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I'm starting this topic for a Mars Ascent Vehicle (MAV) using NTO (N2O4) oxidizer and MMH (Mono-Methyl Hydrazine) fuel vs NTO oxidizer and HAN (Hydroxyl Ammonium Nitrate) storable liquid propellants as alternatives to cryogenic liquids such as LOX/LCH4, LOX/LH2, and LOX/LCO. I think it's highly probable that the first exploration missions to Mars use legacy storable liquid propellants because the mission mass increase and technical challenges associated with ISPP to produce cryogenic liquids from scratch and then store them for many months will likely require some experimentation and refinement on Mars to ensure that everything works properly. To date, no serious funding has been allocated to ISPP. If ISPP tech is placed on "the critical path" to future Mars missions, then ISPP will be treated as another excuse as to why we still haven't sent humans to Mars.
Arguments For Storable Propellants
1. Respectable Isp in both pressure-fed (up to 330s) pump-fed (340s) engines
2. Ability to be stored almost indefinitely in chemically compatible tanks
3. Very high propellant bulk density, greater than liquid water, so propellant tank volume for Total Impulse is much smaller than cryogenic fuels
HAN Fuel Notes:
HAN fuel, when used as a substitute for MMH, eliminates most of the undesirable qualities of MMH.
1. HAN is mildly acidic and will corrode certain metals over extended storage duration, especially ordinary Carbon steel, but otherwise benign enough that humans can and do handle the fuel without extreme precautions taken to avoid skin, eye, or respiratory tract contact. A pair of goggles, Nitrile gloves, and ordinary lab clothing is considered to be sufficient protection. If any HAN fuel is spilled, cleanup is similar to diesel fuel cleanup. Handling of Hydrazine-based fuels requires a fully sealed hazmat suit with a positive pressure respirator.
2. HAN fuel provides a modest 7s Isp improvement over MMH when combusted with NTO. However, HAN's Density Impulse is a 50% improvement over MMH. HAN fuel tank volume for equivalent Total Impulse is half that of MMH.
3. HAN fuel can be frozen solid and then reheated in the tank prior to use without any risk of decomposition and thermal runaway, unlike MMH. This means no electrical power needs to be provided to keep the fuel warm during extended storage periods.
4. HAN fuel does not require double check valves typically used to prevent any accidental release of Hydrazine-based fuels, so fuel handling hardware is lighter and simpler. Research and testing, including testing in space by NASA and the US Air Force, has shown that HAN is broadly compatible with existing hardware commonly used to store and inject Hydrazine-based fuels, with the noted exceptions about metals that corrode easily.
5. It's conceivable that a staged combustion engine burning NTO/HAN, operating in a vacuum or near-vacuum and using a suitable nozzle expansion ratio, could achieve 360s of Isp. That would place overall engine performance in the same range as LOX/RP1, but with significantly greater Density Impulse. A lot of hay is made over the fantastic Isp of LCH4 and LH2, but not enough consideration is given to how greater propellant volume affects dry mass and cost. Using more materials to enclose a greater volume of propellant, even when that propellant is lower mass than more readily storable lower-Isp propellants for a given payload performance, especially for fully reusable spacecraft that must perform reentries, greatly affects monetary cost and developmental timeline. If a vehicle using cryogenic liquids must literally be twice as large as one using storable liquids, then the larger vehicle is likely to cost more and have a higher dry mass fraction as a result. There are no LH2 powered rockets that cost less to develop and operate than comparable RP1 powered rockets for that reason. Unfortunately, materials don't get any stronger as overall size increases. The cube-square law partially compensates for this, but any vehicle twice as large as an alternative design is almost certain to have greater dry mass or require stronger but more expensive lighter materials to compensate. Mindless fixation on Isp, as important as Isp is to rocketry, is driving up the monetary cost and developmental timelines for both rockets and landers.
Example Pump-Fed NTO/MMH Engine
Engine Designation: Aestus II / RS-72 (German-American collaboration)
Application: Ariane V orbital transfer stages
Propellants: NTO/MMH
Mixture Ratio: 1.9:1
Cycle: Gas Generator, Pump-Fed (XLR-132 pumps)
Thrust: 5,647kg-f
Isp: 340s
Chamber Pressure: 60bar
Nozzle Expansion Ratio: 84:1
Max Burn Time: 600s
Length: 2.29m
Engine Nozzle Exit Diameter: 1.31m
Weight: 138kg
Human Landing System Architecture Options Utilizing the XLR-132 Rocket Engine
NASA’s Artemis lunar exploration program aims to return humans to the moon over 50 years after the Apollo 17 crew lifted off from the surface. This time around, the goal is not only to return to the lunar surface, but to stay and explore, learning and preparing for later human exploration missions to Mars. In late 2019, Aerojet Rocketdyne (AR) performed an architecture trade study to evaluate several vehicle configurations and propulsion options that would satisfy the current mission requirements set forth by NASA. This study was completed under a NextSTEP-2 Appendix E: Human Landing System (HLS) Studies, Risk Reduction, Development, and Demonstration contract with NASA. The overarching observations from this study are still relevant as NASA and its HLS contractors are in the midst of designing the vehicles that will ultimately land on the lunar surface. The architecture trade study included an examination of a range of HLS configurations, launch vehicle options, concept of operations (CONOPS) options, main propulsion options, and other subsystem design options. Architectures were evaluated based on cost, schedule, reliability, extensibility, and performance attributes. Architecture configurations were scored and ranked through a utility analysis methodology. High scoring configurations that spanned the design space were further studied with alternate attribute weightings and Monte Carlo uncertainty analyses. Based on the assumptions made during the definition phase of the study, architectures using pump-fed storable propulsion received the highest overall scores across each set of weightings and uncertainty analysis. Architecture configurations that utilized the XLR-132 pump-fed storable engine scored particularly well indicating the XLR-132 to be an attractive main propulsion candidate for HLS elements. This paper provides a summary of the HLS architecture study performed by AR, and emphasizes 3-element and 2-element architecture configurations that take advantage of the XLR-132 pump-fed storable engine. Background details of the XLR-132 engine are also provided.
Arguments Against Storable Propellants
1. Cryogenic LOX/LCH4 (370-380s) and LOX/LH2 (450-470s) offers significantly higher Isp
2. Advocates of ISRU / ISPP would like to make some or all of the propellants from Martian-sourced CO2 and H2O
3. NTO and MMH are highly toxic to humans and require specialized and expensive propellant loading facilities here on Earth
LCH4/LH2 Notes:
Provided that NASA's new Zero-Boil Off (ZBO) cryogenic storage technologies are flight proven in space, cryogenic oxidizers and fuels provide clear performance advantages, which is why virtually all orbital launch vehicles use LOX and RP1, LCH4, and/or LH2. In addition to a ZBO demonstrator mission with LOX and LCH4 or LH2, further demonstrator missions must be performed to evaluate ISRU / ISPP mission hardware elements for durability and reliability in their intended operating environments. Lives cannot be staked upon the proper function of this new tech until exhaustive testing has been performed. Long term bulk storage of cryogenic liquids is a challenging mission requirement, even though there is ongoing refinement of ZBO and ISPP tech in the works. ZBO development is much further along than ISPP at the present time. ZBO and ISPP both radically increase the mission power requirements, so appropriate nuclear and solar power sources must be added to the mission hardware set. This tends to increase the mass of the mission elements, and it's undeniable that making rocket propellants from scratch, using indigenous natural materials, confers greater complexity to any mission relying upon ISPP. However, the propellant mass reduction leverage that ISPP brings to the mission is a compelling reason to pursue it anyway.
1. As fuels, the primary desirable characteristics of LCH4 and LH2 are no coking of engine internals so engine restart and reusability are greatly improved, autogenous pressurization makes it possible to remove heavy and potentially dangerous COPVs, and these fuels provide the highest Isp values amongst all practical chemical fuels.
2. Liquid cryogens are fantastic for removing waste heat from engine components prior to being fed into the combustion chamber. This property makes it possible to use high thermal conductivity metal alloys that are easy to work with, at least for modern automated 3D printing machines. Since not melting engine components is a high priority if you want to use them again, this feature is particularly appealing.
3. There's an enormous body of research, development, and implementation of LCH4 and LH2. We paid dearly for the performance benefits that cryogenic liquid fuels deliver, but all that effort ultimately made interplanetary missions feasible to begin with. The upper stages of the Saturn V used LH2 because that was the only way to deliver the demanded payload performance. If the LH2 fueled J-2 engines were not available, a significantly more powerful rocket would be required. The Soviet N-1 rocket was exclusively powered by RP1, required 5 stages, and could only throw 33t to TLI. In contrast, Saturn V only required 3 stages because both upper stages were LH2 fueled, so it could throw over 52t to TLI as a result. Vacuum LH2 fueled engines have Isp values of up to 470s. On a per mass basis, no other fuel can compete with LH2. Since Hydrogen is also the lightest element on the periodic table, LH2 is also the least dense fuel available by a very wide margin. This tends to make storing large amounts of Hydrogen difficult. However, LCH4 still has a high Isp of up to 380s and is far easier to store in a much smaller tank volume.
4. Since LCH4 and LH2 burn so cleanly, inspecting the engines for signs of damage after a landing burn, prior to a later liftoff and ascent to orbit, is a much simpler proposition. Since virtually all ISPP schemes make a little extra propellant, ground vehicles for surface exploration can also be powered by these propellants. Specialized in-space piston spark-ignited engines have already been developed that run exclusively on injected O2 and H2 or CH4, for example. One such engine has been developed as an APU for a LH2 fueled orbital transfer stage to reduce power generation and propellant pressurization mass over batteries and COPVs for longer duration LH2 storage between the time the rocket ascends to orbit and the time that the upper stage restarts to inject the payload into a higher orbit or escape trajectory. There are a variety of turbine-based APUs that very briefly run on Hydrazine fuels, but no such engine has been used to power ground vehicles as far as I'm aware, because long term engine durability is problematic. Every different kind of engine that exists has operated reliably on H2 or CH4, often for many years. Modern life as we know it would not be possible without fuels containing Hydrogen and Carbon.
5. There are no toxicity concerns with CH4 and H2. Even fairly extreme exposures to these fuels are not known to cause any lasting damage to the human body. Short of inhaling enough gaseous Methane or Hydrogen to be asphyxiated, it's not going to kill you. Virtually all storable oxidizers and fuels are capable of producing at least some long term damage from either massive exposure or chronic exposure.
Cryogenic liquids are staple propellants used by space faring nations. Virtually all high utilization rate orbital launch vehicles use LOX, for example, because it's pretty close to the best oxidizer available, very dense, very cheap to produce, and easy to store as cryogens go. LOX usually constitutes the majority of the propellant mass. While Chlorine and Fluorine are even better oxidizers, they're also incredibly toxic and not generally used outside of solid propellants that bind them to another chemical.
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HAN (Hydroxylammonium nitrate) is itself an oxidizer, and acts as such in many propellant formulations, according to the National Technical Reports Library. Specifically, HAN is used as an oxidizer in conjunction with fuel components, like alkylammonium nitrates, in an aqueous solution. It can also be used in hybrid rocket systems, where it is combined with other oxidizers like hydrogen peroxide, according toScienceDirect.
Here's a more detailed breakdown:
HAN as an Oxidizer:
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HAN is a highly acidic aqueous-based liquid oxidizer. It's oxygen-rich and acts as the oxidizing agent when combined with a fuel in a propellant.
Fuel/Oxidizer Combinations:
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HAN is often used in conjunction with fuel components, such as triethanolammonium nitrate (TEAN), which is a preferred choice for HAN-based liquid propellant guns, according to the Defense Technical Information Center.
Green Propellant Potential:
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HAN-based propellants are considered "green" or environmentally friendly options due to their potential for reduced toxicity compared to conventional hypergolic propellants.
HAN with Other Oxidizers:
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In hybrid systems, HAN can be combined with other oxidizers, like hydrogen peroxide (H2O2), to create high-performance hypergolic fuels, according to ScienceDirect.Successful development of HAN based green propellant
Hydroxyl Ammonium Nitrate (HAN) was first used as rocket propellant in the 1960s in the USA and was employed as a solid oxidizer i...ScienceDirect.com
Janus-type hypergolic fuels for hybrid systems using hydrogen ...
Feb 15, 2023 — Highlights * • Novel hypergols with H2O2-HAN oxidizer formulations were prepared and evaluated. * Top performing hype...ScienceDirect.com
Hypergolic Characterization of HAN-Based Ionic Liquids for ...
Propellants based on hydroxylammonium nitrate (HAN), and high-test hydrogen peroxide (HTP) can be less toxic compared to these con...PSU-ETD
The technology efforts have focused on a family of monopropellant formulations consisting of an aqueous solution of hydroxylammonium nitrate (HAN), which serves as the oxidizer, and a fuel component. HAN usually composes the majority of the formulation and dominates the characteristics of the monopropellant.
HAN-Based Monopropellant Technology Development
NASA (.gov)
https://ntrs.nasa.gov › api › citations › downloads
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FeedbackHydroxylammonium nitrate
Wikipedia
https://en.wikipedia.org › wiki › Hydroxylammonium_...
As a solid propellant oxidizer, it is typically bonded with glycidyl azide polymer (GAP), hydroxyl-terminated polybutadiene (HTPB), or carboxy-terminated ...Successful development of HAN based green propellant
ScienceDirect.com
https://www.sciencedirect.com › science › article › pii
by T Katsumi · 2021 · Cited by 41 — Green oxidizers (chlorine-free) are usually nitrates such as HAN, ADN (Ammonium Dinitramide), and HNF (Hydrazinium Nitroformate); thus, new green propellants ...
HAN-Based Monopropellant Propulsion System With ...NASA (.gov)
https://ntrs.nasa.gov › api › citations › downloads
by RS Jankovsky · 1997 · Cited by 15 — HAN is oxygen rich, and is commonly referred to as the oxidizer, the other salt is fuel rich and is referred to as the fuel. Variations on these formulations.12 pages
HAN (Hydroxylammonium Nitrate) Based Liquid Gun ... - DTIC
Defense Technical Information Center (.mil)
https://apps.dtic.mil › sti › pdf › ADA195246Since LGP 1845 is currently considered a Class B explosive, while its components HAN and TEAN are classified only as oxidizers, it has been suggested that the ...
62 pages
Development of a novel hydroxyl ammonium nitrate based ...Harvard University
http://ui.adsabs.harvard.edu › abs › abstract
by JH Fontaine · 2006 · Cited by 1 — Hydroxyl Ammonium Nitrate is a highly acidic aqueous based liquid oxidizer. Therefore, in order to achieve efficient combustion of a propellant using this ...
Hydroxylammonium nitrate (HAN)-based green propellant ...ScienceDirect.com
https://www.sciencedirect.com › article › abs › pii
by R Amrousse · 2017 · Cited by 165 — It is the main ingredient in the propellants for regenerative guns. HAN is also considered as an oxidizer for hybrid rockets, because hybrid rockets are ...
HAN/HN-Based Monopropellant Thrusters株式会社IHI
https://www.ihi.co.jp › afieldfile › 2023/06/18
by B FUKUCHI · Cited by 43 — HAN/HN ( Oxidizer ). Detonation area. Selected composition. H2O. 400. 450. 300. 350. 250. Fig. 1 HAN/HN-based monopropellant composition with ...
7 pages
Is HAN-water propellant "green" enough to ever be used in ...Space Exploration Stack Exchange
https://space.stackexchange.com › questions › is-han-w...
Oct 29, 2016 — According to Wikipedia NH3OHNO3 is a fuel/oxidizer blend, also known as AF-M315E. It would likely be used in a spacecraft as an aqueous solution ...
3 answersTop answer:
It probably couldn't be used on a cubesat, but not primarily for environmental reasons. ...
Hypergolic Characterization of HAN-Based Ionic Liquids ...Penn State University
https://etda.libraries.psu.edu › catalog
by D Over · 2024 — From all powders tested, lithium aluminum hydride, sodium aluminum hydride, sodium amide, and sodium hydride showed best reactivity with HAN based oxidizer, ...Hydroxylammonium nitrate synthesis
Hydroxylamine nitrate formula
Hydroxylamine nitrate empirical formula
Hydroxylammonium sulfate
Hydroxylammonium nitrate SDS
LMP-103S propellant
ASCENT propellant
Green propellant
The comments flowing from the links may themselves contain useful insights, but the quality of the volunteered input varies.
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An analysis of storable chemical propulsion stages for lunar and Mars missions from DLR:
High-Thrust in-Space Liquid Propulsion Stage: Storable Propellants
It makes use of the ATV and Aestus-II (Aerojet-Rocketdyne RS-72, which is based upon the XLR-132) since it's of European origin.
Edit:
Abstract
In the frame of a project funded by ESA, a consortium led by Avio in cooperation with Snecma, Cira, and DLR is performing the preliminary design of a High-Thrust in-Space Liquid Propulsion Stage for two different types of manned missions beyond Earth orbit. For these missions, one or two 100 ton stages are to be used to propel a manned vehicle. Three different propellant combinations; LOx/LH2, LOx/CH4 and MON-3/MMH are being compared.
The preliminary design of the storable variant (MON-3/MMH) has been performed by DLR. The Aestus II engine with a large nozzle expansion ratio has been chosen as baseline. A first iteration has demonstrated, that it indeed provides the best performance for the storable propellant combination, when considering all engines available today or which may be available in a short- to medium term. The RD-861 K engine has been proposed as alternative to reduce the development duration of the high-thrust stage. Structure analyses and optimisations have converged towards a common bulkhead architecture with a Whipple shield, similar to the one used on the ATV, to protect the main propellant tanks against perforations caused by meteoroids and space debris. The propulsion system has been built around six Aestus II engines equipped with TVC and placed on a circular engine thrust frame. The RCS, the thermal system, and the power system have also been included in the preliminary design, and they have been sized for the most demanding mission. The performance of the high-thrust stage, resulting from the preliminary design, has been assessed for both missions taken into consideration.
Last edited by kbd512 (2025-05-20 13:45:39)
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