4.3 In-situ Propellant Production and Mars Launch Vehicle
Adam Butt
Nomenclature
g = Gravitational constant, m/s2
mo = Initial mass, tonnes
mf = Final mass, tonnes
min = Inert mass, tonnes
mpl = Payload mass, tonnes
mp = Propellant mass, tonnes
Isp = Specific impulse, s
Dv = Velocity change, m/s
l = Propellant mass fraction
¦ = Inert mass fraction
4.3.1 Introduction
One of the major design constraints for our mission is the necessity of producing all the propellant necessary to return to Earth, via the utilization of Martian resources. The reason for producing all return propellant on Mars, as opposed to bringing from Earth, becomes evident after a short analysis. We see the benefits through the ideal form of the rocket equation (1):
(4.3.1)
In equation (1) Dv is total change in velocity required, g is the gravitational constant (equal to 9.81m/s2), Isp is a measure of the overall performance of the propulsion system, and mf and mo are the final and initial masses of the vehicle, respectively. In reality we would have to take into account losses by both gravity and drag, but for our purposes neglecting these is a safe assumption.
Now lets look at a test case to compare the benefits of producing our propellant in-situ on Mars (Case A), as opposed to bringing it from Earth (Case B). Assuming that we are using the same system for both cases, the only variables that will differ between the two cases are mo, mf, and Dv. The reason for this is that g is a constant, and Isp depends on the system (which in this instance is the same for both Case A and B). The initial mass mo, is fully described in equation (4.3.2),
mo = min + mpl + mp (4.3.2)
where, min is the inert mass that consists of everything that doesn’t burn (tanks, engines, fairings, etc.) except the payload, mpl is the mass of the payload, and mp is the mass of the propellant. The final mass mf (4.3.3) is the mass after all the propellant is expended,
mf = min + mpl (4.3.3)
The Dv for Case A will have to include the total change in velocity necessary to go from the surface of Earth to Mars, and back to Earth. On the other hand, the Dv for Case B will only have to include the change in velocity to go from the surface of Earth to the surface of Mars. The reason for this difference is that the return trip for Case B will be accounted for by a separate system, incorporating the propellant produced on Mars (in-situ). We can now look at the two cases and make the comparison, which is seen in Table 1.
Table 1: Comparison of two test cases to show how much of a savingsin-situ propellant production can be.
Case A
/Case B
Dv to get to Mars / 13km/s / 13km/sDv to get back to Earth / 5.6km/s / 0
Total Dv from Earth / 18.6km/s / 13km/s
mpl / 65+185=250tonnes / 65*2=130tonnes
Approximate Price / $20 billion / $10.4 billion
Percent Difference / 48% Higher
Its now plain to see that a mission in which everything necessary to get to Mars and back, involving everything necessary to get back (Case A), would not be a good choice. The obvious reason for this is a price tag that is more than 50% higher than the in-situ mission (Case B). The main price difference between the two comes from the fact that the initial weight that launches off of Earth, basically determines your cost. Estimations made show that each tonne launched from Earth would cost approximately $80 billion (see section 1). Therefore all the extra mass that you would have to carry with you to Mars (Case A) would result in a much more costly mission. Not only would the higher launch weight cost more, but the developmental costs for such a mission would also be much higher. The reasoning behind this is that there is no current launch vehicle with the payload capacity for Case A. Any time you have to develop an entirely new launch vehicle you’re talking about tens of billions of dollars in research and development.
Now that we have finally decided that in-situ propellant production is the way to go, a vehicle must be designed to incorporate this propellant and to return the astronauts home. This vehicle, dubbed the Mars Launch Vehicle (MLV), must also be designed with the knowledge in mind that it will be transported to Mars within a predetermined volume (see section 4.2). The following sections detail the research and design of both the in-situ propellant production and the MLV.
4.3.1 In-situ Propellant Production
As we can see from the introduction, not only is the in-situ production of all return propellant one of our mission directives, but it has been shown to be a necessity for a realistic and cost effective manned mission to the Red Planet. Therefore the next step in the process is to analyze the Martian resources and determine which can be used to provide an adequate combination of fuel and oxidizer.
The most evident and abundant resource is the Martian atmosphere. Mars’s atmosphere, though almost 100 times thinner than Earth’s, is very valuable in terms of usable resources. It is comprised of more than 95% CO2, which by itself is no rocket propellant, but has much potential. Via the use of a chemical process called the Zirconia Cell Process, which will be discussed in a later chapter, the CO2 can be broken down into CO and O2. The combination of CO/LOX is a viable rocket propellant. The only requirement the we have to bring from Earth is a relatively small production plant and the necessary power. Two other possibilities exist for different propellants combinations, but would require us to bring a feedstock of liquid hydrogen (H2), as well as the production plant an power. They include a methane (CH4)/liquid oxygen (LOX) combination, and a methanol (CH3OH)/LOX combination. Each of the three combinations has intrinsic pros and cons, and will be discussed in the next section.
There are other resources on Mars that may have potential for further missions, but for our attempt at a low cost, high reliability first mission, atmospheric CO2 has the most potential. Other resources include the possibly abundant amount of ground water that could be used to create LOX as well as H2 feedstock for the other reactions.
The following sections will detail the researched areas of in-situ propellant production, as well as the final selections for propellants and production methods.
Propellant Selection
After all our preliminary research is conducted to assess the Martian resources for possible propellant combinations, we perform a detailed analysis of the pros and cons of each combination. The first thing that we have to do is to figure out what we will use to compare the various choices of propellant combinations. A number of various “pros” and “cons” are now compared in order to finalize the propellant selection. Table 4.3.2 below is a summary of this process:
Table 4.3.2 Summary of different factors used to decide which propellant to produceFactors / Fuels
Methane (CH4)* / Methanol (CH3OH)* / Carbon monoxide (CO)*
1. Isp [s] / 365 / 340 / 250
2. Optimal mixture ratio / 3.5 / 1.1 / 0.5
3. Density [kg/m3] / 423.5 / 794.4 / 1140.4
4. Approx. amount of power to produce one kilogram of fuel [W] / 0.18**
5. Cryogenic? / Yes / No / No
6. Reactant required for fuel production? / Yes, H2 / Yes, H2 / No
*Note: Each of the fuels compared above use LOX (cryogen) as the oxidizer, and use the same chamber and nozzle conditions to determine Isp
**Estimation based on referencewash
The factors in Table 4.3.2 are considered for the following reasons. First of all the Isp of each of the propellant combinations is compared. Arguably, Isp is the most important factor, as it is a measure of overall system performance. It isn’t the only consideration however due to other constraints, such as maximum volume and power available. The next factor considered is the optimal mixture ratio, which is defined in equation (4.3.4):
(4.3.4)
where r is the mixture ratio, and mox is the mass of the oxidizer, and mfuel is the mass of the fuel. The mixture ratio is important because it predicates how much LOX must be produced, relative to the various fuels. Therefore, a lower mixture ratio is desired, because for each kg of fuel produced, less LOX needs to be produced (as compared to a higher mixture ratio). The next factor considered is the density if the fuels. The reason this is important is because of the volume constraints. These volume constraints come from the maximum size of the fairing that is determined in section 4.2. Therefore a higher density is desirable, basically because this equates to smaller tanks. Because smaller tanks, which are brought from Earth, equate to less mass launched from Earth, this leads to a less costly mission. The fourth factor considered is the (approx.) amount of power required to produce 1 kg of each of the three fuels. Obviously for this case the lowest power requirement is desired because of the power constraints (see section 8.2). The fifth factor considered is whether or not the fuel is cryogenic. Cryogenic storage requires a relatively large amount of power, and the additional supporting systems add weight. The final factor considered is whether or not a reactant is required to produce the propellant (see the following section for specifics on production methods). The importance of this again is the additional weight that we must bring from Earth if a reactant is needed.
Based on the six factors summarized in Table 4.3.2, the methanol/LOX combination is chosen. The CO/LOX combination was thrown out because the Isp was simply too low, and the resulting amount of propellant that would have to be produced would be too large. Methanol is chosen over methane for a few reasons. The first reason is because methanol’s higher density becomes a very important factor in fitting the MLV within the size constraints (see section 4.2). The second reason is that the rovers can more easily use the methanol/LOX combination, than the methane/LOX combo. The third reason is that the optimal mixture ratio for the methanol/LOX combination is more than three times smaller than that of the methane/LOX combo. This results in much less LOX that needs to be produced, and thus less hydrogen feedstock that needs to be brought from Earth, further reducing our mission weight and cost. Another reason is that methanol does not require the cryogenic storage that methane does. The following section details the actual methods of propellant production researched.
Methods of Production
The Sabatier Process. This is the primary process for the production of methane. The following reaction, equation (4.3.5), details the formation of methane:
CO2+4H2 à CH4+2H2O (4.3.5)
In this reaction, the CO2 is sucked in from the atmosphere with an absorption pump, filtered, and then reacted with H2 feedstock under the presence of a catalyst to produce methane and water. Next the water is split into H2 and O2 via water electrolysis. Finally the O2 is stored, while the H2 is cycled back into the reaction until none remains, and the necessary amounts of CH4 and LOX are produced.3 The resulting products are combusted by a rocket according to equation (4.3.6):
1/2CH4 + O2 à CO2 + H2O (4.3.6)
Under the sample conditions of a chamber pressure = 2000psi, and expansion ratio of 50, an Isp value of 365s is achieved. The advantages of this production method are the resultant and comparatively high Isp, the exothermic reaction helps power the water elctrolysis, and the by-product is water. The main disadvantage of this reaction is the necessary hydrogen feedstock that we must bring from Earth. The other disadvantage of this reaction is that it dosen’t produce enough LOX to achieve the optimal mixture ratio of 3.5. Therefore, in order to compensate, another process must be used to produce the deficient amount of LOX.
Zirconia Cell Process. This process can serve a number of primary functions including: producing the additional LOX needed for the methane/LOX reaction, producing CO/LOX, and providing the first phase of the methanol/LOX production. A sample Zirconia cell reactor is seen in Figure 4.3.1.
Figure 4.3.1 Sample Zirconia cell reactor courtesy of JPL – Advanced Propulsion Concepts website.3
The Zirconia cell process begins similarly to the Sabatier process. Atmospheric CO2 is first sucked into the plant via a sorption pump. The CO2 then enters the main reaction chamber. As an electrical current is passed over the Zirconia cells the O-2 ions dissociate from the CO2 and pass through the cells to be collected and compressed into LOX, as seen in equation (4.3.7):
CO2 à CO + O2 (4.3.7)
The remaining products, including CO and unreacted CO2, are then separated and the CO is then stored. Alternately, if the CO is not desired (if this process was being used to produce the deficient LOX from the Sabatier process), it can be vented. The resulting products are combusted by a rocket according to equation (4.3.8):
CO + 1/2O2 à CO2 (4.3.8)
Under the same sample conditions as above of chamber pressure = 2000psi, and expansion ratio of 50, an Isp value of 250s is achieved. The greatest advantage of this system is that we can produce all of the propellant (CO and LOX) with in-situ resources. The only thing that we need to bring from Earth is a relatively light production plant, and supply the necessary power. The other advantage is that this process can be used to produce excess LOX, which can be used as an oxidizer or for human breathing - all this with Martian resources. There are however a few disadvantages. The first, and most important, is that the CO/LOX combination provides such low performance (Isp). The other disadvantages are that the Zirconia cells are fairly weak and brittle, so a number of replacements and safeguards must be utilized, and also CO isn’t very dense and would require large storage tanks.