
• Adds methane tank, helium tank, fluids, and tank connections for transfer demo
• Requires addition of second set of avionics for instrumentation and data transmission

AES Mid-Year Review April 2014

Mission Modes: ΔV Map

Launch

TLI

Moon

Surface

Earth

3200 m/s

2500

900

730

DSG

Lunar Orbit

LV/US

US/L

US/L

L

Potential Elements to Perform Maneuvers

LV = Launch Vehicle

US = Upper Stage

L = Lander

O = Other

3000

L/O

1800

L

Multiple Sites

Landing

20 m/s (DOI)

1640 m/s (Braking)

220 m/s (Approach)

50 m/s (Vertical Drop)

NRO / DSG

TCM’s

30 m/s

Notes:

* All DVs except for landing are ideal/impulsive.

1930 m/s

L

Landing ~ 65 min

TLI + 89.6 days

LLO

178 m/s

US/L

TLI + 4.1 days

Lunar

Flyby

NRO Arrival

250.5 m/s

US/L

TLI + 5.1 days

NRO Departure

250.5 m/s

US/L

TLI + 61.1 days

LOI

648.4 m/s

US/L

TLI + 61.6 days

Segment

TCM’s

10 m/s

Lander-CFM Demo Options

Lander-Only Demo

Lander w/ Payload Demo

Mission Portfolio

• Various lunar mission profiles are assessed for delta-V budgets and timelines
• Lunar mission profile consists of launch profile, lunar arrival mode and landing profiles
• Payload is then a fallout calculation from sizing propellant loads
• Getting to the Surface
• Polar Access: Achievable anytime from a polar orbit
• Global Access: Achievable from a polar orbit with a loiter of up to 14 days
• Once on the Surface
• Crater Lander: Carry additional ΔV for landing
• Hopper: Carry additional ΔV for traversing to secondary landing sites
• Return from Surface: Perform ascent to carry payloads back to orbit
• Reusable Lander: Refuel the lander for multiple landing missions

Lander Performance Example

Launch Vehicle Delivers Lander to TLI; Lander Performs Orbit Insertion

Launch Vehicle Delivers Lander to Lunar Orbit; Lander Performs Landing Only

Reference Case Thru NRO

1000 kg

Reference Lander Thru LLO

2000 kg

Some Mission Performance Cases

Summary & Findings

• A viable lander concept has been developed that leverages cryogenic propulsion technologies
• Inclusion of cryo propulsion increases performance and generates flight data for future applications

​

• Active cryo fluid management supports significant mission flexibility
• Longer duration missions (hopping, return, reuse) will require active CFM
• More ambitious missions with higher ΔV budgets will benefit from the higher performance offered by LOX/LCH4 propulsion

​

• Mission flexibility and performance make this an appealing concept for commercial partners
• System supports a viable CFM demonstration mission

Future Work

• Structures and Configuration
• Examine load configurations with payloads on top of lander instead of “underslung” configuration
• Propulsion
• Refine design of electric pump-driven MPS including power storage & distribution
• Thermal
• Assess environmental heat loading for various loiter trades in LLO vs NRHO
• Power
• Assess alternative battery concepts for reducing battery mass
• Look at kits for alternative mission profiles w/ long-duration surface stays
• Avionics
• Look at kits for various mission profiles featuring AR&D
• CFM Demo Payload
• Trades on LLO vs NRHO testing periods
• Analysis Plans
• Extended portfolio analysis
• Mission Portfolio – Technology mapping exercise

​

​

BACK UP

Mission Modes

Launch

TLI

Moon

Surface

Earth

3200 m/s

2500

900

425

730

1800

730

DSG

DSG

Lunar Orbit

Lunar Orbit

LV/US

US/L

US/L

US/L

US/L

L

L

Potential Elements to Perform Maneuvers

LV = Launch Vehicle

US = Upper Stage

L = Lander

O = Other

3000

L/O

1800

L

Multiple Sites

Varying mission modes by incorporating other mission elements can free up lander propellant for alternative uses. Can be applied to carry additional payload or enable mission profiles with additional ΔV.

Propellant Transfer & TVS Demonstration

8 Week NRO Coast

• 4 Week Payload Active Cooling
• Transfer Demonstration

​

​

​

​

​

​

• 4 Week Payload Passive Storage
• Demonstrate Pressure Control
• Payload Tank at 75% Liquid Level
• Pump Based Mixing with Axial Jet or Spray Bar
• Thermodynamic Vent System (TVS)
• ~ 0.51 kg/day Propellant Loss

4 Week LLO Coast

• 4 Week Payload Active Cooling
• Transfer Demonstration

​

​

​

​

​

​

• Expel propellant from Payload prior to DOI burn

​

​

w/ Demo Payload if Available

AES Mid-Year Review April 2014

CFM Tech: Lander vs Demo Payload

LANDER CONCEPT:

• Two 1.84m Spherical LCH4 Tanks
• Two 1.84m Spherical LOX Tanks
• Long Duration Storage Required
• Actively Cooled

PAYLOAD CONCEPT:

• One 1.5m X 1.5m Cylindrical Tank with Elliptical Domes
• Working Fluid: Methane
• Utilizes Lander Cryocooler

CFM Tech: Lander vs Demo Payload

Test Objectives not Covered by Lander Concept:

• Propellant Tank Chilldown (#15)
• Thermodynamic Vent System (#19)
• Transfer Operations (#20)
• Effects of Scaling in micro-g
• Passive Storage

CFM Tech on Demo Payload:

• Helium Pressurization Capability (#5)
• High VAC MLI (#8)
• PMDs/LADs (#10)
• Low Conductivity Structures (#11)
• Pump Based Mixing (#16) with Axial Jet or Spray Bar
• Tube-On-Tank BAC (#22)
• Unsettled Mass Gauging (#23)
• Valves, Actuators, and Components (#24)
• Propellant Tank Chilldown (#15)
• Thermodynamic Vent System (#19)
• Transfer Operations (#20)
• Effects of Scaling in micro-g
• Passive Storage

CFM Tech Required for Lander Concept:

• Autogenous Pressurization (#2)
• Helium Pressurization (#5)
• High Eff & Cap 90K Cryocooler (#7)
• High Vac MLI (#8)
• PMDs/LADs (#10)
• Low Conductivity Structures (#11)
• Pump Based Mixing (#16)
• Tube-On-Tank BAC (#22)
• Unsettled Mass Gauging (#23)
• Valves, Actuators, and Components (#24)

CFM Tech Mapping

By baselining active CFM, we are able to future-proof the lander, enabling other fallout missions that would follow the first demo mission