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.

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

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.

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