Trade-off Studies
3.1 Comparison of Propellants
Chapter 3
3.1.1 Propellants Selection Criteria
Specific Impulse
The specific impulse (Isp) is given by the exhaust velocity (ve), divided by the standard gravity (g0
= 9.81 m/s2), and is a common performance measure. Two values are given for each combination of propellants: vacuum and sea level. The sea level value will be used on the evaluation of the performance, since most of the flight will be atmospheric. In spite of the final performance also depending on the combustion efficiency of the engine, this is one of the most important parameters so a 25% weight was set to its relative importance. From Table 3.1 it is easy to compare the performance of each propellant combination and assign the scores.
Tanks Structural Mass Estimation
The tanks structural mass can be, in a first approximation, estimated using Mass Estimating Relations (MER’s) [37]. Knowing the operational requirements, presented in Table 1.1, the densities of the various fuels and oxidizers and their mixture ratios, i.e., the oxidizer to fuel ratios, the volume of the tanks can be computed using Equations (3.1) to (3.5).
Fuel Mass= Total Propellant Mass
Mixture Ratio+ 1 (3.1)
Oxidizer Mass=Total Propellant Mass−Fuel Mass (3.2) Fuel Volume= Fuel Mass
Fuel Density (3.3)
Oxidizer Volume= Oxidizer Mass
Oxidizer Density (3.4)
Total Volume=Fuel Volume+Oxidizer Volume (3.5)
According to the MERs, it is then possible to estimate the tanks structural mass.
For liquid hydrogen tanks:
Tank’s Mass= 9.09×Fuel Volume (3.6)
For other propellants:
Tank’s Mass= 12.16×Fuel/Oxidizer Volume (3.7)
The final mass will be the sum of both tank’s mass. Dividing by the dry mass, given in Table 1.1, the ratio between the tanks weight and the dry mass of each stage is obtained.
Along with the performance, this parameter is important since it represents the available mass for systems other than the structures. With this in mind, the weight attributed was again of 25%. Looking at Table 3.2, the scores regarding the tanks structural mass can directly be decided, taking into account that the pressure-fed system will result in heavier tanks, since they will require thicker walls.
Table 3.2: Tanks mass estimation results
First Stage Second Stage
Tanks structural mass estimation [kg]
Tanks structural mass to total dry mass ratio (%)
Tanks structural mass estimation [kg]
Tanks structural mass to total dry mass ratio (%)
LOX / LH2 172 28.7 26 25.9
LOX / Kerosene 71 11.9 11 10.7
LOX / Methane 91 15.2 14 13.7
N2O4 / Aerozine 50 60 10.1 9 9.1
N2O4 / MMH 61 10.1 9 9.1
N2O4 / UDMH 62 10.3 9 9.3
Complexity
This is used to evaluate the complexity of the resulting system using each propellant. The cryogenic propellants will have a more complex system than the others because they require a low operational temperature, which can result in some problems as was explained in Section 2.2.1. In addition, the injector of hypergolics can be very simple and no igniter is needed.
Regarding the combination LOX/Methane, although it is a cryogenic one, both fuel and oxidizer have similar handling temperatures (90 K for LOX and 111 K for Methane) so the infrastructure used for the liquid oxygen can be adapted for the methane. On the contrary, the combination of LOX/Kerosene has very different handling temperatures, so two different handling methods have to be used. Furthermore, extra care is needed in the case of structural tanks to ensure that the LOX is not heated due to the heat transfer from the kerosene tank, since the materials commonly used are heat conductors, thus increasing the system complexity.
Finally, the pump-fed system will also represent an increase in complexity when compared to the pressure-fed one. Simplicity is a desirable characteristic of any system and will result in a more robust and cheaper design, therefore a weight of 10% was given here.
Availability
This parameter is related to the ease of acquisition of the propellants, i.e., if they can be easily obtained or not. An easy access represents a saving in total costs as it decreases the manufacturing and transport fees, so a 10% weight was chosen.
It was found that it is possible to commercially acquire LOX and LH2 from multiple suppliers in Portugal, making these two components the most available. Besides, the place from where the vehicle is expected to launch already has a LOX supply. Methane is also widely available and can be extracted from natural gas by a simple process. In contrast, the kerosene commonly used in launch vehicles is the RP-1 type, which is highly refined and used only for the space industry, therefore having a low supply, which is expected to be even further decreased in the future. In the case of the other propellants, they are chemical compounds that need to be fabricated and may require safety procedures owing to their toxic nature. Because all of the three hypergolic fuels are derivatives of hydrazine they are all given the same score.
Safety
Because some propellants are toxic, corrosive or known carcinogens, they represent greater risk and cost of handling, as well as an increase in environmental pollution, should any accident occur. This was given a 10% percentage because the risk and difficulty of handling can be very harmful to the mission.
The distinction is then made between toxic and non-toxic, i.e., if the propellants are toxic they are assigned with the lowest score, otherwise they get the highest.
Cost
When it comes to the cost, as it is difficult to find recent prices publicly available, the prices NASA was paying between 1980 and 1990 were used for this comparison, with the exception of methane.
Because of environmental regulations, the prices presented in Table 3.1 for the hydrazine derivatives are higher than their production cost, due to their toxicity [38]. With the current environmental protection laws, it is expected that the fees for the toxic components will still be high, which makes these fuels and oxidizers more expensive.
In the case of methane, the price is about 1 Euro/kg [39], which is around 1.24 USD/kg. To be able to compare the various prices, the inflation rate was applied to update the 1990’s values. Then, using the price of each component and their mixture ratio, the final price per kg of each propellant combination was determined, and is presented in Table 3.3.
Table 3.3: Propellants cost Propellant Cost [USD/kg]
LOX / LH2 1.21
LOX / Kerosene 0.21
LOX / Methane 0.37
N2O4 / Aerozine 50 -
N2O4 / MMH 18.48
N2O4 / UDMH 21.40
Despite prices being updated with the inflation rate, a direct comparison between them would not be adequate because there are other parameters that can greatly influence the price variation of the propellants, such as demand or even the environmental regulations. However, with extra care, these are useful to establish a comparative Likert scale. For example, with the low supply of RP-1 kerosene and its increasing demand, along with an expected further reduction in its availability, an increase in its price is predicted. In fact, in 2002 it was reported that its price was three times higher than that of methane, according to a non-disclosed company study.
This parameter was also given a 5% weight since it does not present any structural constraints and this is usually a small fraction of the total launch cost, which includes the vehicle, use of infrastructures and manpower required.
Boiling Point and Insulation
The boiling point was included in this analysis because it represents the ease of vaporization of the propellants. The higher the boiling temperature, the higher the propellant temperature allowed before it vaporizes.
The boil-off rate is defined as the amount of mass that is lost due to vaporization, in a given time period. In the case of low boiling points, like the ones of cryogenics, the vaporization happens more easily so these propellants will be more prone to boil-off. In addition, in the case of a pump-fed system, the easier vaporization will favour the cavitation phenomenon.
Insulation is related both to the boiling point of the propellant, and to the difference between the working temperatures of the fuel and the oxidizer (in the case of a common bulkhead tank or connected structural tanks). The lower the temperature required to maintain the propellant liquid is, the more in- sulation will be needed. In the case of a common tank wall, should the tanks be kept at considerably different temperatures, there also arises a need for extra insulation between them, resulting in an in- creased structural mass.
Although the cryogenic propellants require more insulation due to their low working and boiling tem- peratures, hypergolic propellants will also need it to some extent, because the boiling point of the oxidiser N2O4 is also relatively low (21oC).
Like the performance, the boiling point is also directly evaluated by the data presented in Table 3.1.
Regarding the insulation needed in the tanks, the cryogenic propellants have a greater need than the others. As it was mentioned, the combination LOX/Methane does not require as much insulation as the LOX/LH2, since the temperatures of both fuel and oxidizer are similar, so little to no insulation is required between the tanks. The combination LOX/Kerosene, despite the kerosene not needing insulation, is still a semi-cryogenic mixture, so there is a temperature difference between the tanks. There is no significant distinction between the other combinations. These parameters combined will weigh 10%.
Reliability
To evaluate the reliability of the propellants, the failures of all the currently active launch vehicles were studied, to determine whether or not they were related to the propellants or the feed system (Ap- pendix A). However, the results, which are shown in Table 3.4, were inconclusive since some propellant combinations were used many more times than others, leading to an inadequate comparison.
Based on this data, the Technology Readiness Level (TRL) was chosen to replace the reliability in the trade-off study, so a proper comparison could be made. The propellant combinations of which no information is available were removed from this analysis so the remaining all have maximum TRL, with the exception of methane, which does not have flight heritage. In the case of liquid methane, its TRL is 6 for the pump-fed system, since it has already been successfully used in hot fire tests, but 5 for the pressure-fed system, because these tests were conducted in smaller engines.
This factor is important in the decision-making process, seeing as it influences the probability of failure. However, it was not considered a decisive one so a 5% weight was assigned.
Table 3.4: Reliability of each propellant combination
Feeding System Propellants Launches Reliability
Pressure-Fed System
LOX/LH2 - -
LOX/Kerosene - -
LOX/Methane - -
N2O4/Aerozine 50 153 1
N2O4/MMH 6 1
N2O4/UDMH 7 1
Pump-Fed System
LOX/LH2 396 0.992
LOX/Kerosene 2600 0.999
LOX/Methane - -
N2O4/Aerozine 50 - -
N2O4/MMH - -
N2O4/UDMH 826 0.996
3.1.2 Propellant Selection
The last step of this process is to create the selection matrix, where each row represents a propel- lant combination, divided according to the feeding system, and each column a parameter used for the evaluation. The scores (from 1 to 5) are given in each category to each pair and, in the end, the total scores are determined using a weighted average and a combination chosen. The trade-off study results are shown in Table 3.5.
Table 3.5: Propellants selection matrix
Specific impulse
Tank structural
mass Complexity Availability Toxicity CostBoiling point/
Insulation TRL Final Score
Parameter Weight 0.25 0.25 0.1 0.1 0.1 0.05 0.1 0.05
Pressure fed system
LOX/Methane 4 2 3 5 5 5 2 2 3.35
N2O4/Aerozine 50 3 4 5 2 1 2 4 5 3.30
N2O4/MMH 2 4 5 2 1 2 4 5 3.05
N2O4/UDMH 1 4 5 2 1 2 4 5 2.80
Pump fed system
LOX/LH2 5 1 1 5 5 4 1 5 3.15
LOX/Kerosene 3 4 3 3 5 3 3 5 3.55
LOX/Methane 4 3 2 5 5 5 2 3 3.55
N2O4/UDMH 1 5 4 2 1 2 4 5 2.95
Looking at the matrix, it can be concluded that the combinations LOX/Kerosene and LOX/Methane with a pump fed system are the best options. As Omnidea is interested in exploring their own and new solution, and since this study has not excluded the LOX/Methane combination, this was the combination selected.