This chapter addresses the third of four topics in the committee’s study charge, namely to assess the likelihood that spacecraft reaching the lunar surface will transfer volatiles to polar cold traps and, thereby, pose a potential threat of organic and biological contamination. The contamination impact of volatiles generated by spacecraft on permanently shadowed regions (PSRs) can be evaluated by comparing the amount and distribution of volatiles from both natural and spacecraft sources. This information can then be used to determine the planetary protection measures needed to avoid harmful contamination of these unique lunar regions. The sources of volatiles expected from upcoming lunar missions include products of spacecraft propulsion, out-gassing from spacecraft, and volatiles produced by life support systems. Contrary to the scarce knowledge of existing volatiles on PSRs (their type, amount, concentration, and horizontal and vertical distribution), the likely volatiles from spacecraft systems can be well quantified, and their distribution over the polar regions of interest can be bounded by modeling ahead of every mission. Further modeling, laboratory measurements, remote sensing, and in situ data of the behavior of other molecules, including organic molecules, would be useful to planetary protection.
This information can be compared to a variety of possible scenarios of what is expected to be found on the PSRs, based on existing data collected by remote sensing, the LCROSS mission, current understanding of volatiles in the early solar system, and from ground truth measurements as earlier investigations at the poles are performed. Mission planners and planetary protection officials can then determine whether the level of volatile contamination expected from these missions is of concern for scientific investigations of primordial chemistry and what approaches are appropriate for its mitigation.
DEPOSITION OF ROCKET EXHAUST
While every landing on the Moon will deposit rocket exhaust on the surface, the types and quantities of combustion products that will be deposited are of interest to planetary protection in order to assess whether this deposition compromises prebiotic and astrobiological studies.
In this section, the committee focuses on the distribution of possible prebiotic and organic molecules (see Table 3.1 and Table 3.2). Molecular mass, vapor pressure, and other specifics will affect how the molecules behave, including their desorption activation energy, and how similar their dynamics will be to H2O. Prem et al. (2020; see Box 3.1)1 provide important insight into the distribution and deposition of combustion products on the Moon. Further modeling of other molecules, including organic molecules, would be useful to planetary protection.
1 P. Prem, D.M. Hurley, D.B. Goldstein, and P.L. Varghese, 2020, The evolution of a spacecraft-generated lunar exosphere, Journal of Geophysical Research (Planets) 125: e06464, doi:10.1029/2020JE006464.
TABLE 3.1 Combustion Product Species for CH4-O2 Engines
|Species||Mass Fraction in Combustion Products|
|CO2||4.6 × 10-1|
|H2O||4.5 × 10-1|
|CO||8.5 × 10-2|
|H2||4.5 × 10-3|
|OH||2.0 × 10-5|
|H||2.1 × 10-6|
|O2||1.3 × 10-7|
|O||1.2 × 10-8|
|HCOOH||6.6 × 10-9|
|COOH||2.1 × 10-9|
|HCO||7.0 × 10-10|
|HCHO||3.7 × 10-10|
NOTE: All products with a mass fraction greater than 1 × 10-10 are listed. Modeling conditions: saturated cryogenic temperatures, mixture ratio = 3.6, chamber pressure = 20 MPa, expansion ratio = 100, equilibrium flow. Over 100 combustion products, including larger molecules, were considered in this calculation. For cryogenic H2-O2 landers, ~97 percent of the exhaust is H2O, ~3 percent is H2, and ~3 × 10-9 H and OH are present.
NATURE AND QUANTITY OF COMBUSTION PRODUCTS
Current interest in the Moon includes scientific exploration, human exploration, and resource utilization. However, the exact timeframe of each type of mission, and the frequency and mass of landers is uncertain. Baseline theoretical calculations can be used to provide an order of magnitude understanding of the potential impact of upcoming missions. The composition and mass of combustion products for different size and type of landers are discussed below.
Two major propellant types are planned for upcoming lunar missions and are likely to remain the propellants of choice: (1) hypergols for smaller missions and (2) cryogenic propellants for larger missions.2 Smaller landers are likely to use hypergols because of their relatively simpler engineering design. Larger robotic landers and human exploration missions are likely to use cryogenic H2-O2 and CH4-O2 combinations because of the significantly higher performance in spite of the more complicated engineering associated with these cryogenic propellants.
2 Hypergols are propellants that spontaneously ignite on contact, and cryogenic propellants are liquid only at very low temperatures, usually below 150 K.
TABLE 3.2 Combustion Products of an MMH-NTO Hypergolic Thruster
|Species||Mass Fraction in Combustion Products|
|N2||4.2 × 10-1|
|CO2||3.0 × 10-1|
|H2O||2.1 × 10-1|
|CO||3.9 × 10-2|
|H2||2.7 × 10-2|
|CH4||6.1 × 10-4|
|NH3||1.2 × 10-5|
|HCOOH||1.8 × 10-9|
|HCN||1.5 × 10-9|
|HNCO||8.3 × 10-10|
|HCHO||3.0 × 10-10|
NOTE: MMH is CH6N2, NTO is N2O4. Modeling conditions: 300 K propellants, mixture ratio = 1.6, chamber pressure = 7 MPa, expansion ratio = 100, equilibrium flow. Over 150 combustion products, including larger molecules, were considered in this calculation but were below the 10-10 mass fraction threshold.
Upcoming NASA missions to the Moon include the Commercial Lunar Payload Services (CLPS) landers and the Human Landing System (HLS) missions. Examples of CLPS missions include Mission One (Peregrine lander) and VIPER Rover (Griffin lander) from Astrobotic Technology, Inc., both of which use hypergolic propellants. Currently, three proposed HLS landers are being studied, which have been proposed using H2-O2 and CH4-O2 propellants. Approximately15 lunar lander missions from multiple countries are currently planned over the next 4 years. While details for many of these are not known, the majority are expected to use hypergols and have a landed mass of less than 2,000 kg, with many below 1,000 kg.
The results of complex simulations of rocket engine combustion dynamics depend on the details of the combustion process and propellant properties. The committee used the validated and publically available NASA Chemical Equilibrium with Applications (CEA) code3 with representative combustion parameters to provide insight into the composition of combustion products for the propellant combinations in upcoming missions. Table 3.1 and Table 3.2 show examples of combustion products from oxygen-methane and hypergolic (usually containing only C, N, H, O atoms) combustion, respectively.
The composition of combustion products noted here is in general agreement with literature4 for MMH-NTO, although the calculation in literature appears to show less complete combustion (likely partially due to the lower nozzle expansion ratio) and thus yields more CO and less CO2. More importantly for this discussion, however, literature noted that any higher order molecules are <0.01 mole fraction. Combustion products for other hypergolic propellant combinations are similar. As an example, for a MON-3 and M20 hypergolic thruster5 the main differences are that more H2O and N2, and less CH4, CO, CO2, and H2 are produced.
These calculations are snapshots of their respective engine designs, but they represent reasonable configurations and representative results to first order. Improving the performance of rocket engines explicitly requires the optimization of the combustion process toward complete combustion. Combustion characteristics will change as a function of propellant mixture ratio, nozzle expansion ratio, degree of propellant mixing, chamber pressure, and other parameters.
The combustion products in these calculations assume pure propellants. While propellants are held to high standards of purity, they do contain other molecules. This may produce additional combustion products, although they are expected to occur at very small mass fractions, commensurate with their small fraction in the initial propellants. The propellant composition used for any lunar mission, including the measurement of the isotopic composition, can be characterized at the launch site.
In summary, the efficient combustion process in rocket engines inherently produces mostly water, carbon dioxide, and also nitrogen for hypergols, with a few percent of carbon monoxide and hydrogen. More complex molecules such as HCOOH or larger are at mass fractions of order 10-9 or lower. Thus the majority of the contamination from landers will be molecules that are already at high concentrations in the PSRs on the Moon, and therefore will result in a very small contribution to their measured signal, likely within the measurement uncertainty.
4 K.H. Lee, 2017, Comparison study of exhaust plume impingement effects of small mono- and bipropellant thrusters using parallelized DSMC method, PloS ONE 12(6): e0179351, https://doi.org/10.1371/journal.pone.0179351.
5 MON-3 = 3% NO + 97% NTO (N2O4) by weight, M20 = 20% MMH (CH6N2) + 80% hydrazine (N2H4) by weight; mixture ratio of 1.2 was used.
The rocket equation is used to calculate the propellant mass required to land on the lunar surface as follows:
where Mlanded is the mass at landing [kg], Minitial is the vehicle mass at the start of the burn (landed plus propellant mass) [kg], ∆v is the velocity change that must be accomplished by the engines [m/s], Isp is the specific impulse of the engine [sec], and g0 is standard Earth gravity (9.81 m/s2 ; from the defined relationship between effective vehicle exhaust velocity and Isp).
The delta-V from low lunar orbit (LLO) to the surface is about 1.7 km/s; however, the majority of that burned propellant will escape to space. Following the calculations performed in literature,6 the committee assumes that approximately 25 percent of the exhaust products from LLO have the potential to interact with the surface.
The mass of propellant that needs to be burned to land a representative exploration lander (~1 metric ton) and a representative large human-class lander (~100 metric tons) are shown in Table 3.3. The propellant and Isp used are matched to the lander type.
The calculations in Table 3.3 are representative of the types of missions and masses of landers that will be sent to the Moon. Landers using hydrogen-oxygen propellants are similar to the methane-oxygen system, but they will use about 20 percent less propellant due to the higher engine Isp. These results are intended as order of magnitude values. Details, such as propulsion system characteristics, trajectory, mass of the lander, exhaust dynamics and reactions with the lunar surface (including high activation energies keeping the combustion products more localized), area of lunar cold traps (e.g., including recent discovery of “micro cold traps” extending the total area to ~40,000 km2;7 and others will alter the results.
TABLE 3.3 Representative Mission Types for Missions Using Hypergolic (Smaller/Exploration Landers) and Cryogenic Propellants (Larger/Human-Class Landers)
|Propellant type||Hypergols (e.g., MMH-NTO)||Cryogenic (e.g., CH4-O2)|
|Delta-V (25% of ∆v from LLO)||0.42 km/s||0.42 km/s|
|Isp (conservative)||300 s||360 s|
|Landed mass (illustrative)||1,000 kg||100,000 kg|
|Propellant mass required for landing (and may be deposited on lunar surface)||160 kg||12,800 kg|
|If all combustion products distributed over all PSRs (17,000 km2; Prem et al., 2020a)||9 × 10-9 kg/m2||8 × 10-7 kg/m2|
a P. Prem, D.M. Hurley, D.B. Goldstein, and P.L. Varghese, 2020, The evolution of a spacecraft-generated lunar exosphere, Journal of Geophysical Research (Planets) 125: e06464, doi:10.1029/2020JE006464.
NOTE: Propellant quantities are to land the spacecraft; return of the spacecraft back to lunar orbit will require additional propellant (same order of magnitude but lower masses).
6 Prem et al., 2020.
Net Deposition of Molecules on the Lunar Surface
The calculations above provide a first-order estimate of the total mass of contaminants that is available to be deposited on the lunar surface, and thus it is a maximum value. As noted in Prem et al (2020; see Box 3.1), photolysis plays an important role, which is molecule-specific, and for water—only up to 20 percent of the water is deposited in permanently shadowed areas.
Combining the total amount of propellant burned and the mass fractions of the combustion products, the maximum average surface densities if all exhaust products were deposited at the permanently shadowed regions for a 1 metric ton hypergolic lander are as follows:
- < 2 × 10-9 kg/m2 of H2O
- < 2 × 10-13 kg/m2 NH3
- < 2 × 10-17 kg/m2 HCOOH
- < 8 × 10-18 kg/m2 HNCO
For a human-class, 100 metric ton CH4-O2 lander, the maximum average surface densities if all exhaust products were deposited at the permanently shadowed regions are as follows:
- < 4 × 10-7 kg/m2 H2O
- < 2 × 10-15 kg/m2 COOH
- < 5 × 10-15 kg/m2 HCOOH
- < 3 × 10-16 kg/m2 HCHO
For a human-class, 100 metric ton H2-O2 lander, the maximum average surface densities if all exhaust products were deposited at the permanently shadowed regions are as follows:
- < 6 × 10-7 kg/m2 H2O
- < 3 × 10-8 kg/m2 H2
- < 3 × 10-15 kg/m2 H
- < 7 × 10-16 kg/m2 OH
As a comparison point, the weight percent of volatiles present in PSRs as observed in the LCROSS mission was 5.6 percent H2O, 0.32 percent NH3, and 0.16 percent CH3OH.8,9 In trying to directly compare these values with the contamination from combustion products, the collected sample characteristics must be considered. As an example, if an instrument were to collect a 0.5 cm thick layer of material with a density of 1500 kg/m3, the H2O from combustion products would be <8 x 10-6 weight percent, compared to the 5.6 weight percent that would be expected to already be present in the sample based on LCROSS data. Collecting a thicker layer or a higher density material would decrease the weight percent of the combustion products.
These engine exhaust estimates do not include processes that will affect the total quantities—for example, a significant fraction of the propellant will be destroyed by photolysis, molecule-specific sticking coefficients will affect the spatial distribution, and the effective cold trapping area is different for each molecule (the areal value used in Table 3.3 is for H2O).
8 A. Colaprete, P. Schultz, J. Heldmann, D. Wooden, M. Shirley, K. Ennico, B. Hermalyn, et al., 2010, Detection of water in the LCROSS ejecta plume, Science 330: 463, doi:10.1126/science.1186986.
9 A. Colaprete, et al. 2012, An overview of the Lunar Crater Observation and Sensing Satellite (LCROSS), Space Science Reviews 167: 3-22, doi:10.1007/s11214-012-9880-6.
The possibility of distinguishing between contamination and indigenous volatiles on the surface may also be further enhanced by using standard witness plates on all landers to record the contamination. The witness plates could provide a time history of deposited material. 10
In summary, combustion products will be distributed over much of the lunar surface. The spatial distribution will be unique for each species, will be affected by details of the landing profile and location, and will need to be predicted by modeling processes which are not always well constrained. If a uniform distribution over PSRs is assumed, the robotic exploration missions will result in water contribution of 10-10 or less mass fraction in the collected samples, which is likely to be within the measurement uncertainty. For more complex molecules, the expected mass fraction of 10-18 or less in collected samples is likely to be below the detection limit of instruments. The molecules produced by the combustion process are also readily found in the Earth environment in which the science instruments are built, and the sampling instruments themselves may be contaminated with these molecules, inherently contributing to uncertainty in the measurements.
These estimates point to what appears to be a small level of likely contamination, but further information is required to determine whether this constitutes “harmful contamination,” as required for planetary protection considerations. Upcoming planned missions are small landers with relatively low total combustion products. Once the larger human-class landers arrive on the Moon, they will burn more propellant to land but utilize cryogenic propellants that produce only traces of organic molecules in their exhaust. Exhaust products will deposit on the surface and are expected to remain there—molecular diffusion at the low temperatures of the surfaces of permanently shadowed regions11 and impact gardening (top few centimeters every 81,000 year12) operate at rates that are not relevant on the timescale of lunar exploration. Many of the experts consulted by the committee did suggest that much of the high-value astrobiologically oriented science is expected to be in the subsurface, where higher concentrations of prebiotic material are probably found, while vehicle exhaust only covers the surface
During this study, the committee did not have time to delve deeply into detailed descriptions of the types and sensitivities of scientific investigations that are planned at the PSRs. Understanding the scientific goals and capabilities of planned scientific investigations and, in particular, the mass fractions at which natural and lander-plume molecules of interest can be detected, would allow for a more objective assessment of whether propellant contamination could compromise prebiotic studies. Obtaining the first clean samples from PSRs will be even more valuable.
10 Honniball et al. (2020) describe SOFIA observations that demonstrate that much of the hydration observed by earlier missions could be in the form of H2O (rather than OH). The authors describe this H2O as likely produced during meteoroid impacts and trapped in glass, thus generally immobile. The total concentrations are small (100-400 ppm). As it pertains to the committee’s work, the paper suggests that water delivered via lander exhaust will be distinct from water produced at the surface of the Moon. Water from exhaust could also become trapped, but the time scales are probably quite different (exhaust plume migration versus meteoroid flux entrapment). The SOFIA results represent one observation and more coverage will be needed (spatially and temporally) at 6 microns to fully understand implications. The paper’s authors also cannot rule out chemisorbed water, which has already been detected/inferred from LAMP data (though in greater abundances perhaps because the longer wavelengths probe more deeply into the regolith). See C.I. Honniball, P.G. Lucey, S. Li, S. Shenoy, T.M. Orlando, C.A. Hibbitts, D.M. Hurley, and W.M. Farrell, 2020, Molecular water detected on the sunlit Moon by SOFIA, Nature Astronomy, doi:10.1038/s41550-020-01222-x.
11 N. Schorghofer and G.J. Taylor, 2007, Subsurface migration of H2O at lunar cold traps, Journal of Geophysical Research (Planets) 112: E02010. doi:10.1029/2006JE002779.
12 E.J. Speyerer, R.Z. Povilaitis, M.S. Robinson, P.C. Thomas, and R.V. Wagner, 2016, Quantifying crater production and regolith overturn on the Moon with temporal imaging, Nature 538: 215-218, doi:10.1038/nature19829.
Finding 5: There is a lack of, and need for, studies to characterize the chemical composition, transport, and level of contamination of volatiles that would be harmful to future investigations of prebiotic chemical evolution to be pursued at the PSRs. This information is necessary determine whether to establish a planetary protection requirement for missions to these areas of the Moon, such as a requirement for reporting the inventory of propellants, combustion products, and potential off-gassing volatiles from a spacecraft.
During landing, regolith will also be lofted and may cover PSRs along with volatiles coming from the rocket exhaust. While this regolith material may have implications for science investigations, it is not expected to have any biological impact on the lunar surface due to its lack of organic material. Thus, displacement or transport of regolith is not considered to be a planetary protection concern and was not considered in this study. Finally, for the lander modeling analysis discussed in this chapter, the committee considered only spacecraft landing outside the PSRs and not directly in these craters (discussed in Chapter 2), which may occur in future years for scientific missions or when ice extraction operations are conducted.