1
Sample Return Mission
1 Mission Overview
1 Introduction – Matt Maier
One 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 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.
2 Mission Timeline – Matt Maier
The 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.
2 Launch of Rovers
1 Release of Landers – Allison Bahnsen
After 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 Fig. 4.2. 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.
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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. 4.3, 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. 4.3 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 periapsis altitude of 100 km. We can see this trajectory in red in Fig. 4.3, and can transfer to it for a cost of 1.17 m/s. These results are obtained using the MATLAB code in Appendix G.
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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 Δv’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 Δv’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.
2 Cruise Stage – Andy Kacmar
The “cruise stage” is the configuration of the aeroshell for transport between the spacecraft and Mars. Fig. 4.4 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.
3 Atmospheric Entry / Touchdown
1 Landing Sites – Allison Bahnsen
Nomenclature
MER = Mars Exploration Rover
MGS = Mars Global Surveyor
TES = 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.[?] 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.[?]
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. [?] Fig. 4.6 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.[?]
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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 planet[?] 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.3 We can see the site along with its landing ellipse, with dimensions of 152 km x 16 km, in Fig. 4.7.[?]
In addition to the presence of hematite, this site is appealing because as we can see in the elevation map in Fig. 4.7, 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.[?] 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. 4.8[?]. 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.
2 Entry Trajectory
1 Mission Timeline – Ayu Abdullah
We 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 Fig. 4.9.
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2 Aerocapture – Ryan Whitley
The 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 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. 4.10 above 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.[?],[?]
3 Aeroshell Design – Ayu Abdullah
Nomenclature
BC = ballistic coefficient (kg/m2)
m = entry trajectory mass, Rover + Lander + Aeroshell (kg)
CD = drag coefficient
A = 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);[?] the lift to drag ratio (L/D) and the ballistic coefficient, BC. BC is defined in Eq. 4–1:
|BC = [pic] |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.[?]
▪ 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.[?]
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 data[?] from previous Mars missions.
To find the shape of the Aeroshell, we use following ratios in Table 4.3:
Table 4.3 Sizing ratios used
| |Rn / 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. 4.11 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:
|[pic] |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.
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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. 4.12 FEM analysis.[?] 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).
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 |
[pic]
We also plot the Aeroshell’s atmospheric entry trajectory using data from an integrated code[?]. 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 parameters
|Parameter |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.
1 Monte Carlo Analysis – Ayu Abdullah
We conduct a Monte Carlo Analysis[?] 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 Fig. 4.13 below shows the possible velocities at time of deployment.
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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 Matlab[?] and SODDIT[?], [?] codes discussed in 3.6.7. Fig. 4.14 shows the scheme developed after optimizing the mass and thermal properties of the materials. [pic]
Table 4.6 is the thicknesses used on the Mars Lander’s heat shield.
[pic]
We use an ablative material to greatly reduce the heat loads. The ablator protects the spacecraft by absorbing energy while chemically decomposing.[?] 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.[?] Specific analysis of the heat shield can be found in Appendix I.
4 Parachute Systems — Andy Kacmar
Nomenclature
DO = nominal canopy diameter
L(SL,R ) = length of lines
N(SL,R,G ) = number count
q = dynamic pressure
SO = canopy surface area
W(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 4.7. 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.
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 4.8 shows all the dimensions of the two parachute systems as well as system masses and packing volumes.
5 Retro Rockets
Nomenclature
c* = characteristic exhaust velocity
cF = thrust coefficient
Dstop = distance required to stop Lander
ΔV = total change in velocity of Lander
ε = expansion ratio
F = thrust per rocket in the direction the rocket is facing
gm = acceleration due to gravity on Mars
go = acceleration due to gravity on Earth
Isp = specific impulse
Lcham = combustion chamber length
Lnoz = nozzle length
mi = initial Lander mass at beginning of rocket firing
mf = final Lander mass at end of rocket firing (does not include use of lateral motion fuel)
Rcham = time index during navigation
Rexit = radius of rocket nozzle exit
Rthroat = radius of rocket throat
1 Purpose of the Retro Rockets – Frankie Hankins
We 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.
2 Configuration
We place 4 retro rockets evenly spaced around the Lander. The positions of the rockets are shown in Fig. 4.15. 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. 4.16 and Fig. 4.17 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.
3 Retro Rocket Specifications
We designed the retro rockets with the considerations shown in Table 4.9.
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 mfinal. 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 4.10. 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.
|[pic][pic] |4–3 |
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 analysis[?] gives us the data in Table 4.11 for each rocket.
A to scale view of one of the retro rockets is given in Fig. 4.18. 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.23 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.[?] 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.
4 Lander
1 Introduction - Dan Nakaima
As 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.
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 4.12 and Fig. 4.19 show the dimensions of the Lander. Fig. 4.20 shows how the Lander accommodates the SRV and the Rover.
Table 4.12 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 |
[pic]
Fig. 4.19 Side and top view of the Lander (exaggerated for explanatory reasons)
[pic]
Fig. 4.20 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.
3 Design Specifics
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.[?] 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 4.13shows 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 4.13 Lander’s legs sizes and masses
|Leg |Number |Length (m) |Diameter (cm) |Mass (kg) |
|A |4 |0.95 |5 |14.7 |
|B |8 |1.0 |5 |15.4 |
| | | |Total Mass (kg) |182.0 |
2 Rover Deployment – Andy Kacmar
The rover rests parallel to the Sample Return Vehicle, as can be seen in Fig. 4.20, while fixed within the Lander. The side panel parallel to the Rover, the upper most panel in Fig. 4.20, 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. 4.21 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.
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 of the SRV. Fig. 4.22 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 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 guideline[?], it can be estimated that radio-isotope system that produces 400 Watts for one year has a mass of 75 kg[?]. 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. 4.23.
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 overshadowed by other systems that are being brought to Mars and should not be considered at this stage.
5 Mars Lander Communication – Leigh Janes
The 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 4.14.
[pic]
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 4.15.
[pic]
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 Fig. 4.25. Once the Lander is on the Martian surface the antennas are deployed. Fig. 4.25 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.
5 Rover
[pic]
1 Mission Design - Masaaki Atsuta
The 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.
2 Design Specifics
1 Structure - Masaaki Atsuta
As 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).[?] 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.
2 Rover Design - Masaaki Atsuta
The 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).[?],[?],[?],[?]
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.[?] 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.[?] NASA has also considered Radioisotope Power Supplies as a power system option for their 2009 Mars Science Laboratory (Rover) Mission.[?],[?]
3 Components - Masaaki Atsuta
Our 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.[?]
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.[?]
4 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).35,36 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. 34
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.[?],[?],[?]
Thus, our rover can aggressively explore Mars and collect interesting samples for return to earth
5 Power - Ben Phillips
The 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 extrapolating[?] [?]. 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.
6 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 4.16 contains the specifications for the UHF antenna located on the rover. The placement of the UHF antenna on the Rover is shown in Fig. 4.27.
[pic]
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 4.17 contains the specifications of the high gain antenna. The high gain antenna placement is shown in Fig. 4.27. 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.
[pic]
6 Sample Return Vehicle
1 Design Characteristics
With the surface mission completed we launch the sample return vehicle seen in Fig. 4.28 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 4.18.[?],[?],[?],[?]
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 ΔV 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. 4.29, 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.44,45,46,47
[pic]
2 Structure - Daniel Nakaima
Nomenclature
P = internal pressure
R = cylinder radius
t = cylinder thickness
E = Modulus of Elasticity
[pic] = stress
SRV = Sample Return Vehicle
1 Introduction – Daniel 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.[?] 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 4.19 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.
2 Pressure 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:
|[pic] |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. 4.30 shows how the thickness varies with increasing internal pressure for different materials. Once we have the thickness, we calculate the 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. 4.31 we see how the thickness varies with radius.
[pic]
Fig. 4.31 How thickness varies as the radius increases, when P = 3 Mpa
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:
|[pic] |4–5 |
|[pic] |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 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. 4.32 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.
[pic]
Fig. 4.32 How mass varies as the pressure increases, when R = 0.25 m
3 Power - Ben Phillips
The 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 kg[?]. 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 Propellant Production - Matt Maier
Nomenclature
CH4 = chemical formula for methane
H2O = chemical formula for water
O2 = chemical formula for oxygen
H2 = chemical formula for hydrogen
CO = chemical formula for carbon monoxide
CO2 = 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.[?],[?]
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).
|[pic] |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.50,51,[?],[?]
|[pic] |4–8 |
|[pic] |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 has been proven to be very efficient, some designs operated at 99% efficiency.50,51,52,53
|[pic] |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 (Φ) of 4. Recalling from section 4.6.1 a Φ of 2.99 is required for our rocket engines; therefore this is an acceptable method of acquiring the needed propellant.50,51,52,53
We also note that the components applied for this process are quite simple (Fig. 4.33). 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 4.20.50,51,52,53
[pic]
6 Optimizing the Launch of the SRV – Allison Bahnsen
Nomenclature
EOM = Equations of Motion
FBD = Free body diagram
γ = Flight Path Angle
SRV = Sample Return Vehicle
TPBVP = 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 Fig. 4.34. 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 γ. 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 Fig. 4.35 with the black curve representing the launch, the red ellipse being the Hohmann-like transfer, and the blue ellipse being the orbit of the Transport Vehicle. [pic]
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.[?] Marchand also wrote a second tutorial[?] 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 [pic], are the traditional EOMs obtained from breaking the acceleration into components and [pic] and [pic] are the co-states obtained from the Euler Lagrange Equations.
|[pic] |4–11 |
Table 4.21 shows the boundary conditions:
[pic]
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 Fig. 4.36. Fig. 4.37 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 γ actually becomes negative. While this seems counter-intuitive, as we can see in Fig. 4.36 the altitude continues to increase. This decrease in γ 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.
[pic][pic]
Table 4.22 highlights some of the optimized SRV parameters and the launch data.
[pic]
7 Docking of SRV and Retrieval of Mars Sample – Allison Bahnsen
Nomenclature
DART = Demonstration of Autonomous Rendezvous Technologies
EVA = Extra Vehicular Activities
ISS = International Space Station
RCS = Reaction Control System
RMS = Remote Manipulator System
SRV = 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.[?] 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.[?] 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 Fig. 4.38, 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 Fig. 4.39. 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 Fig. 4.40.
[pic]
[pic]
The lengths of the cables are the diameter of the cylinder, as we see on the left in Fig. 4.41. 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 Fig. 4.41. Once the probe enters, the tip hits a push-button activator located in the back of the receptacle. This 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.[?]
[pic]
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 Fig. 4.42. 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.
[pic]
[1] Mass provided is only for the individual panel and not all four.
[i] Moomaw, Bruce, “Uncovering The Meridiani Formation,” Space Daily, 04/02/01,
[ii] Martel, Linda, “Grey Iron Oxide in Meridiani, Mars”, PSRD Discoveries, 03/13/03,
[iii] Gulick, Dr. Virginia, “Prime Landing Sites for MER-A and MER-B”, 09/23/03,
[iv] Malik, Tarig, “NASA's Mars Rovers Perched on Crater Rims, Extended Mission Ahead,” 03/26/04,
[v] Astrobiology Magazine Staffwriter, “Mars: Upstairs, Downstairs,” 01/29/04
[vi] Burr, Devon, “Recent Eruption of Deep Groundwater into Athabasca Vallis”, 03/02,
[vii] Mars , “NASA Mars Picture of the Day: Athabasca Vallis Circles,” 06/27/03,
[viii] Melton, Melanie, “Homing in on Landing Sites for Mars 2003 Rovers,” The Planetary Society,10/26/01,
[ix]Delavault, 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. 513-528
[x]Haw, Robert J. “Approach Navigation for a Titan Aerocapture Orbiter,” 39th AIAA Joint Propulsion Conference, Huntsville, AL, July 21-23, 2003. AIAA Paper 2003-4802
[xi] East, Robin A., “Atmospheric Re-entry”, Department of Aeronautics and Astronautics, University of Southampton.
[xii] Spencer, David A., Blanchard, Robert C., Thurman, Sam W., Braun, Robert D., Peng, Chia-Yen, Kallemeyn, Pieter H., “Mars Pathfinder Atmospheric Entry Reconstruction”.
[xiii] Sermeus, K., “Applications of Steady Perfect Gas CFD on Unstructured Grids”, Eurovia/Mission to Mars Symposium.
[xiv] Prabhu, Ramadas K, Lockheed Martin Engineering & Sciences Company, Hampton, Virginia.
[xv] Barua, D., AAE 450, School of Aeronautics and Astronautics, Purdue University.
[xvi] Whitley, R. and Manning, R., AAE 450, School of Aeronautics and Astronautics, Purdue University.
[xvii] Whitley, R., AAE 450, School of Aeronautics and Astronautics, Purdue University.
[xviii] Soddit Matlab code written by Damon Landau and modified by Matthew Branson
[xix] Sandia One-Dimensional Direct and Inverse Thermal Code (SODDIT), Sandia National Laboratories, Albuquerque, New Mexico, 1990
[xx] Professor Steven Schnider, Associate Professor Purdue University
[xxi]
[xxii] Charles D. Brown, Elements of Spacecraft Design, AIAA Education Series, Castle Rock, CO, 2002
[xxiii] Humble, Ronald, W. and Henry, G. N., and Larson, W. J., Space Propulsion Analysis and Design, McGraw-Hill, 1995, Chap. 5.
[xxiv]
[xxv] Hexcel Composites
[xxvi] Larson, Wiley J., and Wertz, James R., Space Mission Analysis and Design, Microcosm Press, Torrence, California, pg 407-427.
[xxvii] Larson, Wiley J., and Wertz, James R., Human Spaceflight Mission Analysis and Design, pg 643-665.
[xxviii] Lee, Darlene S., “Design and Verification of the MER Primary Payload Mars Exploration Rover Primary
Payload Design and Verification”, Spacecraft & Launch Vehicle Dynamics Environment Workshop Program, NASA, Jet Propulsion Laboratory, California Institute of Technology, Pasadena, CA, 06/17/03
[xxix] 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
[xxx] Mahaffy, Paul R. and 15 co-authors (2003), The Organic Contamination Science Steering Group, NASA, Jet Propulsion Laboratory, California Institute of Technology, Pasadena, CA, 12/02/03
[xxxi] “Tyvek®”, DuPont, Wilmington, DE
[xxxii] “Planova ® filters are designed for virus removal”, Asahi Kasei America Planova Division, Buffalo Grove, IL
[xxxiii] “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
[xxxiv] Cataldo, Robert L. “Power System Evolution: Mars Robotic Outposts to Human Exploration”, Power System, NASA Glenn Research Center, Cleveland, OH, AIAA Paper 2001-4592
[xxxv] Arvidson, Raymond “NASA Mars Exploration Program: Mars 2007 Smart Lander Mission”, Science Definition Team, NASA, Hanover, MD, 10/11/01
[xxxvi] 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-27202
[xxxvii] Maxon Precision Mortor“Maxon DC Motor”, Burlingame, CA
[xxxviii] 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, 01/19/04
[xxxix] 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, 2001
[xl] “Mars Exploration Rover Mission”, NASA, Jet Propulsion Laboratory, California Institute of Technology, Pasadena, CA
. marsrovers.jpl.home/index.html
[xli] Savage, Donald, Webster, Guy and Brand, David “Mars Exploration Rover Landings Press Kit January 2004” NASA, Hanover, MD, Kit January 2004
[xlii] Larson, Wiley J., and Wertz, James R., Space Mission Analysis and Design, Microcosm Press, Torrence, California, pg 407-427.
[xliii] Larson, Wiley J., and Wertz, James R., Human Spaceflight Mission Analysis and Design, pg 643-665.
[xliv] Longuski, James M. “Optimization in Aerospace Engineering” Lecture Notes. West Lafayette. 2004
[xlv]
[xlvi] Larson, Wiley. “Human Spaceflight: Mission Analysis and Design.” McGraw Hill
[xlvii] Sutton, George P. “Rocket Propulsion Elements.” New York, NY 2001
[xlviii] Hexcel Composites
[xlix] Larson, Wiley J., and Wertz, James R., Human Spaceflight Mission Analysis and Design, pg 643-665.
[l] Zubrin, Robert. “ A comparison of Methods for the Mars Sample Return Mission”. AIAA-2941. 1996
[li] Zurbrin, Robert. “The case for Mars”. New York. 1997.
[lii] Zubrin, Robert. Baker, David. Gwynne, Owen. “Mars Direct: A Simple, Robust and Cost Effective Architecture for the Space Exploration Initiative” AIAA-0328. 1991
[liii] Larson, Wiley. “Human Spaceflight: Mission Analysis and Design.” McGraw Hill
[liv] Marchand, Belinda, “ODEBVP – A Matlab Two-Point Boundary Value Problem Solver,” Purdue University AAE 508: Optimization in Aerospace Engineering Course notes, 02/23/98
[lv] Marchand, Belinda, “ODEBVP: Minimum Time Launch into Orbit for the Flat-Moon Problem,” Purdue University AAE 508: Optimization in Aerospace Engineering Course notes, 03/04/98
[lvi] NASA Facts, “DART Demonstrator to Test Future Autonomous Rendezvous Technologies in Orbit,” Marshall Space Flight Center, 09/03,
[lvii] Golightly, Glen, “Docking Zvezda: Tricky Space Ballet Takes Practice,” , 07/12/00,
[lviii] Thomas, Linda, “EVA Contingency Operations Training Workbook: CONT OPS 2102,” NASA Johnson Space Center, 03/95, pp 4-22 to 4-28
-----------------------
[pic]
Fig. 4.5 Two-D Drawing of Cruise Stage – Created by Rebecca Karnes
[pic]
Fig. 4.14 Layering scheme for Mars Lander heat shield
[pic]
Fig. 4.8 Landing Sites8
2nd Landing Site
[pic]
Fig. 4.7 Map of Second Landing Site6
[pic]
Fig. 4.6 Hematite Distribution Map3
[pic]
Fig. 4.16 Lander Configuration Top View
[pic]
Fig. 4.15 To Scale Lander Configuration Side View
Created by Ben Toleman
Table 4.10 NASA Thermochemistry Code Data
|Isp |cF |c* |
|364 s |1.915 |1865 m/s |
[pic]
Fig. 4.17 Lander Configuration Side View
[pic]
Fig. 4.21 Rover Deployment – Created by Ben Toleman
[pic]
Fig. 4.18 Retro Rocket Image
Created by Ben Toleman
Table 4.11 Rocket Design Data
|F |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 |
[pic]
Table 4.9 Initial Design Considerations
|”V |mfinal |tb |Pc |µ |
85 m/s1575 kg40 sΔVmfinaltbPcε
|85 m/s |1575 kg |40 s |3 MPa |30 |
[pic]
Fig. 4.4 Full Cruise Configuration
Created by Ben Toleman
Table 4.1 Material thickness for Mars Lander heat shield
Table 4.6 Material thickness for Mars Lander heat shield
|Material of Each Layer |Thickness (cm) |
|Graphite Ablator |0.1 |
|Carbon-Carbon Composite |0.1 |
|Glass Reinforced Polyimide Honeycomb |10 |
Table 4.8 Parachute Dimensions and Sizes
|Parameters |Design Values |
| |Drogue |Lander |
|SO [m2] |170 |385 |
|DO [m] |10.4 |16.7 |
|NSL |48 |48 |
|LSL [m] |16 |23 |
|NR |1 |5 |
|LR [m] |5 |3 |
|NG |48 |48 |
| | | |
|Volume [m3] |.021 |.039 |
|Total mass [kg] |17 |32 |
Table 4.7 Specific Weights of Parachute Materials
|Variable 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 |
Table 4.21 Boundary Conditions
|Initial Conditions |Final Conditions |
|to |yf = rc = 100 km |
|xo |vxf = vc = 4.91 km/s |
|yo |vyf = 0 |
|vxo | |
|vyo | |
[pic]
Fig. 4.35 Launch into Hohmann-like Transfer
[pic]
Fig. 4.37 Optimal Steering Law
Table 4.22 Optimized Rocket Parameters
|Parameter |Numeric Value |
|Altitude [km] |100 |
|Range [km] |732 |
|X-Velocity [km/s] |4.91 |
|Hohmann speed at 100 km [km/s] |4.91 |
|Burn Time [s] |307 |
|Thrust [N] |13,000 |
[pic]
Fig. 4.25 Lander UHF Antenna and High Gain Antenna in Stowed Position on Lander– By Ben Toleman
[pic]
Fig. 4.25 UHF Antenna and High Gain Antenna on Martian Surface – By Ben Toleman
Table 4.15 Lander to Transport Vehicle Link Budget
| |Lander to Transport Vehicle |
|Frequency |21.2 GHz |
|Diameter Receiving |2 m |
|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 |229,700 km |
|Data Rate |10 Mbps |
|Diameter Transmitting |0.32 m |
|Power |10 W |
|Mass |0.4 kg |
Table 4.17 High Gain Antenna Link Budget
| |Mars Rover to Transport Module |
|Frequency |21.2 GHz |
|Diameter Receiving |2 m |
|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 |229,700 km |
|Data Rate |10 Mbps |
|Diameter Transmitting |0.32 m |
|Power |10 W |
|Mass |0.4 kg |
[pic]
Fig. 4.27 UHF and HGA Antenna Placement on Rovers – Created by Ben Toleman
Table 4.20 Propellant Production Summary
|Methane | |Oxygen | |Component | |
|Mass Needed |185 [kg] |Mass Needed |550 [kg] |Required Hydrogen |47 [kg] |
|Production Rate |.616 [kg/day] |Production Rate |2.46 [kg/day] |Production |20 [kg] |
| | | | |Equipment | |
|Time |300 [days] |Time |223 [days] |Power Required |400 [kw] |
Table 4.16 Rover UHF Link Budget
| |Rover to Lander |
|Frequency |0.41 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 |0.22 mW |
|Mass |0.0374 kg |
Table 4.18 Sample Return Vehicle Componets and Performance
|Component | |Component | |Component | |
|Overall Height |3.02 [m] |Take Off Mass |950 [kg] |Ispvac |344 [s] |
|Max Radius |0.48 [m] | Dry Mass |200 [kg] |Mix Ratio |2.99 |
|Tank Height |2.42 [m] | Payload |10 [kg] |Chamber P |300 [psi] |
| Radius |0.48 [m] | Fuel |740 [kg] |Area Ratio |15 |
|Nozzle 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 |6064 |
|Cargo Bay Height |0.10 [m] |Burn Time |306 [s] | | |
|Docking Probe |0.20 [m] |Equivalent ΔV |5.2 [km/s] | | |
|Length |0.20 [m] | | | | |
[pic]
Fig. 4.29 Payload Cruise Stage
By Toleman & Maier
[pic]
Fig. 4.33 Prolellant Production Unit
By Toleman
[pic]
Fig. 4.28 Sample Return Vehicle
By Toleman
[pic]
Fig. 4.34 FBD of SRV
[pic]
Fig. 4.36 Trajectory of Optimized SRV Launch
[pic]
Fig. 4.22 SRV Launch Configuration –Created by Ben Toleman
Table 4.14 Lander UHF Antenna Link Budget
| |Lander 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 |
[pic]
Fig. 4.23 Lander RTG Power System - Created by Ben Toleman
[pic]
Fig. 4.3 Lander Trajectory
Protective Doors
[pic]
[pic]
Fig. 4.2 Side View of Landers in Transport and Protective Doors
Created by Ben Toleman and David Goedtel
[pic]
Fig. 4.1 Created By Ben Toleman
[pic]
Fig. 4.38 Petals Opening and SRV Docking Mechanism – created by Ben Toleman and Matt Maier
[pic]
Fig. 4.39 Airlock (left) and Docking Receptacle with SRV mated (right)– created by David Goedtel and Ben Toleman
[pic]
Fig. 4.40 Docking Receptacle, Collars
Inner Collar
Outer Collar
[pic]
Fig. 4.41 Docking Receptacle, Cables
Docking Receptacle
[pic]
Fig. 4.42 Airlock and Docking Receptacle – created by David Goedtel
Airlock
Velocity = .3208 km/s
[pic]
Fig. 4.10 Lander trajectory, altitude vs. velocity
[pic]
Fig. 4.26 CAD Images of our Rover - By Ben Toleman
[pic]
Fig. 4.9 Mission timeline
[pic]
Fig. 4.11 Aeroshell
[pic]
Fig. 4.12 FEM analysis
[pic]
Fig. 4.13 Failure analysis for drogue deployment
Table 4.19 Lander’s legs sizes and masses
|Leg |Number |Length (m) |Diameter (cm) |Mass (kg) |
|A |4 |0.95 |5 |14.7 |
|B |8 |1.0 |5 |15.4 |
| | | |Total Mass (kg) |182.0 |
[pic]
Fig. 4.30 How thickness of cylinder varies as pressure increases, when
R = 0.25 m
Table 4.2 Mission Timeline
|Time (sec) |Altitude (km) |Event |
|0.0 |100.0 |Aeroshell 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. |
Leg A
Bottom Panel
Leg A
Side Panel B
Leg B
Side Panel A
Leg B
[pic]
Top Panel
................
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