“Barsoom Inception: A mission for humanity”members.marssociety.org/.../Barsoom_InceptionReport.pdf · 2019. 9. 6. · “Barsoom Inception: A mission for humanity” 3 1. Propulsion:

“Barsoom Inception: A mission for humanity”

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Table of content Page no.

1. Abstract …………………………………………………………………………… 02

2. Introduction ………………………………………………………………………. 02

3 .Overall structures and design …………………………………………………….. 03

4. Nozzle selection ………………………………………………………………….. 07

5. Trajectory and launch window for mission ……………………………………… 08

6. Capsule design and module planning ……………………………………………. 14

7. Total mission planning …………………………………………………………… 15

8. Total mass analysis ……………………………………………………………….. 16

9. Re-entry speed and heat shield …………………………………………………… 19

10. Several subsystems ……………………………………………………………… 22

11. Radiation protection and safety system …………………………………………. 25

12. Detailed system safety design …………………………………………………… 28

13. Cost analysis. …………………………………………………………………… 33

14. System and mission engineering and landing …………………………………… 36

15. Choices for a Mars Orbit Base Location ………………………………………… 38

16. Descent Technologies …………………………………………………………… 40

17. Re-entry and finally to earth …………………………………………………….. 45

18. Conclusion ………………………………………………………………………. 48

19. References ……………………………………………………………………….. 48

20. Team profile ……………………………………………………………………… 49


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Abstract:

This paper investigates means for achieving human expeditions to Mars utilizing existing or

near-term technology. Mission plans are described here.Those are accomplished with tandem

direct launches of payloads to Mars using the upper stages of the heavy lift booster used to lift

the payloads to orbit. No on-orbit assembly of large interplanetary spacecraft is required. In

situ-propellant production of hydrogen propellant on the Martian surface is used to reduce return

propellant and surface consumable requirements, and thus total mission mass and cost.

Chemical combustion powered ground vehicles are employed to afford the surface mission

with the high degree of mobility required for an effective exploration program. Data is

presented showing why medium-energy conjunction class trajectories are optimal for piloted

missions, and mission analysis is given showing what technologies are optimal for each of the

missions primary maneuvers. The optimal crew size and composition for initial piloted

Mars missions is presented, along with a proposed surface systems payload manifest. The

back-up plans and abort philosophy of the mission plans are described. We have chosen radiation

magnetic protection system. The most significant idea of our concept is that we have tried not to

show “reusable” the usual meaning. We will show even mars ice can be used as hydrogen fuel

for mission cost effectiveness. It’s concluded that the plans offer viable options for robust

piloted Mars missions employing near-term technology.

Introduction:

Getting to Mars with a Lunar Class Transportation System:

Figure 1: spacecraft interior(developed by catia)

Our dreamt requirements for this mission in a nutshell given below-


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1. Propulsion:

Abort Motor, Attitude Control Motor, and High Burn Rate Propellant for Solid Rocket Motors,

chemical hydrogen based propellant

2. Navigation:

Atmospheric Skip Entry, Autonomous Rendezvous and Docking, Fast Acquisition GPS

Receiver, High Density Camera Sensors

3. Avionics:

Algorithmic AutoCAD Generation, ARINC-653 / DO-178 Standard Operating System,

Baseband Processor, High Speed/High Density Memory Devices, and Honeywell HX5000 North

star ASIC

4. Communication:

C3I - Standard Communications, Communication Network Router Card, Digital Video Recorder

5. Power:

Column Grid Array Packaging (CGA), Direct Energy Power Transfer System

6. Thermal Protection System:

Ablative Heat shield with Composite Carrier Structure

7. Life Support & Safety:

Backup and Survival Systems, Closed Loop Life Support, Contingency Land Landing, Enhanced

Waste Management, Environmental Control, Hazard Detection, Isolation and Recovery

8. Structures:

Composite Spacecraft Structures, Human Rated Spacecraft Primary Structures Development

Overall structures and design:

Interdependency of Manned Mars Entry Vehicle

Types with Booster Diameter

• You cannot assemble a re-usable entry vehicle with an integral aero-

Shell in Earth orbit (no factory equipment and no manpower), so such

vehicles must be launched intact from the surface.


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• Two types of Mars Entry Vehicle concepts exist:

– 1. Wide base – Blunt body (capsule shaped)

– 2. Narrow Body – lifting body or cylindrically shaped

• Wide body (up to 15 meters wide at the base) landers are more stable

And can carry more cargo since they need less fuel due to entry drag,

but they need an HLV with a 10 meter or wider diameter.

• Narrow body landers carry less cargo but they can be launched on

Some currently projected HLV boosters with a 7-8 meter diameter.

• The booster’s launch cost must be affordable for dozens of launches

Per year to support a continuing Mars exploration program.

• We can choose a vehicle design based on the booster available OR we

Can pick a booster design to FIT the needs of the payload (the lander).

Capsule shaped blunt body lander:

Figure 2: carbon –carbon tile configuration for 10 and 15 meter diameter vehicle


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Types of Re-usable first Stages for HLVs ordered by increasing development cost:

Cluster of Boosters which separate and are recovered

Individually from the water.

• Cluster of Boosters where each one separates and

Individually flies back to a landing strip.

• Single Large Rocket-powered airframe which flies back to a

Landing strip with jet engines.

• Single very Large Rocket-powered cone-shaped airframe

Which lands vertically on its own rockets.

• Single fly-back rocket powered vehicle which captures its

own LOX supply during flight & for the second stage engine.

• Fully air-breathing (Hypersonic) Booster which flies itself

back to a landing strip with scramjets.

• Highest development cost = lowest operating cost.

• Operating costs usually far exceed development costs.


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Cluster base power booster planning:

Figure 3: phase 3B super HLV (54-140) mt Leo

Expected near term HLV feature:

• Re-usable first stage or first stage segments (required).

• Air breathing engine to increase payload mass.

• Minimizing refurbishment to recovered stages, such as a stage that

flies back and lands like an airplane.

• Flexible payload mass/size if a cluster.

• Very Wide payload capability to accommodate wide aero-shells, re-entry shields and vehicles

(minimum 33 feet (10 meters) wide or more,up to 15 meters). Wider payloads can be launched

with an inverted conical fairing, creating a “hammerhead” payload configuration, up to 50%

wider diameter as the booster.7 meter (23 ft.) wide booster can launch a 10 meter wide payload

8.4 meter booster (ET) can launch a 12.6 meter wide payload 10 meter wide (33 ft.) booster can

launch a 15 meter (49 ft.) wide payload

• Large payload shroud volume to hold large integral structures with low

density like habs.


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• Ability to recover and re-use the second stage if possible.

Nozzle selection:

Two-Step Nozzles. Several modifications of a bell-shaped nozzle have evolved that allow

full or almost complete altitude compensation; that is, they achieve maximum performance

at more than a single altitude. Figure 3-15 shows three concepts for a two-step nozzle,

one that has an initial low area ratio AZ/At for operation at or near the earth's surface

and a larger second area ratio that improves performance at high altitudes. See Ref.3-5. The

extendible nozzle requires actuators, a power supply, mechanisms for moving the

extension into position during flight, fastening and sealing devices. It has successfully

flown in several solid rocket motor nozzles and in a few liquid engine applications,

where it was deployed prior to ignition. Although only two steps are shown, there have

been versions with three steps; one is shown in Fig. As yet it has not made the change

in area ratio during rocket firing. The principal concerns are a reliable rugged mechanism

to move the extension into position, the hot gas seal between the nozzle sections, and the

extra weight involved. The droppable insert concept avoids the moving mechanism and

gas seal but has a potential stagnation temperature problem at the joint. It requires a reli-

able release mechanism, and the ejected insert creates flying debris. To date it has little

actual test experience. The dual bell nozzle concept uses two shortened bell nozzles

combined into one with a bump or inflection point between them, as shown in Fig. 3-15.

During ascent it functions first at the lower area ratio, with separation occur- ring at the

inflection point. As altitude increases and the gas expands further, the flow attaches itself

downstream of this point, with the flow filling the full nozzle exit section and operating

with the higher area ratio at higher performance. There is a small performance penalty

for a compromised bell nozzle contour with a circular bump. To date there has been little

experience with this concept.

Figure: two step dual bell nozzle


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A CAD model of dual bell nozzle given below:

Trajectory and launch window for mission

Planning suitable trajectory and launch window for a space mission specially, when it is an

interplanetary mission is very important task. It is one of the biggest challenges for a mission

design. For manned mission in mars this task must be done perfectly to make the mission safe

and cost effective. While designing trajectory for manned mission in mars some factors are

arisen are choosing suitable launch window, position of earth and mars in their orbits, Δv budget,

time of flight, departure energy etc. Here some idea for designing a suitable trajectory for

manned mission in mars in 2018 is given.

Why 2018


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The planets don’t follow circular orbits around the sun. They’re actually travelling in ellipses.

Sometimes they’re at the closest point to the Sun (called perihelion), and other times they’re at

the furthest point from the Sun (known as aphelion). To get the closest point between Earth and

Mars, it is needed to imagine a situation where Earth and Mars are located on the same side of

the Sun. Furthermore it is a situation where Earth is at aphelion, at its most distant point from the

Sun, and Mars is at perihelion, the closest point to the Sun. When Earth and Mars reach their

closest point, this is known as opposition. In the fig-1 the position of earth and mars at

opposition is shown.

And theoretically at this point, Mars and Earth will be only 54.6 million kilometers from each

other. But here’s the thing, this is just theoretical, since the two planets haven’t been this close to one

another in recorded history. The last known closest approach was back in 2003, when Earth and Mars

were only 56 million kilometers apart [1]. And this was the closest they’d been in 50,000 years.

Here’s a list of Mars Oppositions from 2007-2020[2]

Dec. 24, 2007 – 88.2 million km (54.8 million miles) Jan. 29, 2010 – 99.3 million km (61.7 million miles) Mar. 03, 2012 – 100.7 million km (62.6 million miles) Apr. 08, 2014 – 92.4 million

km (57.4 million miles) May. 22, 2016 – 75.3 million

km (46.8 million miles) Jul. 27. 2018 – 57.6 million

km (35.8 million miles) Oct. 13, 2020 – 62.1 million

km (38.6 million miles)

Figure 4: Position of earth and mars at

opposition

So 2018 is more suitable for manned mission in mars cause, mars will be nearer to earth and the

date is Jul.27.2018.


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Trajectory plan

Two major types of human Mars missions have been proposed based upon their

interplanetary trajectories and associated Mars stay times. They are known as conjunction-class

(or long-stay) and opposition-class (or short-stay) missions.1 Conjunction-class missions are

characterized by long stay times on Mars (order of 400 to 600 days), short in-space durations

(approximately one year total for the Earth-Mars and Mars-Earth legs), and relatively

small propulsive requirements. Opposition-class missions have significantly shorter Mars stay

times but longer stay time in space. Now which will be the best choice for this mission is the

question. Since this is manned mission safety issue for astronauts is very important. Longer stay

in space will cause longer exposure of

spacecraft in direct radiation from sun or

other celestial bodies. So conjunction

class trajectories will be better option for

the mission. [3] Figure-2 shows a simple

representation of conjunction class

trajectory

Figure 5: Conjunction class trajectory.

There are two types of conjunction class trajectories. They are type I and type II. If the

spacecraft travels less than a 180° true anomaly, the trajectory is termed type I. If the spacecraft

travels more than 180° and less than 360°, then it is a type II transfer. Type I trajectories are

preferable for manned mars mission.

Another important consideration for trajectory design is C3 (departure energy). C3 is not

constant. Variations in C3s can be due to many causes: the relative positions of the planets, the

plane change required into the transfer orbit, the velocities of the planets, and the eccentricities

of the orbits. However, this relies on the superposition of two synodic variations. The first

synodic period occurs every 2.14 years, or 25.6 months, and refers to the angular positions of the

two planets. The second is due to the eccentricity of Mars orbit (e = 0.093). The planets nearly


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return to their original relative heliocentric position every 7– 8 oppositions, or every 15–17

years. [4]

Figure 6: Piloted optimal departure energies, 2009– 2024.

Figure-3 shows the piloted departure C3s for minimum initial departure mass in LEO (Low Earth

Orbit) for 180-day outbound mission flights. [4]

Now finding the optimized trajectory can be done using software like Mission Analysis

Environment (MANE) software tool, STK Astrogator by AGI, trajectory planner, MATLAB

based programs etc. Here using trajectory browser website developed by NASA some

trajectories are found for manned mission in mars. Figure-4 represents graphical representation

of trajectories for round-trip rendezvous mission in mars for departure from earth in 2018.


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Figure 7: Trajectories for manned rendezvous mission.

In figure-4 the right column shows total Δv in km/s, horizontal line represents earth departure

and left side vertical line shows duration of the mission (Earth departure to earth reentry) in

years.

Figure 8: Details of the trajectories for manned rendezvous mission in mars.

Figure-5 gives the detail information of possible trajectories for the mission. In the figure

injection C3 is the departure energy for earth departure, Abs DLA is declination of the launching

asymptote, injection Δv is the change in velocity required for earth departure from 200 km LEO


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(Low Earth Orbit). Analyzing the information given in figure-5 two trajectories can be chosen.

These two trajectories are indicated by green rectangle (say trajectory-1) and yellow rectangle

(say trajectory-2). Departure C3s of trajectory-1 and trajectory-2 are same and smaller than

others and injection Δvs are also smaller. So less propulsive thrust is necessary that is

economically important for the mission. Considering the post injection Δv, trajectory-1 is better

than trajectory-2 and earth reentry velocity of trajectory-1 is less than trajectory-2. But the

mission duration of trajectory-1 is more than trajectory-2. Trajectry-1 allows the crew 48 days

more stay on mars but 32 days less stay in space. Stay more in space causes more radiation

exposure in space and more stay in mars surface allows the crew more exploration time. The

earth reentry velocity of trajectory-1 is less than that of trajectory-2. So trajectory-1 can be

selected for the mission. The earth departure date for this trajectory is May 10 2018 and arrival

date to mars is December 04 2018. The mars departure date is August 01 2019 and earth reentry

date is September 20 2020.

The orbital view of trajectory-1 is shown in figure-6.

The above discussion shows not a final trajectory for the mission but it gives an idea to find a

perfect trajectory for rendezvous mission to mars in 2018 with an example. There are so many

trajectories for mission to mars in a finite interval of launch window that can be found using

various trajectory finding software (some of them are mentioned before) and various


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computational method. From these trajectories the final one can be selected according to the

requirements of the mission such as Δv budget, reentry velocity, mission duration, mars stay

time, departure energy, safety of the crew, total cost of the mission etc. Keeping the cost of the

mission minimum is a challenge for the mission. An interplanetary mission requires burning of

enormous amount of fuel in less than one hour to give necessary velocity to the spacecraft. So Δv

should be minimum as much as it possible. Departure energy C3 should also be kept minimum.

For this mission window is an important consideration. It is also necessary to keep in mind that

the reentry velocity to earth atmosphere or entry velocity to the mars atmosphere should not be

so high that causes overheating than the limit of the spacecraft. Total mission duration should

also be considered. Longer mission duration more effort to prevent radiation is necessary that

causes more cost. All these consideration must be focused to plan the final trajectory.

Capsule design and module planning:

Figure 10: module planning

In capsule we will use the usual planning but we are choosing the Orion planning. We are

planning for some lunar module. That’s picture is given below:


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Figure 11: looner module

Total mission planning:

Figure 12: mars concept of operation

The Concept of Operations (ConOps) resulting from the hybrid architecture solution. The IM

Vehicle Stack would be lifted into LEO by the SLS rocket with the DUUS, the only launch

vehicle with enough throw mass for the job. Then the crew would rocket into orbit using the

services of a Commercial Crew provider. Once on orbit the Commercial Crew vehicle would


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dock with the IM Vehicle Stack, transfer the crew and depart. Afterwards, the DUUS would

perform the trans-Mars injection burn to send the spacecraft and crew on their way to

Mars.

The Martian flyby alters the trajectory of the IM Vehicle Stack sending it on a path to intercept

Earth. Just prior to entry interface the crew would transfer over to the ERP and separate from the

Habitat Module. The ERP would then carry out the Entry, Descent and Landing portion of the

mission ending with a splashdown in the Pacific Ocean.

Total mass analysis:

In calculating the throw mass: Using the throw mass numbers we generated a curve of C3 as a

function of throw mass. We first estimate the excess ∆V from a 200 km parking orbit to each of

the trajectories described in TABLE below:

Figure: ∆V EXCESS FROM 200 KM LEO

Then we can feet a curve on this data. That is given below.


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Figure: LEO ∆V Excess vs. Payload Mass

For a C3 of 38.8 km2/sec2, the ∆V excess is 4.86 km/sec.From these curves, we could further

generate a payload mass vs. C3 curve for the Falcon Heavy, shown in Fig. This curve was

generated by adding the excess velocity from table to the state of a 250 km LEO in

STK/Astrogator, and then calculating the resultant C3 for each point. From this curve you can

see that the estimated mass to a C3 of 38.8 km2/sec2is 9,800 kg.We rounded this to 10,000 kg

for the purposes of this study.


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Figure: Falcon Heavy Payload Mass vs. C3

For TMI finite burn modeling:

We use estimated parameters for the Falcon 9 second stage based on a Falcon Data Sheet [13]

(This is the same stage used for the Falcon Heavy.) This stage, the V1.1 Stage 2, will use the

Merlin 1D engine, which is estimated to have a vacuum thrust of 400,000 N, a vacuum ISP of

340 seconds, and a dry mass of 4.7 tonnes. To estimate the fuel usage for stage 2above the 200

km circular LEO baseline, we made an estimate of what fuel would be required for the

appropriate ∆V. using our STK/Astrogator-based numerical integration, we were able to produce

the data in TABLE VI. For the finite maneuver, we then calculated the burn duration to be 420

seconds. We estimated 15 tonnes of payload (including about 4.7tonnes for stage 2 dry mass)

and derived the fuel load for a finite TMI maneuver to require roughly 50 tonnes of fuel beyond

LEO. This compares to the reference data[13] of a73.4-tonne total fuel load for the stage. Note

that while the SpaceX references claim that the Falcon Heavy can lift 53 tonnes to LEO, this

refers to 53 tonnes of payload mass to a 200 km circular LEO. In our case of targeting a

Mars Escape C3 of 38.8 km2/sec2, we are only assuming a payload mass of 10 tonnes. This

leaves excess propellant in stage 2 in LEO for our case. TABLE VI shows an estimated 50

tonnes excess mass in LEO that we consider as excess propellant for the TMI maneuver. (We

assume that there is zero propellant excess when inserting 53 tonnes into LEO.)

Another table is given below:

Launch vehicle and payload value:

To determine if a launch vehicle can throw a space capsule to an Earth departure C3 of 38.835

km2/sec2we need a value for the payload mass. In this study we use a value of 10,000 kg, which

is based on an estimated SpaceX Dragon that has a dry mass of 4,200 kg and a cargo mass of

6,000 kg [7].For a representative launch vehicle we chose to use the Falcon Heavy. According to

a press release by SpaceX [8] in April, 2011, the first Falcon Heavy launch should be in 2013

or2014. According to version 7 of the SpaceX brochure, the Falcon Heavy will be able to

deliver (from Cape Canaveral)29,610 kg to a 28.5 degree inclination 185 km circular Low Earth

Orbit (LEO), and 15,010 kg to a 185 x 35,788 km .Geostationary transfer orbit

(GTO)[7]. Version 12 of this document[9] states the Falcon Heavy will be able to

deliver53,000 kg to a 28.5 degree inclination orbit (no altitude mentioned), and 19,000 kg to


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a 28.5 degree inclined GTO orbit(no specific orbit dimensions are given.) Another reference,

however, on the SpaceX website itself, says that the Falcon Heavy will be able to deliver

53,000 kg to a 28.5 degree inclination LEO, and 12,000 kg to a 27 degree

inclinationGTO[10].

Re-entry speed and heat shield:

By targeting the altitude of perigee to be in the range from 56.5 km to 61.5 km. The numbers in

the figure refer to the perigee altitude that caused that trajectory. A perigee of 62km or greater

escapes, and 56 km or less performs a direct reentry. We used the 1976 US Standard

Atmospheric Model(developed jointly by NOAA, NASA, and the USAF) and a five-tonne mass

with a heat shield area about 10.2 m. We modeled simple atmospheric drag without lift, with

a constant Coefficient of Drag (Cd) of 1.3 in the numerical integrator. (This value is similar to

Apollo.) Fig.7 shows that the velocity (dashed line, with right-hand axis) during reentry peaks at

about 14.2 km/sec velocity. This could possibly be reduced by changing the launch trajectory or

Mars flyby conditions, which we will look at in future studies.

Figure: Aero brake and Reentry Trajectories for Various Perigee altitude

Again,


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Figure: Altitude and Velocity during Direct Re-entry

The purpose of the aero capture before the reentry is to reduce the g-loads on the capsule and the

crew. Studies looking at reentry trajectories from Mars have attempted to define acceptable

effects on the crew, and in particular, the g force astronauts can withstand after long duration

spaceflight. [14] Next figure shows the g force due to atmospheric drag during aero capture for

the range of perigee altitude from 56.5 km to62 km. The lowest perigee has the highest g force,

and the peak load ranges from just under 6 g’s to just over 9 g’s. Of course, an aero capture

increases the length of the mission, up to an additional 10 days or so if the apogee reaches to

lunar orbit. Because the service module of the capsule would be released before aero braking, the

power system of the capsule would likely rely on batteries. The post-aero braking orbit must be

optimized with these considerations, as well as with the reentry conditions.


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Figure: G-force during aero braking

After jettisoning structures that must be released before reentry, the spacecraft is estimated to be

5,000 kg. The reentry of a 5,000 kg, 3.6 m diameter spacecraft into Earth’s atmosphere

present some challenges from an aerodynamics, aerothermodynamics, and thermal protection

system (TPS)perspective. The mission calls for both an aero capture maneuver and a

reentry at Earth. To date, no aero capture maneuver of this type has been attempted either at

Earth or other planetary destinations. The atmospheric entry speed forth aero capture is estimated

at 14.2 km/sec, which would make it the fastest reentry of any manned vehicle by far. The

fastest, successful reentry of a man-made, but unmanned, vehicle to date was the sample return

capsule for the NASA Stardust mission, which reentered at 12.6 km/s with an estimated total

stagnation heat flux of 1,200 W/cm2 [14, 15, and 16]. Assuming that a baseline vehicle

architecture for reentry Isa Dragon capsule, several TPS related issues that require more study.

First, we need to predict with higher fidelity the reentry trajectory and aero thermal loads. The

details of the reentry will have a significant impact on the thermal loads. The total heat load

during Earth reentry will determine how thick the TPS material will need to be to protect the

underlying structure. The peak heat flux will drive the selection of the actual TPS material.

The Dragon capsule heat shield uses a variant of Phenolic Impregnated Carbon Ablator (PICA),

which has been demonstrated in actual flight environments up to 1,200 W/cm2[17]. Some

very preliminary aero thermal predictions of the Mars return aero capture maneuver show peak

stagnation heating of over 3,500 W/c, dominated by shock layer radiation heating. To date there

has not been any ground testing of PICA at those heating heat flux levels. However, there is no


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data indicating that PICA would not perform as predicted under those conditions, using

existing analytical models. Ground testing, such as arc-jet testing, would need to be conducted to

verify PICA's performance. Second, currently operational manned spacecraft have been designed

to operate in LEO conditions for missions that are nominally up to six months to a year in

duration. The micrometeoroid and orbital debris (MMOD) and thermal environments for these

missions are well characterized and systems are in place to mitigate or control their effects. Some

investigation of the space environments during the 500-dayMars mission would be required in

order to predict the impacts that these environments may have on the TPS and determine any

necessary design changes. Third, TPS performance during reentry is influenced by two primary

factors: peak heat flux and total heat load, as well as other factors. Trade studies will need to be

performed to determine if using an aero capture maneuver to spread out the heat load is an

optimal approach. The longer soak times associated with the aero capture may lead to

even more stringent TPS requirements than a direct reentry. Using a direct skipping entry

trajectory with banking maneuvers to bleed off energy, may work just as well as the aero

capture/reentry combination. This study would include trajectory optimization, predictions of

aero thermal environments, and TPS sizing calculations. And lastly, the PICA TPS material has

not been used or tested at the high heat fluxes that can be expected on the Earth reentry from

Mars. As mentioned earlier, ground based arc-jet testing would need to be performed to

verify that PICA performs acceptably at those conditions. This test data would also be used to

validate the material response models used its sizing analyses. TPS thickness is the primary

adjustable design parameter available to satisfy TPS requirements. Higher heating conditions

usually lead to thicker TPS. Due to its manufacturing processes, PICA can be fabricated in

thicknesses up to about ten inches. If sizing predictions indicate that the required PICA thickness

is more than that, another TPS option may be needed. In addition, the thicker the material gets,

the more difficult it is to accommodate mechanical loads. PICA is not a mechanically strong

material, and induced strains in the material due to flexure of the underlying structure or

differential thermal expansion, may lead to mechanical failure (fracture) of the material.

In some cases these mechanical loads on the TPS material can be minimized by using a strain

isolation pad or by increasing the stiffness of the heat shield structure.

Several subsystems:

ECLSS is made up of several subsystems: air, water, food, thermal, waste and human

accommodations (often referred toes crew systems). Human accommodations takes into

consideration the crew’s personal, exercise and medical provisions, radiation protection, how

much free volume should be required for each person, and pressure suit needs. Figure below

shows primary ECLSS subsystems and their interfaces.


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Figure: integrated class interface

The cabin atmosphere assumes 101kPA (14.7psi) with adnominal makeup of 78% Nitrogen

(N2), 21% Oxygen (O2), and1% other. Subsystem components and stored gasses are

included for makeup gasses for leakage and gas equivalent storage for one cabin

repressurization, though no pressure suits). We reviewed numerous technology options to

revitalize the cabin air, and determined that an electrolyzer and Sabatierwork in tandem to

provide necessary air revitalization needs. The electrolyzer produces O2 and H2 from water. The

SabatierprocessesCO2, but needs H2 to do so, which it receives from the electrolyzer. The

amount of water needed to fully replenish all O2 requirements was included in the water


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subsystem calculations, and the difference in H2 production and H2requirements are

balanced to determine the stored H2 required for this subsystem. The water subsystem assumes

basic metabolic requirements for drinking, hygiene, and food processing. Water is added to

accommodate O2 production. Various configurations include no recycling, recycling

condensation from the atmosphere, and recycling water from the biological waste. Hygiene water

is a variable that can be used to reduce mass for this subsystem, and water recycling is

assumed for reduced launch mass, discussed further in section 8, discussing ECLSS sizing.

The SOA technology for condensate recovery utilizes multifiltration, ion exchange, and catalytic

oxidation. Food estimates include packaging and storage factors, as well as a galley and food

processing equipment such as heating, preparing, and disposing of waste. While it is probable

that all food is consumed, factors to account for food adhesion to the packages are included. The

thermal system is based on a redundant loop radiator, and is sized for the ECLSS power as

estimated for the different configurations, plus an assumed 2kW of heat coming from air-cooled

components such as avionics. The waste system provides the interface for collecting and recycle

water from the liquid waste are included in the water system as noted above. Personal provisions

are limited to items such as clothing and hygiene products. Free volume estimations are based on

minimum NASA requirements, and are only used to calculate makeup gasses that are required

for the full duration. No consideration of privacy, or separate sleeping quarters was contemplated

for this study. The mass and volume associated with free volume need not included in the

ECLSS mass and volume, as we assume that they will be carried at the overall vehicle

level, as is typical. Long term exposure to microgravity has demonstrated that deconditioning

must be counteracted with routine vigorous exercise so we assume the use of resistive

exercise equipment that provides full range of motion exercise of all primary muscle

groups. Medical equipment and supplies are provided to address emergencies. In addition to

spacecraft materials of construction, radiation protection is provided by a water shield made up

of water in storage for other subsystems. Other options reviewed were Hydrogen-Impregnated

Carbon Nan-fibers, and a Liquid Hydrogen Shell. Further study needs to be done to find creative

solutions for radiation protection, including the amount of radiation and the level of risk of a

high radiation event deemed to be acceptable. Evaluating these risk factors will include an

exploration of crew age, gender, and various exposure types; and will affect future ECLSS trade

studies.

Class sizing:

The crew’s metabolic rates drive parameters such as O2 consumption and CO2 production.

TABLE below shows interface values for each subsystem based on an average crewmember

(CM) of 70 kg and metabolic rate of 11.82 MJ/d (17).As demonstrated, the most significant

consumable is water at kg/CM-d of dry food. We are using a Water Lean approach, with a

reduced water requirement of 1.2 kg/CM-d as shown in TABLE:


25

We can also share a master equipment list. Given below:

Radiation protection and safety system:

HT mission design needs to be carried out with the aim of reducing radiation exposures in line

with ALARA. The application of ALARA requires that the space radiation environment is well

known – outside and inside a spacecraft. This requires knowledge of the external radiation

environment with its various components of the spacecraft/habitat construction, and of the

results of transport calculations modelling the internal radiation environment. The design

of the spacecraft/habitat requires the use of radiation transport codes to compute dose

equivalents. As described in Chapter 5 the computer codes may be one- or three

dimensional, deterministic or based on Monte Carlo (MC) methods. The construction of a

spacecraft/habitat should include areas where the dose rates are lower than elsewhere in the

spacecraft. There should be area monitors with visual displays of dose rates.


26

To reduce uncertainties further improvements are needed in the models of the galactic cosmic

radiation, the solar energetic particles, and the trapped radiation to allow the accurate forecasting

of the fully integrated model of the radiation environment incident on the spacecraft/ habitat.

Models have been developed for each of the radiation components. These models suffer several

shortcomings: (i) the GCR models inadequately characterize the solar cycle dependency and the

scaling with heliocentric distance; (ii) the SPE models have an incomplete understanding of

the acceleration mechanism of the transport through the Heliosphere and a lack of prediction

capability; (iii) the radiation belt models no longer reflect the current state of the Earth’s

magnetosphere and lack the ability to properly describe the dynamic behavior of the trapped

particles. The forecast models require an improved understanding of the physical processes on

the Sun; the transport and acceleration of the solar wind through the heliosphere; the

processes in the magnetosphere (wave-particle interactions, source and loss processes, and

acceleration mechanisms). The space environment is highly variable on very different time

scales as a result of the variability of the Sun. In general all aspects of the space environment are

affected, but SPEs and CMEs are the most dramatic radiation events and may constitute for

several missions a serious hazard. All the radiation components (including GCR and

trapped) are also modulated by SPEs(For bush decreases in the GCR fluences, for example).

An accurate prediction of SPEs and CMEs would allow for a more effective approach in

the shielding strategy. Forecasting through real time observation and propagation modelling

should be improved. Astronauts are particularly vulnerable during EVAs, when they should be

monitored with active dosimeters. Real-time space weather predictions and remote satellite and

areas Instrumentation will assist in EVA activity. The real-time measurements will provide

guidance, and can suggest changes in mission scheduling to maintain the total risk below

predefined limits. The development of shielding requirements and strategies is important

for the achievement of ALARA. The reduction in exposure can be made by reducing the

exposure time or by passive shielding. Passive shielding may cause an increased risk by

increasing the dose equivalent from any generated secondary particles, projectile and target

fragments (including neutrons). For shielding effectiveness, the use of a shielding material with a

low mean atomic mass is generally better. Information about radiation transport codes is

important and the strengths and weaknesses of the codes should be investigated in detail

via benchmarking procedures against experimental data, including data obtained with

advanced anthropomorphic phantoms exposed at accelerators. The physics at the basis of the

particle transport and cross sectional data tables must also be improved to further develop the

codes.

Area monitoring:

Area monitors at well selected locations in the spacecraft can determine the environmental

conditions and are appropriate for an immediate warning about changing exposure

conditions. Instruments are required to determine the radiation environment in terms of particle


27

type, fluency rate, energy, and direction distributions and, in some instances, dose quantities.

Dose quantities used to assess doses to astronauts and to monitor radiation at a number of

locations should give values of the dose rate. These data can be used to implement ALARA.

Area monitors at well selected locations in the spacecraft can be appropriate for immediate

warning about changing exposure conditions. This can be of importance before or during SPEs,

electron belt enhancements, and EVA. Real-time calibration of instruments should be explored.

If appropriately designed and accurately calibrated instruments are used, it may be that a

quantity measured in fixed position in a spacecraft can, along with appropriate occupancy

data, provide the basis for an adequate assessment of doses to an astronaut or of doses to the

local skin or the extremities. While in principle this procedure may be applicable to astronauts in

space, the large variation of the radiation field in intensity and composition of radiation types

inside a spacecraft, and its variation with time together with the flexibility of the astronaut´s

position, has the consequence that area monitoring is not sufficient to completely

substitute individual monitoring, especially considering the high individual doses to astronauts

and the interest in providing a basis for individual risk estimates.

Individual monitoring

The assessment of organ and tissue absorbed doses, together with radiation quality factors, of

individual astronauts can be accomplished by calculations using anthropomorphic phantoms or

by measurements using personal dosimeters.

One method of calculation of organ and tissue absorbed doses and radiation quality factors does

so directly for a standard male or female phantom for various locations in a spacecraft

with appropriate shielding. The phantoms can be adjusted to approximate a particular

astronaut. The results are normalized using readings of area monitors and personal dosimeters.

Another method requires knowledge of particle fluence and applies conversion coefficients

from particle type, energy and direction distribution of fluence to organ and tissue absorbed

doses and corresponding radiation quality factors for uniform irradiation of an astronaut.

Individual monitoring is mostly performed using personal dosimeters worn at the surface

of the body. The personal dosimeter serves as the dosimeter of record. A single dosimeter

system is, however, not sufficient to provide an assessment of the absorbed dose at the surface of

the body weighted by radiation quality. The broad range of different types of particle requires a

minimum two detectors, one sensitive to low-LET radiation and the other to high-LET radiation.

Because of a possible anisotropy of the exposure in the spacecraft due to variations of shielding

properties, it may be useful to wear more than one dosimeter. Also care needs to be taken

regarding low-energy electrons and particles which are stopped in the skin and, therefore,


28

contribute only marginally to organ doses other than the skin dose, but may induce a large signal

in an external dosimeter.

The use of adequate active personal radiation detectors would enable improved 4036

characterization (input energy, nuclear abundance, fluence rate, direction) of the radiation

field both on the body of the astronaut as well as in the environment. The measurement of

dose-rate can contribute directly to ALARA.

The results of bio-marker measurements can be additionally used to estimate individual

radiation exposure. The determinations can be collaborative and provide all the

experimental radiation information and relative codes needed to achieve an efficient risk

assessment, minimizing the uncertainties in the final risk estimates.

Dose recording

Astronauts in space are exceptionally exposed and the assessment of their individual doses

should be part of the radiation protection programmer for space flights. Astronauts should

be informed of their doses and risk assessments as soon as possible. Their doses should

be regularly registered and a long term registry for all missions should be maintained. The dose

record is the formal statement of the crew member’s exposure and should be kept as a

confidential medical record. The record should contain the history of the exposure and all

the calculation and experimental results, including all information on the particle type

energy and direction distributions of fluence; computer codes; conversion coefficients and

weighting factors; area monitor, personal dosimeter, and biomarker results.

Detailed system safety design:

The system safety activities needed to support detailed system design are presented in Figure

This phase is different from the concept development and early design phase because the length

of time required to complete the detailed design makes it likely that new risk issues

requiring resolution will surface as design details are being developed. These new risks often

involve the necessity of having to reallocate subsystem masses while ensuring that

allocated safety requirements, codes, and standards continue to be satisfied. The lower level

safety objectives for this phase consist of a combination of objectives listed in Figure of S under

the phrases “Design the system to ...” and “Build the system to ...” This results in an

expansion lower-level objectives under the ASARP objective to five, as shown in Figure

The broadening of the early design safety objectives reflects the fact that during detailed

design, new information, both positive and negative in character, is being generated.

Sometimes this information takes the form of new risks and challenges. Other times the new

information takes the form of new opportunities to improve safety.


29

Figure: Principal System Safety Activities and Related Processes during Detailed System

Design, and their Interfaces with the Safety Objectives

Radiation protection (passive protection):

The first protection is constituted by the mechanical structure of the spacecraft. This, in general,

will be very elective for cosmic rays with no more than about 30 MeV/nucleon of E, For

example, the average thickness on the MIR space station is1 g/cm2of equivalent aluminum,

reaching 7 g/cm2in particular places of the spacecraft. For a more generalized example, let us

consider a spacecraft whose external dimensions match the Shuttle capability, i.e. with a radius

of 2 m and several meters in length. Its external vessel should have a mass of1020 kg for each 3

mm of equivalent aluminum and each 10 m in length. This thickness of aluminum can stop, by

ionization losses, all protons with less than 25 MeV of E, Adding a further absorbing material

around a more restricted volume, where the astronauts are supposed to spend most of their time,

would result in the addition of a significant mass to the spacecraft. The protection of a section of

our hypothetical cylindrical spacecraft, 1.5 m in radius and 3 m long, requires 115kgfor each mm

of aluminum, i.e. 3450 kg for stopping protons up to 100 MeV. We could do better by

concentrating the absorber around a smaller volume, where the astronauts could shelter in case of

radiation alarm'. The mass of aluminum around a cylindrical volume 0.5 m in radius and 2 m

long would be 12 kg for each mm of thickness, i.e. 360 kg for stopping protons up to 100MeV,

which rises up to 1160 kg for stopping protons up to 200 MeV.


30

A magnetic lens for shielding the spacecraft from the solar part of the cosmic ray

radiation:

The good directionality of the solar cosmic rays allows one to conceive a cylindrical magnetic

lens for defocusing all the positively charged258 P. Spillantini et al. /Nuclear Instruments and

Methods in Physics Research A 443 (2000) 254}2631The industrial production of NbTi

superconducting wires is already at a critical current density of J#"5000A/mm2 atB"2.8 T and

4.2 K, so we consider to work at 70% of the critical current. Particles away from the axis of the

lens. The need to minimize the mass limits the choice for the magnetic field production: it must

be realized by superconducting coils, and without the use of iron yokes. This implies a

refrigeration system to keep the superconductor at the low temperature necessary for its

operation. For what concerns the techniques that can be envisaged for the coil, in ten years or

more from now, it is probable that the NbTi material presently used for the industrially produced

superconducting cables could be substituted by theNbSn material that is already used in magnets

for high field spectroscopy and, since a few years, is also used by industry for construction of

large coils (for fusion research and specialized projects).The use of NbSn material will certainly

reduce the mass of superconductor material needed forth coils and increase both the field

intensity and the operating temperature while also increasing testability limits. Actually, beyond

NbTi and NbSn, one can envisage the use of ceramic high-Tc superconductors (HTS), like Ag-

stabilizedBi2223tapes;in fact, within 2 or 3 years the electiveness of such material in real size

coils may be proved industry for big projects (SMES and other electric applications). The fast

progress of cry generators for space and other applications allows one treasonably envisage that

in the future their reliability, mass and power consumption will match the request for their use in

interplanetary sights, thus reducing drastically the volume and mass of the helium supply. Surely,

it should be much better to use cry refrigerating systems capable of maintaining the temperature

of the coils at a few Kelvin without requiring liquid helium. Presently, the needed mass and

power consumption rule out this solution. However, also for this kind of device, the technical

progress promises to be quick and a suitable R&D activity devoted to their use in long duration

space sights could bring their parameters in the useful range. The following evaluation for the

magnetic system will be based on the use of NbTi superconducting material whose performances

are well known especially for the intensity of the field and the needed refrigeration power. No

special hypotheses will be made for the cooling system at this stage of the discussion. A

magnetic lens can be obtained by a toroidal configuration: an electrical current running along a

cylindrical conductor, parallel to its axis, with the electric current circuit closed in a return

circuit. The simplest scheme forth return circuit consists of a cylindrical conductor at a larger

radius carrying the return current, connected by radial conductors to the inner cylinder. Because

at this stage is not worthwhile to go into detail for the geometry of the lens, we will assume that

the innermost and outermost cylinders, as well as the connecting radial connectors, are

continuous surfaces, and the total cross-section of the conductor is constant along the whole

circuit. The diameter of the inner cylinder is assumed to be 0.5m and its length 1 m. The radius

of the external cylinder can be as large as we wish, since its mass does not change with the


31

radius, and only the mass of the connecting radial connector’s changes. However, since the

magnetic lens must be transported to space in an already assembled state, its external diameter

cannot exceed that of a big rocket or that of cargo bay of the shuttle vehicle (the diameter of

cargo bay of the present shuttles is 4.6 m). We will assume 4 m for the external diameter of the

lens, and we will maintain the length of 1 m also outside (see Fig. 5). The region inside the inner

cylinder could be well protected against the incoming radiation by the mass of the technical

devices servicing the coil, adding a further absorber only in the parts not well protected by them.

For an approximate evaluation, let us assume that we deviate by 453thecosmic rays arriving

parallel to the coil axis at a radius slightly exceeding 0.25 m. For 500 MeV kinetic energy

protons, slightly less than 3 T] nonintegrated magnetic field is then required. The total current

required to obtain such a field is about 3.5MA. We assume to have a current density of3500

A/mm2 in the inner cylindrical conductor, which is a safe working point for a NbTi

superconductor.1This choice brings to an average current density of 700 A/mm2 #owing in the

whole coil. Hence, the mean cross section of the s.c. coil turns out to be 50 cm2, accounting for a

total volume of the s.c. of 50 cm2]560 cm"27 840 cm3 and a mass of 103 kg. The average

thickness of the s.c. in the inner cylinder is 3.2 mm and in the outer one is0.4 mm. With this coil

the resulting trajectories of500 MeV protons arriving at some selected distances from the axis of

the lens are reported in Fig. 6.A look at this sketch of trajectories indicates that the deflection

power of the lens should be reinforced around 1 m of radius in order to maximize the volume of

the fully protected part of the space craft. A circular-shaped coil can obtain this maximization.

Such a choice represents also a better geometrical situation for the mechanical structure that

should support the magnetic stresses produced by the high field inside the lens. With this choice

for the shape of the coil and assuming the same parameters as before, with the same maximum

field atR"0.25 m, the trajectories of 500 MeV protons are those sketched in Fig. 7. The fully

protected volume significantly increases in comparison to the case of the rectangular-shaped coil

of Fig. 6. Furthermore, if the coil could be realized by a nearly continuous surface, the coil could

be nearly self-supporting against the magnetic stresses, in spite of its thin average thickness: in

fact, the magnetic stresses are directed outside the volume of the coil and nearly perpendicular to

the conductor, and their elect is equivalent to an internal pressure that probably requires only

relatively light external structures. Therefore, the total mass of the magnetic lens will require the

addition of the only masses of the thermal shields and of the cooling system. Another important

advantage of constructing a nearly continuous surface coil is that the magnetic field will be fully

contained inside the volume of the coil, with no external-fringing field.

Comparative discussion of passive and active protection from the cosmic ray radiation:

We begin by comparing the mass of the magnetic system based on the circularly shaped coil of

Fig with the mass of the absorber needed to have the same protection. Also, the absorber can

protect from the directionality of the arriving cosmic rays, and should consist therefore of a disk

of the same diameter as the external diameter of the coil. The internal part of the disk (up to 0.25


32

m in radius) is common to the two devices and is kept out of the comparison. The surface of the

disk will be indeed12.5 m2. The total mass of the magnetic lens has tube evaluated adding to the

mass of the coil the masses of the thermal shields, of the cooling system and of the supporting

structures. In "rst approximation, these masses can be considered proportional to the mass of the

coil itself. The proportionality factor is completely arbitrary without the consideration of a

specific design, and it will be considered arbitrarily equal2to 1. However, it is reasonable to

assume that all the elements servicing the coil cannot be lighter than some minimum mass that

will be arbitrarily3assumed to be 50 kg.The results of the comparison are reported in Table2 and

in Fig. 8 for several values of the maximum kinetic energy of the particles. It is clear that as we

move up from the proton energy region of a few tens of MeV, the magnetic system is much

superior.

Figure: Comparison between the directional magnetic lens and an equivalent Al absorber

Figure: Weights of active and passive shield vs. cutoff energy triangles: passive shield;

diamonds: magnet mass of the active shield; squares: total mass of the active shield


33

Figure: Monitored cosmic rays 1967}1972 by IMP IV and IMP V experiments. The differential

fluxes are measured in the energy range between 20 and 80 MeV. In the same energy the galactic

cosmic ray flux is less than 10~4(cm2s sr MeV) ~1

COST ANALYSIS:

1. Cost Breakdown Structure

2. Lander module

3. Earth return vehicle

4. Mars habitation module

5. Mars Surface Elements

6. Earth Ground Infrastructure


34

7. Prime Contractor Programme Level

8. Space Agency Programme Level

8. Operations


35


36

System and mission engineering and landing:

Rationale for using Low Earth Orbit Propellant Depots with HLVs:

Most sources now show that Mars landers will need to have very wide diameters or bases: 10

meters (33 feet) or more. To use a lander on Mars, FIRST it needs to get to LEO. Current ELV


37

(small-diameter) launchers cannot launch such wide payloads into LEO. We need Mars landers

that can carry very large payloads to the surface - protected from re-entry heating. We can launch

much larger landers DRY than when WET. If we launch them dry, we need orbital Propellant

Depots to accumulate propellants for Transit and Mars landing (EDL).Without Depots,

cryogenic propellants will sometimes boil off before a crew can reach the vehicle to use it .With

Depots, the propellant in the first vehicle is not lost. Building Propellant Depots should NOT be

used as a rationale for not building large Diameter HLV’s. Buildup of Mars Fleet in LEO using

HLVs and Use of Propellant Depot as a ‘Vehicle Accumulator ‘Vehicles and equipment are

launched dry to LEO. Mars Ferries may be able act as a second stage and put themselves in orbit

if fully fueled. Space tugs move orbiting vehicles and cargo to depot area and dock at adjacent

assembly base. External (non-integral) aero-capture shields are attached to all Mars bound

Transiting vehicles such as cargo carriers, crew Earth return vehicles and two Depots: LEO to

LMO (Low Mars orbit).Ferries have integral aero-shells (as part of their vehicle structure).Cargo

is loaded aboard cargo transit vehicles. All Transit vehicles are fueled from large LEO Depot.

Sun-Shielded, Insulated Propellant Depot:

VASIMR – a possible Transit Alternative:

VASIMR is a very efficient plasma rocket system that, given a 200 Megawatt electrical power

source, could reduce Mars Transit time to ~50 days from 6-8 months, greatly reducing crew

exposure to solar radiation. This would save a huge mass of Mars Transit propellants.VASIMR

could widen windows for Mars Transit missions. It reduces mission risk from damage to liquid

fuel engines and loss of liquid propellant accidents. It could also move Mars bound vehicles and

cargo from LEO to GEO before departure from GEO and maintain a Mars base orbit. A current

version is rated in tests at 250 KW. For Mars Transit purposes, development should start now on


38

a much larger and light weight space rated power supply needed to power VASIMR, such as a

compact nuclear reactor, ultra-light solar panels, etc.

PROVISONS:

This System is NOT useful for Mars without the large power supply. The use of chemical

(cryogenic) propellants for Mars transit missions is practical without waiting for a compact

VASIMR power source.

Using aero capture at Mars:

Aero capture uses a very large diameter rigid aero-shield or integral aero-shell to slow all

spacecraft arriving at Mars by flying through the atmosphere once to brake down to orbital

speed. This saves a huge mass of liquid propellant. Non-Cryogenic OMS Propellants may be

used for the orbit trim maneuver and rendezvous with the Low Mars Orbit base. An Aero-shield

is not integral to the spacecraft, but is much wider than it and partly surrounds it during the aero

capture maneuver. Aero-shields may be able to be assembled from sections in LEO for use, (so

they will fit in an HLV cargo space). They can be kept extended (out of the way) on a boom in

front of Mars- bound vehicles until arrival at Mars. The Aero-shield is retracted and locked

before arrival. The Aero-shield can also be used during return to Earth. Aero-capture may not be

compatible with use of VASIMR propulsion, which could eventually replace it.Mars Ferry

vehicles would use an integral aero-shell which is part of the vehicle’s structure to accomplish

the aero-capture.

Choices for a Mars Orbit Base Location:

The base should be high enough to avoid frequent orbit re-boosts which use up fuel (about 400

km high or more – Low Mars Orbit .If the orbit is elliptical, it will reduce the number of

opportunities for landings and takeoffs compared to a circular orbit. A High Mars Orbit (HMO)

would require a lot more fuel to reach from the surface and for landings than a LMO.A near-


39

equatorial orbit will maximize the equatorial eastward speed of 240 meters per second to reduce

fuel use for landings and takeoffs. This is 6% of takeoff delta-V and about 25% or more of

powered landing delta-V requirements. A high inclination or polar orbit would increase the

propellant mass needed for landings and takeoffs considerably. It would also reduce the number

of opportunities for landings and takeoffs compared to a near-equatorial orbit.

Other Equipment Needed in LMO Base:

2 or more Independent Crew Habitats with Solar Radiation Refuges (same as in Transit

Vehicles), each capable of supporting all crew members until return to Earth. Multiple spare

replaceable equipment modules and parts. Redundant food, water and consumables for the

crew.2 Cryogenic Propellant Depots (shielded & active cooling) Sets of Mars cargo landers (one-

way) and Mars Ferries.Intra-system crew vehicles or tugs to explore Phobos etc.Tele-operation

equipment to run surface robots. All equipment needs to be optimized to minimize crew time to

operate and maintain it, including accessibility and modularization. This is a major lesson

learned from the space station.

Supersonic Decelerators:

Ballutesand Hyper cones a Balluteis like an inflated semi-rigid parachute used to decelerate at

the end of re-entry. Work is beginning on a Ballute-like system called a hyper cone that would

deploy after speed dropped below about 3250 mph. For a 60 ton vehicle, the wide end of the

hyperconewould be about 100 feet across. The hyperconefabric still has to be able to withstand

heating caused by friction with the air. The Hyperconeis hard to deploy and control.Ballutes and

Hyperconesare expendable only.

Combinations are required if Supersonic Decelerators are used.

Once speed drops below Mach 1, a subsonic parachute could be deployed.

Below about 1 kilometer, landing rocket engines would be needed to set the lander down gently.

This design requires 2 expendable systems, the hyper cone and the subsonic parachute. There is

no way to recover and prevent a high speed crash if either the Hyperconeor the parachute fails to

open properly. A failed chute could also endanger separation of a crew cabin/escape capsule for

abort to surface mode.

Supersonic Retro Propulsion

This method is called Supersonic Retro Propulsion (SRP).It is now being taken seriously. It

requires rocket thrust firing directly through the heat shield and against the supersonic flow of air

pressing against the base of the vehicle as it decelerates. The rocket engines must be fixed in

position with the nozzle ends flush with and sealed to the heat shield and thus they cannot gimbal

for steering. Steering during re-entry and landing must be done by varying the thrust of a set of

engines or by using side-mounted small Vernier engines.


40

Descent Technologies:

Deceleration for Mars EDL from 400 km orbit.

We should use:

Banish the “Expendable Mentality “Think Re-Usable’’Current scenarios for Manned Mars

landings envision a very large lander which has, inside it , just like a nested Russian doll

(Matryoshka), another entire vehicle for the ascent with its own engines, tanks, controls,

structure, etc. This means that every trip to the surface requires an entire additional pair of

vehicles with all of the descent propellant brought from Earth. It also wastes all of the perfectly

good equipment in the descent vehicle or lander. This kind of architecture is only good for the


41

kind of Mars expedition we do not want: the “Flags and Footprints “style mission, which is

financially unsustainable, and leads to the “one-way “Mars trips currently being proposed by

those desperate to see any kind of Manned Mars Mission during their lifetime.

Designing a Re-Usable Mars Ferry:

Re-usable vehicles using SRP can use their engines for:(1) Initial de-orbit burn: (7560 mph –

7400mph) ~ Mach 13 (1) “Passive “Atmospheric Entry -Deceleration (Mach 13 –Mach 6) (2)

Supersonic Retro-propulsion: (~Mach 6 -Mach 0.9) (3) Subsonic Deceleration Mach 0.9 –Mach

0.2 (4) Final descent and landing: (Mach 0.2 / 100 mph to surface) (5) Ascent back into Low

Mars Orbit. With a fully re-usable vehicle, nothing is thrown away. Hydrogen and oxygen

propellants can be created from Mars ice and volatiles, using a nuclear power source and carried

back to orbit for use on the next trip down, since there is little cargo other than propellant that

needs to go up. Wide base vehicles with lower density slow down more during re-entry and thus

need less propellant to land than a narrow base vehicle. All of the propellants needed for Crew

and Cargo Descent can be carried UP on ascent. 5 tons of extra propellant can be loaded back

into the Propellant Depot in Low Mars Orbit from each Cargo Ferry trip UP for use in Earth

Return

A Single Stage to Orbit and Back Vehicle (SSTOAB) for Mars

The Mars Ferry is essentially an SSTO for Mars. Mars gravity is 0.38% of Earth’s, so achieving

low orbit is much easier than on earth -about 2.5miles per second. This takes only about ¼of the

energy to reach L.E.O. Mars has 1/10 Earth mass, and 8 times lunar mass.If there is no staging,

then there is no first stage to recover –the entire vehicle goes to the Orbital “base”and back to the

surface base -intact. Much less fuel is needed to land than take off to orbit since normal

atmospheric re-entry sheds up to the first 4310 mph (1.7-2.4 km/sec) of speed. A cargo ferry

would carry 25 tons of modules and equipment down to the surface and 20 tons of propellants

back to orbit (15 tons to use for descent and 5 tons for the Depot). Acrew ferry (with its 5 ton

crew cabin) would carry a crew with 20 tons of cargo down to the surface, and a crew and 15

tons of propellant back to orbit.

Discovery of Widespread Sub-Surface Ice on Mars:

makes propellant production a Non-Exotic operation 20 years ago, we had no knowledge of the

widespread existence of Wateron Mars in the Form of Sub-surface ice deposits (ice regolith or

permafrost), some of them fairly close to the equator. Mars Direct (1989) and related concepts

assumed we would bring the hydrogen to make methane fuel from all the way from Earth. Now

the hydrogen can be obtained from Mars ice in large enough quantities to use as a fuel directly.

Producing propellants at a Mars surface base is NOT an exotic zero-gravity technology that still

needs to be developed. All of the individual steps are already performed on Earth every day. This

avoids the need to first develop a technology to do it in micro-gravity, and allows a direct

process of developing the hardware to do it under Mars conditions. The extraction equipment


42

and cryogenic storage system would be built in “package plant”modules so that it could be set up

easily by tele-operated robots, before crew members descend to the surface. Any Surface Base

Site would be influenced by the availability of near-surface ice posits that can be mined with

simple excavation equipment.

Descent Engine Placement for Lander –similar to Ferry using S.R.P.:

Mars Ferry: Launch Weight - 125 Tonnes:

Both Cargo & Crew designs mass ~ 70 mt at descent and 125 mt at lift off, with 20 tons total

ascent payload and 25 tons for descent. Both vehicles have a base heat shield diameter 14 meters

wide. Both Ferry versions have a cargo bay to hold up to 25 tons of large cargo with cargo bay

doors that open on the side below the fuel tanks. Both Ferry versions have oversize tanks that can

hold 95 tons of propellant (about 75 tons for ascent, about 15 tons for the subsequent descent,

and about 5 tons extra for deposit in the propellant depot. (Cargo ferry only)

CREW AND CARGO FERRY DIFFERENCES:

The Crew Ferry has a 5 ton crew cabin / escape capsule on top. It would never separate from the

Ferry except in an emergency. The cargo Ferry has no equivalent capsule on top. The Crew Ferry

would carry 15 tons of propellants UP as payload in its tanks during an ascent mission, and 20

tons of cargo DOWN during a descent mission. The Cargo Ferry carries (as payload) 20 tons

(fuel) UP in its tanks and 25 tons (cargo) DOWN in the cargo bay. Total Ascent and Descent

Masses are the same for both Ferri.


43


44


45

For safe landing and better descending we will use sky crane and supersonic parasuit.

Figure: sky crane

Re-entry and finally to earth:

Most Mars Sample Return scenario involve two launches

One with an orbiter One with a lander, providing a rocket with a sample storage container.

The rocket performs a rendezvous with the orbiter, then the orbiter returns toward Earth.

This type of mission is both costly (two launches) and risky (Mars orbit rendezvous), hence the

interest of single launch. Single launch mission profiles Electric propulsion is a key enabler

for single launch missions. Two options are presented: A short term mission with a

rendezvous in Mars orbit, A more advanced mission with EP, Cryogenic propulsion, ZBO and

a heavy lander with direct return toward Earth.


46

Most Mars Sample Return scenario involve two launches:

One with an orbiter One with a lander, providing a rocket with a sample storage container.

The rocket performs a rendezvous with the orbiter, then the orbiter returns toward

Earth. This type of mission is both costly (two launches) and risky (Mars orbit rendezvous),

hence the interest of single launch. Single launch mission profiles Electric propulsion is a key

enabler for single launch missions. Two options are presented : A short term mission with a

rendezvous in Mars orbit, A more advanced mission with EP, Cryogenic propulsion, ZBO and

a heavy lander with direct return toward Earth.

We will return using ep. process given below:

Mission profile:

Orbiter is separated from the main spacecraft (lander and MAV) after Mars orbit insertion, EP

is used to circularize orbit.

The orbiter waits for rendezvous with the MAV capsule.

Then orbiter EP is used to increase orbit apoapsis and eventually inject orbiter into

interplanetary orbit. EP is used to reach Mars – Earth transfer orbit.

Launch mass 3905 kg (18% L/V mass margin)

Includes EP orbiter + DM + MAV + Carrier

EP orbiter wet mass: 625 kg.

Total xenon used 354 kg, EP Orbiter S/C dry mass 271 kg

Mission duration: 57.5 months (4.8 yrs.)

Operational mission at Mars: 107 days

Descent to nominal orbit 690 days and return up to escape 396 days

Delayed transfers very well suited for dual (simultaneous) launch of electric low thrust and

chemical vehicles.

1PPS®1350 operating at any time (2 for the required lifetime)


47

Decisions:

EP orbiter yields higher mass margins

Longer mission duration allows for significant reduction of

∆V needs ( 10500 hrs. of operations. > 7000 ON-OFF cycles (incl. 50 cold)

Nominal operating point

1500 W / 4,28 A / 350 V

88 mN / 1650 s / 50% total eff.

4,4 kg

Variable power feature demonstrated on

Smart-1

Can allow for cluster operations with


48

Thrust steering capability.

Conclusion:

This is a dreamt mission. A mission for humanity. If there would be no humanity it would be a

failure though it may have so much success. Because making earth livable is more important

than making mars livable.

References:

1. R. Zubrin and D. Baker, "Humans to Mars in 1999,"Aerospace America, August 1990.

2. D. Baker and R. Zubrin, "Mars Direct: Combining Near-Term Technologies to Achieve a

Two-Launch Manned Mars Mission," JBIS, Nov. 1990.

3. R. Zubrin, D. Baker, and O. Gwynne, "Mars Direct: A Simple, Robust, and Cost-Effective

Architecture for the Space Exploration Initiative," AIAA 91-0326, 29th Aerospace Science

Conf., Reno NV Jan.1991.

4. M. Duke, Private communication, Oct. 1992.

5. D. Weaver, "Mars Study Team Reference Mission Overview," Presented at Mars Working

Group conf. NASA Ames Research Center, May, 1993.

6. T. Stafford et-al, “America at the Threshold: Report of the Synthesis Group on America’s

Space Exploration Initiative," U.S. Gov. Print. Off. May 1991

7. K. Joosten, Bret Drake, D. Weaver, and John Soldner, "Mission Design Strategies for the

Human Exploration of Mars," IAF-91-336, 41st Congress of the International Astronautically

Federation, Montreal, Canada, Oct. 1991.

8. R. Zubrin and T. Sulmeisters, "The Application of Nuclear Power and Propulsion for Space

Exploration Missions,” AIAA-92-3778, AIAA/ASME Joint Propulsion Conference, Nashville,

TN, July 1992.

9. A. Cohen et al "the 90 Day study on the Human Exploration of the Moon and Mars,” U.S.

Government Printing Office, 1989.

10. A.Gonzales, L.Harper, E.Dunsky, and B.Roberts, “Mars Surface Mission Life support


49

Team profile:

Team leader: Nazmul Ahsan Saikat; Email: [email protected]

Co- Team leader: Yeasir Mohammad Akib; Email: [email protected]

Other team member: Md. Nur-e-alam Siddiky; Email: [email protected]

Every one of our team member is 3rd year graduate student.

University: Rajshahi University of Engineering & Technology (RUET), Rajshahi, Bangladesh.
mailto:[email protected]

Documents

“Barsoom Inception: A mission for humanity”members.marssociety.org/.../Barsoom_InceptionReport.pdf · 2019. 9. 6. · “Barsoom Inception: A mission for humanity” 3 1. Propulsion: