Introduction

Assuming your spacecraft is not a freaking torchship, a Mars expedition with the entire spacecraft landing then lifting off is going to demand about ten times as much delta V than it has to spend. This is why pretty much all NASA designed crewed Mars missions have the main spacecraft loiter in orbit while the explorers use a tiny Mars Excursion Vehicle (a "lander") to ferry them to and from the surface.

A noteable exception is the Hercules single-stage reusable.

The landers here assume that the planet they are visiting are wilderness worlds, that is, they do not have local starports equipped with booster rockets or anything like that. A couple of Mars expedition designs try to edge around that. They have prior unmanned missions to land robot factories utilizing the Sabatier reaction that manufacture rocket fuel from the Martian atmosphere. This wonderfully lowers the delta-V requirements.

The Lunar landers listed here will also probably work on any airless body in the solar system, with the possible exception of the planet Mercury. That planet has the dubious honor of having the highest orbital velocity of all the airless bodies. This means Mercury is the most delta-V costly world to land/lift-off from. The planets with more gravity than Mercury have an atmosphere suitable for aerobraking, providing free delta-V.

The Mars landers will work on Mars, but no guarantees on them working with any other planet. Most of them require aerobraking, so they only work on planets with atmospheres. And the planets with more gravity than Mars require more delta-V for lift-off than the landers have.

Turning to some science fiction speculation, an exploration starship with a huge on-board power plant might assist their landers. The mothership can use large lasers to send power to the landers to help with landing and lift-off.

THREE SHIP TYPES

The traveling-public gripes at the lack of direct Earth-to-Moon service, but it takes three types of rocket ships and two space-station changes to make a fiddling quarter-million-mile jump for a good reason: Money. The Commerce Commission has set the charges for the present three-stage lift from here to the Moon at thirty dollars a pound. Would direct service be cheaper?

A ship designed to blast off from Earth, make an airless landing on the Moon, return and make an atmosphere landing, would be so cluttered up with heavy special equipment used only once in the trip that it could not show a profit at a thousand dollars a pound! Imagine combining a ferry boat, a subway train, and an express elevator.

So Trans-Lunar uses rockets braced for catapulting, and winged for landing on return to Earth to make the terrific lift from Earth to our satellite station Supra-New York.

The long middle lap, from there to where Space Terminal circles the Moon, calls for comfort—but no landing gear. The Flying Dutchman and the Philip Nolan never land; they were even assembled in space, and they resemble winged rockets like the Skysprite and the Firefly as little as a Pullman train resembles a parachute.

The Moonbat and the Gremlin are good only for the jump from Space Terminal down to Luna . . . no wings, cocoon-like acceleration-and-crash hammocks, fractional controls on their enormous jets.

From SPACE JOCKEY by Robert Heinlein (1947)

Aeronutronics MEM

Aeronutronic Mars Excursion Module

This is from Proceeding of the Symposium on Manned Planetary Missions 1963/1964 Status.

It is used on the 10-Meter Mars Mission Orion. The mission carries two of these, the preferred "tail-sitter" version. The "canted" version has problems, and doesn't fit as well on the Orion. It may also have been used on the Mars Expedition Spacecraft.

Sadly the design assumed a Mars surface atmospheric pressure of 85 millibars. The discovery by the Mariner 4 probe that the actual value was one tenth of this invalidated the design. This is discussed by David S. F. Portree, where he talks about the successor to the Aeronutronics MEM.

The Aeronutronic MEM was sized for a 40 day stay on the Martian surface with three explorers.

The fuel was a devil's brew of the appallingly corrosive, toxic, and carcinogenic monomethylhydrazine (MMH) mixed with the ever-popular but beyond-insanely-dangerous FLOX. At least it is a re-startable rocket. MMH is hypergolic with any oxidizer, and FLOX is hypergolic with anything.

The reason for this fuel is they needed a specific impulse of at least 375 seconds, but liquid hydrogen fuel just takes up too much blasted room. The designers of the successor to the Aeronutronics MEM had the same problem, so they were forced to use FLOX as well.

CIRA/AAS EAGLE

This is from Project Moonlight, a modular lunar mission profile proposed to NASA in 2006 by an Italian team consisting of engineers and scientists from the Italian firms of Alcatel Alenia Space (AAC) of Rome, CIRA of Capua Italy, and the Microgravity Advanced Research and Support (MARS) Center in Naples. The AAC team was lead by Luciano Miccichè, the CIRA team was lead by Gennaro Russo, and the MARS team was lead by Dr. Giuseppe De Chiara.

Project Moonlight proposes a modular lunar exploration architecture that uses many smaller launch vehicles and does not require a heavy lift launcher such as the NASA VSE Ares V CaLV. The proposal relies on on-orbit cryogenic propellant transfer and component dockings in both LEO and Lunar Orbit.

While the Moonlight architecture requires more initial launches of smaller components, it aims to provide a reusable lunar transportation infrastructure consisting of Earth-to-Moon Tugs, reusable landers, and Crew vehicles. This should significantly reduce the cost of recurring lunar operations, instead of launching an entire infrastructure for each lunar mission as NASA's VSE intends.

The lunar infrastructure eventually consists of a small space station in lunar orbit, called the Lunar Orbital Operations Platform (LOOP) logistical station. The lunar station is largely built using existing ISS modules, such as the MPLM for a habitat, a Z1 element with gyros, a docking port node, and an ISS-style truss.

The lunar station would hold multiple docked reusable landers, fuel transported from unmanned Lunar Tanker modules, and provide a staging area for the crew vehicle.

The final phase of the proposal would use the lunar transportation architecture and the reusable modular design of the lunar lander to land crew habitation modules and other cargo and facilities, forming a permanent or semi-permanent moon base.


The two designs that are landing spacecraft are the EAGLE (European Advanced Gear for Lunar Exploration) and the Lunar Outpost.

There is a variant design, the EAGLE C which is an uncrewed cargo version of the EAGLE. The EAGLE C is basically an EAGLE with the control compartment replaced with a large cargo port and loading ramp.

Yes, the EAGLE does bear a slight resemblance to the fictional Space 1999 Eagle Transporter. But so does NASA's LUNOX lander proposal. Not to mention the Scorpion. Form follows function, and all are belly landers. So given the same task, it is not surprising that all of the designs resemble each other.

The EAGLEs and the Lunar Outpost used hypergolic fuels, specifically Unsymmetrical dimethylhydrazine (UDMH) + nitrogen tetroxide (N204 or "NTO").


EAGLE

Since the design was relatively simple, I couldn't resist trying to model it in Blender 3D. It wasn't too hard, though there were a few difficult spots. Mine is not as nice as Marco Gavazzeni, but he has more talent and probably spent more time on it than I did.


Lunar Transportation System

The EAGLE was originally created by Dr. Giuseppe De Chiara for his graduate thesis in 1996. Called the LTS (Lunar Transportation System) project, it had a crewed lander (4 crew) ferried to and from Lunar orbit by an unmanned orbiter. The crew was retrieved in LEO by the STS.

Dr. Chira's LTS was re-used in the Italian Industrial "Moonlight Scenario" in 2006, as a response to NASA's Constellation program. Dr. Chira was professionally involved in the development.


Transitional Versions

These diagrams show some transitional steps as Dr. Chira evolved the LTS into the Eagle.


EAGLE C

The uncrewed cargo version of the EAGLE. The control cabin is replaced by a large cargo hatch and loading ramp.


EAGLE LUNAR OUTPOST

Lunar Outpost
Crew4 to 8
Length10 m
Width13.5 m
Height5 m
Habitable Volume200 m3
Payload10,000 kg
Total Mass38,000 kg
Power Systemsolar panels
Propellanthypergolic
Enginesx2

Eagle Engineering Lunar Lander

EAGLE ENGINEERING LUNAR BASE SYSTEMS STUDY

This study looked at the problem of building a lunar lander to support a small lunar surface base using one lander that could either land 25 mt (28 t), one way, or take a 6-mt (7-t) crew capsule up and down. The initial idea was to build a reusable lander, suitable for minimizing the transportation cost to a permanent base, and use it from the first crewed mission on, taking some penalty and perhaps expending expensive vehicles early in the program to avoid building multiple types of landers while focusing the effort on a space-maintainable, singlestage, reusable vehicle. A four-engine design for a multi-purpose vehicle, with total thrust in the range 155,688 to 177,929 N – 53,379 to 57,827 N per engine – and a throttling ratio in the 13:1 to 20:1 range was proposed. Initial work indicated a regeneratively cooled, pumpfed engine would be required due to difficulties with regenerative cooling over wide throttling ranges with pressure-fed systems. Three cases of interest were studied. The first scenario assumed the lunar lander was used only to place a payload on the surface and was called the “Cargo Down” case. In this case, the lander did not have propellant to ascend to orbit after delivering its payload; it, therefore, stayed on the lunar surface until refueled. The second case also placed a payload on the surface, but it carried enough propellant to return its inert mass to orbit, and was called the “Inert Returned” case. The third scenario described a case in which the lunar lander carried a crew module down to the surface and then back to orbit. This case was called the “Crew Module Round Trip.”

All three scenarios focused on a single-stage, reusable lander using nitrogen tetroxide/monomethyl hydrazine (N2 /O4 MMH) propellants. While the N2 /O4 lander is considerably heavier than an LO2 /LH2 lander in the previous section, it is much smaller, due to higher propellant density. However, features in both landers are essentially the same. The propellant capacity of either version of the lander was 35 mt (38.6 t) divided into four tanks of 16 m3 (565 ft3) each. The tank diameter was 2.5 m (8.2 ft) for all tanks.

Important features included the following:

  • An airlock/servicing tunnel down the center of the lander to allow easy access on the surface and pressurized volume for Line Replaceable Units. Many engine connections could be made and broken inside the pressurized volume.
  • A removable crew module. The lander was flyable without the crew module.
  • The lander fit in a 9-m (30-ft) heavy-lift vehicle shroud with landing gear stowed.
  • The landing gear had electromechanical shock absorbers.
  • Emergency ascent with one or two crew members was possible without the crew module. In that case, the crew would ride in suits in the airlock/servicing tunnel.
The figure shows this lander being serviced on the lunar surface and illustrates how the airlock/servicing tunnel allowed pressurized access to a surface vehicle. An engine is being removed in the figure.

Mass breakdowns are included below for the multipurpose versions of all three cases, using both LO2 /LH2 and N2 /O4 /MMH propellants.

Mass Breakdown – Multi-Purpose Lander Using LO2 / LH2 Propellant
Cargo DownCrew Module
Round Trip
Inert Returned
Delta-v, Ascent (km/s)02.28*2.28*
Delta-v, Descent (km/s)2.102.102.10
kgkgkg
Structure1,6811,6811,681
Engines822822822
RCS Dry411411411
Landing Systems784784784
Thermal Protection2,0172,0172,017
Tanks3,0253,0253,025
Data Management
System/GN&C
150150150
Electrical Power**478478478
Airlock/Tunnel455455455
Inert Mass9,8239,8239,823
Ascent Prop.011,3347,240
Descent Prop.22,59718,13720,486
Unusable Prop. (3%)678884832
FPR Prop. (4%)9041,1791,109
Usable RCS858689778
Unusable RCS433439
FPR (20%)172138156
Total Propellant Mass25,25232,39530,640
Deorbit or Gross Mass
(less Payload)
35,07542,21840,463
Payload, Descent25,0006,00014,000
Payload, Ascent06,0000
Deorbit or Gross Mass
(with Payload)
60,07548,21854,463
Mass Breakdown – Multi-Purpose Lander Using N2O4/MMH Propellant
Cargo DownCrew Module
Round Trip
Inert Returned
Delta-v, Ascent (km/s)02.28*2.28*
Delta-v, Descent (km/s)2.102.102.10
kgkgkg
Structure1,9551,9551,955
Engines956956956
RCS Dry478478478
Landing Systems912912912
Thermal Protection1,0061,0061,006
Tanks1,5091,5091,509
Data Management
System/GN&C
150150150
Electrical Power**478478478
Airlock/Tunnel455455455
Total Inert Mass7,8997,8997,899
Descent32,86130,66531,927
Ascent015,7029,406
Unusable (3%)9861,3911,240
FPR Prop. (4%)1,3141,8551,653
Usable RCS990923961
Unusable RCS504648
FPR (20%)198185192
Total Propellant Mass36,39950,76745,427
Deorbit or Gross Mass
(less Payload)
44,29858,66653,326
Payload, Descent25,0006,00014,000
Payload, Ascent06,000*0
(Inert Mass
returned to LLO)
Total Mass at Deorbit69,29864,66667,326
* Delta-v = 1.85 + 0.43 km/s for a 15-deg plane change in a 93 km circular orbit.
** Electrical power provided for 3 days only, (2 kW). 100% redundant fuel cells/tank sets.

Hercules Single-Stage Reusable

Hercules Single-Stage Reusable
EngineChemical
Methalox
Diameter5.99 m
Height
(starting at engine exit plane)
17.83 m

This is from Hercules Single-Stage Reusable Vehicle supporting a Safe, Affordable, and Sustainable Human Lunar & Mars Campaign (2017), Hercules Single-Stage Reusable Vehicle (HSRV) Operating Base (2017), Long-Term Cryogenic Propellant Storage on Mars with Hercules Propellant Storage Facility (2017), and Lunar and Mars Ascent and Descent/Entry Crew Abort Modes for the Hercules Single-Stage Reusable Vehicle (2018).

This is a concept designed to support future Lunar and Mars campaigns aimed at establishing self-sustaining human presence beyond Terra orbit. Amazingly this is a mere chemically-powered rocket which is both single-staged and reusable. It also has features allowing full coverage aborts during liftoff and landing from either Luna or Mars, which will bring a smile to everybody's face.

Re-usability is a game-changing feature, which most rocket companies grudgingly admit after SpaceX has rubbed their nose in it, multiple times. With non-resuable Mars landers, every outgoing Mars spacecraft will have to lug along a fresh lander. Which will savagely cut into the spacecraft's payload, making it difficult to grow the Mars base. But with Hercules, you just have to transport one or two of the landers. The rest of the spacecraft visits can have payloads that are 100% base infrastructure.

The other sine qua non of the industrialization of space is in-situ resource utilization. In this case, it mainly means using the magic of the Sabatier reaction to convert the Martian atmosphere into rocket fuel. This will give the The Tyranny of the Rocket Equation a brutal kick in the gonads with steel-shod boots, the fondest wish of all rocket designers. So the Hercules will use methane-oxygen rockets, even though it only has 3,700 m/s exhaust velocity, instead of LH2/LOX's 4,400 m/s. But rocket designers don't care. They will gladly pay the 700 m/s performance hit in exchange for seeing the Tyranny writhing in agony moaning "OW! My Balls!" the vastly increased payload capacity.


The Hercules Transportation System will be a family of vehicle configurations, built in the same framework (the technical term is "outer moldline" or OML).

  • HMTV: Hercules Mars Transfer Vehicle [cannot land]
    • Interplanetary Crew
  • HPDV: Hercules Payload Delivery Vehicle [cannot land]
    • Interplanetary Cargo
  • HSRV: Hercules Single-Stage Reusable Vehicle [lander]
    • Mars Cargo
    • Mars Crew
    • Lunar Cargo
    • Lunar Crew
  • HCRV: Hercules Crew Rescue Vehicle [section that rockets crew to safety in case of abort]

The outer moldline was designed to allow aerodynamic entry into the Martian atmosphere.

In addition to the main engines at the base (called the Ascent/Descent System or ADS), there are secondary engines at the top attached to the crew compartment (called Abort/Terminal Landing System or ATLS). These are canted 30° outboard from vertical (cosine thrust loss reduces thrust to 87%). The secondaries are used for aborts and for terminal landing. Design-wise you want the abort engines attached to the crew module so it can propel the module away from the rest of the vehicle. And you want the landing engines angled away from the surface to avoid landing in a self-made crater and/or sand-blasting equipment already on the ground (the report calls it avoiding "Surface/Plume Interaction").


HSRV Lander (Lunar Crew Configuration)

Lunar Crew Configuration
SubsystemDescent Mass
(kg)
Ascent Mass
(kg)
Predicted Mass14,27214,272
Structures5,0885,088
Thermal Protection00
Landing Legs & Actuation1,0361,036
Ascent Propellant Tank1,9361,936
Descent Propellant Tanks318318
ADS Propellant Feed418418
ADS Engines277277
ATLS Propellant Tanks725725
ATLS Propellant Feed543543
ATLS Engines511511
RCS Thrusters114114
Power & Avionics850850
Growth/Margin2,4452,445
Propellant Mass125,83842,350
Ascent Tank105,69922,211
Descent Tanks10,50110,501
ATLS Tanks9,6389,638
Payload Mass5,5005,750
Cargo (Crew Habitat)5,0005,000
Crew & Suits500500
Samples0250
TOTALS145,61062,372

The HSRV can deliver up to 20 metric tons of cargo or a crew of 4 to the lunar surface from the Deep Space Gateway (in Near Rectilinear Halo Orbit around Luna), assuming the DSG supplies the cargo/crew and tops off the propellant tanks.

The Lunar HSRV does not have a thermal protection heat shield. That is only used for aerobraking, and there ain't no air around the Moon.

Design Features

  • Primary Structure: composite; sections designed to separate
  • Thermal Protection: not required
  • Propellant Architecture: liquid oxygen and liquid methane (O2/CH4) common across complete system; interconnected propulsion systems

Nose Section

  • Abort/Terminal landing System (ATLS)
    • CH4 Tanks: 2 x 2 m diameter at 500 psia
    • O2 Tanks: 2 x 2 m diameter at S00 psia
    • Engines: 8 pressure-fed engines installed with 30° cant angle delivering ~60 kN each at min effective Isp 300 sec.
    • RCS Thrusters: 12 pressure-fed thrusters
  • Power Generation: 2 internal combustion engines burning O2/CH4 delivering 3 kWe at idle, 40 kWe at max throttle
  • Habitat Adapter: supports 5.5 t crew hopper habitat
  • Crew Support: standard docking system; pressurized tunnel to habitat

Payload Section

  • Crew hopper habitat supporting 4 crew for 3 days
  • Available Volume: 109 m3
  • Door Clearance: 4.5 m wide x 3.8 m tall

Ascent/Descent Section

  • Ascent Tank: common-bulkhead storing O2/CH4 at 30 psia
  • Descent lanks: dedicated for terminal descent propellant
    • CH4 tanks: 2 x 2 m diameter at 30 psia
    • O2 Tanks: 2 x 7 m diameter at 30 psia
  • Ascent/Descent Engines: 1 pump-fed engine installed at vehicle base delivering ~245 kN at min effective Isp of 360 sec
  • Body Flap & Actuation: primary control for atmospheric flight
  • Landing legs: deployable/retractable

HPDV Uncrewed Interplanetary Cargo Delivery

The HPDV can deliver 40 to 60 metric tons of cargo to Low Mars Orbit using a Hohmann transfer and aerobraking at Mars. The spacecraft departs from the DSG, having been loaded with cargo and propellant tanks topped off.

The first cargo will be a small space station called an "orbital node." This will be be a place to accumulate subsequent cargo shipments, and as a crew transfer point. The node will be in a 500 km circular LMO at an inclination allowing access to the selected surface base site. The uncrewed node will have multiple docking ports and have autonomous or semi-autonomous robots for in-space assembly and servicing. These robots are used to:

  • construct and maintain the node
  • facilitate capture, berth and dock of incoming vehicles
  • facilitate transfers of payloads between the HPDV (cargo vehicle) and the HSRV (crewed vehicle)
  • facilitate propellant transfers from the HSRV to various vehicles at the node or to the node itself

The orbital node will also have a large propellant tank used to accumulate methane and oxygen transported from the Mars surface Sabatier factory.


HMTV Interplanetary Crew Transport

The HMTV delivers a crew of 4 to Low Mars Orbit using a 90 to 120 day fast-transfer with aerocapture as Mars. This reduces the crew's exposure to galactic cosmic rays, Hohman transfers take from 180 to 300 days. The spacecraft departs from the DSG, having been loaded with cargo and propellant tanks topped off.


HSRV Lander (Mars Cargo Configuration)

The cargo HSRV can deliver 20 metric tons of cargo from the orbital node to the Mars surface base. At the base it unloads, refuels and travels back to the orbital node. Upon arrival it will have 5 metric tons (5,000 kg) of propellant remaining. This will be added to the orbital node's supply, used to refuel new arrivals.


HSRV Lander (Mars Crew Configuration)

Mars Crew Configuration
SubsystemDescent Mass
(kg)
Ascent Mass
(kg)
Predicted Mass18,89818,898
Structures5,7015,701
Thermal Protection2,0802,080
Landing Legs & Actuation1,0361,036
Ascent Propellant Tank1,9361,936
Descent Propellant Tanks318318
ADS Propellant Feed474474
ADS Engines1,3831,383
ATLS Propellant Tanks725725
ATLS Propellant Feed543543
ATLS Engines511511
RCS Thrusters114114
Power & Avionics850850
Growth/Margin3,2163,216
Propellant Mass138,17112,959
Ascent Tank121,7140
Descent Tanks8,2866,949
ATLS Tanks8,1716,010
Payload Mass5,5005,750
Cargo (Crew Habitat)5,0005,000
Crew & Suits500500
Samples0250
TOTALS162,56937,607

The crew HSRV can deliver 4 crew from the orbital node to the Mars surface base. At the base it unloads and refuels. At the end of their stay, the 4 crew travels back to the orbital node. Upon arrival it will have 4 metric tons (4,000 kg) of propellant remaining. This will be added to the orbital node's supply, used to refuel new arrivals.


Design Features

  • Primary Structure: composite; sections designed to separate
  • Thermal Protection: mechanically-attached ACC6 hot structure with opacifed fibrous inslation
  • Propellant Architecture: liquid oxygen and liquid methane (O2/CH4) common across complete system; interconnected propulsion systems

Nose Section

  • Abort/Terminal landing System (ATLS)
    • CH4 Tanks: 2 x 2 m diameter at 500 psia
    • O2 Tanks: 2 x 2 m diameter at S00 psia
    • Engines: 8 pressure-fed engines installed with 30° cant angle delivering ~60 kN each at min effective Isp 300 sec.
    • RCS Thrusters: 12 pressure-fed thrusters
  • Power Generation: 2 internal combustion engines burning O2/CH4 delivering 3 kWe at idle, 40 kWe at max throttle
  • Capsule Adapter: supports 5.5 t separable crew capsule
  • Crew Support: standard docking system; pressurized tunnel to habitat

Payload Section

  • Crew hopper habitat supporting 4 crew for 3 days
  • Available Volume: 109 m3
  • Door Clearance: 4.5 m wide x 3.8 m tall

Ascent/Descent Section

  • Ascent Tank: common-bulkhead storing O2/CH4 at 30 psia
  • Descent lanks: dedicated for terminal descent propellant
    • CH4 tanks: 2 x 2 m diameter at 30 psia
    • O2 Tanks: 2 x 7 m diameter at 30 psia
  • Ascent/Descent Engines: 5 pump-fed engines installed at vehicle base delivering ~245 kN each (1,225 kN total) at min effective Isp of 360 sec
  • Body Flap & Actuation: primary control for atmospheric flight
  • Landing legs: deployable/retractable

Human Lunar Return Lander

The Human Lunar Return study was desperate attempt to put astronauts back on Luna at a bargain-basement price. It was one of the last gasps of NASA's Faster-Better-Cheaper design approach.

Yes, I can understand NASA's panicked need for space projects that cost only a few hundred million dollars instead of billons. But if the Apollo Lunar Excursion Module was an antarctic exploration tractor, the Human Lunar Return lander was a unicycle. Blasted thing didn't even have a hull, it was basically a rocket engine with two chairs welded on the top. The contraption is only slightly more sophisticated than the North American Rockwell Moon Hopper.

1996-HUMAN LUNAR RETURN

HLR Option C. This alternative fallback option eventually evolved into HLR's final 1996 baseline after NASA Administrator Dan Goldin criticized the comparatively high cost of the effort at a November 1995 briefing to the Administrator. Goldin wanted "incredible breakthroughs" costing at most a few hundred million dollars rather than billions. The HLR team responded by focusing on bare-minimum lightweight concepts such as Option C's "small lander". Anything that wasn't absolutely necessary (e.g. the LEO cryo fuel depot or lunar orbit station) was to be deleted.

To save weight, HLR would use an unpressurized open-cockpit lunar landing vehicle weighing just 4,565kg with fuel. The vehicle is 3.9 meters tall and 5.6 meters wide. The space-suited crew of two receives oxygen and other life support consumables via umbilicals from the LLV. In the illustration here, arrows indicate foot restraints and ladder. [What a ride that would be!]

The Lunar Orbit Stage (formerly known as Transtage) is protected by a 9.144-meter diameter aeroshell, which is launched in seven segments to save space. The aeroshell is assembled before rendezvous with ISS and then moored to the Space Station. A second Shuttle flight delivers the crew and propellant for the lunar vehicles. The refueling operations are simplified since the LOS and LLV utilize storable hypergolic propellants, which require no new propellant transfer technologies. The 15.6-tonne LOS vehicle only carries enough propellant for lunar orbit insertion and trans-Earth injection; two expendable 20-tonne propulsion modules (derived from the Russian "Breeze" upper stage and launched on two Proton rockets) perform the translunar injection burn. The LOS carries a small unpressurized Lunar Landing Vehicle and a 2.5 meter long Command Module capable of supporting two astronauts for up to 19 days during the Earth-Moon transfer. Date of departure from ISS: August 24, 2001.

Human Lunar Return Lander: 9508-HLR-1

NASA Administrator Dan Goldin initiated the “Human Lunar Return” (HLR) study in September 1995 to investigate innovative, fast-track approaches for crewed spaceflight. The HLR team worked through two initial concepts in an effort to produce the ultimate cut-rate faster-better-cheaper human lunar mission. The 1996 baseline design was a bare-minimum lightweight concept in which anything not absolutely necessary (e.g., the LEO cryo fuel depot or lunar orbit station) was deleted. It consisted of a Lunar Orbit Stage (LOS), the Lunar Landing Vehicle (LLV), and the Habitat.

The LOS was protected by a 9.144-m diameter aeroshell since it would be aerobraking back into LEO when returning from the Moon. The shell was to be launched in seven segments to save space, assembled on orbit, and moored to the International Space Station (ISS) pending integration with the Lunar Orbit Stage. A Space Shuttle flight would deliver the crew and propellant for the lunar vehicles to the ISS, and the LOS and LLV, together with the crew, would depart for the Moon. On the return trip, the LOS would again dock with the ISS after which the crew was to return to Earth via the Space Shuttle.

The 15.6-mt LOS vehicle carried only enough propellant for lunar orbit insertion and trans-Earth injection; two expendable 20-mt propulsion modules (derived from the Russian “Breeze” upper stage and launched on two Proton rockets) performed the Trans-Lunar Injection (TLI) burn. The LOS carried a small, unpressurized LLV and a 2.5-m long Command Module capable of supporting two astronauts for up to 19 days during the Earth-Moon transfer.

The open-cockpit LLV weighed just 4,565.3 kg, including fuel, and was 3.9 m tall by 5.6 m wide. The space-suited crew of two received oxygen and other life support consumables via umbilicals from the LLV. After landing, they were to live in the inflatable Surface Habitat that had been delivered prior to their arrival. Following departure from the lunar surface, the crew reboarded the LOS and the LLV was jettisoned prior to trans-Earth injection.

Baseline for the Expendable Crewed Lander
Vehicle
  • Open cockpit sized for two crew in Extravehicular Mobility Unit (EMU)/Portable Life Support System (PLSS)
  • LOX servicing at pad pre-launch; vent and servicing interfaces integrated into upgraded non-toxic Orbiter
Structures
  • Truss frame structure, composites where feasible
  • Single-stage, four-leg landing gear with load attenuation
  • Cockpit frame/payload box provides structural interface to PLSS
Propulsion
  • Single-stage pressure-fed LOX/RP1
  • 14,679 N, 4:1 throttling main engine w/no gimbal
  • 200 N, 6 DOF Reaction Control System (RCS)
Guidance
Navigation
& Control
  • Auto-rendezvous and Auto-Land w/Redesignation
  • Daytime hazard detection (shadow based)
  • Star Tracker, Deep Space Network, Inertial Navigation System,
    Laser altimeter, Beacon, LIght Detection And Ranging (LIDAR)
Power
  • Two low-mass Proton Exchange Membrane (PEM) fuel cells (load sharing/redundant)
Avionics /
Communication
  • Non-Commercial Off-the-Shelf reduced mass
  • Computer, S-Band, Ultra-high Frequency, video (descent and ascent)
Life Support
  • New EMU with amine swing bed CO2 removal
  • EVA resources via umbilical during ascent/descent
Thermal
  • Multi-Layer Insulation, Passive with heaters/radiators
Mission
  • dV’s: Descent = 1,910 m/s, Ascent = 1,822 m/s; RCS = 40 m/s
  • Land and surface operations during lunar day only
Mass Breakdown
Item(kg)
Primary Structure404.8
Payload Box/Seat28.8
Landing Gear99.3
Propulsion252.5
TPS/Protection113.5
Power125.7
Avionics120.1
Life Support56.0
STAGE INERT1,200.7
Residual Propellant (3%)98.0
LOX Boil-off42.8
FINAL STAGE INERT1,341.5
Usable Propellants3,223.8
GROSS STAGE MASS4,565.3
2 Crew + EMU/PLSS423.3
Payload50.0
TOTAL LUNAR LANDING VEHICLE
MASSTO LOW EARTH ORBIT
5,038.6

Mars DRA 5.0 Lander

This is from Human Exploration of Mars Design Reference Architecture 5.0, Addendum #2.

The Mars Design Reference Architecture needed standard Mars landers, so the study authors brain-stormed many concepts and did extensive comparisons using various metrics to winnow out the top designs. I am only going to cover the top couple of landers, you can read all about the others in the report.

The lander is composed of:

  • Aerobraking Method: either a rigid aeroshell or an inflatable heat shield. This burns off as much velocity as possible by using Mars atmosphere instead of costly rocket fuel
  • Mars Descent Module (MDM): in either horizontal or vertical configuration. A rocket platform used to gently land the payload.
  • Payload: see here. There is one payload inventory for the uncrewed precursor landing and one for the crewed landing.

The two major differences that distinguish the designs are Aerobraking Method and Payload Orientation.

Aerobraking method can be either Mid lift-to-drag ratio aeroshell or Hypersonic inflatable aerodynamic decelerator.

Payload orientation can be either Horizontal Configuration or Vertical Configuration.

So there are four possible combinations: mid-L/D horizontal, mid-L/D vertical, HIAD horzontal and HIAD vertical. The orginal DRA used mid-L/D horizontal.


AEROBRAKING: MID-LIFT-TO-DRAG RATIO (mid L/D)

Lift-to-drag ratio (L/D) is the amount of lift generated by a vehicle, divided by the aerodynamic drag it creates by moving through the air. It tells you if your aircraft has the glide capacity of a brick.

A high L/D aircraft is like an aircraft that can glide for miles. A low L/D aircraft is like a brick, meaning it has trouble flying at all. The Apollo command module has a low L/D, it just falls ballistically and hopes it slows to parachute speeds before the heat shield erodes away.

For a Mars mission, where every gram counts, designers figured they needed to encase the lander in a rigid aerodynamic shell with a medium L/D. A high L/D shell eats up too payload mass because the shell weighs too much. A low L/D shell cannot slow the lander enough so it goes splat! and makes a smoking crater. A mid L/D is just right. Basically the aeroshell acts like a lifting body.

The lander is packed in the mid L/D aeroshell. This deorbits and aerobrakes, while steering to the designated landing site. Once it burns off enough velocity, the aeroshell is shed and discarded like a banana peel. The lander proper ignites its rocket engines and uses them to soft-land on Mars.

Advantages over HIAD: very volume efficient in the booster rocket faring since it is almost an inner layer of the faring; aeroshell can be steered right to the desired landing site.

Disadvantages compared to HIAD: HIAD has much less penalty mass compared to an aeroshell;


AEROBRAKING: HYPERSONIC INFLATABLE AERODYNAMIC DECELERATOR (HIAD)

Aerobraking by using heat shields is tried and true technology. The trouble is that [a] heat shields are heavy which eats into your payload mass, and [b] since the Martian atmosphere is so darn thin you need an extra-wide heat shield, which multiplies the heat shield penalty weight.

The designers try to do an end run around those two problems by making the heat shield an inflatable balloon. A remarkably fireproof balloon, but a balloon none the less.

Advantages over mid L/D: has much less penalty mass since it is a balloon.

Disadvantages compared to mid L/D: deflated balloon takes up lots of room in booster rocket faring; it basically cannot be steered.


PAYLOAD ORIENTATION: HORIZONTAL CONFIGURATION


PAYLOAD ORIENTATION: VERTICAL CONFIGURATION


PAYLOAD

Uncrewed precursor landing

  • Mars Descent Module (MDM): This is the rocket propelled platform that lowers the rest of the payload to a soft landing on Mars. Since it would be more or less impossible to equip it with chemical engines cranking out enough delta V to land unassisted from orbit, some sort of aerobraking has to be used.
  • Mars Ascent Vehicle (MAV): Small rocket propelled vehicle that boosts the Mars astronauts back into orbit at the end of the surface exploration part of the mission.
  • ISRU plant: for certain choices of chemical fuel used by the MAV it will be possible to manufacture said fuel from the Martian atmosphere by using in-situ resource utilization equipment. As always ISRU dramatically increases payload mass since you don't have to lug all that heavy fuel from Terra. If such fuel is not used, the ISRU plant is removed from the payload inventory. Its mass is assumed to be included in the MDM mass, and it is carried as disassembled parts inside the MDM body.
  • Fission Surface Power System (FSPS): nuclear reactor supplying electricity to the crew habitat and ISRU plant.
  • Crew Surface Mobility Vehicle (Rover): some sort of Mars groundcar used to extend the range of the crew's exploration. From the illustrations it appears to be some species of NASA Space Exploration Vehicle (SEV)

Crewed landing

  • Mars Descent Module (MDM): as above.
  • Explorer Crew: four to six, depending upon the design. Study assumes each crew is 98.5 kg (largest US male, per SSP 50005). 4 crew is 394 kg; 6 crew would be 591 kg
  • Crew Habitat: habitat module for the explorers to live in for the duration of the surface mission.

MARS ASCENT VEHICLE (MAV)

The ascent vehicle transports the Mars surface explorers up to the orbiting spacecraft at the end of the surface segment of the mission. In some designs it is transported to Mars without the heavy oxygen fuel, relying on the ISRU unit to generate it from the Martian atmosphere using power from the nuclear reactor.

The study looked at a 4 crew and a 6 crew ascent vehicle. To get an idea of the lower limit, they looked at a 4 crew "taxi" version. This was from a prior lunar ascent vehicle and was a spartan bare-bones design. The main difference is that the 4 and 6 crew have a life support system, the taxi does not. The 4 and 6 allow the crew to do a pre-flight check out in a shirt-sleeve environment, and use the MAV as a supplemental/back-up habitat module. Not so the Taxi, the crew will have to do check out and fly while wearing suits.

The 4 crew has a 2.35 meter inner diameter, the 6 crew has 2.7.

North American Rockwell MEM

North American
Rockwell
Mars Excursion Module
Diameter9.6 m
Height8.8 m
Deorbit Stage
Mass
3,357 kg
Descent Stage
Mass
29,207 kg
Ascent Stage
Mass
16,874 kg
Total Mass
large mission
49,437 kg
Mission Parameters
Mothership
Circular
Orbit
500 km
Mothership
Elliptical
Orbit
300 km by
66,900 km
e = 0.9
Small Mission2-crew/4-day
Large Mission4-crew/30-day
Deorbit Stage
PropulsionChemical
(beryllium solid)
Isp325 s
Thrust133,500 N to
204,600 N
Burn Time48 seconds
Total Mass3,357 kg
Deorbit ΔV200 m/s
Deorbit T/W0.4
Descent Stage
PropulsionChemical
(FLOX/CH4)
Isp383 s
Descent
ΔV
1,070 m/s
(mom circular) or
1,450 m/s
(mom elliptical)
Descent
Thrust
623,000 kg
Descent
T/W
1.5 to 0.15
MASS SCHEDULE
Jettisoned Structure2,109
Retained Structure2,880
Lab Structure612
EPS1,016
Communication168
Guidance &
Control
5
ECLSS739
RCS1,193
Landing Gear1,256
Science Payload1,905
Contingency1,393
Tanks & System1,179
Engine916
Propellant13,835
TOTAL DESCENT
STAGE
29,207
Ascent Stage
PropulsionChemical
(FLOX/CH4)
Isp383 s
Ascent ΔV4,880 m/s
(mom circular) or
6,200 m/s
(mom elliptical)
Ascent Thrust156,000 kg
Ascent T/W1.0
RCSChemical
(ClF5/MHF-5)
Rendezvous
ΔV
100 m/s
Rendezvous
T/W
(RCS)
MASS SCHEDULE
ASCENT CAPSULE(2,386 kg)
Structure445 kg
EPS227 kg
Communication95 kg
Guidance &
Control
102 kg
ECLSS608 kg
RCS240 kg
Return Payload
(Geological samples)
136 kg
Crew
(90 percentile)
318 kg
Contingency215 kg
STAGE II(4,277 kg)
Tanks & System313 kg
Engine222 kg
Propellant3,742 kg
STAGE I10,210 kg
Tanks & System730 kg
Propellant(9,480 kg)
TOTAL ASCENT
STAGE
16,874 kg

The Mars Excursion Module is from a 1966 study by North American Rockwell. This was the first Mars lander designed after the bombshell from Mariner 4 that astronomers had drastically over-estimated how dense the Martian atmosphere was. They had figured it was a useful 85 hectopascals (hPa), in reality it was an almost worthless 6 hPa (just slightly better than a vacuum). By way of comparison Terra's atmospheric pressure at sea level is 1013 hPa.

The poor prior design that was rendered obsolete by the low atmospheric pressure was the Aeronutronic MEM

The low atmospheric blow Mariner 4 dealt to the scientist was just the cherry on top of the sundae. Much more serious was the photographs. The scientists knew there could be no chance of images of scantily-clad Barsoomian princesses, but they were hopefull there would be some lakes and maybe even a canal or two. But nothing but a bunch of freaking craters? The scientists got a sinking feeling in their stomachs, as they could almost see the Mars exploration program go swirling down the toilet right before their very eyes. Once the taxpayers saw these photos the NASA tax dollars would dry up. Mars looks like the freaking moon, for cryin' out loud! And NASA has already been to the moon. Been there, done that, got the T-shirt. No need to go to Moon part deux.

But NASA put a brave face on things, and proceeded to plan for a Mars mission anyway. Sadly, they were right. As I write this it is fifty years later and the movie The Martian is still science fiction, not a documentary. That furry "whumph" noise you hear is RocketCat doing a facepalm.


Given the pathetic whisp of Martian atmosphere, NAR went with a classic gum-drop shape much like the Apollo command module for aerobraking purposes. For one thing all the expertise obtained from Apollo could be leveraged. Plus there was Terra's atmosphere conveniently located for heat shield test purposes.

Though I did read a recent report suggesting that even with the gum-drop design the Martian atmosphere is not up to the task of aerobraking the MEM before it splats into the ground at hypersonic velocities. The report suggested that entirely new technologies are needed.


In a genius move, NAR made the design modular. If you needed a lean and mean mission, you could remove some internal compartments, ascent propellant, and surface supplies to get the total lander mass down to 30 metric tons. Or you could max it out. Or anything in between.

The price of a low mass lander is that it could only support two crew for four days, and the mothership had to be in a low circular Mars orbit for both departure and return to the mother. The high mass lander needed lots more delta V from the mothership, but it could support four crew for thirty days, and the mothership could be in a high elliptical Martian orbit.

You can also make a lander with no ascent stage at all. This can be used to land supplemental equipment, such as an extended-stay shelter, nuclear power module, or a huge Mars mobile lab with fuel supply.


  1. MEM has a mass of 49,437 kg when it separates from the mothership. The deorbit motors fire for about 200 m/s ΔV. Deorbit motor has a thrust-to-weight-earth of 0.4. The MEM starts falling out of orbit, and the deorbit motors are jettisoned. The MEM now has a mass of 46,078 kg.

  2. The MEM enters the Martian atmosphere at an angle of attack of 147°. It starts aerobraking, subjecting the crew to about 7 g's. When it slows to a mere Mach 3.5, it pops a hypersonic drogue chute to stabilize then inflates a 18 meter diameter ballute. This will slow the MEM down to Mach 1.5.

  3. Once the MEM lowers to 3 kilometers of the surface, it jettisons the ballute. The plug in the heat shield over the descent engine is jettisoned. The descent engine is canted about 13° off center, because the MEM center of gravity is off center, because due to design consideration the MEM is not radially symmetric. Mostly because of that pesky crew quarters and laboratory.

  4. The conical section of the heat shield is jettisoned. The descent engine ignites and burns for 1,070 m/s to 1,450 m/s ΔV, depending upon whether the mothership was in a circular or elliptical orbit when the MEM detached. At this point the engine will have a thrust-to-weight-earth of 1.5 to 0.15.

  5. The design managed to squeeze in enough extra propellant for about two minutes of hovering (about 457 m/s of ΔV). Which could be a life-saver if the landing site unexpectedly turned out to be full of jagged boulders or something. Instead of hoving, the extra propellant can move the MEM laterally about 6.7 kilometers to an alternate landing site. The design had six landing legs. In concert with the incredibly stable gum-drop shape, they could manage a ground slope of up to 15°. Actually the shape is similar to a no-spill coffee mug, and for the same reason.

  6. The crew then frantically does as much Martian scientific research as they can cram into 30 days. The pressurized volume is 21.6 m3 (14.4 is laboratory/living quarters, 7.2 is ascent capsule). 20% is taken up by equipment, leaving barely 4.3 m3 per crewperson (right at the ragged limit before claustrophobia strikes).

  7. When it is time for departure, the descent stage becomes the launch pad (which stays behind on Mars), and the center becomes the ascent stage. It brings the crew and 136 kilograms of Martian geological samples back to the orbiting mothership. It launches as Ascent Stage I.

  8. When the Stage I tanks run dry, they are jettisoned. The ascent stage continues as Ascent Stage II. The two stages have a combined ΔV of 6,200 meters per second. The ascent has five components.
    1. Initial burn to 19 kilometer altitude (mothership circular: 4,206 m/s ΔV, mothership elliptical: 4,286 m/s ΔV)
    2. Coast to 185 kilometer altitude
    3. Burn to circularize orbit (23 m/s ΔV)
    4. At appropriate time, burn to ascend for rendezvous (mothership circular: 168 m/s ΔV, mothership elliptical: 1,327 m/s ΔV)
    5. Rendezvous with mothership at apoapsis
    6. Total ΔV: mothership circular 4397 m/s, mothership elliptical 5,635 m/s

  9. The ascent stage docks with the mothership using its Reaction Control System (RCS). It has 100 meters per second of ΔV left for the docking at this point. The rest was burnt during the descent and ascent phases.


The deorbit stage uses a beryllium solid rocket fuel with a specfic impulse of 300 to 325 seconds, a thrust of 133,500 to 204,600 Newtons, and a burn time of 48 seconds.

The reaction control system was supposed to use Chlorine pentafluoride (ClF5) oxidizer with Mixed Hydrazine Fuel-5 (MHF-5). The latter is a devil's brew of monomethylhydrazine, unsymmetrical dimethylhydrazine, diethyline triamine, acetonitrile, and hydrazine nitrate. Which is just as vile as it sounds. It has a specific impulse of 336 seconds.

The space shuttle used a more modern mix of monomethylhydrazine fuel with nitrogen tetroxide oxidizer. Still toxic but nowhere near as bad. It also has a specific impulse of 336 seconds.


The one joker in the deck was the specified fuel for the descent and ascent stages. It seems they couldn't quite get the ΔV they needed out of conventional liquid oxygen (LOX) and liquid hydrogen (LH2). With the mass ratio the design had (i.e., the tiny fuel tanks), LOX/LH2 could not even manage the 4,880 m/s ΔV required to reach the mothership in a low circular orbit, much less the 6,200 m/s ΔV required if it was in a high elliptical orbit. The problem was that LH2 takes up a lot of room, but the MEM's fuel tanks are cramped. There wasn't room for enough LH2 even if the entire area was converted into one gigantic tank.

So they used FLOX and liquid methane (CH4) instead. That can do 6,200 m/s ΔV easy because liquid methane is more than six times as dense as liquid hydrogen. So you can cram six times as much liquid methane mass into the same sized tanks. Using FLOX instead of LOX makes up for the lower energy in methane. FLOX/CH4 has a specific impulse of 383 seconds, compared to LOX/LH2's specific impulse of 449 seconds. LOX/CH4 is lucky to get a pathetic 299 seconds.

What is FLOX I hear you ask? Why, just a simple mixture of liquid oxygen, and Liquid Fluorine.

Fluorine is beyond insanely dangerous. It is incredibly toxic, and will corrode almost anything (some explosively). They don't call it "The Gas of Lucifer" for nothing. The pious hope of the MEM designers was to contain the FLOX in tanks lined with nickel or something similar that would form a passivation layer.

The FLOX mix is 82.5% fluorine and 17.5% oxygen. Mixing liquid fluorine and liquid oxygen is actually relatively safe. For some odd reason those two will not chemically combine without some coaxing. If they do combine, however, you get the dreaded compound Dioxygen Difluoride. This is the compound with the chemical formula FOOF, which coincidentally is the sound your laboratory will make as it blows up. This is the most famous compound in Derek Lowe's hysterical list of Things I Won't Work With (take a minute to read it, the article is a scream).

Another concern is that in a tank the fluorine and oxygen might separate. Then the engine would periodically be sucking pure fluorine, which certainly will not be doing the engine any good.

Carrying entire tankfuls of ultra-corrosive flaming explosive death to Mars seems to be a questionable decision, to say the least. If the MEM lands a trifle hard and the tanks rupture, you won't have the basis for a re-make of The Martian. More like a large melted crater with a few odd pieces of corroded metal and polished skeleton bits at the bottom.

At least the MEM designers saved mass on the ignition system. You don't need any. FLOX/CH4 is hypergolic (because fluorine is hypergolic with almost anything). This is also a help when the ascent stage does staging, you can easily re-start the engines in mid-flight.


I'm doing more research, but apparently the MEM design is so popular, that it was later redesigned just a bit to remove the need for liquid fluorine oxidizer. This would involve removing equipment and increasing the size of the fuel tanks.

North American Rockwell PEM

These are from Technological Requirements Common to Manned Planetary Missions: Appendix D by the space division of North American Rockwell (1965). They detail a s family of Planetary Excursion Modules (PEM).

Here I will focus on the retrobraking PEMs, that is, the ones that use retrorockets to land because the target planets have no atmosphere to allow aerobraking. These particular PEMs were intended for landing on Ceres, Vesta, Ganymede and Mercury. They are based on the Apollo Lunar Excursion Module with an Ascent Stage on top of a Descent Stage. They are designed for a 30 day stay on the planetary surface, before returning to orbit.

They do, however, have Rockwell's fixation on using the insanely dangerous FLOX as an oxidizer, and the only somewhat dangerous Monomethylhydrazine (MMH) as fuel. The reason is that liquid hydrogen requires preposterously huge tanks, but if you use anything else the specific impulse goes way down. Unless you take the mad-scientist step of using FLOX oxidizer to make up for it.

In the diagrams below, the Ascent Stage is pink, the Descent Stage is green, and the mission crew quarters is gold.

3-Crew Ceres-Vesta PEM

North American
Rockwell
3-Crew Ceres-Vesta PEM
Total
Wet Mass
14,036 kg
Descent Stage
Descent
Dry Mass
9,000 kg
Descent
Wet Mass
11,350 kg
Descent
Propellant Mass
2,350 kg
PropulsionChemical
FLOX/MMH
Propulsion Isp333 s?
Descent Engine
Thrust
8000 N
(835 kg-force)
Descent Engine
Mass
1.85 kg
Ceres
Gravitational
Acceleration
0.28 m/s2
Descent Start
Acceleration
2.0 Ceres
gravities
Ascent Stage
Ascent
Dry Mass
2,280 kg
Ascent
Wet Mass
2,686 kg
Ascent
Propellant Mass
406 kg
Ascent Engine
Thrust
3,780 N
(384 kg-force)
Ascent Engine
Mass
0.85 kg
Ascent Cabin
Volume
6.3 m3
Ascent Cabin
Diameter
2.44 m
Ascent Cabin
Long
1.83 m
Crew Quarters
Volume
14.0 m3
Crew Quarters
Diameter
2.44 m
Crew Quarters
Wide
3.96 m
Airlock
Volume
1.1 m3
Airlock
Diameter
0.915 m3
Crew Quarters
Volume
(inclu. airlock)
15.1 m3
Total Crew
Volume
21.2 m3
Consumables
(30 day, 3 crew)
MMH54.4 lbs/ft3
FLOX90.0 lbs/ft3
Nitrogen12.30 ft3
295 kg
Oxygen10.00 ft3
325 kg
Water8.86 ft3
244 kg
Food5.63 ft3
175 kg

The mission crew quarters is merged with the ascent stage. Note the Descent Stage Dry Mass does not include the mass of the ascent stage.

10-Crew Ceres-Vesta PEM

North American
Rockwell
10-Crew Ceres-Vesta PEM
Descent Stage
Descent
Dry Mass
18,100 kg
Descent
Wet Mass
kg
Descent
Propellant Mass
4,700 kg
PropulsionChemical
FLOX/MMH
Propulsion Isp333 s?
Descent Engine
Thrust
N
(1,675 kg-force)
Descent Engine
Mass
3.7 kg
Ceres
Gravitational
Acceleration
0.28 m/s2
Descent Start
Acceleration
Ceres
gravities
Crew Quarters
Volume
m3
Crew Quarters
Diameter
m
Crew Quarters
Height
2.14 m
Airlock
Volume
1.41 m3
Airlock
Diameter
0.915 m3
Crew Quarters
Volume
(inclu. airlock)
66.0 m3
North American
Rockwell
10-Crew Ceres-Vesta PEM
Total
Wet Mass
27,800 kg
Ascent Stage
Ascent
Dry Mass
4,530 kg
Ascent
Wet Mass
kg
Ascent
Propellant Mass
815 kg
Ascent Engine
Thrust
N
(760 kg-force)
Ascent Engine
Mass
1.68 kg
Ascent Cabin
Volume
11.2 m3
Ascent Cabin
Diameter
3.06 m
Ascent Cabin
Long
1.94 m
Total Crew
Volume
3
North American
Rockwell
10-Crew Ceres-Vesta PEM
Consumables
(30 day, 10 crew)
MMH54.4 lbs/ft3
FLOX90.0 lbs/ft3
Nitrogen980 kg
Oxygen1,080 kg
Water815 kg
Food590 kg

A page appears to be missing from my copy of the document

3-Crew Ganymede PEM

North American
Rockwell
3-Crew Ganymede PEM
Wet Mass17,300 kg
Descent Stage
Descent
Dry Mass
9,500 kg
Descent
Wet Mass
kg
Descent
Propellant Mass
3,740 kg
PropulsionChemical
FLOX/MMH
Propulsion Isp333 s?
Descent Engine
Thrust
27,000 N
(2,800 kg-force)
Descent Engine
Mass
6.15 kg
Ganymede
Gravitational
Acceleration
1.428 m/s2
Descent Start
Acceleration
Ganymede
gravities
North American
Rockwell
3-Crew Ganymede PEM
Ascent Stage
Ascent
Dry Mass
2,280 kg
Ascent
Wet Mass
kg
Ascent
Propellant Mass
3,010 kg
Ascent Engine
Thrust
13,000 N
(1,340 kg-force)
Ascent Engine
Mass
2.95 kg
Ascent Cabin
Volume
6.3 m3
Ascent Cabin
Diameter
2.44 m
Ascent Cabin
Long
1.83 m
Crew Quarters
Volume
14.93 m3
Crew Quarters
Diameter
2.44 m
Crew Quarters
Wide
3.96 m
Airlock
Volume
1.55 m3
Airlock
Diameter
0.915 m3
Crew Quarters
Volume
(inclu. airlock)
m3
Total Crew
Volume
21.2 m3
North American
Rockwell
3-Crew Ganymede PEM
Consumables
(30 day, 3 crew)
MMH54.4 lbs/ft3
FLOX90.0 lbs/ft3
Nitrogen12.30 ft3
295 kg
Oxygen10.00 ft3
325 kg
Water8.86 ft3
244 kg
Food5.63 ft3
175 kg

The mission crew quarters is merged with the ascent stage.

10-Crew Ganymede PEM

North American
Rockwell
10-Crew Ganymede PEM
Descent Stage
Descent
Dry Mass
28,100 kg
Descent
Wet Mass
kg
Descent
Propellant Mass
7,500 kg
PropulsionChemical
FLOX/MMH
Propulsion Isp333 s?
Descent Engine
Thrust
55,000 N
(5,550 kg-force)
Descent Engine
Mass
12.3 kg
Ganymede
Gravitational
Acceleration
1.428 m/s2
Descent Start
Acceleration
Ganymede
gravities
Crew Quarters
Volume
66 m3
Crew Quarters
Diameter
6.1 m
Crew Quarters
Height
2.14 m
Airlock
Volume
m3
Airlock
Diameter
0.915 m3
Crew Quarters
Volume
(inclu. airlock)
m3
North American
Rockwell
10-Crew Ganymede PEM
Total
Wet Mass
45,800 kg
Ascent Stage
Ascent
Dry Mass
4,536 kg
Ascent
Wet Mass
kg
Ascent
Propellant Mass
6,050 kg
Ascent Engine
Thrust
26,000 N
(2,670 kg-force)
Ascent Engine
Mass
5.9 kg
Ascent Cabin
Volume
11.2 m3
Ascent Cabin
Diameter
3.3 m
Ascent Cabin
Long
1.98 m
Total Crew
Volume
m3
North American
Rockwell
10-Crew Ganymede PEM
Consumables
(30 day, 10 crew)
MMH54.4 lbs/ft3
FLOX90.0 lbs/ft3
Nitrogen985 kg
Oxygen1,090 kg
Water815 kg
Food590 kg

The mission quarters is merged with the descent stage, instead of the ascent stage as is the case with the 3-crew vehicles.

LUNOX Phoenix

LUNOX MASS BREAKDOWN
Uncrewed Lander
ItemMass
(kg)
Phoenix Uncrewed Lander (dry)4,717
Cargo Package12,454
Propellants16,578
Trans-Lunar Injection Stage (dry)6,130
Propellants43,930
TOTAL MASS IN LOW EARTH ORBIT83,809

LUNOX is a concept designed by engineer Kent Joosten of the Johnson Space Center in 1993. You can find more details here, here, and here.

The basic idea was to harness the awesome might of In Situ Resource Utilization (ISRU) to dramatically cut the cost of establishing a lunar outpost.

At the time hydrogen was not believed to be readily available on Luna, but oxygen certainly was. Freaking Lunar regolith was 45% oxygen, the stuff was everywhere! There would be a fantastic cost savings if spacecraft could refuel with both hydrogen and oxygen (instead of having to transport the fuel needed to return home). But even refueling just with oxygen alone (and lugging along the extra hydrogen needed to burn it) would reduce the cost by a whopping 50%! Such is the grim mathematics of Every Gram Counts.

OK, so there is lotsa oxygen there. Can it be easily extracted? Engineer Joosten pointed to no less than fourteen methods of lunar oxygen (LUNOX) extraction are known. He favored a hydrogen ilmenite reduction technique patented by the US-Japanese Carbotek/Shimizu consortium. Joosten figured that an automated/teleoperated plant using solid-state high-temperature electrolysis could produce 24 metric tons of cryogenic liquid oxygen per year (assuming an extraction efficiency of 4% after benefication). This would require somewhere between 40 and 80 kilowatts of continuous electricity, which argues for small nuclear reactor. Such a reactor could also charge up teleoperated mining drones and supply astronauts with power when they were on site.


The plan was for a series of four uncrewed landers (on one-way trips) to deliver the lunar outpost elements. The landers are belly-landers instead of tail-sitters, in order to:

  1. Reduce the chance of toppling over and destroying itself and the cargo when landing

  2. So that later astronauts can access the cargo without needing a crane with a freaking 18 meter long cable

Flight one delivered the oxygen extractor plant and the nuclear reactor. The reactor was on a teleoperated cart so the radioactive thing could be moved far far away from the plant. The plant just stayed in the lander.

Flight two would deliver teleoperated Loaders (to dig up regolith and transport it to the oxygen plant), Tankers (to transport the liquid oxygen produced), and Haulers (for the astronauts to use, when they arrive). All the teleoperated vehicles are electrically powered, periodically recharged by the reactor.

The loaders can collect and sort about 500 kilograms per hour of ilmenite-rich regolith.

Once the plant has produced enough liquid oxygen to refuel a crewed mission (so that the crewed mission has enough fuel to return the crew home), flights three and four will deliver equipment needed by the astronauts. There is no sense going to the expense of sending the astronaut gear if it turns out the oxygen plant or something is broken and cannot refuel the crewed mission.

LUNAR OUTPOST ELEMENTS
Flight Item Mass (kg)
1Lunox Plant7,269
1Nuclear Reactor (40-60kWe)5,110
2Tanker #11,471
2Tanker #21,471
2Loader #11,728
2Loader #21,728
2Hauler #1962
2Hauler #2962
3Pressurized Rover #15,150
4Pressurized Rover #25,150
3Mobile Power Unit #11,544
4Mobile Power Unit #21,544
3, 4Science Payload2,000
5Airlock/Node Support Vehicle11,010
6Logistics & Spares12,454
Total:59,553

LUNOX MASS BREAKDOWN
Phoenix Crewed Lander
ItemMass
(kg)
Apollo-type Crew Module5,935
Crew and Support
(x4 crew)
609
“Phoenix” Crewed Lander
(4 x 31,150 KN thrust engines)
5,505
Cargo2,000
Trans-Lunar Injection Stage
(3 x RL-10-A4 engines)
6,130
ELEMENTS20,179
Crew Module Propellant199
Lander Propellant16,944
Liquid Hydrogen Fuel for Return Trip2,492
Trans-Lunar Injection Stage Propellant43,930
PROPELLANT63,565
TOTAL MASS IN LOW EARTH ORBIT83,744
Lunar Oxygen Required10,165
TOTAL MASS, INCLUDING LUNAR OXYGEN93,909

When it is assured that the ISRU site has mined enough oxygen to refill the Phoenix spacecraft's tanks, the four astronauts board the ship and head for Luna. The Phoenix is boosted into orbit and uses a Trans-Lunar Injection Stage to send it into a trajectory to the moon. After it lands, a teleoperated tanker rover transports 10,165 kilograms of LUNOX to the crewed ship and refills its empty oxygen tanks.

Yes, the Phoenix does resemble the fictional Eagle Transporter from the TV show Space 1999. It even has the Eagle's downward pointing chin bubbles. But this is more a case of "form following function", certainly NASA engineers didn't crib spaceship design features from a scifi TV show.

The crew can now explore Luna in style, with the tons of equipment transported to the site. The two pressurized rovers can provide power, communication, thermal control, life support, and habitation for 14 days on the lunar surface. They have a minimum exploration range of several hundred kilometers. Each rover obtains mobile power and water for life support from a mobile power unit. These power units are trailers containing fuel cells. The cells produce water and electricity by combining LUNOX with hydrogen brought from Terra.

But the big advantage is the ISRU has cut the project cost in half.

After the mission and departure from the Lunar surface, upon approaching Terra the conical crew section separates and reenters the atmosphere by aerobraking much like an Apollo Command Module (which is essentially what it is). Joosten recommends designing a steerable parasail-type parachute so that the capsule can come to rest on dry land instead of the ocean, because Apollo-style splashdown and Navy water recovery is so hideously expensive. The rest of the spacecraft is aimed at a remote spot of the ocean and (mostly) burns up like a huge meteor.

Atomic Rockets notices

This week's featured addition is ANTIMATTER-DRIVEN SAIL

This week's featured addition is PROJECT MALLAR

This week's featured addition is THE HABITANK CONCEPT

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