Conceptual Design of a Crewed Lunar Lander

Extensible Design of a Lunar Lander Robert Guinness & Arthur Guest International Space University 28 August 06

Overview Introduction Internship Overview Our Starting Point Our Initial Concept Our “Second” Initial Concept Refining Our Initial Concept Creating Subsystem Sizing Tools Sizing our Concept Modifying our Concept The Final Result Recommendations Rob Arthur

Internship Overview Our Goal: To develop a conceptual lunar lander in parallel to NASA’s Lunar Lander Preparatory Study Learn about spacecraft systems engineering by creating subsystem sizing tools

Apollo Moon Landings of 1969-1972 were of profound significance for the human community and the evolution of our civilization. Motivation Although important fragments remain, the “Apollo Generation” has largely retired from active work, leaving current work to the current generation of space explorers. But the job was left unfinished… Neil Armstrong and Harrison Schmitt at NAC meeting ISU Students with Capt. John Young Apollo 10 Pilot, Apollo 16 Commander

NASA and Constellation Program Organization Manages Agency-wide VSE Program Manage Agency-wide Programs Level 1 Level 2 Level 3 This office not yet “stood up.”

Advanced Projects Office and Lunar Lander Pre-Project Providing an “international perspective” on the lunar lander design ?

LLPS Requirements Crewed Sortie Mission

LLPS Requirements Crewed Outpost Mission ~180 days (south pole)

LLPS Requirements Cargo Outpost Mission

“Conceptual Design” Subset of the propulsion tradespace

Methodology Three Cycles of Design : Size system elements. Study layout options. Refine subsystem sizing/perform trades. Choose layout and evaluate its strengths and weaknesses. Repeat. 2 3 1 Trajectory analysis also performed to validate staging concept. Layout Evaluate Size Refine

Our Starting Point We studied various designs and documents in order to get an overview of different lander concepts and what is involved in designing a concept ESAS LSAM – “The Baseline” Phase One Conceptual Designs Other Related Studies Eagle Engineering “Lunar Lander Conceptual Design” First Lunar Outpost AIAA Papers NASA Papers AIAA LaRC Phase 1 All images from public domain, except LaRC image used with permission

Concepts from LLPS Phase One Minimized Ascent Stage Staging of the LOI and descent maneuvers Cargo Access to the Surface Horizontal Landers Split Landers Images courtesy of LaRC, GRC, MSFC, and GSFC

Break lander up into its simplest, minimal functional parts At minimum, must bring crew to the surface and return them safely to orbit. Our Approach View additional capabilities as mission-specific “cargo elements” Sortie Missions : Surface habitat for crew of 4, ~7 days. Relatively small cargo capacity for rovers, science packages, etc. Outpost Missions : Larger, more capable habitat. Large cargo capacity for pressurized rovers, power stations, ISRU facilities, etc. Outpost mission supported by uncrewed, cargo version of lander.

Main Design Objectives Minimize Ascent Stage in order to maximize cargo capacity. Slope of sensitivity line 1.52 – 1.85 Maximize extensibility of sortie mission lander to outpost mission crew and cargo versions. Purpose is to minimize life-cycle costs. Minimize distance of crew and cargo from lunar surface. Reduces risk and exhaustion of crew for egress/ingress. Maximizes ease of cargo offloading.

The Truck Analogy Crew Cabin Generic Storage Fuel Engine Image of Mack Truck used with permission courtesy of Bob Young

Initial Results of Cycle 1 Vertically-oriented “truck” with minimal ascent stage on top of descent stage propellant tanks. “Truck bed” is oversized to be compatible with outpost missions (crew and cargo) with little redesign. “Sortie habitat” would be replaced with larger “outpost habitat” to be delivered on cargo mission. Kept cargo bed on lowest level for ease of unloading Crewed Sortie mission version

Refining Our Conceptual Design Transitioned from a vertically configured lander to a horizontally configured lander Moved the ascent stage from on top of the habitat to beside the habitat Raised the problem of configuring the stack so the thrust vector is always through the centerline (TLI, LOI, Descent) Needed to find a solution that allowed us to keep our design philosophy from the original concept (standardized stages) Solution: Use a separate stage for the LOI burn Habitation Module Ascent Stage

Refining Our Conceptual Design Use a modular approach to allow similarity in stages Lunar Capture & Initial Descent Stage (LCIDS) Final Descent & Landing Stage (FDLS) Lunar Ascent & Rendezvous Stage (LARS) Lunar Sortie Habitat Module (LSHM) Small Lunar Outpost Cargo Module (SLOCM) Large Lunar Outpost Cargo Module (LLOCM)    Cargo Outpost     Crewed Outpost     Crewed Sortie LLOC Module SLOC Module LSH Module LAR Stage FDL Stage LCID Stage

Refining Our Conceptual Design Our design allowed for investigation into the following: 2-Stage Descent using LCIDS (Lunar Capture and Initial Descent) Baseline is having the LCIDS do 75% of the descent burn. Finding the optimal portion of descent discussed in next section. Minimal Ascent Stage Horizontal Lander Additional trades to study: Inboard vs. outboard engines Ascent stage propellants: cryogenic vs. storable Central power system vs. separate power systems in each module Cryo-cooler vs. passive TCS Habitat size and shape

Descent Trajectory SORT used to model powered descent trajectories Assistance from Flight Mechanics Lab (Ron Sostaric and Ryan East) Gravity turn steering used as first cut guidance approach. Simulation constraints: 2 m/s descent rate at final approach altitude of 30 m Additional ~100 m/s delta-V allowed for vertical landing phase and hover. At LCIDS doing ~76% of powered descent, jettison takes place at about 8.4 km altitude. Jettisoned stage impacts ~2 km downrange. Will impact 42 sec prior to landing. What is the minimal safe separation distance? (More details in back-up slides.)

Strengths and Weaknesses of Cycle 1 Design Strengths: Large cargo space possible for outpost missions (crewed and cargo versions). Minimal modifications between different mission types Weaknesses: Access to and from the ascent stage requires the astronauts to cover a large vertical distance. Operational difficulty and issue of dealing with an incapacitated astronaut Complexity of the plumbing due to the separation between the propellant tanks and engines High of center of gravity Would be difficult to land on sloped terrain Could present landing stability issues

Subsystem Sizing Tools Examined and learnt about each subsystem Created sizing tools for the various components Used Microsoft Excel with Visual Basic We had to choose which components to examine in depth Structures 10 Command & Data Handling 9 Guidance & Navigation 8 Communications 7 Extra Vehicular Activities 6 Crew Accommodations 5 Environmental Control & Life Support 4 Power 3 Thermal 2 Propulsion 1

Sources of Information We used two approaches to learning about spacecraft subsystems We read textbooks and papers to get a top-down look Human Spaceflight Mission Analysis and Design SMAD, SPAD, etc. NASA Papers AIAA Papers We looked at the subsystems and components of other designs to get a bottoms-up look Apollo Shuttle ESAS LSAM Phase One concepts

Propulsion Subsystem Propellant Mass Main Trade - LOX/LH 2 or LOX/LCH 4 or MMH/NTO Calculated the propellant masses for various propellant types for each maneuver and stage Propellant Storage Tanks Sizes Created a tool to size the propellant tanks based on total propellant mass (including unusable and extra propellant mass for other uses) Designed from “Space Propulsion Analysis and Design” Boil-off and Cryo-coolers Calculated boil-off for propellant tanks based on average temperature of the environment Calculated mass of one-stage & two-stage cryo-coolers based on information from GSFC Pressurant System Sizing Feed System Sizing Engine and RCS Thruster Sizing Apollo RL-10 Engine Used with permission from NASA

Propulsion Trade 3268 kg 2888 kg Power-adjusted Total System Mass 0 kg ~30 kg Mass to produce Power for 99 days 0 W 78 W Power for Cryo-cooler 3268 kg 2858 kg System Mass Subtotal 2553 kg 1797 kg Total Usable Propellant 964 kg 257 kg Usable Fuel Mass 1590 kg 1540 kg Usable Oxidizer Mass 714 kg 1061 kg Inert Mass NTO/MMH LOX/LH2 Propulsion System Type

Thermal Subsystem Sized both active thermal control systems and passive thermal control systems Multi-Layer Insulation Radiators Developed from discussion from GSFC thermal team Based on overall heat load (electronics, crew, environment) Sized for horizontal OSR (second surface mirror) radiators Coldplates Point design from HSMAD Fluid Evaporator Systems Plumbing Multi Layer Insulation Used with permission from NASA - MSFC

Power Subsystem Split the subsystem into primary power source and PMAD (power management and distribution) components Focused on primary power source calculations Main Trade - Fuel Cells or Batteries Fuel Cell Sizing Calculations for sizing of hardware based on Shuttle Fuel Cells Mass and volume per kilowatt of electricity produce Calculations for amount of reactants needed and water produced based on efficiency and chemical reaction Battery Sizing Sized Lithium-Ion Batteries using unit masses from the ESAS LSAM and guidelines from HSMAD Power Management and Distribution Components Sized based on amount of power in the system Followed guidelines from HSMAD Space Shuttle Fuel Cell Used with permission from NASA

ECLS Subsystem Atmospheric Management Monitoring System Carbon Dioxide Systems Trace Contaminant Control System Oxygen and Nitrogen for leakage, re-pressurizations, respiration Water Management Calculated water requirements for the crew Used the fuel cell water produced to meet these requirements Fire Suppression and Detection Point masses, volumes, and powers for: Smoke Detectors Fixed Suppression System Portable Fire Extinguishers

Crew Accommodations Subsystem Food System Waste Collection System Personal Hygiene Operational Supplies Crew Health Care Sized based on Information from HSMAD Information from ESAS LSAM Sometimes there were discrepancies; for example: HSMAD toilet – 45 kg ESAS toilet – 25 kg Subsystem left similar to ESAS LSAM

Extra Vehicular Activities Subsystem EVA Suits Actual Suit Mass Suit Spares Maintenance Tools Airlock Sizing Pressure Vessel Atmospheric Management Support Systems Umbilical and Consumable Sizing Consumption based on values from HSMAD for an 8-hr EVA Extrapolated for umbilical sizing per hour Mark-III Lunar Surface Suit Photo used with permission from NASA

Avionics Subsystems Communications Subsystem Ka-Band S-Band Point design based on MSFC, GSFC, ESAS Guidance & Navigation Subsystem Point design based on MSFC, GSFC, ESAS Redundancy Command & Data Handling Point design based on MSFC, GSFC, ESAS Redundancy Examining the three other designs for avionics, showed discrepancies within NASA for various components. Avionics Components Photos used with permission of NASA

Structures Subsystem Most difficult subsystem to size Unique to this design Very difficult to size for any conceptual design Must either use parametric data or design a full structures system We used the parametric equations from NASA’s Design Mass Properties II Pressure Vessel Mass = 1.27 * (Surface Area) 1.15 Unpressurized Structures Mass = 0.71* (Surface Area) 1.15 The problem is that the parametric data is based on numerous spacecraft but only one lander (Apollo) and NO horizontal landers Also sized tank support structures, landing gear, LIDS, and windows.

Sizing our Conceptual Design kg 53599.95 kg 45000.52 kg 45000.16 Launch Mass kg 20654.14 kg 2140.92 kg 564.07 Cargo-Launched kg 566.76 Inert Mass LLOCM kg 7041.09 Inert Mass SLOCM kg 5314.02 Inert Mass LSH kg 2182.53 kg 2153.80 Propellant kg 3284.94 kg 3277.66 Inert Mass LARS kg 2986.70 kg 2176.46 kg 1796.54 Propellant kg 5036.57 kg 4484.38 kg 4485.02 Inert Mass FDLS kg 21265.41 kg 20371.06 kg 23737.88 Propellant kg 3090.38 kg 3319.15 kg 3671.16 Inert Mass LCIDS Cargo Outpost Mission Crewed Outpost Mission Crewed Sortie Mission

Conceptual Problems of Cycle 2 Design Major issue: Horizontal Center of Gravity during descent Side mounted ascent stage (~5.1 mT) led to difficulty of ensuring the center of gravity of the landed mass was in line with the thrust from the descent stage engines. Centerline of Thrust Ascent Stage 5.1 mT x 3.5 m = 17.9 mT-m Habitation Module 5.3 mT x 0.5 m = 2.7 mT-m

Modifying the Conceptual Design To solve the problem: FIRST STEP We moved the ascent stage towards the middle of the lander New Problem #1 Ascent engine and the descent engines overlapped . To solve this problem: Integrated the ascent engine into the descent engines and removed two of the descent engines. New Problem # 2 Positioning of our 5 meter long habitat Solution was to split the habitat into two parts 3 meter long habitation module and 2 meter long airlock Used the ascent stage to connect the two parts

The Final Result Horizontal lander with centrally-located ascent stage Cargo volume sized for outpost missions Sortie mission version shown

Recommendations Lunar Capture and Initial Descent Stage Reduces volume (and mass) of landed propellant tanks, which greatly enhances layout options and reduces overall height of components from ground. Combined descent and ascent propulsion systems Final descent stage small enough to allow commonality of engine type between descent and ascent. Single engine type reduces life-cycle costs Provides mass savings and packaging savings “Green propellants” for all stages LOX/LH 2 chosen for high performance One-stage cryo-cooler for descent stage, two-stage for ascent One-stage cryo-cooler makes hydrogen boil-off manageable with little power cost or mass penalty. Two-stage cryo-cooler makes all boil-off near zero; significant but manageable power cost.

Thank You Any Questions (or comments)?

Important SORT Parameters 35571 lbm Optimization Variable OPTNAM Landed mass 379 s Control INDNAM(3) Time of start of vertical landing 279 s Control INDNAM(2) Time of first burn shutdown 25.0 km Control INDNAM(1) Initial altitude 2 m/s Constraint DEPNAM(2) Final descent rate 30 m Constraint DEPNAM(1) Final height above surface 9854 kg Input SWGT(3) Total Propellant Mass 3671 kg Input SWGT(2) Inert mass of jettisoned stage 15794 kg Input SWGT(1) Landed mass (inert mass plus ascent stage and cargo) Mass Statement 10 seconds Event criteria TPHASE CRITR (@ Event 115) Time of free-fall after stage jettison -179.9 degrees Input DALPHA Guidance angle of thrust vector 0.66667 Input XKCMD (@ Event 115) Throttle command setting for first burn segment 0.92 Input XKCMD (@ Event 100) Throttle command setting for first burn segment 32000 lbf Input LTHR01 Maximum Thrust of Engines 460 seconds Input LSI01 Specific Impulse of Engines Value Units Type Variable Name Parameter

Descent Trajectory: Full and Close-up

Time Plots of Important Parameters

Time Plots of Important Propulsion Parameters

Conceptual Design of a Crewed Lunar Lander

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