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Chapter 4 – Sample Return Mission

4 Sample Return Mission

4.1 Mission Overview

4.1.1 Introduction – Matt MaierOne of the most important scientific tasks we are conducting during this mission will be

the return of a Martian soil and rock sample. This mission lasts for the duration of the astronauts

stay in Martian orbit. The astronauts have the task of using the two different rovers to collect the

sample and perform other scientific measurements. Note that the rovers are located on opposite

sides of the planet for maximum communication time. This mission will play a significant role

AAE 450 Senior Spacecraft Design Spring 2004167

Fig. 4.1 Created By Ben Toleman


for future manned missions to the surface of mars. We would like for a human mission to use a

maximum amount of resources found on Mars, this would reduce the mass and cost of putting

the first human on Mars. In conjunction with the data collected from the Mars Exploration and

Pathfinder missions the soil and rock data from our robotic missions will help choose an

appropriate landing site for such a mission. Another benefit we gain from the sample return

mission is the demonstration of producing the required propellant for a Mars to orbit launch. This

is a very important technology that must be proven before a human landing is possible. Other

technological benefits such as precision landing will are also demonstrated in our rover missions.

The rock and soil samples once returned to Earth will provide researchers with data that would

have taken numerous Mars rover mission to accomplish.

4.1.2 Mission Timeline – Matt MaierThe two rover landers are launched shortly after the aero-capture maneuver for the

spacecraft has been completed. Two landers are sent to the surface to ensure the success of the

sample return mission in the event that one fails. These failures include but are not limited to

unsuccessful landing, improper rover or sample return vehicle (SRV) deployment, complications

in propellant production or unfavorable weather conditions. We target the landers at two

different landing sites on different sides of the planet. We need two landing locations for two

different reasons; variety of samples and communication. In the event that both sample return

missions are successful it is beneficial to future missions to have very in-depth analysis of two

different landing sites. Placing the rovers on opposite sides of the planet allows for the design of

our spacecraft’s orbit to ensure that the astronauts are always in contact with at least one landing

site. After the landers touchdown and deploy the rovers a subsystem of the lander starts

producing the propellant for the sample return vehicle using in-situ production processes (section

4.6.4). It is necessary that the SRV should employ this technology not only for the reduction in

mass but also to prove these techniques for future manned missions. During this time the

astronaut controlled rovers collect up to ten kilograms of samples and perform other important

scientific duties. Once the SRV has been fueled it is launched to rendezvous with the spacecraft

orbiting. The rest of this chapter discusses the details of these components and procedures.



4.2 Launch of Rovers

4.2.1 Release of Landers – Allison BahnsenAfter the Transport Vehicle performs aerocapture and the periapsis-raise maneuver, and

prior to the apo-twist maneuver, we release the two landers that venture to the surface of Mars.

The side, cross-sectional profile of the landers in the Transport Vehicle is shown on the

left of Error: Reference source not found. As we can see, the landers are housed within the body

of the main spacecraft. The image on the right shows a top view of the Transport Vehicle. The

protective, hexagonal doors covering the two landers are indicated by arrows. Prior to release,

these doors slide open to reveal the landers.

We release the landers when the Transport Vehicle is traveling as slowly as possible to

reduce propellant costs. The slowest point in the Transport Vehicle’s orbit, seen in blue in Fig.

4.2, occurs at apoapsis, where we release the first lander. This release at apoapsis costs 1.05 m/s,

and places the first lander on the green trajectory in Fig. 4.2 with periapsis altitude at 100 km.

Since the second landing site is on the opposite side of the planet we wait half a sol (half a

Martian day) to release the second lander. Now that the spacecraft is no longer at apoapsis, we

must find the orbit that intersects the current location of the Transport Vehicle and has a


Fig. 4.2 Side View of Landers in Transport and Protective Doors Created by Ben Toleman and David Goedtel

Protective Doors

169


periapsis altitude of 100 km. We can see this trajectory in red in Fig. 4.2, and can transfer to it

for a cost of 1.17 m/s. These results are obtained using the MATLAB code in Appendix G.

Transport Trajectory

Trajectory of 2nd LanderTrajectory of 1st LanderTransport Trajectory

Trajectory of 2nd LanderTrajectory of 1st Lander

Fig. 4.2 Lander Trajectory

For the above calculations we assume that the Transport Vehicle is in the same plane as

the landing sites: the equatorial plane. In actuality the Transport Vehicle is in the ecliptic plane

when we releases the landers, and thus we would have to wait until the entry point in the

atmosphere above the landing site is a node between the ecliptic and equatorial planes. This

plane change would cause the release v’s (or change in velocity) to be three-dimensional, but

their magnitudes would not be much larger than those of the aforementioned values. Solving for

these three-dimensional v’s and the implementation times of hitting the two selected landing

sites are not trivial matters by any means, and thus are out of the scope of this study. This in no

way affects the feasibility of the mission, it simply adds to the complexity of time lining the

overall mission when it comes to fruition.

4.2.2 Cruise Stage – Andy KacmarThe “cruise stage” is the configuration of the aeroshell for transport between the

spacecraft and Mars. Fig. 4.3 shows the full cruise configuration. The cruise stage resembles the

Mars Pathfinder and Exploration mission designs. The major differences that arise come from

the fact that these missions were designed for the system to travel from Earth to Mars where our

stage is only going to transport the aeroshell to the Martian atmosphere and ensure the proper

entry point.



The structure affixed to the

aeroshell is approximately 5.0 m in

diameter and .5 m thick. With a

mass of 235 kg, the structure consists

of a basic aluminum frame with an

inner and outer ring for support. The

top surface of the stage is lined with

solar panels to supply power once

detached from the spacecraft and the

outer ring is lined with radiators to

dissipate any heat build up from the

solar radiation and electronics on

board. For navigation, there are three sun sensors (for redundancy), one star scanner, and an

onboard positioning system coupled with the antenna to relay position and information back to

the HAB.

For correctional maneuvers, the

maximum Δv the system needs is less than 2.0

m/s. This amounts to about 5 kg of fuel when

accounting for departure from the spacecraft,

the correctional maneuver, and excess

propellant left in the two aluminum lined

tanks. The cruise stage consists of two

thruster clusters of four thrusters each running

off of hydrazine propellant running through a

catalyst bed. The clusters allow for

corrections in any direction to ensure a safe

insertion into the Martian atmosphere.


Fig. 4.3 Full Cruise Configuration Created by Ben Toleman

5 m

Thrusters Prop. Tank

Sun Sensor

Solar Panels

Star Scanner

Heaters

Fig. 4.4 Two-D Drawing of Cruise Stage – Created by Rebecca Karnes


4.3 Atmospheric Entry / Touchdown

4.3.1 Landing Sites – Allison Bahnsen

Nomenclature MER = Mars Exploration RoverMGS = Mars Global SurveyorTES = Thermal Emission Spectrometer

One of the main scientific objectives of this mission is to return a Martian rock sample

back to Earth for analysis hopefully leading to many new discoveries, including if life once

inhabited Mars. A major indication that water, the building-block of life, once existed on this

hostile planet is the presence of an iron oxide mineral called hematite. On Earth this mineral is

usually formed in a large body of water in which iron is dissolved and gradually oxidized into

hematite. This insoluble mineral is then precipitated out and mixes in with the lake bottom

sediment which eventually hardens into rock. The hematite deposits on Earth are also one of the

best rocks to serve as home to microscopic fossils of microbes that were trapped in the sediment

before it hardened into rock.1 The presence of crystalline gray hematite on Mars was first

observed by scientists analyzing the Thermal Emission Spectrometer (TES) data obtained from

early phases of the Mars Global Surveyor (MGS) mission.2

Knowing that finding hematite could be the next step to discovering if life once existed

on Mars, the presence of this mineral in the landing sites is a necessity. The first landing site we

select is located in the Terra Meridiani region of Mars, with the exact coordinates of 1.98° S,

6.18° W and a landing ellipse with dimensions of 81.5 km by 11.5 km. 3 Error: Reference source

not found shows a photo mosaic of this region from Viking which is superimposed with data

from the MGS TES. We can see that the exact site, marked with an arrow, is located in an area

with approximately 15% hematite. This landing site is also the location of the Mars Exploration

Rover (MER) Opportunity, which landed there on January 26, 2004. Already a few months into

the mission, this site has proven to be a jackpot in the eyes of scientists containing the largest

concentration of hematite that they have ever seen.4



Aside from having a large distribution of hematite, this site also boasts low wind shear, a

low abundance of boulders and low slope angles in the craters, all of which are positive attributes

when looking to land and operate a rover. The low wind shear in combination with the relatively

low amounts of dust compared to other parts of the planet5 make this site not only a very good

scientific candidate, but also very environmentally appealing.

We select the second landing site on the opposite side of the planet in the Athabasca

Valles at 8.92° N, 205.21° W. One of the main reasons to choose the second site to be on the

opposite side of Mars is for communication issues. This guarantees that one rover will always be

on the side of the planet that is facing the Transport Vehicle, which gives the astronauts the

maximum time to control the rovers. This site was also one of the back-up sites for the MER

mission.Error: Reference source not found We can see the site along with its landing ellipse,

with dimensions of 152 km x 16 km, in Error: Reference source not found.6


Fig. 4.6 Hematite Distribution Map3

173


In addition to the presence of hematite, this site is appealing because as we can see in the

elevation map in Error: Reference source not found, the site is in a large channel system that

could have possibly been cut out by catastrophic floods or some other type of flowing water.

This location is also the seed of a great debate between geologists concerning the age. Some

think it is a geologically young site, while others think it is an ancient site that has just recently

been exhumed.7 Therefore, obtaining a rock sample from this site could settle the dispute.

We can see both of the landing sites on a map of Mars in Fig. 4.58. During the design

process, concerns were expressed with regards to communication and the difference in

inclination between the equatorial landing sites and the 63.4° inclined Transport Vehicle orbit.

These concerns have been addressed and dispelled in full in Appendix G.


Fig. 4.7 Map of Second Landing Site6

2nd Landing Site


4.3.2 Entry Trajectory

4.3.2.1 Mission Timeline – Ayu AbdullahWe present our mission timeline for the Aeroshell containing the Mars Lander and Rover

in Table 4.2 below. This timeline begins at first point of entry into the atmosphere, taken to begin

at 100 km altitude. A graphic timeline is also provided in Error: Reference source not found.


Rover 1: 1.98° S, 6.18° W Rover 2: 8.92° N, 205.21° WRover 1: 1.98° S, 6.18° W Rover 2: 8.92° N, 205.21° W

Fig. 4.5 Landing SitesError: Reference source not found


Table 4.2 Mission TimelineTime (sec) Altitude (km) Event

0.0 100.0Aeroshell with rover enters the atmosphere of Mars at 4.896 km/s and begins the landing sequence of events. Entry, descent and landing (EDL) takes approximately 6.8 minutes.

261.9 9.0 Drogue deploys (304m/s).262.2 8.9 Drogue fills.267.2 7.9 Aeroshell bolts are fired (200m/s). Heat shield separates.272.2 6.9 Parachute attached to lander deploys, releasing it from backshell.272.8 6.8 Parachute fills.

368.9 1.7 Lander altimeter returns information on altitude, rocket-assisted deceleration engines (retro-rockets) fire (85m/s). Bridle cable is cut.

408.9 0.0 Rover lands softly on surface of Mars.

4.3.2.2 Aerocapture – Ryan WhitleyThe equations of motion delineated in Section 3.5.1 also apply to a more general reentry.

Thus, we propagate an entry trajectory for the Lander using these same equations. The Lander’s

desirable trajectory ends at parachute deploy altitude, and will hit the ground if left unchecked.

We choose a parachute deploy altitude of 9 km. Thus, an optimized trajectory contains an initial


Fig. 4.9 Mission timeline

176


flight path angle with the smallest velocity at this altitude, occurring at -8.6250 degrees.

Although this is the ideal angle for obtaining the smallest velocity with nominal initial

conditions, it is not the optimal angle to fly if there are uncertainties. For reentry, a corridor also

exists, and it is desirable to not be near the bounds. Unfortunately, the small angle (-8.6250

degrees) is very close to the shallow skip out angle. Fortunately, the second bound, the upper

constraint on the final velocity, is lax. The final speed increases as the flight path angle becomes

steeper. However, the parachute would deploy successfully at a speed less than or equal to 0.5

km/s. Even with much larger flight path angles, the speed did not come close to this value. Thus,

to accommodate for skip out losses and because the parachute is suitably strong, the nominal

flight path angle is -11 degrees. The nominal trajectory is shown in the following plot:

Fig.

4.6 abov

e shows the

nominal

trajectory, arriving at a speed of .3208 km/s at the specified 9 km altitude. A Monte Carlo

simulation was run to test for mission success. We used the same types of variables that we used

for aerocapture. However, all uncertainties specified in the table (see section 3.5) are increased

by a factor of 10. It is anticipated that the available navigation will be significantly worse than

that available for the transport module. This discrepancy is verified in two navigation articles.9,10


Fig. 4.6 Lander trajectory, altitude vs. velocity

Velocity = .3208 km/s


4.3.3 Aeroshell Design – Ayu Abdullah

Nomenclature BC = ballistic coefficient (kg/m2)m = entry trajectory mass, Rover + Lander + Aeroshell (kg)CD = drag coefficientA = cross-sectional area of Aeroshell (m2)q = maximum heating rate (W/cm2)ρ = density (kg/m3)Rn = nose radius of Aeroshell (m)Rs = shoulder radius of Probe (m)D = diameter of Probe (m)

The critical component of the Mars Lander and Rovers’ atmospheric entry is the

Aeroshell. The Aeroshell encases the Lander and Rover during entry into the Mars atmosphere.

Hence, our Aeroshell design must protect the Lander and Rover from extremely high heat loads.

Our Aeroshell design will also define the entry trajectory.

The three major constraints in our Aeroshell design are heating, deceleration and

accuracy of landing. Deceleration and accuracy of landing are described by entry trajectory.

Deceleration is a major concern as the vehicle and its payload will have to withstand the

maximum deceleration during the entry trajectory. Accuracy of landing is defined as landing in a

certain footprint on Mars, a constraint met by adjusting trajectory.

We find that in any atmospheric entry only two design parameters define the entry

performance (which describes heating and trajectory);11 the lift to drag ratio (L/D) and the

ballistic coefficient, BC. BC is defined in Eq. 4–1:

BC = 4–1

In designing our Aeroshell, we find that the most major concern is its shape, which is

described by the above two parameters.

During atmospheric entry, a blunt-shaped vehicle is more desirable than a more pointed

vehicle for two main reasons:



A blunt vehicle experiences more drag and hence decelerates more rapidly than a pointed

vehicle. Increasing nose bluntness also decreases the maximum stagnation point heating

rate.12

In hypersonics, a blunt vehicle has a detached shock wave, rather than an attached shock

wave. This means that the blunt vehicle distributes heat over a larger volume and overall is

subjected to less maximum heat loading than if it were pointed and had a shock wave

attached.13

Past Mars atmospheric entries (such as the Viking, Pathfinder and Spirit missions)

employed Aeroshell configurations of 70º spherically-blunt cones. Aerodynamic performance is

virtually impossible to obtain theoretically as there are no governing equations available and

computational fluid dynamics (CFD) has not yet progressed to successfully analyze hypersonics.

Hence, our analysis is obtained only from empirical data. We decide that this mission’s

Aeroshell shall also employ the 70º spherically-blunt cone configuration as aerodynamic

performance for this configuration is available from historical data.

Our configuration is a ballistic shape, which means the L/D ratio is zero value at trim

conditions (at zero angle of attack). As for the drag coefficient, CD, we determine this value to be

1.69 after we analyze aerodynamic data14 from previous Mars missions.

To find the shape of the Aeroshell, we use following ratios in Table 4.3:

Table 4.3 Sizing ratios usedRn / D Rs/Rn

Ratio Value 0.25 0.1

We design our Aeroshell shape and size after determining how much internal volume is

needed for the Lander and Rover. Once we know the volume needed, we make a basic CATIA

drawing (using the ratios in Table 4.2) of the Aeroshell with a model Lander and Rover placed

inside. Fitting the Lander and Rover inside, we obtain the dimensions of the Aeroshell from the

drawing. We obtain the surface area, A = 30.59 m2. We use this surface area, A to find BC and

accordingly, entry trajectory.

We now have an Aeroshell shape and size shown in Fig. 4.7 Aeroshell below. We must

now determine its mass. The major concern involved in determining its mass is heating. Using

the following equation, the Aeroshell is subjected to maximum heating rate:



4–2

Maximum heating is 256.13 W/cm2. We use this value to determine the Aeroshell’s

thickness as well as materials used for the Aeroshell. The Aeroshell thickness must withstand

this maximum heating. This analysis is found in section 4.3.3.2.

Once we find the thickness, we incorporate this into the CATIA drawing of the

Aeroshell. We also insert the material properties into CATIA to obtain the mass. We obtain the

total mass of the Aeroshell to be 595 kg.

We also conduct a Finite Element Methods (FEM) stress analysis on the Aeroshell to

confirm structural integrity. We must ensure that the Aeroshell’s structure can withstand forces

during entry. This FEM analysis is shown in Fig. 4.8 FEM analysis.15 The Aeroshell is made of

three layers, one of which is the honeycomb layer (Analysis in section 4.3.3.2). We conduct the

analysis on the honeycomb layer of the Aeroshell, as this layer is designed to withstand most of

the structural loads. We find (Table 4.4Table 4.2) that the stresses the Aeroshell is subjected to is

below the honeycomb’s yield stress (σy = 6.89 106 N/m2).


Fig. 4.7 Aeroshell


Table 4.4 Aeroshell maximum stresses and displacement Parameter Maximum value

von Mises stress 2.24 104 N/m2

Displacement 4.62 mm

Compressive stress 2.14 104 N/m2

We also plot the Aeroshell’s atmospheric entry trajectory using data from an integrated

code16. After we analyze and plot this data, trajectory and other parameters are found. Entry

trajectory parameters are as in the Table 4.5 below:

Table 4.5 Aeroshell entry trajectory parametersParameter Value

BC 49.07 kg/m2

Maximum G-loading 5.03 Earth G’s

Estimated cross range 727 km

Analysis and trajectory is found in Appendix G.


Fig. 4.8 FEM analysis


4.3.3.1 Monte Carlo Analysis – Ayu AbdullahWe conduct a Monte Carlo Analysis17 where 5000 test cases are run at a nominal flight

path angle to study the possibility of different velocities at time of drogue deployment. The

drogue can withstand a maximum velocity of 510 m/s. Hence, a Monte Carlo analysis is

conducted to determine the probability that the drogue will withstand velocity at altitude of

deployment. We run test cases varying density and dust levels, with low navigation accuracy to

provide a worst case scenario.

From the 5000 test cases, we find that only four cases are above the maximum velocity.

This is a 0.08% failure rate and these failures are all skip-out angle failures. Skip-out angle

failures are when the flight path angle is too shallow for the Aeroshell to penetrate into the Mars

atmosphere and will result in the Aeroshell bouncing off the Mars atmosphere. This failure is

defined as a mission failure as bouncing off the atmosphere results in failing to land the Rover on

Mars. This analysis then yields a 99.92% mission success rate. The Error: Reference source not

found below shows the possible velocities at time of deployment.



4.3.3.2 Material Analysis – Matthew Branson

Analysis of the Mars Lander heat shield is similar to the analysis done on the aerocapture

aero shell even though the layering scheme is different. The heat data discussed earlier is used in

conjunction with the Matlab18 and SODDIT19, 20 codes discussed in 3.6.7. Error: Reference

source not found shows the scheme developed after optimizing the mass and thermal properties

1 Moomaw, Bruce, “Uncovering The Meridiani Formation,” Space Daily, 04/02/01, http://www.spacedaily.com/news/mars2003-01a3.html2 Martel, Linda, “Grey Iron Oxide in Meridiani, Mars”, PSRD Discoveries, 03/13/03, http://www.psrd.hawaii.edu/Mar03/Meridiani.html5 Astrobiology Magazine Staffwriter, “Mars: Upstairs, Downstairs,” 01/29/04 http://www.astrobio.net/news/modules.php?op=modload&name=News&file=article&sid=81112 Spencer, David A., Blanchard, Robert C., Thurman, Sam W., Braun, Robert D., Peng, Chia-Yen, Kallemeyn, Pieter H., “Mars Pathfinder Atmospheric Entry Reconstruction”.13 Sermeus, K., “Applications of Steady Perfect Gas CFD on Unstructured Grids”, Eurovia/Mission to Mars Symposium.14 Prabhu, Ramadas K, Lockheed Martin Engineering & Sciences Company, Hampton, Virginia.


Fig. 4.13 Failure analysis for drogue deployment

183


of the materials.

Error: Reference source not found is the thicknesses used on the Mars Lander’s heat shield.

We use an ablative material to greatly reduce the heat loads. The ablator protects the

spacecraft by absorbing energy while chemically decomposing.21 The heat absorption capacity

greatly out weighs the mass density of the graphite ablator.

We select Glass Reinforced Polyimide Honeycomb (GRPH) for the main insulator with a

layer of carbon-carbon reinforced composite (C-C composite) on the outside since GRPH is not

strong enough to withstand the anticipated loads.22 Specific analysis of the heat shield can be

found in Appendix I.


Fig. 4.14 Layering scheme for Mars Lander heat shield

Table 4.6 Material thickness for Mars Lander heat shield

Material of Each LayerThickness (cm)Graphite Ablator0.1Carbon-Carbon Composite0.1Glass

Reinforced Polyimide Honeycomb10

184


4.3.4 Parachute Systems — Andy Kacmar

Nomenclature DO = nominal canopy diameterL(SL,R ) = length of linesN(SL,R,G ) = number countq = dynamic pressureSO = canopy surface areaW(C,SL,RT,R ) = specific weight of design material

Given the weight of the aeroshell, along with its shape, the shell will continue to slow

while plummeting through the atmosphere, but will not slow enough for a direct deceleration

with a parachute; the opening force is too great. For this reason, we break the procedure into two

separate maneuvers. The First stage deploys a single, 10 m drogue from the backshell and slows

the entire system down to about 200 m/s. Explosive bolts fire once the shell reaches terminal

velocity and releases the Lander from the heat shield. The Lander drops from the shell and a

second parachute fires to slow the Lander’s decent down to 85 m/s.

The components of the parachute and the construction materials are shown in Table 4.6.

The canopy material weight scales with the surface area while the suspension lines, radial tape,

and risers all scale with the force they are designed to withstand along with their respective

lengths. A Nylon/Kevlar blend is chosen for the canopy because of its strength and low weight

characteristics while Kevlar lines connect the canopy to the body to ensure the parachute doesn’t

disconnect during to the large force upon opening.

Error: Reference source not found


Table 4.6 Specific Weights of Parachute MaterialsVariable Name Material Specific Weight

WC (canopy) Nylon/Kevlar .0115 lb/ft2

WSL (suspension lines)

Kevlar .0035 lb/ft/1000 lb strength

WRT (radial tape) Kevlar .0035 lb/ft/1000 lb strength

WR (riser) Kevlar .0035 lb/ft/1000 lb strength


We designed both parachute systems using the approach outlined in Appendix G. The

opening force on each system is approximately 40,000 N, but we built the parachutes to

withstand an opening force of 100,000 N. Due to the large fluctuations in the atmosphere, the

opening velocity or opening density could cause the opening force to be greater than the

predicted value. We hold the design limit at 100,000 N because there are preexisting parachutes

specifically designed for the Martian atmosphere built to withstand an opening force of

approximately that magnitude. We took the maximum atmospheric fluctuations into account

when designing the parachutes, but strengthening them to withstand the limiting opening force

does not add a significant amount of mass. Table 4.7 shows all the dimensions of the two

parachute systems as well as system masses and packing volumes.


Table 4.7 Parachute Dimensions and SizesParameters Design Values

Drogue LanderSO [m2] 170 385DO [m] 10.4 16.7NSL 48 48LSL [m] 16 23NR 1 5LR [m] 5 3NG 48 48

Volume [m3] .021 .039Total mass [kg] 17 32


4.3.1 Retro Rockets

Nomenclature c* = characteristic exhaust velocitycF = thrust coefficientDstop = distance required to stop LanderΔV = total change in velocity of Landerε = expansion ratioF = thrust per rocket in the direction the rocket is facinggm = acceleration due to gravity on Marsgo = acceleration due to gravity on EarthIsp = specific impulseLcham = combustion chamber lengthLnoz = nozzle lengthmi = initial Lander mass at beginning of rocket firingmf = final Lander mass at end of rocket firing (does not include use of lateral motion fuel)Rcham = time index during navigationRexit = radius of rocket nozzle exitRthroat = radius of rocket throat

4.3.1.1 Purpose of the Retro Rockets – Frankie HankinsWe employ a system of four retro rockets to assist in the descent of the Mars Landers.

These rockets run on Methane/LOX fuel. We place the rockets at an angle of –45o from the

horizontal. This angle allows the rockets to have the robust ability to move the Lander laterally

if necessary as well as slow the Lander’s descent. The Lander may need to move laterally to

avoid poor landing areas on the surface of Mars. Such areas may damage the Lander on

touchdown, cause it to be tilted, or make the deployment of the rover difficult or impossible. It

is not expected to be a difficult task to find a suitable landing site on the surface of Mars, mainly

due to the relative flatness and lack of large rocks in the selected landing areas.



4.3.1.2 Configuration

We place 4 retro rockets evenly spaced around the Lander. The positions of the rockets

are shown in Fig. 4.9. We place the rockets toward the bottom of the Lander as shown. The four

legs are attached to the empty sides of the octagonal Lander. In this arrangement, the rockets do

not adversely affect the legs during the firing and there is space for both to connect to the

Lander.


Fig. 4.9 To Scale Lander Configuration Side ViewCreated by Ben Toleman


Fig. 4.10 and Fig. 4.11 show the propellant (red) and oxidizer (green) lines. All

propellant and oxidizer originates from the tanks within the sample return vehicle (SRV) shown

in the Fig.s. We can draw all the rocket fuel from the SRV because the SRV tanks will be

refilled while on the surface of Mars, so there is no need for separate tanks for the Lander

propulsion system. Using fewer tanks saves a large amount of mass and space. We pass the fuel

lines outside the SRV in two locations and place them along the bottom of the Lander. We

position them so that they are out of the way of the Rover and go around all other components

inside the Lander.


Fig. 4.10 Lander Configuration Top View

Fig. 4.11 Lander Configuration Side View


4.3.1.3 Retro Rocket Specifications We designed the retro rockets with the considerations shown in Table 4.8.

The ΔV is the terminal velocity provided by the parachutes that are deployed before the

rockets fire. We get the mfinal from the masses of the components that will be on the Lander

when the rockets fire. Components such as the aeroshell, cruise stage, and parachutes will have

been jettisoned before the rockets are fired and are therefore not included in m final. We also

included 50 kg of spare propellant in this mass for the lateral movements described previously.

We set the burn time based on the previous similar mission, Viking. The chamber pressure is a

typical number for rockets of this type. A higher chamber pressure will give better performance.

We used this expansion ratio because better performance is realized in near vacuum conditions

with larger expansion ratios.

We enter the Methane/LOX fuel combination and chamber pressure into the NASA

thermochemistry code to obtain the data in Table 4.9. From the given Isp, we can now employ

the rocket equation (or Tsiolovsky equation), Eq. 4–3. The rocket equation gives the initial-to-

final mass ratio. With this value and the final mass, we find the necessary propellant mass to

provide the required ΔV.

4–3


Table 4.8 Initial Design ConsiderationsΔV mfinal tb Pc ε

85 m/s 1575 kg 40 s 3 MPa 30

Table 4.9 NASA Thermochemistry Code DataIsp cF c*

364 s 1.915 1865 m/s


The rocket equation gives an initial mass of 1698 kg and a propellant mass of 123.35 kg,

this propellant mass does not include the extra propellant for lateral movements. Altogether we

have a propellant mass of 173.35 kg.

Further analysis23 gives us the data in Table 4.10 for each rocket.

A to scale view of one

of the retro rockets is given in

Fig. 4.12. We chose the

material for the chamber to be

Columbium, a typical Nickel-

based thrust chamber material.

The density is 8600 kg/m3 and

the tensile strength is 310

MPa.Error: Reference source

not found The material for the

nozzle is a Carbon-Carbon

composite that has a density of

1680 kg/m3 and tensile strength

of 67.6 MPa.24 The nozzle can be made of a lighter material because it has less stringent

requirements in the areas of tensile strength and temperature resistance. These values give the

masses of the chamber and nozzle as 0.1411 kg and 0.0192 kg respectively per rocket.

Therefore, the total mass of the 4 rockets together is 0.641 kg. While 0.641 kg may seem like a

very small mass for four rockets, it is reflected in the value of the Dstop parameter. The

parachutes will put the Lander at the terminal velocity of 85 m/s at a very high altitude, which

allows the stopping distance to be large. A large stopping distance allows for the rockets to be of

little consequence in terms of mass.


Table 4.10 Rocket Design DataF Dstop Rthroat Rexit Rcham Lcham Lnoz

1739 N 2408 m 0.0098 m 0.054 m 0.0252 m 0.193 m 0.131 m

Fig. 4.12 Retro Rocket Image Created by Ben Toleman


4.4 Lander

4.4.1 Introduction - Dan NakaimaAs part of the mission, we are to obtain and return up to ten kilograms of Martian sample

(i.e. soil, rock, etc). A lander designed to carry all the tools such as the Martian Rover, the

Sample Return Vehicle (SRV) and other components accomplishes such a mission. The Martian

Rover gathers data, obtains and stores sample. The SRV delivers the sample to the crew in orbit.

Other components include landing, pumps, communication and power systems. Geometry,

volume and mass are the design parameters, but for a successful mission, the Lander also needs

to endure all the loads applied during Earth launch and Mars entry.

4.4.2 Layout - Dan Nakaima We separate the Lander into two parts, the lower and upper body. The upper body stores

most of the Lander components, together with the Rover and the SRV. The lower body includes

the legs and the retro rockets. Table 4.11 and Fig. 4.13 show the dimensions of the Lander. Fig.

4.14 shows how the Lander accommodates the SRV and the Rover.

Table 4.11 Lander’s sizes and masses


Panel Number Length (m) Height (m) Thickness (cm) Mass (kg)

Side A 4 1.3 1.1 2 14.3

Side B 4 1.4 1.1 2 15.6

Top 1 N/A N/A 1 44.5

Bottom 1 N/A N/A 10 444.9

Total Mass (kg) 609.0

192


Fig. 4.13 Side and top view of the Lander (exaggerated for explanatory reasons)


Top Panel

193

Bottom

Panel

Leg A

Leg B

Leg B

Side Panel A

Side Panel

B

Leg A


Fig. 4.14 Fig. shows how the SRV, the Rover and components are accommodated - Created by Ben Toleman

The designed legs not only provide stability for the lander but also prevent the retro

rockets from touching the ground. We chose an octagonal geometry not only due to space

purpose, but because a side panel serves as a ramp for the Rover once the Lander touches

ground. Having the SRV placed in the middle of the Lander gives a more evenly distributed

mass across the Lander and results in a simpler pump system for feeding propellant from the

SRV's tank to the four altitude control rockets.

4.4.3 Design Specifics

4.4.3.1 Structure -Dan Nakaima The mass of the Lander with the aero-shell approximately totals 2,500 kg, and during Mars

entry the Lander is subjected to 4-5 Gs. A big part of the structure design involves the material

selection, which varies from the traditional aluminum to the high-tech composites. During the

design process we considered two materials, Aluminum and Honeycomb composites.

Aluminum exhibits low density and a high Elastic Modulus, but for our purpose Aluminum’s



density is not low enough. Honeycomb Composites have even lower densities ranging from 15-

900 kg/m3, which makes it a great material to save mass.25 We chose Carbon Fabric honeycombs

for being stronger and lighter than some Aluminum honeycombs. Assuming the entire structure

of the body is composed of Carbon Fabric composite, we obtain a mass of 609.0 kg.

The Lander’s legs support and stabilize the entire Lander during landing and throughout

the entire mission. To obtain a stable Lander we design it so that its center of gravity is as close

as possible to the ground. To prevent any unexpected damage to the retro rockets, the nozzles are

not in direct contact with the surface, leaving a clearance of about 30 cm between the ground and

the bottom of the Lander. Table 4.12shows the legs sizes and masses. Each leg can be simplified

into a system of three steel rigid rods with diameter of 5 cm and lengths of 0.95 and 1.0 m. We

chose steel as the primary material for the legs, because of its traditional use in aircraft landing

gear and high Modulus of Elasticity, which yields a small compact system. A structure, located

on each foot, crunches itself and acts as a shock absorber providing a softer landing.

Table 4.12 Lander’s legs sizes and massesLeg Number Length (m) Diameter (cm) Mass (kg)

A 4 0.95 5 14.7

B 8 1.0 5 15.4




4.4.3.2 Rover Deployment – Andy Kacmar The rover rests parallel to the

Sample Return Vehicle, as can be seen

in Fig. 4.14, while fixed within the

Lander. The side panel parallel to the

Rover, the upper most panel in Fig. 4.14,

is hinged and connects to a small motor

that lowers the panel to allow the rover

to exit and reenter the Lander. The rover

is tightly fastened within the structure,

so it has to back up and do a point turn

to exit straight from the Lander side.

Fig. 4.15 shows the Rover exiting the

Lander after the ramp is deployed. The

falling side allows the rover to reach the

Martian surface and find an adequate

sample to return. Once a valuable sample is found, the rover enters the Lander by the same

means it exited, and detaches the storage unit within the SRV compartment.

4.4.3.3 SRV Deployment – Andy Kacmar Once the SRV is fully fueled, and

the sample is secured within its

compartment, the rocket begins the

deployment procedure. The SRV, while in

the Lander, rests on box beam rails to

secure it in place. The rails connect to a

form fitted platform at the base of the

rocket to allow ground clearance for

takeoff. Lifting arms connect each rail to

the Lander body and control the position


Fig. 4.15 Rover Deployment – Created by Ben Toleman

Fig. 4.16 SRV Launch Configuration –Created by Ben Toleman


of the SRV. Fig. 4.16 shows the SRV in launch position. The arms slide the base and rail off the

floor of the Lander and rotate the SRV about the lower edge of the Lander. The lifting arms

raise the rocket into a vertical position and ready the vehicle for launch.

4.4.3.4 Power - Ben Phillips The Martian lander must safely guide the rover to the surface of Mars and then produce

the propellant needed for the sample return rocket to lift-off. These are the mission requirements

for the Martian lander. As is always the case in spacecraft design, we must tailor the power

system to the specific objectives of the mission.

The power needs for the Martian lander are driven by the propellant production. This

single mission requirement outweighs the other power draws by an order of magnitude. The

power needs for the lander are (1) the in-situ propellant production, (2) communication with the

rover and the orbiting hab module, and (3) positioning the sample return rocket into a position

for lift-off.

The power needs of the lander must be examined before a choice can be made on the type

of power system. The estimated power needed for the Martian lander to produce enough

propellant is about 400 Watts for 300 days. This is a very large power need and can only be

reasonably accommodated by using a radio-isotope (RTG) power system.

A radio-isotope power system is the best choice for a number of reasons. The first

concern is the long duration of power that is needed. To produce power for 300 days without

any human interference is a difficult task. This power problem could only be solved with either

a RTG power system or a very large solar array and battery system. However, the mass of the

solar array/battery system would be prohibitively large because of the degradation of the solar

panels and the mass of the batteries. The batteries would be massive because of the number of

charge cycles that is needed for the one-year lifetime on Mars. This is because the amount of

charge that a battery can hold decreases each time a battery is charged and then discharged. To

take this into account, the batteries would be built much larger than they would need to be

initially. The solar arrays degrade with time because of Martian dust that would settle on the

arrays themselves. With time, less and less sunlight would reach the arrays and the power output

would fall. The estimated mass of a solar array/battery system is around 150 kg.



A radio-isotope system is a better choice because of two reasons. The first reason is

because of the mass savings that a RTG system would introduce. By using currently built RTG

systems as a guideline26, it can be estimated that

radio-isotope system that produces 400 Watts

for one year has a mass of 75 kg27. Secondly,

the power that the RTG system produces is

more constant than solar power and will not

degrade as quickly. This can be done by

choosing an isotope with a relatively long half-

life. This means that the Martian lander will

have more time to complete its objectives. This

is good in the event that the in situ propellant

production takes longer than was anticipated.

Once the propellant production is completed, the rocket must be moved to a launch-ready

configuration and once again the RTG system has an advantage. After a year on the ground, the

radio-isotope will still be producing a large amount of power that can be used to move

approximately 1000 kg rocket to a more upright position.

The design specifics for lander’s radio-isotope power system are as follows. The lander

will use a cylinder that has a mass of 75 kg and a diameter of 0.5 meters. The length of the

cylinder is 1.5 meters. This sizing can easily fit within the lander and not interfere with any

other placing requirements for the lander. Images of the RTG system are shown in Fig. 4.17.

The reasonable size and mass of the RTG power system gives it a considerable advantage

over the solar array/battery alternative. The only drawback to RTG power is its use of

radioactive material as fuel. This could have a public reaction consequence, but in this situation

the use of the radio-isotope is acceptable. The propulsion system for the crew habitation module

is carrying a full-fledged nuclear reactor and the consequences of an accident with that system

far outweigh the relatively small radio-isotope power system.

In conclusion, the relatively high power requirements of the Martian lander and the long

lifetime needed led to the choice of a radio-isotope power system. The RTG system has several

advantages including low mass and volume and a long mission lifetime. The only drawback is


Fig. 4.17 Lander RTG Power System - Created by Ben Toleman


overshadowed by other systems that are being brought to Mars and should not be considered at

this stage.

4.4.3.5 Mars Lander Communication – Leigh JanesThe antenna on the Lander intended for communication with the Rover is a dipole, half

wavelength, ultra high frequency (UHF) antenna. The height of the antenna is 33.83 cm. The

antenna transmits at a frequency of 420 MHz and receives at a frequency of 410 MHz. The

difference in frequencies allows for uplink and downlink on the same antenna. The Lander has

only one antenna for UHF frequency transmissions, as opposed to one antenna for receiving

signals and another for transmitting signals. We design the UHF Lander antenna to transmit at a

power of 0.23 mW with a maximum link distance of 1 km. The specifications for the Lander

UHF antenna are given in Table 4.13.

Table 4.13 Lander UHF Antenna Link BudgetLander to Rover

Frequency 0.42 GHz

Efficiency Transmitting 0.65

Efficiency Receiving 0.65

Bit Error Rate 5.00e-6 bps

Link Margin 2 dB

Noise Temperature 300 K

Atmospheric Loss 2 dB

Distance of Transmission 1 km

Data Rate 2.00e-4 bps

Power 0.081 mW

Mass 0.0365 kg

The Lander also has a high gain antenna (HGA) which is located next to the UHF

antenna. This antenna is used for communication with the Transport Vehicle, for purposes such

as monitoring the propellant production for the Sample Return Vehicle (SRV). The high gain

antenna has a diameter of 0.32 m and a transmitting power of 10 W. It transmits on a frequency



of 21.2 GHz and receives on a frequency of 23.6 GHz. Both of these frequencies are Ka-band

frequencies. The high gain antenna on the Lander has the same specifications as those of the

high gain antenna that is located on the Rover, for the convenience of manufacturing. The

complete specifications for the high gain antenna are presented in Table 4.14.

Table 4.14 Lander to Transport Vehicle Link BudgetLander to Transport Vehicle

Frequency 21.2 GHz

Diameter Receiving 2 m




Link Margin 2 dB



Distance of Transmission 229,700 km

Data Rate 10 Mbps

Diameter Transmitting 0.32 m

Power 10 W

Mass 0.4 kg

We placed the UHF antenna and high gain antenna on a platform over the Lander

computers. This placement allows for close proximity to the computers required for operation,

and allows for good transmission conditions. The Lander has walls that would obstruct the radio

signals if the antennas were not elevated. For transport and cruise, the antennas are located in

their stowed positions. The UHF antenna is retracted into the antenna stand to protect it from

damage until the Lander arrives at its final resting place on the Martian surface. The high gain

antenna is stored such that it is parallel to the bottom of the Lander. These stowed positions are

shown in Error: Reference source not found. Once the Lander is on the Martian surface the

antennas are deployed. Error: Reference source not found shows the deployed positions of both



the UHF antenna and the high gain antenna. The high gain antenna has the mobility to position

itself so that it may point towards the Rover devoted antenna on the Transport Vehicle. We

create a directed radio signal by being able to position the high gain antenna in the direction of

the Transport Vehicle.


Fig. 4.25 UHF Antenna and High Gain Antenna on Martian Surface – By Ben

Toleman

Fig. 4.25 Lander UHF Antenna and High Gain Antenna in Stowed Position

on Lander– By Ben Toleman


4.5 Rover

Fig. 4.18 CAD Images of our Rover - By Ben Toleman

4.5.1 Mission Design - Masaaki AtsutaThe requirement of our mission is to collect at least 10 kg of samples such as rocks and

soil from the surface on Mars.

After touchdown, the stowed rover deploys the antennas and raises the mast and releases

the arms. Then, the astronauts on the spaceship communicate with the rover and perform a

health check. After the health check, the rover ventures out from the lander and begins a one

(Earth) year journey on the Martian surface.

When the astronauts find a candidate for the samples, they command the rover to

approach the target and to analyze it by using its science instruments. Once they find samples

interesting enough to return to Earth, the rover picks it up and delivers it to a Radiation Detector.

Only when the Radiation Detector determines the sample is not harmful to people, the rover puts

it into the Sample Container.



Once the rover houses samples in the Sample Container, it starts on its way home. When

the rover gets to the Sample Return Vehicle, it rolls up the ramp and puts the container inside the

vehicle. After the rover replaces a new Sample Container, it rolls down the ramp and leaves for

the next adventure.

During 365 Earth days, the rover must document and collects a set of samples consisting

different types of rocks and soil and ensure at least 10 kg of sample mass.

4.5.2 Design Specifics

4.5.2.1 Structure - Masaaki AtsutaAs we can see in Fig. 1, our rover is almost identical to its predecessor, a Mars

Exploration Rover (MER). The size of the rover is also similar to that of the MER, about 150 kg

in mass, 1.2 m long, 1.0 m wide, and 0.7 m tall with its mast deployed. Like the MER,

the rover has six wheels and two pairs of cameras perched on the end of the long neck.

Despite the appearance, our rover is more powerful and sophisticated than the MER. The

rover uses a Radioisotope Power System so that it can handle its 365 Earth-day mission on the

Marian surface, showing a dramatic increase over the 90 Martian-day (92 Earth-days) life of the

solar-powered MER.

The body of the rover, the Warm Electronics Box (WEB), mainly consists of a 5056

Aluminum honeycomb composites insulted with a high-tech material aerogel so all electronic

components for the rover can survive at cold nighttime temperatures (-96 C).28 Also, all the

electronic components are radiation-hardened and protected against cosmic radiation.

The rover has two identical robotic arms. The right arm is for science interments: a

Raman spectrometer, Alpha Proton X-ray spectrometer (APXS), and Microscopic imager. The

left arm is for sample collection tools: a parallel gripper and scoop.

4.5.2.2 Rover Design - Masaaki AtsutaThe rover for our mission must meet the following requirements:

1) The rover must collect Martian samples of at least one kilogram and put them to the

Sample Return Vehicle (SRV).

2) The development cost must be as low as possible.

3) The rover must run for 365 Earth days.



To meet requirement 1) , our rover needs a robotic arm to collect a sample. At the same

time, our rover has to bring its science instruments close to or on the sample to analyze it. We

equip our rover with two arms: the left arm for sample collection tools and the right arm for

science instruments. So, the left arm can smoothly move a light-weight sample collection tool

such as a parallel gripper and scoop to a target, while the right arm can carry a heavy but

sophisticated science instrument such as a Raman spectrometer. Also, to make sure that a

sample that the rover brings back to the SRV doesn’t cause any biohazards, we equip our rover

with a Radiation Detector and a Sample Container. The Sample Container covers a sample with

a Biobarrier made by Tyvek ® (a fiber sheet , which can protect against bacteria and virus and is

used for a chemical suit) and Planova ® (a virus removal filter).29,30,31,32

To keep costs down, we can use as many elements as past successful missions as

possible. Also, the use of proven hardware increases the reliability of a rover. In fact, NASA

plans to use a Mars Exploration Rover 2003 class rover with the capability to collect rocks for

their first Mars Sample Return Rover Mission.33 We borrow many components from the MER.

To keep a rover alive for a year on Mars, our rover uses a Radioisotope Power System as

a main power source. A Solar Power System, a traditional rover power source, is not enough for

our mission because dust accumulation on the solar panels limits, for example, the life of the

MER to about 90 days. However, a Radioisotope Power System may last for 10 years. The

Viking Landers 1 and 2 used this power system and they had functioned for six years until the

last lander was shut down. Besides, the power production is independent of the day / night cycle

and the distance from the Sun.34 NASA has also considered Radioisotope Power Supplies as a

power system option for their 2009 Mars Science Laboratory (Rover) Mission.35,36

4.5.2.3 Components - Masaaki AtsutaOur rover uses blushed DC motors that can function in the carbon dioxide atmosphere at

temperatures between –120 C and + 45 C and can survive between –120 C and + 110 C.37

The rover’s suspension for wheels uses a rocker-bogie mobility system. This system

doesn’t use springs but rotates its joints to rock the rover’s body up and down depending on the

positions of the wheels to keep the rover balanced on the rough surface. The wheel diameter is

0.25 m and the ground clearance is 0.30 m so that it can easily overcome rocks that is taller than

its wheel diameter.



The computer in the rover runs with a 64-bit RAD750, radiation-hardened version of the

PowerPC750TM. It also runs the reliable and flexible VxWorks, real-time operating system. 8

Inertial Measurement Unit (IMU) provides tri-axial information on its position, allowing

the rover to make precise vertical, horizontal and rotating movements. This rover uses the same

hardware to tell an F-16 pilot which way is up and down to just drive on the ground, because it is

challenging to keep a good idea of its position and which way it is heading on the rough Martian

Surface.38

4.5.1.1 Rover Capabilities - Masaaki Atsuta Our rover has a maximum speed of 5 cm/s. The service life of the wheel actuators is

usually between 1000 hours and 3000 hours.

The Radioisotope Power System produces 4800 watt-hours of energy per Martian day.

This energy is about five times more than the energy that the Solar power system of the Mars

Exploration Rover can produce (900 watt-hours per Martian day).Error: Reference source not

found,Error: Reference source not found The power production of the radioisotope power is independent of not

only the accumulation of dust, but also the day/night cycle and the distance from the sun. Error:

Reference source not found

Our rover also has an ultimate hazard avoidance system, the crew on the spaceship flying

around Mars. Since they can control the rover, their closeness makes hazard avoidance almost

instant.

With this power and hazard avoidance system, our rover will be able to travel at close to

the top speed and for more than1000 hours, even under unfavorable environment condition on

Mars. We expect the rover can reach a total distance about 180 km. We suggest that the rover

travel for a maximum 2.5 hours/day.39,40,41

Thus, our rover can aggressively explore Mars and collect interesting samples for return

to earth

4.5.1.2 Power - Ben PhillipsThe Mars rover requires a robust and reliable power system to survive the year that it will

stay and operate on the Martian surface. Several different options were considered when

planning began on the Martian rover, with solar cells, batteries, fuel cells, and radio-isotope



systems among them. To effectively choose a power source for the Martian rover, we have to

consider the mission requirements.

The Martian rover needs to gather up ten kilograms of Martian dust and rocks and then

return them to the Martian lander. The lander will then send the Martian rocks back up to the

orbiting astronauts via rocket. After the rover returns the Martian rocks to the lander, then it will

move around the surface of Mars, conducting experiments and gathering data for an entire year.

The one-year mission lifetime requirement is the limiting factor when choosing a power

source for the rover. The extreme length of time requiring continuous power rules out several

potential power sources. Batteries and fuel cells are immediately dropped from consideration

because of the very large mass that would be needed to provide power for a year.

This leaves us with two possible power options: a solar cell/battery combination or use of

a radio-isotope power system. Even still, the solar array for the rover would be rather large and

the power that the array would produce would diminish with time because of dust accumulation

on the arrays. Also, the batteries that the rover would need would be relatively massive because

of the number of times that they would be charged and discharged. As a battery is charged and

discharged, the amount of power that it can hold decreases with each charge cycle. This must be

taken into account and can lead to batteries that are prohibitively large.

On the other hand, the rover could also make use of a radio-isotope power system. These

power systems have been used on many previous space missions and can produce power for long

periods of time. Radio-isotope (RTG) power systems also have an excellent safety record; a

mission has never failed because of a RTG power system. A RTG power system takes

advantage of the heat that is given off when certain radioactive materials decay. This heat is

converted into electricity, which powers the Martian rover. The power output of a radio-isotope

system decreases slowly with time, but this decrease can readily be taken into account when

sizing the system.

The main factor that has kept RTG power systems from becoming more widespread is

public concern. Radio-isotope power systems do contain radioactive material that could be

harmful to the environment if an accident should occur. However, this is a possibility that can be

planned for in advance. With careful planning, the chance of the RTG system spilling

radioactive material into the environment can be significantly diminished. Moreover, the RTG

system was refurbished and then used again on another launch. Radio-isotope systems maintain



this level of reliability because they contain no moving parts. The level of safety and robustness

factors heavily in choosing the power system for the rover.

Now that a RTG system has been chosen for the rover, it must be sized so that the

volume and mass can be known. The Martian rover needs 200 Watts of power to operate

continuously, and this can be provided by a RTG system with a mass of 25 kg. This mass was

arrived at by examining RTG systems in use today and extrapolating42 43. The specific plans for

the rover call for two cylinders that will fit inside the body casing. The dimensions of these

cylinders are 0.4 meters in diameter and 0.4 meters in length. Excess heat produced by the RTG

system can be easily dispersed with the use of a heat pump inside the rover. This will insure the

operability of the temperature sensitive electronic equipment that is housed inside the rover.

The radio-isotope system is the only main power source for the rover. No batteries are

carried onboard to provide secondary or backup power. This is because of the extremely long

mission life. If the RTG system failed in any way, batteries could not provide power for any

meaningful amount of time and also remain at a sensible mass.

In conclusion, the long mission lifetime of the Martian rovers led to the need for a radio-

isotope power system. No other power source can provide the same amount of power for the

same duration. The safety concerns for a radio-isotope are also not sensible because the Mars

mission requires the use of a full-blown nuclear reactor. The safety concerns for this system far

outweigh any concerns for a relatively small radio-isotope power system.

4.5.1.3 Mars Rover Communication – Leigh Janes The Mars Rover and Lander communicate on an ultra high frequency (UHF). A dipole,

half wavelength antenna with a height of 34.75 cm is located on the Rover. With a transmitting

frequency of 410 MHz, and a transmitting power of 0.22 mW, the maximum link distance of the

radio transmission is 1 km. This means that the Rover has the ability to move in a circle of a

radius of 1 km, with the Lander as the center point. The antenna is made of aluminum 6061-T6,

resulting in a mass of 0.0374 kg. Table 4.15 contains the specifications for the UHF antenna

located on the rover. The placement of the UHF antenna on the Rover is shown in Fig. 4.19.



Table 4.15 Rover UHF Link BudgetRover to Lander

Frequency 0.41 GHz




Link Margin 2 dB



Distance of Transmission 1 km

Data Rate 2.00e-4 bps

Power 0.22 mW

Mass 0.0374 kg

4.5.1.3.1 Mars Rover High Gain Antenna Communication between the Mars Rover and the Transport Vehicle occurs on a Ka-band

frequency. The Rover has a high gain antenna (HGA) of diameter 0.32 m. The antenna

transmits at a frequency of 21.2 GHz and receives at a frequency of 23.6 GHz, with a maximum

link distance of 229,700 km. The transmitting power of the antenna will be 10 W. Table 4.16


Fig. 4.19 UHF and HGA Antenna Placement on Rovers – Created by Ben Toleman


contains the specifications of the high gain antenna. The high gain antenna placement is shown

in Fig. 4.19. The Rover HGA has the ability to point in the direction of the Transport Vehicle,

for directed radio signals. The purpose of the communication connection between the Rover and

Transport Vehicle is to allow the astronauts the ability to command the Rovers in real time. The

HGA communication link is designed to handle this commanding on the Rovers.

Table 4.16 High Gain Antenna Link BudgetMars Rover to Transport Module

Frequency 21.2 GHz

Diameter Receiving 2 m




Link Margin 2 dB



Distance of Transmission 229,700 km

Data Rate 10 Mbps

Diameter Transmitting 0.32 m

Power 10 W

Mass 0.4 kg

4.6 Sample Return Vehicle

4.6.1 Design CharacteristicsWith the surface mission completed we launch the sample return vehicle seen in Fig. 4.20

for the first launch off the Martian surface. After a short journey we then dock with the

spacecraft in orbit. The sample return vehicle is designed for an optimal launch trajectory to

minimize weight and cost. The driving constraint for this design is the amount of fuel required

for the sample to reach the orbiting spacecraft. We obtain the final version of the rocket by

designing an efficient trajectory and structure. This method is described in Appendix G. The

major components of the design is seen in Table 4.17.44,45,46,47



Our propulsion system requires approximately 740 kilograms of propellant to accomplish

the 5200 meters per second of velocity change. This is the change required for a single-stage

accent to orbit in order to dock with the spacecraft. To achieve this V the rocket is power by

three custom engines using liquid oxygen and methane burning for 306 seconds. This fuel

combination is ideal for its propulsive characteristics and most importantly for ability to be

manufactured on the surface of Mars (section #). Once the burn is complete the payload cruise

section separates from the sample return rocket. This cruise stage includes space for the Martian

sample along with the navigation system and docking mechanism. Also included in the payload

cruise section is a reaction control system employed to stay on course during the orbital

trajectory for the rendezvous with the spacecraft. The payload cruise stage can be seen in Fig.

4.21, this section is separated from the main rocket and its protective shell. The docking

mechanism and payload bay, along with the expunged protective shell is visible in the Fig.. The

SRV has a maximum height of 3 meters and a maximum diameter of .96 meters and a gross lift

off weight of 950 kilograms.Error: Reference source not found,Error: Reference source not found,Error: Reference source

not found,Error: Reference source not found


Fig. 4.20 Sample Return VehicleBy Toleman


Table 4.17 Sample Return Vehicle Componets and PerformanceComponent Component ComponentOverall Height 3.02 [m] Take Off Mass 950 [kg] Ispvac 344 [s]Max Radius 0.48 [m] Dry Mass 200 [kg] Mix Ratio 2.99Tank Height 2.42 [m] Payload 10 [kg] Chamber P 300 [psi] Radius 0.48 [m] Fuel 740 [kg] Area Ratio 15Nozzle Length 0.30 [m] Engines 3 Thrust

Coefficient 1.707 Exit Radius 0.11 [m] Thrust/Weight 4.54 Throat Radius 0.03 [m] Total Thrust 16,400 [N] Characteristic

Velocity 6064Cargo Bay Height 0.10 [m] Burn Time 306 [s]Docking Probe Length

0.20 [m]0.20 [m]

Equivalent V 5.2 [km/s]

4.6.2 Structure - Daniel Nakaima

Nomenclature P = internal pressureR = cylinder radiust = cylinder thicknessE = Modulus of Elasticity

= stressSRV = Sample Return Vehicle

4.6.2.1 Introduction – Daniel NakaimaThe mass of the Lander with the aero-shell approximately totals 2,500 kg, and during Mars

entry the Lander is subjected to 4-5 Gs. A big part of the structure design involves the material

selection, which varies from the traditional aluminum to the high-tech composites. During the

design process we considered two materials, Aluminum and Honeycomb composites.

Aluminum exhibits low density and a high Elastic Modulus, but for our purpose Aluminum’s

density is not low enough. Honeycomb Composites have even lower densities ranging from 15-


Fig. 4.21 Payload Cruise StageBy Toleman & Maier


900 kg/m3, which makes it a great material to save mass.48 We chose Carbon Fabric honeycombs

for being stronger and lighter than some Aluminum honeycombs. Assuming the entire structure

of the body is composed of Carbon Fabric composite, we obtain a mass of 609.0 kg.

The Lander’s legs support and stabilize the entire Lander during landing and throughout

the entire mission. To obtain a stable Lander we design it so that its center of gravity is as close

as possible to the ground. To prevent any unexpected damage to the retro rockets, the nozzles are

not in direct contact with the surface, leaving a clearance of about 30 cm between the ground and

the bottom of the Lander. Table 4.18 shows the legs sizes and masses. Each leg can be simplified

into a system of three steel rigid rods with diameter of 5 cm and lengths of 0.95 and 1.0 m. We

chose steel as the primary material for the legs, because of its traditional use in aircraft landing

gear and high Modulus of Elasticity, which yields a small compact system. A structure, located

on each foot, crunches itself and acts as a shock absorber providing a softer landing.

4.6.2.2P

ressure Loads - Daniel Nakaima The SRV feeds the propellant to the retro rockets, and through the rest of the mission, it

stores the propellant produced in Mars. Internal pressures reach about 3 MPa when thanks are

completely filled. We calculate the thickness required to sustain such pressure from the hoop

stress equation:

4–4

Using internal pressure (P), the radius of the cylinder (R), and the ultimate stress of the

material as parameters, we calculate the thickness. Fig. 4.22 shows how the thickness varies with

increasing internal pressure for different materials. Once we have the thickness, we calculate the


Table 4.18 Lander’s legs sizes and massesLeg Number Length (m) Diameter (cm) Mass (kg)

A 4 0.95 5 14.7

B 8 1.0 5 15.4



volume of the structure. From material’s density and structural volume, we calculate the mass.

Notice that radius of cylinder also affects the thickness (Eq. 4–4). As part of the design we

decide which is the best radius and material to use. From Fig. 4.23 we see how the thickness

varies with radius.



Fig. 4.23 How thickness varies as the radius increases, when P = 3 Mpa


Fig. 4.22 How thickness of cylinder varies as pressure increases, when R = 0.25 m


4.6.2.3 Buckling Loads - Daniel Nakaima During launch the SRV is subjected to axial loads that can lead to buckling effects. To

prevent buckling we could use stringers or thicker panels. Since the radius of the SRV is small,

using thicker panels is more appropriate. Thickening the cylinder panels not only helps to

prevent buckling, also helps to sustain pressure loads. During launch the SRV will be

pressurized. To calculate the thickness required to prevent buckling we have the following

equations:

4–5

4–6

Where R is the cylinder radius, t is the thickness of the cylinder and E is the Modulus of

Elasticity. The critical stress and the Modulus depend on the material and the radius depends on

the design. Since the thickness required due to pressure loads is greater than due to buckling

loads, we calculate the SRV mass based on the thickness due to pressure.

4.6.2.4 Material Selection for SRV - Daniel Nakaima The mass also varies according to the material we choose, not only because thickness of

the cylinder changes, but also because densities differ from material to material. For that reason

selecting the appropriate material is important for the design of the SRV. From Fig. 4.24 we can

see that Aluminum 7975 and Titanium are the best materials to save mass. We chose Aluminum

instead of Titanium since price of Aluminum is cheaper, and because Aluminum is a traditional

material in the fabrication of rockets and propellant tanks.



Fig. 4.24 How mass varies as the pressure increases, when R = 0.25 m

4.6.3 Power - Ben PhillipsThe sample return vehicle (SRV), which will carry the Martian rock sample back to the

astronauts in the habitat module, requires power for approximately seven days. This power is

needed for the vehicle’s navigation while performing rendezvous maneuvers and eventual

docking. The relatively long mission lifetime of this return vehicle limits the number of options

available for power.

There are only two viable options, either a solar array and battery combination could be

used or a radio-isotope (RTG) power system. The difficulty that arises with the solar

array/battery combination is determining where the solar array would be placed on the return

vehicle. The use of a RTG power system is a good choice in this aspect because the entire

system can be placed inside the return vehicle. The effectiveness of a solar power array can be

seen in the appendix. The power that a solar array decreases at a rate of one over the distance to

the sun squared. This means that at Mars a solar array can only produce about half as much

power as it could at Earth.

The mass for a RTG power system needed for the SRV is relatively small. The radio-

isotope system would have a mass of approximately 15 kg49. This value is miniscule when

compared to the approximate 170 kg mass that a battery power system would require to provide



power to the SRV for its seven-day lifetime. A fuel cell that could provide the power would

have a mass of about 500 kg. These mass comparisons show the savings that using a RTG

power system gives in this case. A radio-isotope system would also require less volume, in this

case a cylinder with a diameter of 0.3 meters and a length of 0.4 meters.

Once the SRV has docked with the manned habitat module, the RTG power system will

not present a radiation problem. The chamber that in which the radioactive material is kept

protects the astronauts from radiation, as does the EVA spacesuit that they would wear. Once

the astronaut has retrieved the Martian rock sample, the SRV will be jettisoned along with the

radio-isotope power system.

In conclusion, the radio-isotope power system for the SRV offers several advantages two

of which are savings in volume and mass. The increase in reliability that a radio-isotope system

gives is also a major advantage. These savings help keep the SRV payload mass to a minimum

and thus keeping the sample return rocket to a minimum mass. The difficulties in placing solar

arrays on the SRV are also avoided.

4.6.4 Propellant Production - Matt Maier

Nomenclature CH4 = chemical formula for methaneH2O = chemical formula for waterO2 = chemical formula for oxygen H2 = chemical formula for hydrogenCO = chemical formula for carbon monoxideCO2 = chemical formula for carbon dioxide

For this mission it is necessary to employ the techniques of in-situ propellant production.

This is an important process that must be proven before landing a man on Mars. In this process

we produce oxygen and methane from a supply of hydrogen and the Martian atmosphere (95%

CO2). One of the benefits we gain from this process is reducing the mass needed to be taken from

Earth. For this mission it might seem like a small mass savings, but it is a technology that we

need to be prove for human missions where the propellant produced would be used not only to

bring the astronauts back to Earth but to provide ground based power and other various

resources.50,51

49 Larson, Wiley J., and Wertz, James R., Human Spaceflight Mission Analysis and Design, pg 643-665.



This process requires a supply of H2 and since it is not readily available on Mars we must

import it from Earth. The first step in producing our fuel is a process known as the Sabatier

reaction (Equation 4–7).

4–7

For this analysis we apply the water-gas shift reaction (Eq. 4-8) in conjunction with the

Sabtier reaction to produce (Eq. 4-9). The reaction converts carbon to methane and water by

reacting it with the imported hydrogen; there is also an excess amount of carbon monoxide

produced which we release into the Martian atmosphere. This equation is exothermic therefore in

the presence of a catalyst the reaction requires no net input of power to operate; therefore this is a

favorable method.Error: Reference source not found,Error: Reference source not found,52,53

4–8

4–9

From here the CH4 is stored cryogenically and the water is then reacted using electrolysis

(Eq. 4-10). The O2 is then cryogenically stored and the H2 is reacted again as in Equation 4–. We

repeat this process until the H2 supply is exhausted producing CH4. This is a very economical

process that ideally consumes all of the H2 in producing CH4 and also produces O2. This system

50 Zubrin, Robert. “ A comparison of Methods for the Mars Sample Return Mission”. AIAA-2941. 199651 Zurbrin, Robert. “The case for Mars”. New York. 1997.17 Whitley, R., AAE 450, School of Aeronautics and Astronautics, Purdue University.21 http://encyclopedia.thefreedictionary.com/Heat%20shield22 Charles D. Brown, Elements of Spacecraft Design, AIAA Education Series, Castle Rock, CO, 200225 Hexcel Composites26 Larson, Wiley J., and Wertz, James R., Space Mission Analysis and Design, Microcosm Press, Torrence, California, pg 407-427.27 Larson, Wiley J., and Wertz, James R., Human Spaceflight Mission Analysis and Design, pg 643-665.28 Lee, Darlene S., “Design and Verification of the MER Primary Payload Mars Exploration Rover PrimaryPayload Design and Verification”, Spacecraft & Launch Vehicle Dynamics Environment Workshop Program, NASA, Jet Propulsion Laboratory, California Institute of Technology, Pasadena, CA, 06/17/03http://www.aero.org/conferences/sc-lv/pdfs/lee_mer_03.pdf29 Rummel , John D., Race, Margaret S. DeVincenzi, Donald L., Schad, P. Jackson., Stabekis, Pericles D., Viso, Michel., and Acevedo, Sara E., NASA, Hanover, MD, October 2002, NASA/CP-2002-211842



has been proven to be very efficient, some designs operated at 99% efficiency.Error: Reference

source not found,Error: Reference source not found,Error: Reference source not found,Error: Reference source not found

4–10

With this combination of chemical processes we produce 4 kilograms of CH4 and 16

kilograms of O2 for every kilogram of H2. This is a 20:1 mass savings. This produces a mixture

ratio () of 4. Recalling from section 4.6.1 a of 2.99 is required for our rocket engines;

30 Mahaffy, Paul R. and 15 co-authors (2003), The Organic Contamination Science Steering Group, NASA, Jet Propulsion Laboratory, California Institute of Technology, Pasadena, CA, 12/02/0331 “Tyvek®”, DuPont, Wilmington, DEhttp://www.tyvek.com32 “Planova ® filters are designed for virus removal”, Asahi Kasei America Planova Division, Buffalo Grove, IL33 “Preliminary Report: A Study of Options For Future Exploration of Mars”, Mars Science Program Synthesis Group, NASA, Jet Propulsion Laboratory, California Institute of Technology, Pasadena, CA, 04/18/03 34 Cataldo, Robert L. “Power System Evolution: Mars Robotic Outposts to Human Exploration”, Power System, NASA Glenn Research Center, Cleveland, OH, AIAA Paper 2001-4592 35 Arvidson, Raymond “NASA Mars Exploration Program: Mars 2007 Smart Lander Mission”, Science Definition Team, NASA, Hanover, MD, 10/11/0136 Heninger, R., Sandler, M., Simmons, j. , Muirhead, B., Palluconi, F., and Whetsel, C., “Mars Program: Mars Science Laboratory Mission 2009, Landed Science Payload DRAFT Proposal Information Package”, NASA, Jet Propulsion Laboratory, California Institute of Technology, Pasadena, CA, 11/21/03, D-2720237 Maxon Precision Mortor“Maxon DC Motor”, Burlingame, CAhttp://www.maxonmotorusa.com/38 Neil ,Dan, “Kicking the Tires on Mars: An auto reviewer finds rover Spirit a bit pricey -- $410 million, with destination and delivery charges -- but enthuses it really shines off-road”, the Los Angeles Times, Los Angles, CA, 01/19/0439 Krishnan, Satish, and Voorhees, “The Use of Harmonic Drives on NASA’s Mars Exploration Rover”, NASA, Jet Propulsion Laboratory, California Institute of Technology, Pasadena, CA, Drive International Symposium 2001 , November 19-21, 200140 “Mars Exploration Rover Mission”, NASA, Jet Propulsion Laboratory, California Institute of Technology, Pasadena, CAhttp://www. marsrovers.jpl.nasa.gov/home/index.html41 Savage, Donald, Webster, Guy and Brand, David “Mars Exploration Rover Landings Press Kit January 2004” NASA, Hanover, MD, Kit January 200442 Larson, Wiley J., and Wertz, James R., Space Mission Analysis and Design, Microcosm Press, Torrence, California, pg 407-427.43 Larson, Wiley J., and Wertz, James R., Human Spaceflight Mission Analysis and Design, pg 643-665.48 Hexcel Composites4 Malik, Tarig, “NASA's Mars Rovers Perched on Crater Rims, Extended Mission Ahead,” 03/26/04, http://www.space.com/missionlaunches/rovers_update_040326.html7 Mars Today.com, “NASA Mars Picture of the Day: Athabasca Vallis Circles,” 06/27/03, http://www.marstoday.com/viewsr.html?pid=9625


Fig. 4.25 Prolellant Production UnitBy Toleman


therefore this is an acceptable method of acquiring the needed propellant.Error: Reference source

not found,Error: Reference source not found,Error: Reference source not found,Error: Reference source not found

We also note that the components applied for this process are quite simple (Fig. 4.25).

This is a very attractive quality of the production process. The Sabatier reactor is basically steel

pipes containing a catalyst bed, exercised to jump start the reaction, and a required filter, to keep

Martian dust out of our propellant. The reaction occurs spontaneously if the catalyst is nickel or

ruthenium (noble metals). Other necessary components include a cryogenic cooling system. This

is the main source of the energy requirements for the system. The electrolysis reaction is the

other process that requires a significant amount of energy. A summary of the systems

components is seen in Table 4.19.Error: Reference source not found,Error: Reference source not found,Error:

Reference source not found,Error: Reference source not found

9Delavault, Stephanie and Jacques Foilard. “Optical Navigation for the Mars Premier 2007 Orbiter Approach Phase,” Spaceflight Mechanics 2002; Proceedings of the AAS/AIAA Space Flight Mechanics Meeting. Vol. 1, San Antonio, TX, Jan. 27-30, 2002, San Diego, CA, Univelt, Incorporated, 2002, p. 513-52810Haw, Robert J. “Approach Navigation for a Titan Aerocapture Orbiter,” 39th AIAA Joint Propulsion Conference, Huntsville, AL, July 21-23, 2003. AIAA Paper 2003-480211 East, Robin A., “Atmospheric Re-entry”, Department of Aeronautics and Astronautics, University of Southampton.16 Whitley, R. and Manning, R., AAE 450, School of Aeronautics and Astronautics, Purdue University.24 http://www.goodfellow.com/csp/active/static/A/C_41.HTML3 Gulick, Dr. Virginia, “Prime Landing Sites for MER-A and MER-B”, 09/23/03, http://marsoweb.nas.nasa.gov/landingsites/mer2003/topsites/final/6 Burr, Devon, “Recent Eruption of Deep Groundwater into Athabasca Vallis”, 03/02, http://webgis.wr.usgs.gov/mer/March_2002_presentations/Burr/Burr-Landingsite3.pdf8 Melton, Melanie, “Homing in on Landing Sites for Mars 2003 Rovers,” The Planetary Society,10/26/01, http://www.planetary.org/html/news/articlearchive/headlines/2001/mars4sites.html15 Barua, D., AAE 450, School of Aeronautics and Astronautics, Purdue University.18 Soddit Matlab code written by Damon Landau and modified by Matthew Branson19 Sandia One-Dimensional Direct and Inverse Thermal Code (SODDIT), Sandia National Laboratories, Albuquerque, New Mexico, 199020 Professor Steven Schnider, Associate Professor Purdue University 23 Humble, Ronald, W. and Henry, G. N., and Larson, W. J., Space Propulsion Analysis and Design, McGraw-Hill, 1995, Chap. 5.44 Longuski, James M. “Optimization in Aerospace Engineering” Lecture Notes. West Lafayette. 200445 http://www.airliquide.com/46 Larson, Wiley. “Human Spaceflight: Mission Analysis and Design.” McGraw Hill47 Sutton, George P. “Rocket Propulsion Elements.” New York, NY 200152 Zubrin, Robert. Baker, David. Gwynne, Owen. “Mars Direct: A Simple, Robust and Cost Effective Architecture for the Space Exploration Initiative” AIAA-0328. 199153 Larson, Wiley. “Human Spaceflight: Mission Analysis and Design.” McGraw Hill


mailto:[email protected]


Table 4.19 Propellant Production SummaryMethane Oxygen ComponentMass Needed

185 [kg] Mass Needed

550 [kg] Required Hydrogen

47 [kg]

Production Rate

.616 [kg/day] Production Rate

2.46 [kg/day] Production Equipment

20 [kg]

Time 300 [days] Time 223 [days] Power Required

400 [kw]

4.6.5 Optimizing the Launch of the SRV – Allison Bahnsen

Nomenclature EOM = Equations of Motion FBD = Free body diagram = Flight Path AngleSRV = Sample Return VehicleTPBVP = Two-Point Boundary Value Problem

In order to simulate the launch of the SRV we first set up the basic FBD, which we can

see in Error: Reference source

not found. From this FBD we

can obtain the EOMs by

breaking the acceleration of the

rocket into x and y

components, where the flight

path angle is denoted as . We

can see these components in

the first four equations of Eq. 4–.

We know that we want the rocket to start from zero altitude and velocity and hit a certain

speed at a certain altitude. This speed and altitude correspond to periapsis of the Hohmann-like

transfer that travels out to the apoapsis of the Transport Vehicle orbit. We can see this illustrated

in Error: Reference source not found with the black curve representing the launch, the red ellipse


Fig. 4.34 FBD of SRV


being the Hohmann-like transfer, and the blue ellipse being the orbit of the Transport Vehicle.

Since we have EOMs and boundary conditions, this problem lends itself nicely to

functional optimization and solving a TPBVP. In this problem we want to minimize the launch

time to orbit, which in turn minimizes propellant. We set up the TPBVP and solve it via a

MATLAB code written by Professor Marc Williams through following a tutorial written by

Belinda Marchand.54 Marchand also wrote a second tutorial55 that details how to set up an

optimization of a launch off of the moon, and we will follow her example. The full details of

this analysis can be found in Appendix G.

Below is our well-defined TPBVP. Eq. 4-11 shows the differential equations, where

, are the traditional EOMs obtained from breaking the acceleration into components

and and are the co-states obtained from the Euler Lagrange Equations.

4–11

Error: Reference source not found shows the boundary conditions:


Fig. 4.35 Launch into Hohmann-like Transfer

222


Where rc is the desired altitude and vc is the desired speed at that altitude on the

Hohmann. Another parameter necessary for Professor Williams’ code is the specific thrust of

the rocket, which is set at 4.54 as provided by Matt Maier in Section 4.6.1.

The above well-defined TPBVP once inputted into Professor Williams’s code gives the

optimal trajectory, which we see in Error: Reference source not found. Error: Reference source

not found shows the optimal steering law, which tells us that after launching vertically for a few

seconds from the lander to avoid impacting any surroundings, the guidance rotates the rocket

down to about 30° and will continue to angle the thrust downward until actually becomes

negative. While this seems counter-intuitive, as we can see in Error: Reference source not found

the altitude continues to increase. This decrease in is used to push the velocity in the y

direction to zero, which is one of our final boundary conditions and a requirement to be at

periapsis in a Hohmann transfer.


Table 4.21 Boundary ConditionsInitial ConditionsFinal Conditionstoyf = rc = 100 kmxovxf = vc =

4.91 km/syovyf = 0vxovyo

223


Error: Reference source not found highlights some of the optimized SRV parameters and the launch data.


Fig. 4.36 Trajectory of Optimized SRV Launch

Fig. 4.37 Optimal Steering Law

224


4.6.6 Docking of SRV and Retrieval of Mars Sample – Allison Bahnsen

Nomenclature DART = Demonstration of Autonomous Rendezvous TechnologiesEVA = Extra Vehicular Activities ISS = International Space StationRCS = Reaction Control SystemRMS = Remote Manipulator SystemSRV = Sample Return Vehicle

Once the SRV launches off the surface of Mars following the optimized steering law

detailed in Section 4.6.5, it begins its seven-day journey back to the Transport Vehicle. When it

closes within a few hundred kilometers of the Transport Vehicle, it is well on course due to

continual course monitoring by the onboard guidance system and slight correctional inputs from

the RCS jets. It is at this time that the computer switches on the automated rendezvous software

using technology obtained from DART.56 The DART technology includes collision avoidance

software, and the system uses radar to determine the closing distances and relative speeds of the

two spacecraft, similar to the proven Russian Kurs system used on the ISS.57 This software

commands the RCS jets to fire until the relative velocity between the two spacecraft is

negligible. The software then switches over to the autonomous docking sequence which first

ensures that the SRV is lined up in the general area of the docking receptacles. The petals

revealing the docking probe, seen Error: Reference source not found, have been opened and

jettisoned along with the spent fuel tanks earlier in the mission. Finally, the docking sequence

uses the RCS jets to slowly insert the docking probe on the SRV into the cylindrical docking

receptacles on the Transport Vehicle, which we can see in Error: Reference source not found.


Table 4.22 Optimized Rocket ParametersParameterNumeric ValueAltitude [km]100Range [km]732X-Velocity [km/s]4.91Hohmann speed at 100 km [km/s]4.91Burn Time [s]307Thrust [N]13,000

225


The docking receptacles consist of three overlapping steel cables, each with one end attached to a

fixed outer collar, and the other end attached to the movable inner collar, as we see in Error:

Reference source not found.


Fig. 4.38 Petals Opening and SRV Docking Mechanism – created by Ben Toleman and Matt Maier

226


The lengths of the cables

are the diameter of the cylinder, as

we see on the left in Error:

Reference source not found. Prior

to entry of the probe, the inner

collar is rotated 60° causing the

cables to go slack and allowing for

the probe to enter. We can see this

configuration on the right of Error:

Reference source not found. Once

the probe enters, the tip hits a

push-button activator located in the

back of the receptacle. This


Fig. 4.39 Airlock (left) and Docking Receptacle with SRV mated (right)– created by David Goedtel and Ben Toleman

227

Fig. 4.40 Docking Receptacle, Collars

Inner CollarOuter Collar


activator releases a torsion spring between the two collars that then spins the inner collar back to

its original position, thus securing the SRV to the side of the Transport Vehicle. The concept of

using cables attached to fixed and moving collars to secure payloads has been proven; it is used

in the end effector of the Canadian RMS arm on the Space Shuttle to securely grapple and

transport large pieces of hardware.58

After confirmation that the two SRV’s have successfully attached to the side of the

Transport Vehicle, the astronauts ready themselves for the pre-scheduled EVA. Depending on

when the sample retrieval is placed in the EVA timeline, the astronauts make their way over to

where the SRV’s are secured, which we can see in the overall view of the Transport Vehicle in

Error: Reference source not found. Opening the same hatch that the rover used to place the rock

sample cartridges in the rocket, the astronauts carefully remove the cartridges and place them in

carrying bags. The SRV’s, having completed their mission, are left attached to the side of the

Transport Vehicle.

54 Marchand, Belinda, “ODEBVP – A Matlab Two-Point Boundary Value Problem Solver,” Purdue University AAE 508: Optimization in Aerospace Engineering Course notes, 02/23/9855 Marchand, Belinda, “ODEBVP: Minimum Time Launch into Orbit for the Flat-Moon Problem,” Purdue University AAE 508: Optimization in Aerospace Engineering Course notes, 03/04/9856 NASA Facts, “DART Demonstrator to Test Future Autonomous Rendezvous Technologies in Orbit,” Marshall Space Flight Center, 09/03, http://www1.msfc.nasa.gov/NEWSROOM/background/facts/dart.pdf57 Golightly, Glen, “Docking Zvezda: Tricky Space Ballet Takes Practice,” Space.com, 07/12/00, http://www.space.com/news/spacestation/zvezda_docking_000712.html


Fig. 4.41 Docking Receptacle, Cables

228


58 Thomas, Linda, “EVA Contingency Operations Training Workbook: CONT OPS 2102,” NASA Johnson Space Center, 03/95, pp 4-22 to 4-28


Fig. 4.42 Airlock and Docking Receptacle – created by David Goedtel

Airlock

229

Docking Receptacle

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