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Landing humans on Mars: Virtual prototyping of a piloted Lander

Mark D. Paton

Finnish Meteorological Institute

DRAFT 06.06.07

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Contents

Introduction..............................................................................................................3

1. A history of piloted Mars Landers .......................................................................4

1.1 Thinking big...................................................................................................4

1.2 Rise of the robots ...........................................................................................6

1.3 The case for Mars...........................................................................................9

1.4 End of the space race ...................................................................................10

1.5 Reference Missions: a question of mass? ....................................................12

1.6 Walkabout ....................................................................................................18

2. How on Earth do we land on Mars?...................................................................20

2.1 Aerocapture..................................................................................................21

2.2 Entry, Descent and Landing.........................................................................24

2.2.1 Entry, steering and g levels.................................................................24

2.2.2 Descent and the hypersonic transition problem..................................27

2.2.3 Landing and pin-point targeting..........................................................33

3. Design and test flying a virtual Mars Lander.....................................................35

3.1 The Mars Society paper ...............................................................................34

3.2 Building on a mission architecture ..............................................................36

3.3 The Orbiter space flight simulator ...............................................................37

3.4 Virtual Prototyping a piloted Mars Lander in Orbiter .................................39

3.5 Software for rapid prototyping of EDL systems: A2D................................45

3.6 Aerocapture vehicle design and aerodynamic stability ...............................49

3.7 Entry, Descent and Landing System design ................................................54

3.8 MLAM and the effect of Martian winds on a piloted Lander .....................64

4. Conclusions........................................................................................................68

4.1 More Mars Landers......................................................................................70

4.2 Lessons learned: A piloted Mars Lander v2.0 .............................................72

References..............................................................................................................74

Links ......................................................................................................................76

Acknowledgements................................................................................................77

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Introduction

Landing humans on Mars will be an expensive, technologically difficult, risky and drawn

out affair. According to the NASA Roadmap Team it will take 25 years to develop

human scale demonstration missions. When eventually the crewed space ship departs, the

inward and outward leg of the journey to Mars will still take six months apiece; with a

one and a half year stop over on Mars. At Mars the piloted (i.e. with humans aboard)

Lander will descend to the surface using an ugly amalgamation of technologies. It will be

a scary nail biting experience, taking only a few minutes to transform the Lander from a

aerodynamic hypersonic entry vehicle to a rocket powered, legged Lander. The Martian

base will probably be manned for a year and a half before Mars and Earth are aligned so

they can return to Earth. In the meantime the crew have, what appears to be, an infinite

and desolate desert for company. To keep down the mass launched from Earth and the

associated cost, they'll live off the land, making their own rocket fuel for the return flight

back to Earth. Still the total expense will most likely be well over 10 billion US dollars

even with every cost cutting trick applied to the mission. How can people understand and

support such a seemingly expensive and pointless adventure? One way to educate people

about the complexities of human space travel, and its excitement, is to use the increasing

power of computers, software and the internet.

A few intrepid space explorers have taken up the challenge and implemented various

human mars mission add-ons in a popular, free space flight simulator called Orbiter,

developed by Dr. Martin Schweiger at University College London. It has a good

Newtonian physics engine and a good graphics engine, with close to photo realistic

renderings in some cases. It has proved to be an effective tool in communicating the

technical complexities of a human Mars mission, with two virtual mission projects

presented at the Mars Society Conference in August 2006. One of these presentations was

by Bruce Irving with a paper called “Virtual Prototyping of a Human Mission to Mars”.

Things are evolving and now Orbiter is being used by Bruce, a recently created JPL Solar

System Ambassador as part of his mission to communicate the complexities of space

exploration to the general public. A contributor to the virtual prototyping (VP) paper,

Andrew McSorley has also moved on and is busy prototyping and iterating NASA’s

DRM 3.0 in Orbiter with participation of the Orbiter community. Even Dr. Robert Zubrin

took time out to sit down with MSC presenters and Orbiteers Seth Hollingsead and Cyrus

Phillips to watch his mission Mars Direct flying in the virtual world!

This article focuses on the building of the piloted Mars Lander model for Bruce’s VP

paper by Mark Paton, and the design of its Entry, Descent and Landing trajectory

together with some further work. Mark Paton is a postdoc at the Finnish Meteorological

Institute, currently researching the Martian atmosphere. A presentation of this work was

made at the International Space Development Conference, Dallas, TX, in May 2007.

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1. A history of piloted Mars Landers

This section describes past and present mission plans beginning in the nineteen-fifties and

finishing with some recent initiatives. Many human Mars missions have been studied but

those only with details of Mars Landers are included in this section. Mars Landers have

evolved over time, as more has been learnt about the atmosphere of Mars and new

technologies have become available. The mission objective has changed from (mostly) a

simple excursion to the surface, like with the Apollo LEMs on the Moon, to full scale

exploration of the surface complete with a base able to sustain the crew for up to one and

a half years and pressurised rovers to travel large distances. The political environment

and the money available for space exploration have also had a significant effect on our

plans to send humans to Mars. Consequently this makes for an altogether different kind

of Lander (although there are superficial similarities) than that proposed in the early days

of the space race.

1.1 Thinking big

One of the first serious plans to send humans to Mars was attempted by Wernher von

Braun (Braun, 1952). Das Marsproject described a scheme to send a flotilla of ten 4000

mT ships with a 70 strong crew to Mars. To construct the fleet 950 winged ferry flights

flew the parts into Earth orbit. Three of the ships carried a 177 mT glider for landing on

Mars. The gliders had huge swept wings that allowed an extended glide path reaching

half way around the planet if necessary. The maximum g force from entry into the

atmosphere was calculated to be 0.12 g and a maximum temperature of 649 K. One of the

Landers had to land on a smooth surface at the pole using skids. The crew then trekked to

somewhere around the equator and build a runway to allow the remaining wheeled

gliders to land on Mars. Von Braun based the design of his Landers on a model of the

Mars atmosphere with a surface pressure ten times higher than that measured by robotic

Landers (i.e. by Viking etc). The Martian atmosphere is very thin and consequently the

landing speed for these gliders would be very high. It would be like trying to land the

Space Shuttle on a runway at 40 km altitude on Earth. In Braun’s architecture the gliders

would then be stood on their tails for launch back into orbit after a 400 day stay on Mars.

Total time for the mission would be about 3 years.

Between 1953 and 1959 the US Army Ballistic Missile Agency studied a mission concept

using electric propulsion to send humans to Mars (e.g.Stuhlinger, 1954). Stuhlinger lead

the investigation and published his first paper on the subject in 1954 (who had like von

Braun and other rocket experts been rounded up in Germany at the end of World War II,

see figure 1.2). The 1962 version of the mission involved the use of five ships, each being

150 m long, 360 mT in mass and with a crew of three. A nuclear reactor drove electric-

propulsion thrusters, taking 56 days to spiral out of Earth orbit followed by a 146 day trip

to Mars journey and then another 21 days to spiral down to Mars. Artificial gravity at one

tenth of Earth’s would be generated by spinning the ship. The ship also included a 50 mT

radiation shelter. Three of the five ships each carried a 70 mT Lander. Two of the

Landers carried cargo, one being a backup incase the first failed. The other carried the

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crew to the surface. First a cargo Lander landed followed by the crewed Lander for a 29

day stay on the surface. The 1957 version of Stuhlinger’s mission was portrayed by

Disney in their TV show Walt Disney’s “Wonderful World of Color” in an episode called

Mars & Beyond (See figure 1.1). Here a Mars Lander was used that entered the Martian

atmosphere nose first. A parachute was used to brake its fall followed by a rocket

powered descent. The vehicle had short stubby fins to land on. The Lander is reminiscent

of recent rocket shaped Lander designs such as the Case for Mars II study in 1984 or

more recently NASAs Design Reference Mission 3.0 from 1997. However the DRM 3.0

biconic designs are based on nuclear weapon delivery systems probably not around or in

a very early stage of development in the 1950s.

Jackson and Hammock of the Manned Spacecraft Center (now the Johnson Space Center)

presented a Mars mission study in 1963 (Jackson and Hammock, 1963). Their mission

comprised of two ships one of which was crewed and housed a piloted Mars Excursion

Module (MEM). The other ship was an unmanned Earth Return Vehicle (ERV). The

ERV would be launched 50-100 days before the crewed ship and take 200 days to reach

Figure 1.2 Rocket men. In the foreground is

Hermann Oberth an influential rocket pioneer. On

the right is Wernher Von Braun and on the left is

Ernst Stuhlinger. At the back, in military uniform, is

General Holger Toftoy who was responsible for

rounding up German rocket experts and at the back

in the suit is Eberhard Rees Deputy Director of the

Development Operations Division at the Army

Ballistic Missile Agency. Image credit: NASA.

Figure 1.1 A Mars Lander. This design was

featured in Walt Disney’s TV feature Mars and

Beyond. The rocket entered the atmosphere nose

first and used the parachute for braking. The final

part of the descent used a rocket engine. Image

credit: Walt Disney.

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Mars. The crewed ship would travel using a faster trajectory taking only 120 days to

reach Mars. This meant they would overtake the ERV and land on Mars in the MEM,

spending 10 to 40 days on the surface. Then as the ERV flies by Mars the crew launch in

the ascent stage of their MEM and catch up with the flyby ERV. They would then enter

Earth's atmosphere in an Apollo type command module.

Aeronutronic studied a MEM design for the Center’s mission (Franklin, 1963). Their

design utilised a lifting body to minimise the amount of propellant required and reduce

the mass. The vehicle was 9 m long with a mass of 30 mT. The maximum temperature

during entry would be about 2000 K. At Mach 1.5 between an altitude of 23 and 30 km a

parachute would be deployed from the nose. The MEM would land tail down with

enough fuel for 60 s hover time if necessary. The landing site was chosen to be Cercopia

near the Martian North Pole as it had been theorised that organisms may follow the

retreating ice cap. The MEM would have windows enabling the crew to look out and

evaluate local hazards including unfriendly life forms! The MEM had two stages with the

upper stage used for leaving the surface after a stay of 10-40 days. For their studies

Aeronutronic used a model Mars atmosphere which was composed of 98.1% nitrogen

and 1.9% carbon-dioxide at a pressure ten times less than on Earth, or at the same

pressure on Earth at an altitude of 20 km. Their craft design would not work on Mars as

designed. However lifting bodies have been proposed for use on Mars by Energia, NASA

and others in more recent times for their low g levels and accurate targeting capability.

1.2 Rise of the robots

In the 1950s and early 1960s interplanetary robotic missions were struggling to

demonstrate their use with numerous failures and poor images. However in 1965 Mariner

4 reached Mars and took a swath of images (see figure 1.3) of the surface (about 1 %),

running from north to south and east to west, as it flew by. A large number of these

images showed a heavily cratered landscape bearing a striking resemblance to our Moon.

We now know that this is not representative of Mars. For example the northern

hemisphere landscapes are younger and so have fewer craters than the southern

hemisphere. Mariner 4 was also used to make radio-occultation measurements of the

Martian atmosphere, revealing it to be about a hundred times less dense as Earth. Mariner

4 also helped determine that most of the Martian atmosphere must be carbon dioxide not

nitrogen as previously thought. For future Landers this meant that more propellant would

be needed for the descent and lifting bodies would be of reduced use. Landing Von Braun

gliders would be out of the question. The Mariner 4 results dispelled the idea that

intelligent life may exist on Mars, as the heavily cratered landscape and thin atmosphere,

suggested that Mars must be very inhospitable.

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One of the last big NASA studies for a while was done by Boeing in 1968 and

incorporated the Mariner 4 findings (e.g. see Portree, 2001). Their piloted Mars

spacecraft was 33 m long with a mass of 140.5 mT. There was also a 144 m long

propulsion section weighing 1000-2000 mT that included an 837 kN NERVA engine. A

33 mT Mars Excursion Module (MEM) would be used to descend and support two

people on the surface for four days. Alternatively a 54.5 mT MEM could support four

people for 30 days. The MEM had two stages, one for descent and one for ascent like the

Apollo LEM. The vehicle looked like the Apollo CM (see figure 1.4) with a conical body

and spherical shaped heat shield. The crew would experience seven Earth gravities during

landing. We now know from long stays in microgravity aboard MIR and ISS that the

maximum g deconditioned crews can experience is between three and five Earth

gravities.

Figure 1.3 Mariner 4 imaging of Mars. Image credit: NASA.

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Between 1966 and 1968 a series of unmanned Mars probes was planned as a precursor to

humans on Mars in the 1980s. This was known as the Voyager Mars Program. The

program was to use hardware developed during the Apollo program such as the Saturn V

to launch a pair of probes at the same time. The probes were to consist of a Mariner 9

type Orbiter with a Lander based on the lunar surveyor probe but fitted with a heat shield

and parachutes. In 1968 a cut-price version of the Voyager probes was announced as the

Viking program which included a biology experiment to look for life.

In 1976 Viking landed on the surface of Mars in a desert rock-strewn surface (figure 1.5).

Scraping away at the surface Viking did not find any conclusive evidence for life which

was not good for Mars exploration support. However Viking did find silicon, calcium,

chlorine, iron and titanium, demonstrating Mars was resource rich. Studies into In-Situ

Resource Utilisation (ISRU) were published after the Viking landings. One idea was to

split water to produce hydrogen and oxygen. It was decided that this needs heavy cooling

equipment. Liquid methane / oxygen production was decided to be an interesting

compromise. First water is used to produce oxygen and then hydrogen is reacted with

carbon-dioxide to produce methane. The Viking results generated a surge of interest

amongst scientists, engineers and in sending humans to Mars. In 1978 the first paper was

published exploring the life-support possibilities for humans on Mars. It was called the

“The Viking Results-The Case for Man on Mars” (Clark, 1978).

Figure 1.4 Mars Lander. North American Rockwell’s design for a Mars Lander incorporating

the findings of Mariner 4. Image credit: Boeing Company.

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1.3 The Case for Mars

In 1981 the first Case for Mars conference took place, named after the paper mentioned at

the end of the last section. This conference was a public forum for Mars enthusiasts to

meet and discuss how to explore Mars. In the 1984 conference plans for a permanent

Mars base were developed and presented (CFM, 1984). The plan included cyclers, ISRU

and aerobraking to reduce the mass and optimise reusability. Cyclers are large spacecraft

that transfer the crew to and from Mars. In a resonant orbit with Mars and Earth they

could return to Earth carrying only a small amount of onboard propellant. The crew then

used smaller shuttle crafts, assembled or refurbished in Earth orbit, to chase and catch up

with a cycler. Three shuttles, with a total of 15 crew members docked with the cycler in a

pin wheel configuration. The whole assembly could be spun-up to generate artificial

gravity at one third of Earth gravity. The shuttles, shown in figure 1.6, had a reusable heat

shield enabling them to serve as Mars Landers, Mars Ascent Vehicles and Earth Return

Vehicles (together with a cylcer on its inward leg). The shuttle crafts derived from

nuclear weapon delivery systems could be accurately steered at hypersonic velocities.

The Landers were biconic in shape, 20 m long and had a mass of 28 mT.

On arrival at Mars the shuttles separated from the cycler and aerobrake into orbit around

Mars. Once regrouped and the position of the base established the shuttles landed with

their nose point upwards (i.e. base down), using parachutes and rockets. The cycler then

passed Mars and returned to Earth, picking up the previous base inhabitants. After a two

year stay the crew from the base blasted off the Martian surface using the shuttles,

refuelled using ISRU. They hooked up with the passing cycler, while the new crew

waited in orbit.

Figure 1.5 Viking on Mars. The left image is a self portrait of Viking on the surface of Mars in the 1970s

and the image on the right is an image of Viking taken from orbit in 2006 by MRO. Image credit: NASA

Viking 2

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1.4 End of the space race

Sally Ride charged with planning a new course for NASA after the Challenger Space

Shuttle disaster published a report with four sets of initiatives (Ride, 1987). Among these

was included an initiative to land humans on Mars. The mission design was begun in

1987 by Science Applications International Corporation (SAIC). Their mission described

in the report was based on a study by students at the University of Texas. Orbital Transfer

Vehicles (OTVs) were used to throw the automated one-way cargo vehicle that included

a Mars lander to Mars. A 19.4 mT sprint vehicle included the crews living quarters and

the ERV mounted behind a 24.4 m diameter aeroshield for aerocapture. Once the sprint

vehicle docked with the cargo vehicle in Mars orbit three astronauts accessed the Lander

and descended to the Martian surface for a 10 to 20 day stay. A crew of three remained in

orbit. The cargo vehicle was used to refuel the crewed vehicle (with ERV). After the

surface mission the crew entered the ERV and returned, using aerocapture to enter into

orbit around Earth. The vehicle could then be refurbished for reuse. In 1987 Martin

Marietta was chosen to lead the study (with SAIC assisting) successfully retaining its

contract and producing reports for NASA until 1990.

Figure 1.6 The Case for Mars II mission Landers. From left to right, Aerobraking, landing, setting

up base and the finished base. Image credit: Carter Emmart.

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In 1988 the Soviets sent two 6.5 ton spacecraft towards Mars, the largest spacecraft ever

sent to the Red planet. The probes were called Phobos 1 and Phobos 2 after one of the

two small moons of Mars, which was also a target for investigation. Also at around this

time the Energia heavy lift launch vehicle, capable of putting 80-100 mT into LEO (about

the same as the US Saturn V) was used to launch the Soviet space shuttle, Buran. Energia

were also busy publishing a detailed human Mars mission using their heavy lift Energia

rockets. As if to compliment these accomplishments and studies Titov and Manarov

completed a year aboard the Mir space station, long enough for a round trip to Mars. It

seemed the Soviet Union (with their impressive array of space hardware and experience)

was on its way to sending humans to Mars. However the economic reality was very

different and it would be several years before the Soviet Union was no more.

The counter the Soviet’s apparent mastery of space the Space Exploration Initiative (SEI)

was started in 1989 by the Bush administration in the US. The SEI proved to be a failure

in terms of producing hardware but generated some good ideas. The 90 day study

proposed the Space Station Freedom with fuelling facilities and hangers, a base on the

moon and then Mars, with a heavy dependence on ISRU at the Moon and Phobos. Four

different, detailed, architectures were proposed to send humans to Mars. Also in 1990

Martin Marietta, sponsored by NASA to produce SEI concepts put forward the Mars

Direct mission. This mission (Zubrin et al., 1991) used a clever synthesis of ISRU,

aerobraking and artificial gravity. A 40 mT cargo Lander including heat shield, descent

stage, ERV, ISRU, 5.7 mT of hydrogen and a 100 kW nuclear reactor to be launched

DIRECTLY to Mars without assembly or refuelling. Once on the surface of Mars an

automated rover deployed a nuclear powered reactor from the ERV. A 38 mT piloted

craft would then be launched (and another ERV for back-up and following crew) after the

fuel for the ERV had been produced. A small ERV (figure 1.7) was thought to be a

possible weak link. Mars Direct required a hypothetical Ares heavy launch vehicle (240

mT) derived from shuttle technology (a 130 mT cargo version is now being planned by

NASA, see figure 1.10). The mission was presented by Robert Zubrin, its leading author

and a Martin Marietta engineer, at the Case for Mar IV conference in 1990.

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Figure 1.7 Mars Direct. Habitat and Earth Return Vehicle (ERV) on the surface of

Mars. The Habitat is the squat cylindrical structure on the left. The ERV is the two

stage cone shaped structure on the right. Image credit: The Mars Society.

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1.5 Reference Missions: A question of mass?

Reference Missions (such as Mars Direct, NASAs DRMs, Mars Society reference

mission etc) are used to further develop and explore mission architectures that can take

humans to Mars, land them and return them to Earth. They are not intended to be

complete. Quoting the introduction to DRM 3.0, “First, it is used to form a template by

which subsequent exploration strategies may be evaluated for consideration as alternative

or complementary approaches to the human exploration of Mars. Second the Reference

Mission is intended to stimulate additional thought and development in the exploration

community and beyond.”

A response to Mars Direct was developed by the Mars Exploration Study Team in 1993.

The mission architecture included a cargo Lander (MAV, ISRU, propellant factory and

hydrogen and 40 tons of cargo including a pressurised rover), empty habitat Lander and

ERV Orbiter each being 60 to 70 mT each. These would be sent to Mars before a crewed

habitat Lander is sent. This was to allow time for the production of fuel and consumables

ready for the crew’s arrival. The crewed habitat Lander can be docked with the backup

second habitat on the surface using wheels to move it. The mission design included a

departure from Mars Direct in that a large ERV Orbiter was included for a more

comfortable return journey. Mars Direct used an ERV launched directly from the surface

of Mars after using ISRU to produce its fuel. Due to mass constraints this allowed only a

small living space for the returning crew. With NASA DRM 1.0 the crew blasted of the

surface in a small MAV which then docked with the large ERV in orbit.

The subsequent NASA DRM 3.0 study in 1997 (Drake, 1998) reduced the mass launched

towards Mars. It became clear that a 200 mT HLLV launcher, required for DRM 1.0

would be very expensive to develop. Instead a more realistic 80 mT HLLV was

Figure 1.8 NASA DRM 1.0. Piloted Mars landing. Image credit: NASA.

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envisaged for the DRM 3.0 study. Also the initial uncrewed habitat Lander was dropped

for DRM 3.0. The Lander technology was revised to include biconic shaped Landers.

These types of vehicles, based on nuclear warhead technology, have good targeting

ability and the capability to manage g load due to their relative high lift to drag ratio. On

the surface an inflatable habitat is used to augment the living volume available to the

crew.

Mass considerations are important for a Mars mission when using available launcher

technologies. Reference missions particularly try to reduce the fuel carried into orbit and

beyond. To do this their architectures utilise immature / untested technologies (e.g.

nuclear rockets, in-situ resource utilisation) to make a mission possible. The following

‘tricks’ are used.

• High specific impulse (a measure of efficiency) propulsion such as nuclear or

electric propulsion.

• Aerobraking in Mars atmosphere for landing. The presence of the Martian

atmosphere is essential for helping minimise the fuel taken up for a trip all the

way to the Martian surface. A reduction of >90% in kinetic energy can be

achieved by simply ploughing into the atmosphere with a heat shield. The

remaining energy can then be removed using parachutes and rocket engines.

However for human scale Landers untested large aerobrakes (15 m diameter

heat shield, 30 m diameter parachutes) are required.

• Aerocapture into Mars orbit. Aerobraking has been used successfully on

robotic probes for circularising their orbit after propulsive capture into an

elliptical orbit. The probes dip into the top of the atmosphere at each pass

using their solar panels as drag surfaces. This significantly increases the

payload delivered into Mars orbit and does not require a heat shield as the

Figure 1.9 NASA DRM 3.0. Crew habitat. Image credit: NASA.

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heating load is not very much. It however this technique takes too long for a

human mission and a full aerocapture will be required dipping deep into the

atmosphere. A heat shield will be required. Human mission may use two heat

shields, a heavy one for the high speed aerocapture event and a lighter one for

the lower speed entry. A mass penalty may come from the packaging of

nested heat shields.

• In-situ resource utilisation (ISRU) on the surface of Mars to make fuel for the

return flight to Earth. Sending a precursor robotic mission, with a tank of

hydrogen and a power source, to break down atmospheric carbon dioxide to

created methane, oxygen and water.

• Aerocapture into orbit around Earth. High approach speeds to Earth requires a

heavy heat shield. Also an efficient lifting body is required to keep the g loads

off a possibly deconditioned crew.

• Of course these undeveloped or untried technologies could be side stepped by

using HLLVs capable of lifting >200 mT.

Landing humans on Mars is inherently complicated as it is impractical to launch a single

spacecraft that can transport a crew from the surface of the Earth to the surface of Mars

and back again. Even the Apollo moon landings required two spacecraft, the Command

and Service Module (CSM), to transport the crew into orbit around the moon and the

Lunar Excursion Module (LEM) to get humans onto the lunar surface. Both Apollo

spacecraft were launched on a single heavy lift launch vehicle (HLLV) the now

discontinued Saturn V. To land on the more distant Martian surface and return to Earth

requires two or three spacecraft, launched individually on HLLVs. Therefore a new

HLLV capability will have to be developed.

Figure 1.10 HLLVs. Image credit: NASA.

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It has been shown in DRM studies that about five times the mass landed on Mars has to

be placed in LEO first. Since >30 mT of payload will need to be landed on the surface

DRMs typically invoke launchers capable of launching 100 to 150 mT into low Earth

orbit (for comparison the Saturn V could launch 100 tons into LEO). These high lift

launchers can only be built through a national organisation like NASA due to the high

cost (10 billion US dollars).

Alternatively smaller Medium Lift Launch Vehicles (MLLV) are being developed by

private companies or those already in existence with NASA could be used to launch a

modular design that is assembled in Earth orbit. This may then be cheaper especially if

commercially development of space (e.g. tourism) drives the cost of launchers down

through increased efficiency (small team).

Launchers both expendable and reusable are being developed by private companies and

may one day take us into orbit and beyond, perhaps taking human to Mars for the first

time. Due to the expensive development cost these rocket launchers are being built by a

handful of billionaires, who have already made their money elsewhere. Possibly the first

private company that will begin launching payloads up to 570 kg into orbit will be

SpaceX. The company was founded by Elon Musk after making his fortune with PayPal.

They have already flown their falcon 1 vehicle in March 2006 carrying a satellite but the

engine failed 26 s after launch. However the payload was thrown free when the rocket

impacted the ocean on its side and was recovered with some damage. A second attempt to

launch the falcon 1 will take place in 2007. SpaceX also have plans for a MLLV launcher

Figure 1.11 Medium Lift Launch Vehicles. From left to right are three MLLVs, Ariane 5 -

21 mT LEO increased to 27 mT with improved engines (Iranzo-Greus, 2005), 5.4 m

diameter fairing, Proton 20 mT to LEO, 4.35 m diameter fairing, Delta 4, 26 mT to LEO, 5

m diameter fairing.

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the falcon 9 S9 that could boost 25 mT into LEO. Another billionaire who is developing

rockets is the owner of Amazon.com. The rocket company is called Blue Origin. They

have already flown a prototype of their passenger carrying New Sheppard rocket at low

altitude in November 2006. The New Sheppard will carry fare paying passengers into

space on a suborbital trajectory. The great thing about development of reusable rockets is

that the design can be iterated with the same rocket! This is assuming the rocket isn’t lost

in an accident like the DC-X being developed by NASA and McDonald Douglas in the

1990s. However it demonstrated a turnaround time of 26 hours before being destroyed

The Japanese space agency have had a successful research project iteratively developing

a reusable test vehicle from 1999 to 2003. In a book called “The Rocket Company”

Patrick Steinnon writes about the development of a reusable rocket in great detail and he

also projects into the future when these rockets take human to Mars. He has a patented

design of a unique design and trajectory for a reusable rocket that can, in principle, be

operated under conditions as airlines do today.

The details of (Entry Descent and Landing) EDL technology in DRMs are often vague.

This is a challenging part of the mission to design in terms of how to do it. We basically

know how to return a crew back to the Earth's surface utilising its thick atmosphere. It

was done as part as the Apollo mission to the Moon using a heat shield and three large

parachutes to set it down in the Atlantic ocean at a leisurely 9 m s-1. A vehicle returning

from interplanetary space will be travelling faster so a new type of vehicle may have to

be used based on a more aerodynamic design to give greater lift and reduce the g levels

experienced by the crew. If the spacecraft is particularly large (it may be to house a crew

of 4 to 6 for 6 months) then it may not even descend to the surface but remain in Earth

orbit after using the atmosphere as a brake and dock with a space station.

As well as the Earth we know how to land humans on an airless body like the Moon, as

with the LEM, which put about 10 mT on the surface of the Moon. However we do not

know how to land humans on Mars, which has a thin atmosphere, unlike the Earth or the

Moon. It is a technological problem, especially as the payload masses involved are so

much higher than past robotic missions.

Aerobraking technologies used by past robotic missions to Mars, like the parachutes for

Viking, cannot simply be scaled up. Mass for a solid object will increase approximately

as the cube and surface area will increase as the square of a single dimension like radius

or length. For a spacecraft there is a similar scaling up in the mass to surface area ratio as

you increase the volume. Parachutes for a manned mission to Mars would have to be

much larger than previous parachutes used on Mars and deployed at higher speeds,

possibly creating stability problems and increasing deployment time. In Reference

Missions the authors often make estimates for parachute, heat shield and rocket fuel

massed to land their habitats of excursion vehicles on the surface. However there are

several consequences of this scaling approach that need to be examined. For a human

payload g levels need to be constrained (effecting heat shield and parachute sizing) and

inflation time of larger parachutes may incur a time penalty during the descent. Also pin-

point targeting (necessary for delivery of the crew next to a cargo Lander) needs to be

developed (possibly requiring a large amount of propellant). These should be important

18

considerations for Reference Missions as they control the amount of mass that has to be

launched towards Mars (and landed on Mars).

Entry into the atmosphere at high speed produces very high temperatures reaching a maximum at an altitude of about 30 km. A heat shield is required to protect the spacecraft (situated behind the shield) from the hot plasma. The picture on the left shows a test.

Once the entry phase is over a parachute is deployed at supersonic speeds to slow down the Lander and pull it off the heat shield. This then exposes a terminal descent system such as rockets or airbags. On the right shows a test of the MER parachute.

On the left is a test of the retro rocket for the MER rovers. The retros slow the Lander to zero velocity just above the surface. The rover is then dropped to the surface, cocooned in airbags which deflate once bouncing has ceased.

Figure 1.12 Entry, descent and landing technologies for a robotic mission

(<1mT). Image credit: NASA

19

1.6 Walkabouts

Public interest in Mars was renewed in 1996 with the announcement of evidence for life

found in a meteorite from Mars ALH84001. This was a piece of Mars that had been

thrust into space by an asteroid (0.5-2 km in size) impact at least 16 million years ago and

then made its way to Earth, arriving about 13000 years ago. Only some scientists now

think the space rock shows evidence for life but whatever the truth the announcement

caused some excitement and generated enthusiasm for space exploration.

In 1997 a tiny rover called pathfinder landed on planet and watched by a base, fitted with

meteorological sensors and a camera, trundled around the nearby rocks on the first

Martian “walkabout” in history. The rover’s investigation included looking at

surrounding rocks, thought to be deposited by a catastrophic flood in Mars past. The

rover could peer down to millimetre levels, nowhere near the level to look for

microfossils as seen in ALH84001. However the rover created tremendous interest and

was a very positive mission for NASA. After running a competition the rover was named

Sojourner after an African-American reformist and champion of women’s right.

Sojourner also means a person working in a foreign country with the intention to return to

the homeland after a period of time. However this Sojourner was stuck on Mars. Contact

was lost after three months of “walkabout” on the surface. Could it have given people a

taste for Mars exploration? Humans would, like Sojourner, journey across Mars in their

pressurised rovers.

To help mission planners understand the mass requirement for Reference Missions it was

proposed in 1997 that the surface mission needs to be defined in detail. It was envisioned

that a large group of people try and understand how a small crew would live and work on

Mars.

At present there are two Mars analogue stations, based on NASA DRM and Mars Direct

type habitats, attempting to develop efficient surface mission scenarios. One is F-MARS

at Devon Island and one is the Mars Desert Research Station (MDRS) in the Utah desert.

Figure 1.13 Mars Pathfinder. Image credit: NASA.

20

The habitats comprise of 8 m cylinders about 11 m tall and can house a crew of six on

two floors. A new habitat, EuroMars is to be built in Iceland. Another habitat is planned

to construction in Australia called OZ-MARS. This habitat is basically a long cylinder on

its side. It is based on a bent-biconic structure, that offers high manoeuvrability and

targeting capability compared to the top-hat entry vehicles that MDRS is based on; the

mission architecture is the same as Mars semi-Direct, a cross between the NASA DRM

and Mars Direct. The Mars Society is the principle driver behind all these Mars analogue

stations. The Mars Society was established in 1998 by Robert Zubrin to promote the idea

of humans on Mars. F-MARS became operational in 2001 and MDRS became

operational in 2002.

As well as defining the mass requirement these analogue help mission planners

understand how people will work together on Mars. The Mars Society has three prime

goals for their four Mars Analog Research Stations. These are:

• Optimise the productive exploration of Mars by humans

• Conduct useful field research to understand geology, biology and environmental

conditions on Earth and on Mars.

• Generate public support for sending humans to Mars

To fulfill these goals the Mars Desert Research Station features a habitat fitted with a

field laboratory and tools for exploration of the surrounding area including All-Terrain

Vehicles. These are small open motorized buggies designed for off road use. They are

used for exploring remote areas and sample collection, similar to the methods used by the

Apollo astronauts who briefly wandered on the Moon.

Figure 1.14 MDRS. Image credit: The Mars Society.

21

2. How on Earth do we land humans on Mars?

When we go to Mars the beginning of the mission, for most people, will when the crew

blasts off into space. This is what everyone will see and the pictures will be flashed

around the world to much applause. However at the other end they’ll be no Martian

anticipating the arrival of the intrepid voyagers, like the splashdown of the Apollo

astronauts on Earth or the high altitude fireball made by the space shuttle as it shoots

across the USA towards Florida. At the other end of the Solar System the crew will

plunge into the Martian atmosphere, unrecorded by human eyes. Perhaps their arrival will

be captured by robotic cameras but there will be certainly no journalists (except the

astronauts themselves) to tell us the story of how it all feels.

The arrival at Mars will certainly be eventful for the crew though. The most dramatic

arrival would be with a Mars Direct type mission architecture where the Lander plunges

into the Martian atmosphere travelling at an interplanetary speed of 6 km s-1. This will

create a huge fireball, as the heat shield compresses the atmosphere in front into shock

fronts and the heat ablates material away in effort to keep the crew safe from the

surrounding inferno of hot plasma. The Lander will decelerate very quickly around an

altitude of 30-40 km slowing down enough by 10 km to give its huge supersonic

parachutes a chance to slow it down to subsonic speeds and allow the rockets to deliver

the crew to a pinpoint landing next to a cargo Lander. All these things would happen in a

few minutes, similar to the time scale to reach space in a rocket launcher.

A less dramatic scenario will involve the Lander dipping into the atmosphere for a few

minutes to reduce its speed so it is captured into Mars orbit. The crew will have to hope

their craft is on target because if they hit the atmosphere to steep they will not be able to

pull up and will plunge to a fiery end. On the other hand if they strike the atmosphere at

too shallow an angle they will retain their interplanetary velocity and head off into space,

perhaps circling the sun forever or perhaps they could limp back to an Earth. Once in

orbit the crew can breathe a little easier, eject their hot heat shield and prepare for descent

to the surface. After checking that their second heat shield is intact after ejecting the hot

aerocapture shield they fire up their engines to deorbit. The descent will be very similar

to a Mars Direct entry but the heating and deceleration will be less severe but everything

must work to get the crew safely on the surface.

These descriptions may well represent some future piloted Landing on Mars but the fact

is much of this technology is untested, like aerocapture, nested heat shields, large

supersonic parachutes and pin point landing. How can we test these elements before we

go? Most likely it will involve a large amount of computer simulations and few expensive

experiments in the Earth’s upper atmosphere. This section describes how we could start

to think about these problems and some of the problems that need to be solved.

22

1

2

3

4

5

6

7 8

Atmospheric entry interface

Begin bank angle modulation

Periapsis

End bank angle modulation

Atmospheric exit

Periapsis raise manoeuvre

Hyperbolic approach trajectory

Circularisation manoeuver

2.1 Aerocapture

Aerocapture is a technique whereby an interplanetary spacecraft flies through a planet’s

atmosphere at high speed (>6 km s-1) generating significant aerodynamic drag forces to

slow it down to orbital velocities. Figure 2.1 shows the approach of how such a

manoeuvre may be conducted for a robotic science mission. The circular orbit shown in

figure 2.1 is important for science missions as it allows the surface to be observed from a

constant height.

Aerocapture is an important technique because it reduces the amount of propellant (and

hence mass) carried by the spacecraft so allowing a higher mass payload or a small

launch vehicle. The mass requirement of the heat shield needs to be considered carefully

as there is a trade off between the mass of fuel used for propulsive capture and the mass

of the aerocapture shield. For capture into a highly elliptical orbit around Mars the

increase in payload is only about 5% while a capture into a more circular orbit allows a

15% increase in payload (Keys, 2006). This is much smaller percentage than for Venus or

Titan but is significant especially if the risk is shown to be comparable to an all

propulsive capture.

Figure 2.1 Aerocapture overview. The maneuvers required to but a vehicle into a circular science orbit.

The trajectory begins with a hyperbolic approach, with the periapsis at some point in the atmosphere.

Atmospheric forces slow the vehicle between point 3 and four so at atmospheric exit an elliptical orbit has

been obtained. The periapsis of the orbit needs to be raised out of the atmosphere so a stable orbit at point

7 and to make the orbit more circular. At point 8 a trimming of the orbit can be made. At point 7 and 8

propulsion is used. Apoapsis (opposite periapsis) is the location that is most efficient in terms of using

rocket fuel.

23

Blunt body Slender body

Trailing ballute Attached ballute

Aerocapture has been classed as a high priority for NASA’s in-space propulsion

technology program. This technology can enhance and enable science missions. Four

types of designs have been investigated by NASA’s aerocapture project. These are blunt

body, slender body, trailing ballute and attached ballute designs as shown in figure 2.2.

Blunt body designs involve the use a rigid heat shield like those used by Mars Landers or

by the Apollo Command Module on Earth. This technology is at a high level of maturity

(or readiness/development). Assessment of the aerothermal environment and correct

choice of (Thermal Protection System) TPS will have to be made. A slender body design

(low to moderate maturity) is under consideration because it provides increased volume

and improved packaging advantages for larger spacecraft (like for a human mission). A

ballute is a cross between a balloon and a parachute. The trailing type of ballute is under

consideration because it is applicable to all size and shaped payloads. It is however at a

low state of maturity. An attached ballute has volume and packaging advantage for larger

spacecraft (like for human missions). It is at a low to moderate level of maturity.

For a human mission to Mars aerocapture is an attractive architecture because of the

potential mass savings. A human mission to the surface will either require a very heavy

rigid heat shield for both aerocapture and descent or a pair of nested (rigid) shields, one

for aerocapture and one for descent. There are problems that need to be addressed with

both approaches. During aerocapture there will be high thermal loading of the heat shield.

For a single heat shield the heat absorbed needs to be dumped somehow without

excessive heating loads on the crew habitat. A heavy heat shield makes for an overly

heavy entry system, reducing the surface payload and negating some of the advantage

gained from aerocapture. Using two shields can potentially increase the amount of

payload delivered to the surface and solve the problem of removing the absorbed heat

after aerocapture. Once aerocapture is completed the hot heat shield is ejected leaving

Figure 2.2 Aerocapture designs. Four designs being researched by

NASA. Image credit: NASA.

24

another heat shield (lighter than the single heavy shield) which can then be used for the

descent. However nesting of two heat shields is an engineering challenge as they have to

be connected so as not to damage the TPS. However the mass of the attachment structure

may be larger than the mass of payload gained through using aerocapture negating its

advantage over an all propulsive capture.

One reason why aerocapture would be used for a human mission to Mars is that the g

level loading of the crew is easier to manage. A descent from orbit (i.e. after aerocapture)

compared with direct entry will experience lower g levels due to the lower entry speed.

High g levels are dangerous for the human body, especially if the crew has been

deconditioned from living in a microgravity environment. If large inflatable devices

required for human missions, like the trailing or attached ballutes, are shown to be stable

under hypersonic speeds (and demonstrate aerodynamic lift for precision trajectory

control), then their lightweight structure will be highly compatible with a mission

architecture that uses aerocapture.

25

2.2 Entry, Descent and Landing

Entry, Descent and Landing (EDL) refers to the sequence of events required to deliver a

payload onto the surface of a planet or moon intact. The following discussion divides

EDL into their constituent parts that being, atmospheric entry, descent and landing. For

each section the important issues related to delivering humans onto the Mars surface and

setting up a base to support them are discussed. As each part of an EDL sequence

depends on the preceding operation being executed successfully there will inevitably be

some overlap in the discussion. For instance a pinpoint landing required for setting up a

base, depends on the performance during atmospheric entry and descent so the section on

landing will also discuss the other sections of the EDL sequence preceding it. The section

on descent and how to reconfigure the Lander for landing brings in related issues to from

the atmospheric entry phase. The atmospheric entry phase deals with g levels during

entry and how these can be kept down (by having some steering capability) so the crew

can remain conscious during entry.

2.2.1 Atmospheric entry, steering and g levels

The Atmospheric entry phase requires a vehicle design that can keep the payload safe

from thermal heating and deceleration forces during hypersonic entry into an atmosphere.

Also the vehicle subsystems and ground support have to deliver the vehicle to an entry

point above the atmosphere with the correct flight parameters such as position, velocity

and angle with the horizon. For example if the trajectory at the entry point is too steep,

the thermal heating or deceleration forces will be too great for the vehicle design and it

will break-up. If the trajectory at the entry point is too shallow then the vehicle will not

enter the atmosphere and return to space. Other important requirements for a human

mission are a vehicle with a high degree of steering, for safe recovery, and an ability to

manage the g levels so the crew are not harmed.

A requirement of an EDL sequence is to reduce a spacecraft’s kinetic energy sufficiently

so it can survive an impact with the surface, or in other words have a soft landing. On a

moderately size body in the Solar System like the Earth, the Moon or Mars a spacecraft

needs to land with under a micropercent of their original energy to survive. The Stardust

capsule entered the Earth’s atmosphere at 12.8 km s-1, after visiting comet Wild 2, and

impacted the ground at only 0.005 km s-1. By comparison the Apollo astronauts entered

the atmosphere at 11 km s-1 and impacted the Atlantic Ocean at 0.009 km s

-1 but that’s

still a pretty high decrease in velocity. Most of this decrease occurs during atmospheric

entry and hypersonic deceleration.

As a spacecraft ploughs into the atmosphere it not only experiences high deceleration but

high temperatures due to the severe compression of the surrounding gas. The shock

waves disassociate the molecules forming a plasma containing ions. When the plasma

density is high enough radio waves will be reflected or absorbed causing the well known

phenomena of radio blackout during a spacecraft’s entry into an atmosphere. Peak

hypersonic heating occurs shortly before peak deceleration which is at an altitude of

26

about 50 km on Earth and 30 km on Mars. On Earth a spacecraft will reach subsonic

speeds high up in the atmosphere. For example the Apollo Command Module capsules

used a spherical heat shield for themal protection and a drag surface during hypersonic

deceleration enabling it to reach subsonic speeds at an altitude of 20 km. Parachutes were

used for the final descent from about 7 km. On the Moon, with no atmosphere, the Apollo

LEM used a rocket to first deorbit from an orbital velocity of ~1.5 km s-1 (relatively low

compared to about ~7.5 km s-1 for the Earth), then descend and soft land a 16 mT on the

surface. On Mars, with a thin atmosphere a combination of aerobraking (heat shield and

parachutes) and retro rockets (together with air bags or a rigid landing gear) have been

used so far to successfully land robotic payloads.

Another requirement apart from soft landing, for EDL systems, is some kind of targeting

or steering capability. The ability will manifest itself by contributing to reducing the

landing ellipse. The landing ellipse is an area of targeting uncertainty depending on the

navigation capabilities of the spacecraft, environmental uncertainties such as wind

direction and strength (if landing on the Earth or Mars) and the capability EDLS to

compensate for these uncertainties. For example the Apollo CM had steering capability

during hypersonic flight to bring it close to a recovery ship. Apollo had the first modern

embedded computer system. It was able to guide the Apollo command module to within a

few kilometres of the recovery ship. The LEM on the Moon also had good targeting

capability, particularly as there were no winds to blow them of course. For example the

Apollo 12 LEM landed 160 metres away from Surveyor 3 probe that had previously

landed, or rather bounced and skidded (as can be seen by the pad prints and ploughing of

the regolith), three and half years earlier. The Space Shuttle, a winged spacecraft, can

make a pinpoint landing on a 91.4 m wide and 4.5 km long runway, using its

aerodynamic surfaces to compensate for winds and other navigational errors.

It may be surprising to learn that capsule type spacecraft can fly too! It has aerodynamic

properties and a steering system that can be used to control its trajectory through the

atmosphere. The space shuttle has obvious aerodynamic surfaces and wings that are used

(combined with its RCS system at high altitudes) to control its path. The capsule’s

“wing” is the underside of the heat shield angled at 27° to the direction of flow. Lift is

then obtained from deflection of the “wind” downwards. The CM is inherently stable in

this orientation due to an offset centre of mass (about 30 cm above the z-axis). The offset

centre of mass creates moments around the CM’s centre of pressure (located near the

capsules apex). If the capsule is perturbed slightly from its stable position then

components of the lift and drag forces will act to restore the CM to its normal attitude.

The centre of pressure (COP) is a point where the vectors of aerodynamic forces

coincide. The COP is near the apex due to the spherical shape of the heat shield. Try and

imagine where the COP of a parachute is and how this effects its behaviour.

In a lift up configuration the COM is in the upper position as shown in the figure 2.3 and

2.4. The CM is steered by using the RCS thrusters to bank to the left or right (rotate

around its long axis). This will rotate the lift force vector and the CM will go in that

direction. The CM can be steered to the left or right using this technique. It can even be

positioned in a lift down position where all the lift force is directed downwards. If no lift

27

L COP

D

F

COM

Relative flow

F

z-axis

27°

Restoring moments

provide stability

Heat shield

Rotation around z

axis for steering

L: Lift (upwards)

is required then the CM can rotated slowly so the lift effects cancel each other out. The

Apollo CM was able to splash down within a few kilometres of the recovery ships. For a

manned mission to Mars it is likely that this kind of steering or better (combined with

rocket propulsion for the final approach) will be essential for setting up or landing near a

Martian base.

Figure 2.3 Flying the Apollo command module. This shows an outline

of the CM. The CM has three important aspects that make it into a flying

machine. These are lift, stability and control (steering). The lift is

achieved by flying the CM at an angle to the oncoming flow (similar to

an aircraft wing). Stability is achieved through an offset centre of mass.

Control is achieved by rotating the lift vector as shown in the adjacent

diagram. F represents the average force due to dynamic pressure; D

represents the force due to drag. COP is Centre of Pressure and COM is

Centre of Mass.

28

Lift upwards

Lift left

Lift right

Trajectory with no lift

Trajectory with lift / steering

COM

Peak deceleration is a function of lift over drag. Lift over drag is defined as the ratio of

the lift and drag forces. A high L/D is important for manned vehicles returning at high

speed from the moon or interplanetary trips (like from Mars) as the peak deceleration

without any lift capability would be dangerous. The altitude-time profile and g levels for

the Apollo 6 CM with and without lift are shown in figure 2.5. The Apollo capsule had an

L/D ~0.4, flying at an angle of attack of 27°. The maximum g level for the Apollo 10

lunar orbital re-entry profile was about 7 g. The space shuttle has a hypersonic L/D ~2.0

and flies at an angle of attack that begins at 40° decreasing to 14° towards the end of

hypersonic flight. The maximum g level experienced on the space shuttle has been no

more than 3 g. The low g level compared to the Apollo CM is partly due to the slower

entry speed from orbit but is probably also due to the Space Shuttles high L/D. For a

deconditioned crew, who have spent many months in a microgravity environment aboard

the space station or perhaps returning from Mars will have problems above 5 g. During

microgravity the heart muscles do not have to work so hard and become weak, unable to

cope with high accelerations for extended periods. Designs for return from Mars include

biconic shapes that have a high lift over drag ratio.

Figure 2.4 Steering the CM. The CM can be steered by rotating the lift

vector around its z axis (see adjacent diagram) using its banking

thrusters. If no lift or steering is required then the CM can be slowly

rotated around its axis so the effects of the lift force cancel out. The

“egg” shape in the diagram represents the Earth. The dotted line

represents the CM trajectory without any steering. The solid line

represents upwards lift (coming over the horizon), and then a left turn,

then a right turn.

29

Figure 2.5 The red and blue lines show a reentry simulation of the Apollo 6

Command Module with and without lift. The lift of the CM causes a skip as

shown in diagram A (red). The lift also keeps the CM in the higher part of

the atmosphere reducing the drag and g loading as can be seen in diagram B.

The program used is called A2D and is an aerobraking simulation developed

by the author, details of which can be found in section 3.5.

A

B

30

2.2.2 The descent phase and the hypersonic transition problem

The descent phase begins once hypersonic deceleration is over (less than Mach 5 and

parachutes have been deployed). On a body like the Earth, with a thick atmosphere, an

unpowered descent is possible using parachutes to decelerate to a terminal velocity

(where drag and gravitational force balance) and a safe landing speed (~9 m s-1 for the

Apollo CM splashdown). On an airless body the descent will be controlled entirely using

rocket engines as with the Apollo LEM on the moon and is relatively straight forward.

On a planet like Mars that has a thin atmosphere and a high terminal velocity even with

parachutes (~20-50 m s-1) a Terminal Descent System (TDS) is necessary to kill the

remaining velocity before landing. The TDS may consist of rocket engines as used by the

Viking Lander or the soon to be launched Phoenix Lander. Alternatively airbags can be

used to cushion the Lander on impact as used by the MER rovers. These systems require

that the heat shield be removed to expose the landing gears and retro rockets, normally

facilitated by a parachute. With massive payloads (>1 mT) it is unclear how to decelerate

a Lander so it can deploy its parachutes before hitting the ground and has been referred to

as the hypersonic transition problem.

On Earth the hypersonic deceleration phase ends at a high altitude and the entry vehicle

has sufficient time to brake and permit gravity to turn the trajectory into a vertical descent

(i.e. perform a gravity turn). For the Apollo CM this was at an altitude of about 20 km.

Parachutes can then be used to reduce the terminal velocity for the landing or

splashdown. On Mars hypersonic deceleration occurs much closer to the surface giving

little time to perform a gravity turn and begin the terminal descent phase. Without any

assistance decelerating Viking would have impacted the surface at several hundreds of

metres per second with a significant horizontal velocity component. For an Apollo CM

type vehicle on Mars it would be even higher. How do you stop something so massive

and travelling so fast?

Viking EDL technology cannot simply be scaled up for this purpose is at the heart of the

hypersonic transition problem. An approach to understanding the problem is to examine

the ballistic coefficient. This is a useful way to characterise an objects ballistic behaviour.

It is defined as being equal to the mass divided by the effective area. Here the effective

area is the product of the drag coefficient and the reference area. The reference area is the

cross-sectional area projected into the flow. For a falling sphere the reference area would

be bounded by a circle with a radius equal to the sphere. An object with a high ballistic

coefficient, such as a cannonball, will have a high terminal velocity (where drag and

gravitational forces balance) whereas an object with low ballistic coefficient, such as a

balloon, will fall with a lower terminal velocity. However it is very important to note that

the terminal velocity does not remain the same if you increase the size but keep the

density the same. This is because the mass increases as the cube while cross-sectional

area increases as the square. Therefore a small cannonball, with a low ballistic coefficient

will fall slower than a large cannonball with a higher ballistic coefficient, (i.e. drag is

more effective on the smaller cannonball).

31

Horizon

D irection of motion

Lift force

Drag force

Entry angle

16.5°

Angle of attack 11.5°

A

B

C

D

W ind

Gravity

To the parachute canopy

High drag

Low drag

As the drag area and payload volume (mass) scale up differently for a specified ballistic

coefficient this suggests a human scale Lander, with a similar ballistic coefficient to

Viking, will not look like a scaled up Viking aeroshell during the hypersonic deceleration

phase, and it will certainly not look like an Apollo CM. If the vehicle is to have similar

payload density to Viking then the scale length (radius) of the heat shield will have to be

larger than the radius of the payload (Lander) creating a top hat appearance (as with the

Mars Direct MTSV). A Viking sphere-cone shaped heat shield has a low lift (L/D~0.2)

capability and poor control authority. A spherically shaped heat shield, like used on the

Apollo CM would allow better control authority (L/D~0.4). Alternatively if the L/D is to

Figure 2.6 Entry and descent of the Viking Lander as an example of entry and descent on Mars. Diagram A

shows Viking in its aeroshell entering the atmosphere. The entry angle or direction of motion is 16.5° to the

horizon. The angle of attack is 11.5° to the direction of motion giving it some lift. The entry speed is about

4.5 km s-1 and the entry altitude is 300 km. Diagram B shows deployment of the main parachute. This

occurs at about 6 km altitude and at a speed of about 250 m s-1 or just over mach one. Simulation of the

Viking Lander suggests the angle of motion (velocity vector) relative to the horizon was about 40°. Seven

seconds after main chute deployment the heat shield is released and due to its higher ballistic coefficient

moves off to the side relative to the Lander. At about 1.5 km and a speed of 60 m s-1 the parachute has

turned with the Lander directly below the canopy as in diagram D. At this point the Martian winds are

blowing the Lander to the side. Rocket engines are used to make the following downward and sideways

corrections for a soft landing of about 2.5 m s-1.

32

be increased significantly a biconic slender body design would be used. For a biconic

vehicle a higher ballistic coefficient (atmospheric drag has less effect) can be afforded as

more time will be flying in the atmosphere (so drag has a longer time to act) and the heat

shield can be integrated onto the hull of the Lander.

The difference is scaling between area and volume (mass) also has an effect on the

parachute descent phase. On Mars a parachute plays a pivotal role turning a Lander into a

vertical descent ready for the terminal descent phase. However the terminal velocity on

Mars is relatively high compared to Earth and a parachute (of practical size) alone can

not be used on a human scale Lander for a soft landing as with Apollo CM. A massive

payload (>30 mT), like a manned Lander, requires a giant parachute (~100 m) to reduce

its speed low enough so its terminal descent system (rockets) can take over. There are

however concerns about the deployment and stability of large parachutes (>30 m in

diameter). The largest flight qualified parachute for the Viking project had a diameter of

19.7 m while the Viking Lander actually used a 16.15 m diameter parachute. All US and

European Mars Landers have used Viking heritage parachutes. The expensive parachute

qualification tests have provided engineers with the operationally bounded limits. These

limits are sometimes referred to as the “parachute box” and help in designing an EDL

were Viking heritage parachutes are used, such as Mars Pathfinder or MER.

1. The parachute have to be deployed either in the subsonic or supersonic region. At

Mach 1.13 there are stability problems that prevent the chute from opening properly.

2. Above Mach 2.2 heating will melt the fabric.

3. At low dynamic pressure, less than 240 Pa (high altitude), there is not enough dynamic

pressure to inflate the parachute.

4. At high dynamic pressure, more than 850 Pa (low altitude) there will be too much

force and the parachutes will be destroyed.

A simple investigation highlights the problem for landing a massive payload on Mars.

Figure 2.7 shows results from an EDL simulation program developed by the author.

Viking’s EDL has been simulated together with the larger Apollo CM. The Apollo CM

uses the same entry state as Viking but the L/D is varied from 0 to 0.5 (L/D=0.6 resulted

in a skip back out into space) to try and shoot the trajectory through the Viking

supersonic parachute box. It can be seen that with an L/D of 0.5 (unrealistic for a blunt

body) the CM just passes through the box. This implies that a parachute could slow it

down sufficiently so some TDS can come into operation. A 50 m DGB type parachute is

deployed at Mach 2.2 and an altitude of 3.5 km. The heat shield (848 kg for the CM) is

released landing a 5 mT payload on the surface (at 47 m s-1 with no TDS). The g level at

parachute deployment is about 20g, probably a dangerous (or painful) level of shock for

the crew. This simple example demonstrates that it is possible to land humans on the

surface of Mars but an Apollo CM is a very uncomfortable way to travel on Mars.

33

0

5000

10000

15000

20000

25000

30000

35000

40000

45000

50000

0 1000 2000 3000 4000 5000

Velocity / m/s

Altitude / m

One approach to the hypersonic transition problem is to remove the parachute descent

stage altogether. Mass trade-off studies have shown that an all propulsive descent may be

preferable to using a parachute. Two favoured approaches to exposure of the descent

engines are opening a door in the heat shield or jettison of the heat shield. Both of these

would require new technologies to be developed and tested. Another challenge is to

understand the behaviour of the rocket exhaust under high dynamic pressures (i.e. at

supersonic speeds) and consequently determine the stability of the Lander.

However a parachute plays an integral component of the EDL sequence; a parachute aids

stability through the transonic region. It also serves, in the case of a Viking type Lander,

to pull the Lander off the heat shield and so expose the terminal descent system and

landing gear. Also the release of the shield means less mass has to be landed and hence

less propellant is required for the powered descent. It may be then that a parachute would

be used sometime during an all propulsive descent for these reasons.

Another approach to solving the hypersonic transition problem is to use a very large heat

shield so more of the deceleration is done during hypersonic deceleration, and then

retained as an aerobrake through the descent phase. Such a lightweight heat shield may

be possible using inflatable technology.

Figure 2.7 Parachute operational environment on Mars. This shows the EDL

(thick line) velocity for the Viking all the way from 50 km to touchdown at 0

km. The dark parallelogram shows the operation environment for the Viking

supersonic parachutes on Mars. Also shown are results from flying an Apollo

type capsule in the Martian atmosphere. With a L/D of 0.5 it appears that a

Viking type 50 m diameter parachute could be used. An L/D of 0.6 results in

the CM skipping back out into space and is not shown in the figure.

Viking

Apollo CM no lift

Apollo CM L/D=0.3

Apollo CM L/D=0.5

Viking chute

34

2.2.3 Landing and pin-point targeting

The landing phase is the final step where the payload is deposited on the surface by a

terminal descent system such as rockets. A human mission will require targeting

capabilities of 100 m so the crew are within easy reach of previous landed cargo which

may include a Mars Ascent Vehicle (MAV) or Earth Return Vehicle (ERV).

Wolf et al. (2004) conduct a comprehensive investigation into the problem as follows.

The EDL sequence can be divided into four segments each of which contributes some

uncertainty to the target delivery uncertainty. These are entry approach, hypersonic entry,

parachute descent and powered descent.

The entry approach phase is when the Lander, encased in its aeroshell, is approaching the

entry point. The position of the spacecraft is measured by taking radiometric observations

using Earth-based radio tracking. This data is then used to make final propulsive

adjustments ready for entry. As there is a time delay from Mars there will be a delivery

uncertainty related to this gap in the knowledge. The spacecraft position can be improved

by optical observations by the spacecraft of Phobos, Deimos or Mars itself just before

reaching the entry point. Also observations from other spacecraft can provide input

before entry. The hypersonic phase can be then used to “fly out” the uncertainty. As

discussed previously a spacecraft can control its trajectory during the hypersonic entry

phase in a similar manner to that of Phoenix or MSL and reach a target point above the

Martian surface for parachute deployment with an uncertainty of ~6.5 x 2 km at

parachute deployment

During the parachute phase wind causes a large uncertainty and may carry the Lander

several kilometres off target. This requires some fly back ability to the target for the

Lander. One possibility is to use a steerable parachute to actively compensate for the

winds. Another possibility is to use propulsion, during the powered descent phase to fly

back to the target but his will require extra propellant. However map-tie error (about 100

m in the near-future) will require target relative navigation and there will inevitably some

fly back possibility with future robotic Landers. The winds are poorly characterised on

Mars below about 10 km. Mesoscale modelling of the Martian winds could help decide

how to proceed with the design of the robotic and also piloted Lander’s EDLS.

The targeting capability of a project like MSL or the Phoenix Lander will be something

like 5-10 km. For comparison the targeting capability for the Mars Exploration Rovers

was 35 km. All past robotic Landers have flown uncontrolled entries into the Martian

atmosphere, without exercising any control over their trajectory. Pheonix and MSL will

demonstrate hypersonic guidance using a segment of the Apollo earth-entry program. It is

capable of delivery to ~6.5 x 2 km at parachute deployment. Entry guidance uses density

altitude as its control input and the Martian atmosphere is highly variable in density

perhaps with an uncertainty of 3 km. Some of target uncertainty is then caused by winds

blowing the Lander off course during the descent. Wolf et al. assume an average wind

speed of 50 m s-1 and 74.6 on the parachute they calculate that the displacement will be

3.5 km. They do not explicitly state how the wind velocity is calculated. However

35

multiplication of the wind speed by the time on the parachute gives a displacement of 3.7

km. This approach assumes the parachute instantaneously starts moving at 50 m s-1.

There will be a delay before the parachute is being pushed along at the same speed as the

wind because it has mass. Therefore the displacement will be lower than 3.7 km as Wolf

et al. suggest. However they only consider a 2000 kg Lander. A human scale Lander (~50

mT) will take a much longer time before it is travelling with the wind. Consequently the

displacement will be less. Preliminary simulations by the author suggest the displacement

is at most ~km. See section 3.7 for details. This is still a significant effect though.

36

3. Design and test flying a virtual Mars Lander

A piloted Mars Lander is normally one of two basic types as described in human

missions to Mars. There is the Mars Excursion Module (MEM) which is a two stage

vehicle, like an Apollo LEM, the top stage is used for ascent to orbit after a short stay on

the surface. The second type is a Mars Transfer and Surface Vehicle (MTSV). This type

of vehicle has to transport the crew to the surface, doubling up as a habitat for the journey

to Mars and a habitat on the surface after landing. All these functions make for a very

massive Lander. Consequently it is harder to design.

The piloted Lander described here is an MTSV. It has to support a crew of four for six

months in space and one and a half years on the surface on Mars. It also has to brake into

orbit around Mars by skimming through the top of its atmosphere and then has to descend

to the surface touching down close to the Earth Return Vehicle (ERV). There were two

main requirements for the MTSV in terms of performance that helped to constrain the

design. During aerobraking and descent the g levels must not exceed 5 g for crew safety

and the MTSV must land close to the ERV (~100 m). In terms of properties there were

also two fundamental constraints based on assumed launcher capability (~25 mT). These

were mass (~50 mT spacecraft assembled in orbit using two launches) and diameter,

constrained by the size of the fairing which has a diameter of 4.5 m. The aim of the

modelling exercise, in the case, was to produce a Lander that was within in this

performance and properties “envelope”. Other design issues (such as solar panel

deployment, attachments for artificial gravity tether) were also tackled during the design

process to try and produce, together with the aforementioned constraints, a feasible

design for a piloted Mars Lander and integrate it with the rest of the mission.

Other drivers for the design approach of the Mars Lander was to demonstrate the

capability of the Orbiter space flight simulator (OSFS) for virtual prototyping, for the

presentation of the paper called “Virtual Prototyping of a human Mars mission” at the

Mars Society conference and fun!!

3.1 The Mars Society conference paper

A team for virtual prototyping (VP) of a human Mars mission was assembled by Bruce

Irving from the United States, an optical engineer with a passion for space exploration

and now a JPL Solar System Ambassador. Bruce has been closely involved with OSFS

testing and promotion being a member of the OSFS beta team, forum member and author

of the e-book, Go Play in Space, a guide for flying in the OSFS. Bruce flight tested the

spacecraft models built by co-authors, Andrew McSorley and Mark Paton. He was

responsible for investigating issues relating to artificial gravity generation, in orbit

assembly of the spacecraft elements and creating a base on Mars that looked “lived in”.

Co-author Andrew McSorley is from the United Kingdom and has also been involved

with testing and promoting Orbiter, with his website called Virtual Spaceflight and

contributing to the second edition of Bruce’s e-book. Also Andrew had experience in the

37

modelling of virtual spacecraft (specifically launch vehicles and OTVs) and using OSFS

navigation tools for interplanetary space flight. Andrew modified the Arianne 5 launch

vehicle for the project, adding textures and increasing the fairing size. He also built and

designed the booster stack required for TMI injection of the ERV and MTSV. A story

board was maintained by Andrew during the project on his website.

Mark Paton, a postdoc at the Finnish Meteorological Institute (FMI), Helsinki was

brought in to build the model of the MTSV and ERV and to provide technical input for

the design of the EDL, having previously modelled EDLs in OSFS for various robotic

missions and having also modelled the trajectory of a fireball during his PhD at the Open

University, UK. His background in planetary science provided some technical support for

the Martian atmosphere used in the OSFS. Mark also liaised with Andrew over docking

issues between the ERV and the Proteus booster stack and tether connection.

The fourth co-author was Grant Bonin, a Research Assistant and aerospace engineer at

Carleton University in Canada. Grant is the author of the MarsDrive reference mission

“Mars for Less”. MarsDrive is a web based group promoting human exploration of Mars.

Grant was consulted on technical issues during development of models.

An end product for our VP work was to be a paper to be presented at the Mars Society

Conference in August 2006. Another objective was to release the models for other people

to fly an end-to-end Mars mission in the OSFS. The OSFS was essentially chosen for this

work as it offered a good simulation environment. Probably most importantly it is free,

comprehensively tested by a large web based community and easy to use.

3.2 Building on a mission architecture

The purpose of the VP paper was a number of things, one of which was to demonstrate

OSFS as a powerful tool for prototyping the mission. However there was an opportunity

to iterate the design by e.g. coming up with a realistic proposal for a Mars Lander using

present of near-future technologies, particularly determining a sensible EDL technology.

Also other elements of the mission were iterated as we saw fit. This fitted in with the

overall theme of the paper quite nicely. Grant Bonin agreed to allow his mission concept,

Mars for Less, to be further developed in OSFS after being contacted by Bruce Irving.

The Mars for less architecture is similar to that of Mars Direct but has one essential

difference. The vehicles are not launched directly to Mars by a hypothetical Ares HLLV

(although NASA has plans to build one now) but are assembled in orbit using current

MLLVs like the Ariane 5 or the to be built Falcon 9. The MTSV and ERV are launched

as components (4) and assembled in orbit. A modular design is used to save on costs.

Both the MTSV and ERV have similar Lander components situated at the base including

landing legs and rocket engines. In Grant’s original Mars for Less paper a modular design

was also used for the habitats of the MTSV and ERV. However the design of the ERV

was modified extensively using a biconic shape for the VP paper.

The habitats and Landers rendezvous in orbit and are mated together. Before the human

38

components are launched, the booster stack has to be assembled in orbit. The boosters are

then used in an identical way as the OTVs in section 1.4 giving a kick and staging at each

successive perigee until TMI is achieved. These are the essential elements of the

architecture of the mission. More details on the mission can be found in the Bonin

(2006).

3.3 The Orbiter space flight simulator

The Orbiter space flight simulator was created for fun in 2000 by computer scientist, Dr.

Martin Shweiger from University College London in the United Kingdom. It has proved

to be a very popular amongst space enthusiasts. The software is under perpetual

development by its author with support from an enthusiastic beta testing team. It is free

for anyone to use and has a flexible interface allowing others to create add-ons to the

software. OSFS aims to be a realistic simulator. Gravitational motion and atmospheric

flight are accurately modelled. OSFS uses Newtonian physics and good graphics,

including virtual cockpits allowing a high level of immersion for the user. The software

has been presented by Dr. Martin Shweiger at the 2nd and 3

rd ESA International

Workshop on Astrodynamics Tools and Techniques (2004, 2006) and has been received

very well.

One of the most important add-ons is a module called “spacecraft”. It is now in its third

version “spacecraft3” and its developer goes under the name of Vinka. It allows rapid

modification of spacecraft parameters such as mass, thrust, aerodynamics etc by simply

changing the numbers in a text file. Without this the add-on maker has to recompile their

module each time they make a change to a spacecraft parameter. This was a key to fast

virtual prototyping in OSFS.

39

3.4 Virtual prototyping a piloted Mars Lander using Orbiter

The ERV lands first sufficient time (a couple of years) guaranteeing the production of

rocket fuel for the return journey to Earth. The MTSV is the second Lander to land on the

surface of Mars. In our virtual prototyping project the MTSV was chosen as the first

model to construct to make the most of the time available to us. The time was limited to

about four months, initiated in April 2006 with the conference presentation at the

beginning of August 2006. The MTSV is more challenging as a Mars Lander (than the

ERV) to design because of the crew’s sensitivity to g levels. The ERV’s purpose is to

land cargo on Mars and to safely deliver the crew into orbit around Earth. The

atmosphere of Earth is used for aerocapture of the ERV. The g levels experienced during

this time were investigated but are not reported here.

As well as delivering the crew to the surface the MTSV has to house a crew of four for

the six month trip to Mars and for a one and a half year stay on the surface. The

specification for the MTSV is outlined in Bonin (2006). The upper section is a two level

habitat and the lower section is a Lander module comprising of the rocket descent

Figure 3.1 Screen shot from the Orbiter space flight simulator. This shows a view from

inside the cockpit of a fictional spacecraft called the Deltaglider, which is provided with

the Orbiter base package. The view also shows the blue oceans of Earth and some white

streaks of clouds and on the horizon is a band of atmospheric haze. The virtual cockpit

is interactive; buttons can be activated with the mouse. The knees of the pilot can also

be seen. This image is taken from the Orbiter web site, www.orbitersim.com.

40

engines and landing gear. A heat shield is attached to the bottom of the Lander module

intended for both aerocapture and descent to the surface. In Bonin (2006) the heat shield

is assumed to be rigid and fully deployed in Earth orbit.

To model the MTSV a powerful freeware tool called anim8or (figure 3.2) was used by

the author. As its name suggests it is used to make 3D animations. However it is perfectly

good for shaping and texturing stand alone objects, or meshes. The meshes are saved in

3ds format but can be imported into the OSFS msh format using a handy conversion

program developed by Vinka.

The habitat and Lander are roughly based on a cylindrical design similar to that of Mars

Direct habitat except it is narrower in MFL due to the launcher fairing constraint of 4.5

m. See figure 3.3. The MTSV consists of the Lander, at the bottom, a garage housing a

rover and the habitat on top. The Lander element is 3 m high with a 2 m high garage

above containing a rover. The habitat is about 7 m high, giving ~110 m3, giving about 25

m3 per crew member. NASAs DRM 3.0 specified 90 m

3 per crew member for long

duration spaceflight (Drake, 1998). The MTSV was increased in diameter (6.5 m) to

increase the living space inside the habitat to 232 m3 or 58 m

3 per person. The living

Figure 3.2 The MTSV model in anim8or. This is the finished mesh of the MTSV. The 4x100

kN engines are clearly visible at the bottom. The model consists of many 3D objects made

using anim8or’s lathe. To make an object a line is drawn and then rotated around one of the

axis. To make a cylinder a line is drawn parallel to an axis and just rotated around the axis.

Indented features like the window or the gear section are manually ‘sculpted’ by moving

individual mesh points (not visible here). Meshes can be manually ‘coloured’ or images (like

the foil on the Lander) can be imported to give the mesh texture.

41

Habitat

Garage

Lander

Pilot’s window

Crew

Ladder

Upper floor

Lower floor

Upper floor

space was also augmented, on the surface by an inflatable structure. To launch the larger

habitat the fairing of the launch vehicle (Ariane 5) was modified by Andrew McSorley.

Even though the exterior of the habitat was considered higher priority (i.e. for image and

movie production) some time was spent designing the interior to give some sense of

immersion. The interior habitat was split into two levels with an opening in the dividing

floor for a connecting ladder. The upper level was allocated to personal space and crew

accommodations (galley, beds, and hygiene) and the lower level was allocated to the

airlock, laboratory and storage. During aerocapture and descent to the surface it was

thought that the crew should be on the lower floor. This was primarily a centre of mass

consideration. During entry the closer the COM is the heat shield (fitted to the base of the

Lander below the habitat) the greater stability during flight. Also stability on the surface

was a concern upon landing, especially when the fuel in the Lander has been exhausted.

A lower centre of mass would increase the MTSV’s stability on sloped terrain. A window

was placed on the lower level allowing identification a safe spot to put the MTSV down,

if required. A window on the upper deck would give a greater view due to the

advantageous perspective. A single angled window was inserted in the MTSV hull at a

height that a crew member could stand and look out and down, similar to the Apollo

LEM design. This window was then identified as the “nose” of the vehicle axis used in

OSFS. Basically this would mean that in cockpit view the pilot would be looking through

this window and in the direction of motion when applying thrust to move forward.

Figure 3.3 MTSV in Orbiter Space Flight Simulator. The left hand image shows the MTSV just

after landing on the surface of Mars. The landing legs are deployed in this image. In figure 3.1

the landing legs are in their storage position. Animations are used (by editing a text file that is

read by Vinka’s spacecraft3.dll) in Orbiter to deploy the legs and other. A virtual cockpit is

shown on the right with four crew figures on seats. Mounted on the window are two multi-

functional displays and a head up display to show flight information (just visible in top right

image). The bottom right image shows the ladder and entrance to the top floor.

42

The side of the habitat is tiled to aid thermal management during atmospheric entry. The

habitat outer shell is essentially equivalent to an aeroshell backshell except it is integrated

onto the structure to save mass. To control the attitude of the MTSV four clusters of

attitude thrusters (similar to the Apollo Lunar Lander) were placed at the four “corners”,

one each side of the habitat’s window and two on the opposite side. There was some

angst over where to place these thrusters, due to efficiency issues arising firstly from the

centre of mass position and also from the downward thrusters exhaust impinging on the

heat shield. For a while the thrusters were placed on the end of four solar panels deployed

so they would be positioned over the rim of the heat shield as in figure 3.4. It should be

pointed simulations of the RCS thrusters did not involve measuring any moments of

inertia. Their positioning was assessed from visual inspections alone.

The heat shield proved to be troublesome because it also obscured sunlight from the solar

panels in some orientations. Eventually an inflatable heat shield was substituted for the

rigid one. Before arriving at Mars the heat shield would be stowed beneath the MTSV

(figure 3.5). This then simplifies the RCS design and gives the solar panels unrestricted

access to sunlight. The risk of mission failure due to failure of the inflatable heat shield

could be mitigated by including some redundancy such as extra inflation gas and heat

shield. Calculations of the heating during entry showed the temperatures would be low

enough to support the choice of an inflatable device (Irving et al., 2006).

The solar panels are deployed from below the habitat as described in Bonin (2006) and

shown in figure 3.4. At first it was decided to have four sets of panels, to make up the

required area. This was reduced to two deployed from each side of the habitat window.

During artificial gravity tests (see figure 3.5), using the tether it was found that the extra

pair would be of limited use due to the orientation of the spinning habitat. A ring of

Figure 3.4 Reaction Control System (RCS) design. Figure A shows an early iteration of the

MTSV. Clusters of thrusters were placed on the end of the solar panels to avoid the rigid heat

shield and give greater leverage (like attitude gas control jets on the Viking orbiter). However this

created further issues such as flexing of the panels. The RCS was beefed up to double up as deorbit

engines and forward propulsion on Mars. The final design shown in figure B saw the RCS thrusters

rigidly mounted onto the side of the habitat. The rigid heat shield was replaced with an inflatable

shield deployed just before entry into Mars atmosphere.

A B

43

auxiliary panels were added onto the shell of the garage section to beef up the collection

area of the deployable panels and to give some redundancy in case of redeployment

failure after aerocapture or descent.

During aerocapture the panels have to be retracted, and then extended for reuse in orbit

around Mars. They are also retracted again before entry, descent and landing and then

extended for reuse on the surface. Supporting cables run from the end of the panels to the

top of the MTSV habitat. These are to support the flimsy nature of the panels against

flexing due to artificial gravity generation on the way to Mars and due to the Martian

gravity once landed.

For constant, direct two way communications with Earth two high gain antennas were

placed on the roof of the habitat. These could be folded down against the roof during

entry into the Martian atmosphere and then erected once on the surface. Also on the roof

were two low gain omni-directional antennas and the parachute canisters. All these item

are positioned around a docking hatch. This hatch is used to transfer the crew to the

MTSV from the Crew Expedition Vehicle while in Earth orbit. A tether attachment

device was also housed on the roof. When deployed it was positioned above the docking

hatch and supported by an open framework to give clearance for EVAs if necessary. The

Figure 3.5 Tether for artificial gravity. This image is from Irving et al. (2006) and shows the

MTSV connected to the last stage from the TMI booster stack (model built by Andrew McSorley).

To generate 0.4 g the tether would need to be 1.5 km long completing a rotation about once every

five minutes. The long tether is to reduce gravity gradients enough so they are not sensed by the

crew. Tether experiments were conducted by Bruce Irving to investigate solar panel shadowing and

orientation issues. Solar panels were reduced from four to two as can be seen in figure 3.3. The

tether is an OSFS add-on made by “MattW”.

44

Martian surface access hatch, positioned on the side of the MTSV, would have an airlock

behind it. This would presumably be more convenient for conducting EVAs during the

voyage.

Just before touchdown the landing gears are deployed. These were modelled on the DC-X

type deployable gears. The reusable rocket was destroyed when one of its landing legs

failed to deploy just before touchdown and it toppled over. In hindsight it may be prudent

to deploy the landing gear in Earth or Mars orbit allowing repairs or mission abort.

Recent reusable rocket designs (ISAS/JAXA’s RVT and Blue Origin’s New Sheppard

prototype) feature fixed landing gear.

On the surface the MTSV is a home for the crew and is part of a Martian base (see figure

3.6). For the project the base was named Mandya Arti (suggested by Mark Paton) after a

location in the Australian aboriginal story called How the Hills Came to Be. In this story

the hills are created by a buck euro (similar to a kangaroo) called Mandya when he blows

on a stone he finds in his wound after having a fight with a buck kangaroo called Urdla.

The hills are red because of the blood from Manya’s wound. We thought that this was a

good name because Mars is associated with war (which involves fighting), has a red

surface and in the OSFS add-on developers have to create their own hills using meshes as

planets are modelled as smooth spheres. Also Vallis Dao is situated on the edge of Hellas

Basin, a large deep depression in the southern hemisphere high-lands. As a planetary

feature it is somewhat similar to the Australian continent on Earth.

Figure 3.6 Mandya Arti. Image of the base (Irving et al., 2006). The MTSV with solar panels

extended dominates the scene. There are several figures in futuristic skin tight space suits

(courtesy of add-on maker Greg Burch) dotted around the base. They help show the scale of the

Mars Lander (the MTSV). Note the living space of the rather narrow MTSV is augmented by an

inflatable structure behind the MTSV. On the right is a pressurized rover (Andrew McSorley)

and in the background the ERV (Mark Paton) can just be made out. The base also features domed

greenhouses and a nuclear generator.

45

3.5 Software for analysing EDL systems: A2D

There are well tested high fidelity astrodynamical tools available for prototyping EDL

systems for robotic mission to the planets (e.g. see Balaram et al., 2002). These can

model 6-degrees of freedom movement, simulate spacecraft thrusters and test guidance

systems. However this level of sophistication is not required at the proof of concept stage

for piloted Landers as described in this work. So a bare bones astrodynamical tool called

Aerobrake2D (A2D) has been developed by the author evolving from fireball trajectory

modelling work (Paton, 2005).

A2D simulates the motion of an object under the influence of gravitational and

atmospheric drag forces. Newton’s equation of motion is used, converted into 2D polar

coordinates and expressed as four first order equations. The drag equation is used to

model aerodynamic drag forces. Lift is simulated by adding a force perpendicular to the

drag force and with a magnitude depending on a specified lift over drag (L/D) ratio. The

simulation is coded in FORTRAN 77. The integrated Force 2.0 (Lepsch, 2005) editor and

compiler windows environment is used by the author as it allows rapid editing, compiling

and execution of code in a windows environment, great for development.

The first order equations of motion are iterated during each time step. For each discrete

time step there is a step increase in the following, radial distance from the centre of the

planet, radial velocity from the centre of the planet, angular distance around the planet,

angular velocity around the planet. This approach is fairly accurate for a few orbits

around a planet such as the Earth or Mars. For example a body orbiting the Earth starting

at an altitude of 200 km, an initial speed of 7.9 m s-1 and a velocity vector angle of zero

will increase its semi-major axis by 8 m each orbit when using a time step of 0.01 s. The

program takes about 43 s to run that simulation. The error increases to ~km when using a

time step of 0.1 s but the simulation takes only 5 s. Using a 0.1 s time stop is certainly

accurate enough for modelling a short path such as during aerocapture or EDL. Once a

system has been defined the error due to numerical integration can easily be controlled

using the time step with an error of around one metre at the end of the EDL path.

The atmosphere is modelled by dividing the atmosphere into levels (see figure 3.7). For

each level a temperature and wind speed are imported from the Mars Local-Area Model

(MLAM), a mesoscale simulation of the Martian atmosphere or can be defined manually

via a parameters (P) file. The atmosphere gas constant, surface pressure and surface

temperature are also defined in the P-file. These values are used to calculate the surface

density using the ideal gas equation. The density at each level is calculated from the

surface upwards using the level temperature and using hydrostatic theory. An infinite

number of levels can be defined. However MLAM uses 32 atmosphere levels, which

dictates the number of levels in A2D as well. A2D has the same level values all around

the planet which is unrealistic. MLAM uses gridboxes of kilometre size in the horizontal

directions. An EDL path can run over 1000 km in distance so eventually A2D will need

to eventually mirror this fact in its modelling of the atmosphere.

46

Upper Stratosphere

Troposphere

P=101.3 kPa

T=15°C

P=22.6 kPa

T=-56.5°C

10

20

30

40

0

50

0 20 -100 -80 -60 -40 -20

Altitude (km)

Temperature (°C)

Lower Stratosphere

Level 1

Level 2 Level 3

Level 4

A2D can simulate an entry vehicle, with detachable heat shield, two parachutes and a

Lander with powered descent. The vehicle properties are defined in the same P-file as the

atmosphere. Lander effective area, mass, fuel mass, rocket thrust, isp, powered descent

initiation and target touchdown speed are defined first. Second the shield effective area,

mass, angle of attack (to calculate drag reference area), L/D are defined. Following the

shield properties is a section that simulates bank modulation of the lift vector as used for

guidance and targeting of the Apollo CM. It is simply a list of times a roll angles that is

used by A2D to calculated to vertical component of lift. This was included primarily to

try and validate A2D against Apollo era flight and simulation data. This has had limited

success. Figure 3.8 shows a comparison with simulation data by Young and Smith

(1968). They ran several simulations to explore guided Apollo CM type trajectories. The

altitude with time profile appears to be a good fit. There is a discrepancy of only 1.8

miles at the end of the EDL path. This is the order of error expected from A2D when run

at a time step of 0.1 s. However Young and Smith denote their results in units of nautical

miles. If A2D results are converted into this unit then a discrepancy of well over 100 km

exists. This result is not mirrored in the tightness A2D results follows Young & Smith’s

profile of altitude with time.

Figure 3.7 Variation of temperature with altitude in the U.S. standard

atmosphere. Level numbers are for illustration purposes, in reality 10

levels have been used for Earth and 32 levels have been used for Mars.

47

After the roll commands section the shield properties continues with a definition of the

shield nose radius. This is used to calculate thermal heating of the heat shield. Following

this is a shield release event number. This is used to determine what event (e.g. parachute

deployment, powered descent start) is used to trigger the heat shield release timer. The

time delay between this EDL event and heat shield release is defined next. At heat shield

separation A2D calculates if the dynamic pressure is low enough on the shield to allow

positive separation between the shield and the Lander. If the drag is too high the shield

will not be released and a message will be output indicating this.

Size and mass information for the parachutes together with deployment trigger speeds

follow the heat shield definition section. The atmospheric entry speed, angle with the

horizon and speed is after the heat shield section. This is followed by the numerical time

step and whether the EDL data should be output to file or not.

Further validation of A2D with reconstructed best estimate trajectory data from the

Apollo 6 mission is shown in figure 3.9. While the match is close there are noticeable

Figure 3.8 A2D simulation compared with an Apollo era simulation by Young and Smith (1968). The

top chart shows the roll angle of the command CM. A2D is the red line and approximated the roll angle

with step changes. The lower diagram shows A2D altitude data superimposed on data from Young and

Smith. The entry information is as follows: Speed, 10.972 km s-1, altitude, 121.92 km, angle, -6.5°,

ballistic coefficient, 322 kg m2 and L/D, 0.5. In this simulation the rotation of the atmosphere with the

Earth’s surface was not modeled but can be easily implemented. See figure 3.9.

A2D

A2D

48

deviations of A2D from the data. This is probably due to the approximate way A2D

models the guidance computer roll commands and the fact that a real atmosphere will be

different from the standard model.

Figure 3.9 A2D simulation compared with data from Apollo 6 CM test flight. Apollo 6 was then last

test mission before the manned flight of Apollo 7. Figure A shows A2D placed over the altitude data

during the guidance phase of the Apollo CM. There is clearly a good fit with some deviations. These

are probably due to the approximate way A2D has modeled the roll commands. These are shown in

figure B. The curved sections have been approximated by step functions. The L/D has been modeled

as a function of Mach number starting at 0.3 at Mach 30 rising to 0.45 at Mach 5. This was found to

give a good fit. However the data in the literature suggests an L/D of 0.35 at Mach 30. The drag

coefficient was also approximated as a function of Mach increasing from 1.0 at Mach 5 to 1.28 at

Mach 30. Rotation of the atmosphere was included in A2D for this simulation.

49

3.6 Aerocapture vehicle design and aerodynamic properties

The mission architecture as described in Bonin (2006) requires the use of aerocapture to

decelerate the MTSV into orbit around Mars before descending to the surface. To

investigate the g level loading on the astronauts and determine the vehicle type (blunt or

slender body) during this phase a computer model of the habitat module with aeroshell

was developed by the author in FORTRAN called aerobrake2D. A validation was made

by comparing the results with those in OSFS using the same model of the MTSV. It was

found that neglecting the rotation of the Martian atmosphere in aerobrake2D gave a 10 %

difference in the semi-major axis of the capture orbit compared to OSFS. When the

rotation of the atmosphere was included the difference in the semi-major axis predicted

by aerobrake2D and that in OSFS went down to 0.2%. This was deemed to be acceptable

for setting up test scenarios in OSFS and experiment with the aerodynamic model.

The Martian atmosphere, like the Earth, can be split into distinct isothermal or adiabatic

sections. Within these levels the temperature will either increase or decrease with altitude

(adiabatic) or it will remain constant (isothermal), effecting the thickness (or density)

with height. In OSFS five levels are modeled based on the real Martian atmosphere. In

OSFS the atmosphere cuts to vacuum at 100 km. For an entry vehicle with a high enough

ballistic coefficients, such as the MTSV, this has little effect as significant hypersonic

decelerations start between 60 and 70 km.

A spacecraft approaching Mars on a hyperbolic trajectory has to reduce its velocity

enough so it enters into an orbit. With aerocapture this is done by passing through the

atmosphere and using drag forces for deceleration. For crewed missions the deceleration

must be kept below a certain level. A crew that has been in micro-gravity will suffer from

muscle wastage, including the heart. For a deconditioned crew the maximum is between

3 and 5 g. Therefore it is important to fly the spacecraft along a path that minimizes the

forces on the crew. The forces experienced by the crew will depend on the entry angle,

the entry speed (which depends on the approach speed to Mars and the gravity of Mars),

the desired target orbit (with a small eccentricity for large payload advantage of an all

propulsive capture) and on the aerodynamic properties of the vehicle (specifically the lift

over drag ratio). Implementing artificial gravity during the trip will help preserve the

crew’s physical strength for the descent and also for working on the surface.

To determine if a low lift blunt body type of vehicle could be used for aerocapture the

dependence of corridor width on L/D was investigated. The calculation of the corridor

widths also determined if the decelerations during aerocapture were low enough for a

human crew. The vehicle used was an Apollo type blunt body heat shield with a L/D of

0.4. The mass was 46 mT and the effective area was 242 m2 (CD=1.242 and

diameter=15.75 m). Three values of L/D were used. These were L/D=0 at an angle of

attack of 0°, L/D=0.25 at an angle of attack of 25° and L/D=0.4 at an angle of attack of

40°. The area of the shield was settled upon after performing these tests.

This entry corridor width was determined using A2D. A vehicle with a fixed L/D was

flown several times into the atmosphere of Mars from a start point (or entry point) of 125

50

km altitude and a start velocity of 7 km s-1 which corresponds to a low energy transfer to

Mars. The velocity angle vector or the entry angle (relative to the horizon) was varied

until the vehicle was successfully captured into orbit. Then the entry angle was decreased

(more negative) until the vehicle experienced peak deceleration of 5 g (limit for a

deconditioned crew). This was then the undershoot angle. Note in this case the

undershoot case is not the entry angle for which the vehicle would impact the surface.

Once the undershoot entry angle had been established the entry angle was increased (less

negative) until the vehicle was no longer captured into orbit and flew off into space. The

g level never exceeded the limit so this was then the overshoot angle. There was no

accounting for uncertainties in navigation, atmosphere density or the aerodynamic

properties of the vehicle during this process.

The aerocapture undershoot and overshoot boundaries were plotted as shown in the figure

3.10. An uncertainty of 0.4° in entry angle was assumed (from navigation, atmospheric

and vehicle aerodynamic property uncertainties). This then makes the effective (safe)

corridor 0.8° smaller than the actual corridor as shown in figure 3.6. The information can

then be analysed to determine the best vehicle L/D value to use. The corridor for an L/D

of 0 is only 0.2° which is too small to be certain of a successful aerocapture. It may be

important to note the lower boundary of this particular corridor does not reach the 5 g

limit. This is because the search algorithm used to find the boundaries used 0.1° step, a

little coarse but good enough for this purposes. In any case it is clear that the corridor

widens with increased lift capability. An L/D of 0.25 opens up the corridor wide enough

for a successful capture into orbit. To reach the target orbit (e=0.2) the vehicle has to

reach the entry point at 10.7°. Therefore a blunt body design with an effective area of

~250 m2 (as opposed to a slender body design) could be used for an aerocapture vehicle

into Mars orbit.

51

-12.7

-12.2

-11.7

-11.2

-10.7

-10.2

Undershoot: 3.1 g peak

Overshoot: 2.43 g peak

direction of motion

aoa: 0 Mass: 52 mT Effective Area: 252 m

2

Entry speed: 7 km s-1

Entry height: 125 km

0.2° corridor (not usable)

Heat shield

Lander

-12.7

-12.2

-11.7

-11.2

-10.7

-10.2

Undershoot: 4.92 g peak

Overshoot: 1.41 g peak Lift down: L/D=-0.25

Lift up: L/D=+0.25

direction of motion

2.12° th

eoretical

corrid

or

Target orbit e=0.2 Peak: 3 g 1° usable

corridor

aoa: -25°

direction of motion

aoa: 25°

Lift down

Lift up

-12.7

-12.2

-11.7

-11.2

-10.7

-10.2

Undershoot: 5.0 g peak

Overshoot: 1.41 g peak Lift down: L/D=-0.4

Lift up: L/D=+0.4

2.65° th

eoretical corrid

or

Direction of motion

aoa: 40 °

1.9° usable corridor

Direction of motion

aoa: 40 °

Lift up

Lift down

Figure 3.10 Entry corridors. Entry corridors were generated by flying a simulation of

a blunt-body MTSV with different L/D values. The target orbit can be achieved with

certainty if the L/D is bigger than 0.25.

52

The MTSV consists of four components that are assembled in Earth orbit before traveling

to Mars. These are, from Bonin (2006), the habitat, garage, Lander and aeroshield (heat

shield). In OSFS they are assembled into a whole spacecraft using the docking ports

facility. Each component has its own set of properties such as mass, drag, lift etc. For a

single spacecraft, forces will act on the geometric centre of the mesh. For a multi-

component spacecraft there will be an effective centre where forces will act. Table 3.1

lists the mass of each part of the MTSV together of the positions of each centre of mass

relative to the origin of the heat shield. The centre of mass is calculated to be 5.38 m

behind the centre of the heat shield.

It is essential for an entry capsule to maintain a forward facing heat shield for continuous

thermal protection and for providing an effective drag surface. Stable flight can be

achieved when the centre of mass is forward of the centre of pressure. Therefore it was

desirable to define the MTSV model properties so the centre of mass was in front of the

centre of pressure. The main pressure on the MTSV model in OSFS will be drag

pressure. In reality lift would contribute but to keep the number of variables to a

minimum (also OSFS only has a static atmosphere) only drag is modeled. The drag

depends on the effective surface area (drag coefficient multiplied by the cross-sectional

surface area). The effective surface area for the heat shield, normal to the z (long) axis is

242 m2. The effective surface area of the other MTSV component (Lander module,

garage, and habitat) was kept low so as not to add to the drag forces. The point where the

drag acts on the shield was moved backwards by 10 m (by editing the position for the

vehicles airbrake in the spacecraft3 text file), placing it 4.22 m behind the centre of mass

of the MTSV. The centre of pressure (or drag in this case) was placed at this point where

guided by flight tests in OSFS. Figure 3.11 shows a schematic diagram of the MTSV

model derived from aerocapture flight tests.

Component Total mass / kg Centre of mass along z axis /

m

Habitat 24200 8.43

Garage 7500 4.045

Lander 8500 1.56

Shield 5800 0

Total 46000 5.38

Table 3.1 MTSV model properties in OSFS. The total mass means

empty mass plus fuel mass. The centre of mass is measured from the

origin of the axis of the heat shield. The centre of drag of the heat

shield has been moved back by 10 m to increase stability.

53

Centre of pressure

Centre of mass

A

B

C

D

-z axis for shield

Zshield=0

Zshield =+10 m

Zshield =+5.38

m

RCS thrusters

Pilot’s window

Figure 3.11 Aerodynamic stability. MTSV location of the centre of mass

(COM) and the centre of pressure (COP) for aerodynamic stability in Orbiter.

Locations A, B, C and D show the centre of mass for the shield, Lander, garage

and habitat respectively.

54

Hellas Basin

Southern

hemisphere

Northern

hemisphere

MARS

(upside down)

Aerocapture orbit (e=0.2,

q=100 km)

Descent orbit,

q=30 km

Approaching

Vallis Dao

Vallis Dao & landing site

(Mandya Arti)

Main chute deploy

Heat shield release &

PDS start

Landing

Mars surface

ERV Heat shields

Shadow from chute

MTSV

Deorbit burn

3.7 Entry, Descent and Landing System design

Entry, Descent and Landing Systems have to reduce a spacecraft’s velocity so a

survivable impact or landing can be made within a predefined target area. Also for a

human Lander the g levels should be constrained for human comfort during the descent

(<5g). For our Mars Lander three types of decelerators were used, deployed in a

sequence. These were a heat shield, parachutes and rocket engines. An expected

sequence, say for the robotic Viking Lander, would be peak hypersonic deceleration and

heating at about 30 km (~2.5 minutes after entry), deployment of parachute and release of

heat shield at 6 km (~6 minutes after entry) and initiation of rocket powered descent at

1.5 km (~9 minutes after entry) and landing about a minute later.

Before descent the MTSV was in a fairly circular orbit with an eccentricity of 0.2 and

periaposis of 100 km. This was found to be easily achievable using aerocapture technique

described in the previous section (to within 0.2% of the predicted semi-major axis). To

initiate descent the periaposis was reduced from 100 km to 30 km using the MTSV

rockets at apoapsis. By the time the MTSV reached 30 km it would be experiencing

maximum deceleration. It was found that the hypersonic deceleration load was within

limits for human comfort. The descent sequence is shown in figure 3.12.

Figure 3.12 Descent trajectory. The MTSV ignited its engines to perform a deorbit burn. The MTSV then

begins its atmospheric entry at an altitude of 100 km traveling at 3.5 km s-1, a 30 m DGB parachute is

deployed at about 6 km altitude, at an altitude of 1.5 km the engines are ignited pulling the MTEV off the

heat shield, at about 400 km the parachute is released and the MTSV heads off to land next to the ERV.

The image of the Mars globe is painted with the MOLA elevation map. Blue means low and red means

high elevations.

55

We found during descent experiments that the parachute was not large enough and did

not provide enough drag to pull the Lander of the heat shield. The parachute diameter

was fixed at 30 m as a practical maximum size. The heat shield was also fixed in size

from the aerocapture experiments. It was thought perhaps the parachute area could be

increased just a little bit but the problem was more complicated than it seemed and the

parachute would have to be huge.

This may be realised by taking the Viking Lander as a reference and increasing the

Lander mass while maintaining the ballistic coefficient (mass divided by area) under the

parachute. The radius of drag area scales in proportion with the square root of the mass

when the ballistic coefficient is kept constant. Therefore if the Viking Lander (~800 kg)

is scaled up to a piloted Lander (~50000 kg) its 16 m diameter parachute increases to

~100 m. The disadvantage with a large parachute is that it takes a long time to deploy and

is literally a huge unknown as this size of supersonic parachute has never been tested.

Another approach is to reduce the size of the heat shield so it falls away faster relative to

the Lander. With a 30 m parachute the heat shield should only have to be double the

Viking Lander (~4 m) at 8 m to obtain positive separation between the shield and piloted

Lander at shield release. However this is not large enough for aerocapture. A small heat

shield means that the MTSV has to travel into the denser part of the atmosphere and so

experiences higher decelerations. Also a small heat shield may not fully protect the

Lander from the hot plasma generated during hypersonic deceleration.

The solution to the problem came from Grant Bonin, the author of Mars for Less, who

suggested using the Lander’s rocket engines to help pull the Lander away from the shield.

This was tested and the extra deceleration was enough to pull the Lander off the heat

shield. Figure 3.13 shows the drag forces on the heat shield and the Lander shortly after

starting the engines and heat shield release. It is clear without the extra kick the drag

forces alone cannot separate the MTSV and heat shield.

56

The next task was to determine the optimum altitude for the release of the heat shield and

the beginning of the terminal descent phase. For this purpose A2D was modified to

simulate entry, descent and landing. The program was updated to change the parameters

of the vehicle such as drag area and mass (for parachute deployment and shield release)

triggered by reaching a specified velocity or altitude. For the powered descent a simple

algorithm was implemented. Once ignited the engines were throttled at maximum thrust

until the target touchdown speed was reached. The velocity was then maintained at this

value until touchdown. During the descent the amount of fuel used was also calculated

which manifested itself as a decrease in the thrust during a constant velocity descent.

Several trial flights were performed to find the optimum altitude for engine start and

shield release. A trial flight that produced a constant velocity descent rate from 10 m

above the surface was chosen. From this the velocity and altitude triggers could be

programmed into the MTSV autopilot and flown in the OSFS.

Discrepancies in EDL results from flying the MTSV in A2D and OSFS led the author to

Figure 3.13 Heat shield release. The MTSV engines are fired to enable positive separation of the

heat shield. The 3D arrows superimposed on the MTSV and heat shield are a feature of OSFS

2006. They are visual aids for monitoring the forces and locating the axis. The drag force on the

MTSV with the parachute deployed is 133 kN and the drag force on the heat shield is 38 kN. The

MTSV deceleration without aid from the engines would decelerate slower than the heat shield.

The MTSV weighs in at 46 mT while the heat shield is about 6 mT. With the 400 kN engines

firing this gives plenty of additional kick for positive separation.

57

compare the atmosphere models. This revealed an inconsistency (step change) in the

density-altitude profile in the OSFS. The creator of OSFS was informed and kindly (and

quickly) updated the module plug in (dll) for the Martian atmosphere.

During EDL simulations in aA2D it was noticed that the g level at parachute deployment

was exceeding 5 g (see figure 3.14), possibly to high for a deconditioned crew. To ease

the load on the crew a drogue parachute with half the area of the main 30 m parachute

was added into the EDL sequence before the main parachute. It acted to reduce the

velocity of the MTSV so the load on the main parachute (and g levels) would be lower.

The drogue parachute had to be deployed at Mach 3.1 implying significant amount of

heating and high dynamic pressure. Attempts were made to calculate the temperature of

the parachute and investigate if a parachute could be used at this velocity on Mars.

The stagnation point (in front of the shield nose) temperature of a Mars Pathfinder

aeroshell travelling at Mach 3.2 will be 821 K from a Sutton-Grave correlation, assuming

a nose radius of 0.375 m and with an ambient atmosphere temperature of 195 K. A

parachute is a completely different (inverted) shape and size so this type of analysis may

not be useful. Indeed the Sutton-Grave correlation shows that the temperature is

dependant on nose radius, the smaller the radius the higher the stagnation temperature.

Also the parachute is travelling at speeds in the supersonic velocity region (or very

nearly, the boundary is very nebulous), not in the hypersonic region where the thermal

physics is more complicated. At hypersonic speeds (>Mach 5) heating of surfaces is

complicated as the high temperatures change the gas chemistry. At very high Mach

numbers heating is also due to radiation as well as convection. At supersonic speeds

Figure 3.14 G levels at parachute deployment. The dotted line shows the g levels when using a single

parachute. The solid line shows the g levels when using two parachutes. The inset shows a broad bump

which is due to hypersonic deceleration during entry. The data is from the author’s simulation program

called A2D. The MTSV mass is 52 mT and the parachute areas are 525 m2 (20 m diameter with CD of

1.5) for the drogue parachute and 1050 m2 (30 m diameter with CD of 1.5) for the main parachute

58

(<Mach 5) aerodynamic heating occurs over flat adiabatic surfaces (like wings) when a

gas is bought to rest via friction. Assuming the sides of the parachute are nearly flat (not a

bad approximation for a cross parachute, for example) the temperature at the stagnation

point will be 569 K. A simple analysis can be made by assuming all kinetic energy is

turned into heat. This gives an average temperature of about 1619 K. This would be the

maximum temperature possible.

Clearly the parachute will experience high temperatures. An accurate analysis involves

numerical simulations such as undertaken by LaFarge et al. (1994). They conducted

dynamic pressure and temperature analysis were on a cross parachute at Mach 4. A

stagnation temperature of 915 K was found with temperatures reaching 904 K at the

skirts of the canopy. The majority of the structure experiences temperatures greater than

750 K. Nylon has a melting temperature of 522 K and Kevlar has a melting temperature

of 750 K and both materials are unsuitable. LaFarge et al. decided to only consider a

Mach 3 case. For the Mach 3 case they calculated an average canopy temperature of 373

K, significantly less than the Mach 4 case. However our EDL design demanded a >Mach

3 so some solution had to be found to enable the parachute to operate under high

temperatures.

Inflatable heat shield technology involves impregnating the fabric with an ablative

material around the skirt of the canopy (see figure 3.15). At high temperatures the

ablative material decomposes so limiting the heat entering the capsule’s interior. It was

thought impregnating the cross parachute with ablative material would allow it to operate

at high Mach numbers. The dynamic pressures were not so much of a concern. A

parachute deployment at 14 km altitude generated a loading of about 4 kN m-2. La Farge

et al. (1994) observed peak stresses of approximately 93 kN m-2 in the canopy and 27.6

kN m-2 in the suspension lines. Under these condition they concluded that the cross

parachute should preserve structural integrity.

The main 30 m parachute was released at Mach 2.2. This is inside the Viking heritage

parachute box and so ablative materials are not required for cooling. Keeping the main

parachute attached during the powered descent phase reduced the amount fuel used,

allowing for greater target cross-range and down-range correction capability. The main

parachute was released at some time just before touchdown. The MTSV could then fly

out from under the parachute in search of the ERV before the parachute falls on top of it.

59

The idea for the cross parachute design came after the presentation of the virtual

prototyping paper at the Mars Society Conference. However a few weeks before the

conference, Bruce Irving, suggested that we should use an inflatable heat shield for the

MTSV. An inflatable heat shield is stowed until approaching Mars. This solved a lot of

operational problems, like the positioning of the RCS thrusters and the shadowing of the

solar panels. A thermal analysis was made by the author on a heat shield with a nose

radius and found that the temperature, using a Sutton-Grave correlation and a nose radius

of 10 m is about 900 K. The heating rate is about 17 W cm-2. During an inflatable

technology demonstrator mission by esa (esa, 2005) the maximum heating rate was 35 W

cm-2. It was concluded that such a heat shield could realistically be used. The low

temperature is due to a combination of a large heat shield and low entry speed from orbit

(3.5 km s-1). It may be interesting to note, as a comparison, that the Viking entry was

from orbit around Mars and the peak heat flux was only 21 W cm-2. It was a simple

matter of replacing the mesh of the rigid heat shield with a new mesh that looked like it

was inflatable. The mass and area values were kept the same as the time was limited to

work out the numbers for a new EDL sequence.

We chose our landing site to be in Vallis Dao on the edge of Hellas Basin. The valley

floor is much lower than the surrounding highlands giving a greater distance to decelerate

(and denser atmosphere). In the OSFS it is possible to render landscapes as textured

meshes. Planets are represented as a smooth sphere textured with an image based on

spacecraft images. The mesh then sits on this sphere, the surface of the sphere being a

solid surface. On Mars in OSFS there are no southern hemisphere lowlands or northern

hemisphere high lands as on the real Mars, although it can be done if a mesh of the whole

planet is made. Having 3D terrain went some way to making the final approach and

landing more realistic.

Figure 3.15 MTSV cross parachute. A cross parachute, impregnated with an ablative material (coloured

black), was added when it was realized there may be high temperatures (and high mechanical loads).

60

Modelled inflation sequence

Real inflation sequence

A B C

AIR

Fully deployed

Fully deployed

During a recent examination (while writing this account) of the parachute properties used

for EDL simulations it was noticed the drag coefficient was incorrect for a Disc-Gap-

Band (DGB) parachute depicted in the OSFS models. A DGB parachute has a hole at the

top of the canopy with a gap between the top and bottom of the canopy. The space at the

top of the canopy allows air to escape during inflation and so reducing the loading on the

parachute structure. A parachute canopy without any spaces will have a higher drag

coefficient as the drag area will be larger than a DGB type of the same radius. It was this

type of parachute, with a drag coefficient of 1.5 that was used in the work described in

section 3.7. Also see figure 3.12. The effective area (drag coefficient multiplied by actual

drag area) used in the simulations was 1050 m2. The actual physical cross-sectional area

presented to the air flow is then 700 m2 giving the parachute a diameter of 30 m which is

depicted in the OSFS models. The drag coefficient of a DGB parachute is only 0.55. This

type of parachute will need to be 50 m in diameter to make up the same physical area

used in the EDL simulation. Large parachutes may prove to be impractical on Mars due

to long inflation times and a time penalty (Braun et al., 2006).

To investigate parachute inflation times a simple inflation simulation was conducted by

the author. This was used to find how exactly such a parachute can have a significant

time penalty as stated in Braun et al. (2006). The model assumed that the parachute

inflated as a cylinder of constant height but with an expanding ‘mouth’ that is forced to

open from air rushing in as shown in figure 3.16. The parachute was considered fully

inflated when the radius of the cylinder was equal to the height of the cylinder, the broad

shaped cylinder approximating the shape of a fully inflated parachute canopy.

Figure 3.16 Parachute inflation model. The top row shows the

modeled inflation sequence (from A to C). The bottom row shows

the shape of the canopy during inflation of a supersonic parachute.

61

The model was validated as follows. The model shown in figure 3.16 was used to find the

inflation time for a skydiving type of parachute with a radius of 5 m and drag coefficient

of 0.8. An inflation time of 3 s and a shock loading of 3 g were found. This compares

favourably to a typical inflation time for a sports parachute which is between 2 and 3

seconds with a shock loading of 3-6g. A further comparison was made with results from

POSTII a parachute simulation program (Raiszadeh and Queen, 2002). The parachute

they modelled had a reference area of 178 m2 which corresponds to a radius of 7.5 m

assuming no gaps in the canopy. The drag coefficient for their parachute was 0.46. The

payload used was a 761 kg Mars Lander and the atmosphere density at parachute

deployment was 0.0135 kg m3 (which corresponds to an altitude of 8.4 km). The inflation

time for the parachute modelled in Raiszaheh and Queen (2002) was 0.32 s and the drag

on the parachute force was about 13 kN (~17 g). These numbers compare favourably

with the author’s own model results with an inflation time of 0.34 s and a maximum g

level of just over 16 g.

So it does appear that a simple expanding cylinder model can approximate real parachute

inflation dynamics. This model was then used to understand the dynamics of a large

parachute on Mars as may be used by a piloted Mars Lander. This of course ignored any

engineering issues that may arise from deployment of such a large parachute. So for a

parachute with a 30 m diameter an inflation time of 0.4 s was found and for a parachute

with a 50 m diameter an inflation time of 0.8 s was found. This was travelling at Mach 3

at an altitude of 8 km. Exploring the relationship still further a 100 m diameter parachute

inflation was simulated and found to take 1.3 s to inflate. Travelling at Mach 3 (~750 m s-

1) this means the distance travelled would be about ~1000 m during inflation. Including

10% porosity (to account for the loss of air through the gap in the top of the canopy)

increases the inflation time by ~0.3 s giving a total inflation time of 1.6 s and distance

travelled ~1200 m.

A piloted Lander (50 mT) with a large heat shield (15 m) will decelerate to Mach 3 at an

altitude between 10 and 15 km. Even using an extreme example of a 100 m diameter

parachute it appears there will be plenty of time and distance for deployment. However at

Mach 3 the g level on the crew will be about 20 g, possibly too high. The deceleration

limit for an unconditioned crew is between 3 and 5 g during entry (Condon et al., 1999).

This is over periods of minute or so. A crew maybe able to tolerate higher g levels for

short times such as during parachute deployment. However a shock loading of 20 g may

still be very painful and even dangerous if the crew are not restrained properly.

One way to reduce the g levels is to deploy the parachute at a lower altitude where the

Lander is travelling at a slower speed. EDL simulation suggests a 50 mT Lander will

reach 300 m s-1 (Mach 1.3) at 1.7 km. At this altitude parachute inflation of a 100 m

diameter parachute will take about 5 s and the crew will experience only 5 g deceleration,

which is a tolerable level. During parachute inflation the Lander will then travel ~1.5 km

downwards only leaving a second or so before hitting the ground. Obviously there is then

a significant time penalty for deploying a large parachute close to the surface. However

this does not mean such a parachute cannot be used at higher altitude (except the g levels

will be high). A large parachute, say ~50 m in diameter, deployed at high altitude should

62

be able to decelerate a 50 mT Lander to subsonic speeds (and keep the g levels low) at

several kilometres above the surface. See figure 3.12 for details of the inflation dynamics.

A2D has been updated to model lift as well as drag and results are shown for an MTSV

with an L/D of 0.3 in figure 3.17. Parachute inflation has not been simulated (one effect

would be reducing the deployment g levels) yet. With an L/D of 0.3 it was found that a

single 30 m diameter DGB parachute (now with the correct, lower, drag coefficient) can

be used for the MTSV EDL rather than the staged release used in section 3.7. The

operational environment is inside Viking heritage parachute box. This will help keep

development costs down. Also the parachute diameter is 30 m which is suggested as a

practical parachute size for Mars (Braun et al., 2006). The properties for the latest version

of the MTSV are listed in table 3.2.

Figure 3.12 Parachute inflation simulation results. The parachute was modeled as a cylinder with

constant height (25 m) and one end open to the incoming air. The radius of the opening increases

until it reaches the 25 m. The final shape of the parachute is then a 50 m diameter, 25 m high

cylinder. A real inflated canopy is more like a bloated hemisphere and so will have less volume (not

more than a third less) to fill than a cylinder with the same radius. The inflation time is calculated

here is then probably an overestimate (~10%). The speed of travel used was 750 m s-1 and the density

of the atmosphere used was 0.01 kg m-3.

Parachute radius

g level

63

Figure 3.17 G levels. Results from new simulations featuring an MTSV

with an L/D of 0.3. The g levels are very low. The lift takes the MTSV

through the higher less dense part of the atmosphere. Parachute

deployment is also at a higher altitude. A DGB parachute is used with an

effective area of 388 m2. The previous version of the MTSV EDL used

staged parachute deployments to keep the g levels at about 3 g. Here a

parachute with an effective areas 525 m2 and 1050 m

2 were used

corresponding to DGB parachute diameters of 50 m and 31 m respectively.

All other properties are kept the same for the MTSV model.

New EDL simulation using MTSV with lift

EDL simulation from Irving et al. (2006)

64

3.8 MLAM and the effect of winds on a piloted Lander

Simulation of the Martian atmosphere can be modelled using two types of models. There

is the global General Circulation Models (GCM). This numerically solves fundamental

thermal and mechanical equations for each atmospheric gridbox. Typical resolutions of

the grids are from 2 to 10°. For example the thermophysical GCM has a typical

resolution of 64 x 48 km or 5.5 x 3.25° and vertical coverage of 250 km with 60 layers.

Established GCMs such as Ames GCM or the Oxford Mars GCM can operate over

similar resolutions (e.g. see Sanchez et al., 2006). A GCM may operate with 2° grids for

high spatial resolutions (Rodin and Wilson, 2006).

The second model type is the mesoscale type such as the Mars Limited-Area Model

(MLAM). This also numerically solves fundamental thermal and mechanical equations

for each atmospheric gridbox. However a mesoscale model may operate at higher grid

resolutions but the whole grid only covers a part of the globe say ~1000 km in size. Initial

and boundary conditions are from a GCM. A coarse grid simulation with MLAM may

use 100 points over 100° whereas a fine grid simulation will use 200 points over 50°

(Siili et al., 2006). Mars has an equatorial radius of 3398 km so each degree is about 60

km. A fine grid 0.25° then corresponds to a about a grid size of 15 km. The smallest grid

size is dependant on the hydrostatic equation with the finest grids at 3-5 km (Kauhanen,

personal communication). The vertical grid is 32 levels and the atmosphere has its top at

48 km. The lowest level is at about 1.5 m. This level corresponds to the level of the

altitude at which Viking meteorological sensors were positioned. The next level up is

about 6.5 m. The levels are finely spaced near the surface and expand near the top of the

atmosphere with a maximum spacing of 7 km at the top.

The layers in atmospheric models are defined using sigma coordinates. These are defined

as the ratio of the pressure at a given point in the atmosphere to the pressure on the

surface underneath it. This way a sigma level close to the surface will follow the terrain

at approximately the same altitude. Sigma coordinates are used to simplify the lower

boundary condition. Further up in the atmosphere the levels become independent of the

surface terrain. However with MLAM the sigma coordinates follow the terrain even at

high altitude levels.

MLAM surface and atmospheric data can be viewed by an earth science data graphics

package called the Grid Analysis and Display System (GraDS) or the data can be

exported to a text file. Here the temperature, wind speed, surface pressure and surface

temperature at one particular location on Mars is exported into a parameters text file that

is then read by the author’s EDL simulation, aerobrake2D (A2D). The surface density

can then be calculated by A2D using the pressure and temperature at the surface together

with the atomic mass of the atmosphere. The density for the rest of the atmosphere can

then easily be calculated using the temperature data from MLAM assuming an adiabatic

atmosphere between levels. The calculation of the wind speed between levels is

calculated assuming a linear variation. The gradient is defined by the distance between

the levels divided by the difference of the wind speeds from those levels. The constant

offset for that section is taken from the wind speed at the lower level.

65

Wind speed is output as vector pairs, one south to north, labelled v, the other west to east,

labelled u. The magnitude and direction of the wind can be calculated from these two

vectors. Figure 3.18 shows the u component from the wind data at 0E 53S on the 7th July

2005. The season (LS=250°) was late autumn in the northern hemisphere or late spring in

the southern hemisphere. The location is in the cratered uplands at an elevation of ~2 km

above Martian “sea level”. This location was chosen as it gave relatively high winds

compared to other locations on the MLAM grid (-70W to 70E and -55S to 55N) during a

short search over the grid. It may give an upper limit to magnitude of displacement

experienced by a pilot Lander. It is interesting to note that according to the water map

derived by Mars Odyssey Gamma Ray Spectrometer this location is close to the

boundary between high concentrations of water (towards the pole) and low

concentrations of water (towards the equator) (NASA/JPL, 2006). This makes an

interesting landing site for exploration by humans although the EDLS design will be

challenging due to the high elevation.

To asses the effect of Martian winds on a piloted Lander the winds were applied in a

normal direction to the trajectory. This has the effect of blowing the Lander to the side.

The drag area was calculated by assuming both the heat shield and parachute are

hemispherical in form and taking the side cross section and assuming a drag coefficient

Figure 3.18 Variation of wind speed and displacement with

altitude. The wind speed shows a steady decrease from ~100 m s-1

at 50 km down to ~4 m s-1 at the surface. The sideways

displacement is due to the wind pushing on the large drag area of

the heat shield (15 m) and parachute (50 m). There is a plateau at

33 km altitude in the profile of the displacement because the

Lander flies at that altitude for quite some time.

A. Wind speed

B. Sideways displacement

4 m s-1

66

of unity. This then gives an effective area of ~1000 m2. The resulting displacement is

shown in figure 3.18. During the entry phase the Lander brakes using the drag surface of

its 15 m diameter heat shield. The sideways acceleration on the Lander from high altitude

winds is only about 1 cm s-2. An aero vehicle such as the blunt body entry configuration

described in this work, with an L/D of 0.3 will surely have enough control authority to

compensate for the tiny effect of these winds.

As can be seen in figure 3.19 there is a dramatic increase in sideways acceleration at an

altitude of about 12 km. This is due to the deployment of a large 50 m diameter

parachute. This pushes the sideways velocity of the Lander up to 6 m s-1 at touchdown.

This is even though the surface speeds are only 4 m s-1. This is an interesting result as it

highlights that a high mass Lander is hard to slow down i.e. resists slowing down once

the winds drop down to a lower velocity value than the Lander. In this example the winds

drop below 6 m s-1 below an altitude of 50 m so there is not enough time for the sideways

velocity of the Lander to respond to the lower velocity winds. It is clear that it may not be

safe to assume a Lander will be moving to the side at the same velocity as the surface

winds.

Figure 3.19 Sideways acceleration on the Lander due to the Martian winds

and the sideways velocity of the Lander with altitude. The acceleration

increases dramatically at about 12 km where parachute deployment occurs.

The sideways velocity also increases at this point. Interestingly the sideways

velocity, if not corrected will be 6 m s-1 while the actual surface winds are

around 4 m s-1. The Lander is pushed up to the higher speeds higher up in the

atmosphere and there is a delay (due to the inertia) before the Lander velocity

equalizes with the surrounding winds.

A. Sideways velocity

B. Sideways acceleration

67

The displacement of about a kilometre shown in figure 3.18 was calculated assuming

there is no correction during the hypersonic entry phase. Future robotic (and crewed)

entry vehicles will have substantial control authority to overcome atmospheric

uncertainties and so aid a pin-point landing (Wolf et al., 2006). So in reality the sideways

displacement will be due to the parachute descent phase only. Once the parachute is

released the power of the descent rockets can easily overcome any winds near the

surface.

To investigate the Lander displacement during descent by parachute an identical

simulation was run but this time the winds are only engaged during the parachute phase.

Results suggest a sideways displacement of ~300 m and a sideways velocity at

touchdown of 4 m s-1. This is the same as the surface velocity, although this is probably

coincidence. The sideways velocity of the Lander slowly builds up with decreasing

altitude to 4 m s-1 where as the actual wind velocity drops down quickly to 4 m s

-1 close

to the surface.

Braun et al. (2006) propose that an L/D aeroshell (i.e. blunt body) together with an

autonomous guidance algorithm and an accurate Inertial Measurement Unit (IMU) will

reduce the error at parachute deployment to the navigation knowledge at IMU

initialization. Striepe et al., (2002) noted using Monte Carlo EDL simulations that for a

blunt body the guidance program delivers the entry system right on target at parachute

deployment in navigation space (the space in the Lander’s computer). However due to

knowledge errors and IMU/Navigation error build-up the downrange error will be about 5

km and the cross range error will be about 1 km at parachute deployment. Wolf et al.

(2006) perform similar simulations for the MSL mission. They obtain a downrange

delivery error of ±3 km. Presumably this error will be the same for a vehicle with a high

L/D such as the biconic type vehicles used in NASA DRM 3.0 and other Reference

Missions. IMU errors may be corrected by direct observations of the surface and

comparing with maps. With aircraft, on Earth, IMU drift is corrected using GPS. In the

literature high L/D biconic type vehicles are outlined with superior targeting ability stated

as a rationale for selecting but it is not clear how this is achieved.

Wolf et al. (2006) suggest that angle-of-attack modulation together bank control, like the

Space Shuttle, can help reduce the parachute deployment altitude uncertainty. With a

blunt body the angle-of-attack can be controlled, in principle, via a movable mass. A

biconic vehicle can use body flaps for both angle-of-attack and bank modulation. For a

human scale Lander this may be a more practical approach, due to base architecture

arguments (i.e. Mars-Oz).

68

4.0 Conclusions

A piloted Mars Lander has many sophisticated and interdependent systems for a soft

landing on the surface. Using a combination of virtual prototyping in the OSFS and

detailed EDL simulations it has been shown that the design for a Mars piloted Lander can

be converged upon using an iterative process. A detailed description of the descent to the

Martian surface, using in a piloted Lander, may be as follows.

Entry begins at 100 km altitude at a velocity of 3.5 km s-1 and an entry angle of 4° to the

horizon. The MTSV mass is 52 mT at entry including a 6 mT inflatable heat shield and

12 mT of methane/oxygen fuel. The spherical surface heat shield, with an effective area

of 242 m2 a diameter of 15.75 m and a drag coefficient of 1.28 is flown in a lift up

orientation at an angle of attack of 30° to the flow, giving a L/D ratio of 0.3. A segment

of the Apollo earth-entry guidance program is used to send roll commands to the MTSV

bank thrusters. The MTSV rolls around its axis in response to uncertainties in the Martian

atmosphere. At one point craters and hills on the floor of Hellas basin glides into view of

the crew. The guidance computer soon decides to jiggle the MTSV around a bit and the

crew see the horizon with a thin tenuous orange haze hanging above it. Maximum

hypersonic deceleration occurs 5 minutes after entry and reaches 0.75 g. The MTSV then

flies at a constant altitude of 40 km, covering 350 km in 3 minutes, before continuing its

plunge downwards at 8 minutes after entry. Another deceleration peak of 0.75 g occurs at

about 10 minutes after entry. Two minutes later after the second hypersonic g peak (12

minutes after entry) the Disc-Gap-Band supersonic parachute is deployed while the

MTSV is at an altitude of 6.5 km and travelling at a speed Mach 1.8. Ejected from its

canister the parachute takes about one second to inflate with a peak g force of 0.5 g on

the crew. The MTSV slows down to Mach 0.8 (subsonic) by the time an altitude of 1.8

km is reached. Here the 400 kN thrust engines are ignited pulling the MTSV of the heat

shield. The heat shield falls away and to the side. The MTSV is flying at an angle of

about 45° to the horizon at this point. The valley walls of Vallis Dao can be seen looming

in the distance. After having travelled over 1700 km across the Martian surface since

entry the deceleration from the engines and the parachute combined turn the MTSV into a

vertical descent. The Martian winds have blown the MTSV several hundred metres off

course during the parachute descent. At an altitude of about 400 m the parachute is

released and the MTSV uses it’s beefed up 50 kN RCS to moved forward out from under

the parachute. The ERV location beacon is used to bring the MTSV to within 100 m

where it gobbles the final ton of fuel for the descent, throwing up clouds of red dust past

the crew’s window as it softly touches down. Fourteen minutes after atmospheric entry

thirty four tons of Lander and four human beings are on the surface of Mars.

Well that is one description of a possible scenario for landing humans on Mars. However

more work needs to be done to investigate other mission architectures to understand the

benefits and drawbacks of different technologies, such as an all propulsive EDL or

slender body Landers and to understand how the Martian atmosphere affects these

Landers.

69

Property or component

Notes

Pre-entry mass 46 mT (52 mT for 6.5 diameter hab?)

Landed mass 34 mT

Fuel Originally 6 mT (Bonin, 2006) of methane/oxygen, 2 mT for ~150 m s-1

delta V and 4 mT for ~70 s hover and down-range and cross-range correction (~0.5 km travelling at 10 m s

-1)

Hover engine 4x100 kN, isp 327 s-1, can land on 3 engines if one fails.

Main engine 50 kN thrust, part of a beefed up RCS system, used for deorbit and pin-point targeting on the surface

Landing gear Originally DC-X type deployable gear (x4 pads), perhaps better deployed in Mars orbit after aerocapture. Also deployed outwards to increase slope tolerance.

Slope tolerance 22° calculated from centre of mass and position for 4.5 m diameter MTSV. Moving the landing gear outwards by one metre (i.e. expanding diameter to 6.5 m) will increase the slope tolerance by 8°.

Habitat diameter and height (or length)

6.5 m expanded, 4.5 m nominal, 7 m high

Living space 232 m3 with 6.5 m diameter habitat or 90 m

3 with 4.5 m diameter habitat.

Additional ~100 m3 inflatable on the surface.

Heat shield type 16 m diameter, spherical surface cut from a sphere 20 m in diameter sized from thermal considerations, CD=1.28, L/D=0.3, aoa=30°

Heat shield mass 1x6 mT rigid heat shield, or 2x1.1 mT (with gas, pumps etc scaled up from data in Allouis, 2003) inflatable heat shield plus 2x1.1 mT backup shields

Hypersonic guidance and targeting

Bank modulation based on Apollo guidance program, 2 km x 5.5 km uncertainty at parachute deployment from knowledge errors (assuming extrapolations from Phoenix/MSL work) not implemented into simulation.

Parachutes 1 x 25 m Disc-Gap-Band parachute (388 m

2 effective area) deployed at

Mach 1.8, 6.5 km altitude

Table 4.1 Latest version of the MTSV (March 2007).

70

4.1 More Mars Landers

The Mars Direct project for OSFS is a web based group consisting of a volunteer team of

designers who are implementing the Mars Direct mission in the OSFS. Once the project

is finished virtual pilots will be able to fly the mission to Mars in OSFS. The project is

large taking advantage of OSFS flexibility. Spacecraft have there own dll modules to

control the vehicle properties and animations. This is unlike the Mars for Less

prototyping exercise which was done “fast” using a standard spacecraft module where the

properties of the vehicle were edited via a text file. The Mars Direct models are high

quality and special attention is paid to technical accuracy. Figure 4.1 show a model in

development, of the Earth Return Vehicle (ERV) and launcher.

The Mars Direct piloted Lander, the MTSV, is a blunt-body design like the Lander

prototyped for the Mars for Less project. Mars analogue stations such as MDRS are

based on such a squat cylindrical structure. However a slender entry vehicle like used in

NASA’s DRM 3.0 (see section 1.5) has a higher L/D and has higher manoeuvrability

important for pin-point targeting and landing close to pre-landed cargo. The Mars Society

are planning to extend there analogue research of Mars by building a habitat in Australia.

The Mars-Oz approach is to use a bent-biconic shape that has good control authority

(high L/D) for accurate navigation to a parachute release point and descent to a Martian

base. The Mars-Oz vehicle lands on its side for better stability and ease of unloading

cargo. Andrew McSorley, who is implementing NASA’s DRM 3.0 in Orbiter with

participation from the Orbiter community, has decided to change the landing vehicle. He

is still using a high L/D vehicle but it will land on its side (unlike in NASA’s DRM 3.0

Figure 4.1 ERV model. A model of the Mars Direct ERV by Seth Hollingsead

http://www.eveminer.com/setheden/index.htm

71

where the Lander lands on its end) fitted with wheels as shown if figure 4.2. The wheels

will allow the Lander to be moved and docked with other Landed vehicles. Rough terrain

is smoothed by bulldozers. This enables a large base to be built on Mars.

Figure 4.2 Wheeled cargo Lander. A model under development. A Lander may have wheels,

making construction of a large base easier (in concert with bull dozers. Model her is by Andrew

McSorley. It is part of his implementation and iteration of NASAs DRM 3.0 in OSFS.

http://orbit.m6.net/Forum/default.aspx?g=posts&t=10846

72

4.2 Lessons learned: A piloted Mars Lander v2.0

With knowledge of the requirement for landing humans on Mars, the practical

technologies available and their capabilities it should be possible to design a massive

Lander that addresses the main design concerns after considering the following lesson

learned during the whole project.

• Aerocapture is clearly possibly with a reasonably sized heat shield and modest lift

capability

• A lightweight heat shield is difficult to “drop” in the atmosphere during EDL

• Quick and safe transition from the entry configuration to powered descent

configuration is possible but heat shield release is a critical event

• High g loading on the crew is not a problem even with modest lift during entry

(from orbit)

• Accurate targeting can be enhanced by mitigation of wind drift through careful

consideration of the parachute release altitude

• It is not clear what benefit a high L/D vehicle has with regard to targeting during

the hypersonic phase

• Stable vessel structure on the ground is an important issue with a tall thin Lander

like the MFL one.

Figure 4.3 shows the EDLS for a Mars Lander that tries to address the above issues. The

design is based on the Mars-Oz approach that utilizes the bent biconic shape of the

vehicle. This shape has several good points such as high maneuverability for accurate

spatial targeting (although a blunt body can achieve similar results it seems), stability on

the ground, easy to move to form a base (especially if wheels are added like the design in

figure 4.2) convenient for unloading cargo (same design can be used for a cargo Lander).

The Mars-Oz vehicle however does not drop its shield before landing which incurs a

mass penalty and makes exposure of the engines and deployment of landing gear

complicated. The EDLS in figure 4.3 has a detachable heat shield. A detachable heat

shield allows the engines to be exposed and the landing gear to deploy. It may be wise to

flatten the bottom of the vehicle. If the gear fail to deploy at least there will some

probability of remaining upright. Also body flaps (like on the STS and nuclear warheads)

are added to make hypersonic steering possible.

73

1. Interplanetary 2. Aerocapture 3. Atmosphere entry

4. Main parachute deployment

5. Entry heat shield release

6. Powered descent

7. Pin-point landing

Cargo Lander (rover and MAV)

Figure 4.3 Modified Mars-Oz biconic design. Body flaps and detachable heat shield are the main additional

features. Interplanetary configuration features stored inflatable heat shield taken from Irving et al. (2006) EDL is

also slightly different.

74

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Links

Bruce Irving’s space blog

http://flyingsinger.blogspot.com/

Andrew McSorley’s models

http://www.flickr.com/photos/virtualspaceflight/

Andrews DRM 3.0 development thread

http://orbit.m6.net/Forum/default.aspx?g=posts&t=10846

Andrew’s Virtual Spaceflight website

77

http://www.aovi93.dsl.pipex.com/

Mark Paton’s website

http://www.freewebs.com/markpaton/index.htm

Finnish Meteorological Institute, Space Research

http://www.fmi.fi/research_space/space.html

Acknowledgements

Walter Schmidt and fmi for supporting this work. Thanks to Bruce Irving for the

invitation onto the virtual prototyping (VP) team and getting me interested in landing

humans on Mars. Andrew McSorley for encouragement and assistance during the model

building phase of the VP project. Grant Bonin for letting us play around with his Mars for

Less mission. The Open University for use of a personal computer, for initial prototyping,

and for the use of other supporting facilities. Catherine Maguire for listening to my

ramblings and for support and interest in the VP work. Sini Merikallio for inspiring this

work. Janne Kauhanen for exporting MLAM data into a text file and answering numerous

questions about MLAM.