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I am taking the risk of creating this topic because GW Johnson asked about it.
The subject was discovered by Void in mid-2026.
At this point all I know about the propellant is what Void has posted in one of his many topics. This topic is available for members to contribute details about the propellant, it's ISP, it's advantages and disadvantages, and anything else that would be important for a rocket designer to know.
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This post is reserved for an index to posts that may be contributed by NewMars members.
Index:
Post #3: Google Search report on silicon oxygen propellant
Post #4: Google Search report on traditional self-igniting propellants
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This preliminary report comes from Google Search... it provides a range of ISP for SILOX...
AI Overview
Silicone-based (silox/silicone elastomer) binders in composite solid propellants deliverspecific impulses around 238–247 s,
high thermal stability, flash points over 300 °C, and fast burning rates near 60 mm/s at 10 MPa, making them durable alternatives to traditional polyurethane or polybutadiene systems.
Key Performance Characteristics
Specific Impulse (
): Reaches ~2340 Ns/kg for plain AP/PDMS (polydimethylsiloxane) and up to 2430 Ns/kg with aluminum additive formulations.
Burning Rate: Exceeds 60 mm/s at 10 MPa with pressure exponents ranging stably between 0.5 and 0.6.
Safety and Stability: Exhibits zero detonation sensitivity at 50 mm diameters alongside high chemical inertness and elevated flash points (>300 °C).Material Advantages & Limitations
Thermal Protection: Silicone elastomers naturally feature minimal ablation rates (0.33–0.52 mm/s), offering superior internal heat shielding.
Processing: High solid loadings up to 84% are achievable using multi-modal ammonium perchlorate (AP) fractions, though viscosity management remains critical during casting.Explore property testing of silicone binder formulations in the ResearchGate Study on Silicone Binders.
Would you like to compare these silicone metrics against HTPB-based propellants or focus on a specific silane fuel application?
Fast Burning Rocket Propellants Based on Silicone Binders – New Aspects of an Old System
Here are some properties of AP/PDMS and AP/Al/PDMS propellants: * **Specific impulses** 2340 N s/kg (238 s) for AP/PDMS and 2430 N...
ResearchGateRocket Propellant Characteristics of Silanes/O2
ResearchGate
https://www.researchgate.net › publication › 24543561...
We present a theoretical analysis of the applicability of various silicon hydrides (silanes) as rocket propellants.Note trademark for a product:
https://www.momentive.com › en-us › brands › Silox
SiloxTM can help optimize the performance of polyethylene pipes and considered in both one and two-step extrusion processes.Fast Burning Rocket Propellants Based on Silicone Binders
ResearchGate
https://www.researchgate.net › publication › 24385839...
Basic tests reveal high thermal and chemical stability with flash points over 300 °C and zero detonation sensitivity at 50 mm diameter.
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For Comparison to Post #3, here is a summary of performance of traditional self-igniting propellants:
Per Google Search:
AI Overview
Traditional self-igniting (hypergolic) propellants typically achieve a specific impulse () ranging from 270 to 325 seconds in a vacuum, and roughly 210 to 290 seconds at sea level.
Specific Impulse of Common Hypergolic Bipropellants
Traditional self-igniting combinations rely on a storable liquid oxidizer—most commonly dinitrogen tetroxide (NTO)—paired with hydrazine-based fuels. Their exact efficiency depends on the mixture ratio and the expansion ratio of the rocket nozzle:
Dinitrogen Tetroxide (N₂O₄) + Aerozine-50 (50-50 mix of hydrazine and UDMH):
Vacuum
: ~315–320 seconds
Notable use: Apollo Service Propulsion System (SPS) and Lunar Module Ascent/Descent engines, Titan II ICBMs.Dinitrogen Tetroxide (N₂O₄) + Monomethylhydrazine (MMH):
Vacuum
: ~305–312 seconds
Notable use: Space Shuttle Orbital Maneuvering System (OMS) and Reaction Control System (RCS) thrusters.Dinitrogen Tetroxide (N₂O₄) + Unsymmetrical Dimethylhydrazine (UDMH):
Vacuum
: ~295–305 seconds
Notable use: Proton, Ariane, and Long March upper stages.Nitric Acid (HNO₃ / IRFNA) + Aniline or Hydrazine (Older historical mix):
Vacuum
: ~270–285 seconds
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Because GW Johnson expressed interest in this propellant system, I decided to take it up with Gemini.
We are pursuing this because Void brought this research to our attention in one of his topics.
At first glance, this propellant system would appear to be highly competitive on the Moon or any other airless body far from Earth.
it would appear to be less attractive as a propellant for a Mars lander. However, it might prove highly attractive for the space tug business that GW Johnson has been describing for the past several years.
Here is a detailed report prepared by Gemini as supervised by tahanson43206:
Executive Summary & System Overview
A Silicon–Oxygen (Si–O₂) propellant system—typically structured as a porous matrix or micro-engineered solid silicon grain through which oxygen gas or liquid oxygen (LOX) is fed—operates as a hybrid or micro-propulsion concept. By structuring the silicon at a controlled pore size and density, combustion transitions from an explosive nanoscale detonation into a steady, controlled burn suitable for rocket thrust.
When evaluated against standard storable hypergolic propellants—such as Monomethylhydrazine and Nitrogen Tetroxide (MMH/NTO)—the Si–O₂ system introduces a unique set of trade-offs between chemical performance, environmental safety, and operational complexity.
Technical Parameter Comparison
Typical Vacuum Isp:
• Structured Si + LOX: 210 – 250 s (Limited by high exhaust molecular weight)
• Traditional MMH/NTO: 310 – 325 sFuel Density:
• Structured Si + LOX: High (~2.33 g/cm³ bulk solid)
• Traditional MMH/NTO: Moderate (~0.87 – 1.0 g/cm³)Oxidizer Storage:
• Structured Si + LOX: Cryogenic (90 K / -183 °C)
• Traditional MMH/NTO: Ambient Storable (-11 °C to +50 °C)Toxicity / Safety:
• Structured Si + LOX: Inert / Non-toxic (Safe ambient handling)
• Traditional MMH/NTO: Extremely Toxic / Carcinogenic (SCAPE suits required)Primary Exhaust Product:
• Structured Si + LOX: Silicon Dioxide (SiO₂ solid/liquid particulates)
• Traditional MMH/NTO: Gaseous H₂O, CO₂, N₂ISRU Potential:
• Structured Si + LOX: Exceptional (Abundant in Lunar/Martian regolith)
• Traditional MMH/NTO: Poor (Requires rare volatile hydrogen/nitrogen)Detailed Performance & Systems Analysis
1. Specific Impulse (Isp) and Combustion Physics
Specific impulse is governed by the effective exhaust velocity, which scales inversely with the mean molecular weight of the exhaust gas (M_bar):
Isp ∝ √(Tc / M_bar)
The Challenge with Silicon: The primary combustion reaction, Si + O₂ → SiO₂, produces silicon dioxide. SiO₂ has a high molecular mass (60.08 g/mol) compared to standard rocket exhaust species (H₂O = 18.02 g/mol, CO₂ = 44.01 g/mol).
Two-Phase Flow Losses: At typical rocket chamber temperatures (2,500 – 3,000 K), SiO₂ condenses into liquid micro-droplets or solid particulates during expansion through the nozzle. Liquid and solid particles do not expand gas-dynamically to perform thermodynamic work, leading to severe two-phase flow momentum losses.
Net Result: Si–O₂ delivers a vacuum Isp of approximately 210 – 250 seconds, noticeably lower than the 310 – 325 seconds delivered by standard MMH/NTO thrusters.
2. Storage & Cryogenic Overhead of Oxygen
The key operational difference lies in oxidizer storage:
Storable Hypergolics (MMH/NTO): Can remain in spacecraft tanks at ambient temperatures for decades without active cooling or venting.
Liquid Oxygen (LOX): Requires cryogenic temperatures (90 K / -183 °C). For long-duration in-space missions (e.g., GEO satellites, deep-space probes), using LOX requires:
• Multi-Layer Insulation (MLI) and active cryocoolers (Zero Boil-Off / ZBO systems).
• Tank Venting Systems: Unvented pressure buildup can rupture tanks over long durations.Gaseous Oxygen (GOX) Alternative: Storing oxygen as a compressed gas avoids cryogenics, but drastically increases tank structural mass and volume requirements.
3. Container & Dry Mass Budget
Because silicon is a dense solid (~2.33 g/cm³), the fuel portion of the system occupies a remarkably small tank volume. However, the system's overall dry mass is dominated by the oxidizer side:
High Oxidizer Mass Ratio: Stoichiometrically, burning silicon requires an oxygen-to-fuel mass ratio (O/F) of 1.14.
Tank Mass Penalty: Storing cryogenic LOX or high-pressure GOX requires thicker tank walls or complex vacuum-jacketed insulation compared to thin-walled aluminum/titanium tanks used for ambient MMH/NTO.
Overall Mass Equivalent: Because of the lower Isp (230 s vs 310 s), a Si–O₂ system requires roughly 40% to 50% more total wet propellant mass than an equivalent MMH/NTO system to deliver the same total velocity increment (Δv). This extra propellant requirement offsets the volumetric efficiency of solid silicon.
Critical Considerations for Decision Makers
Key Takeaway: A Si–O₂ system is generally uncompetitive as a primary propulsion engine for standard Earth-launched satellites due to Isp penalties and LOX storage overhead. However, it holds significant advantages in specific niche applications.
Safety & Ground Operations:
Hypergolic propellants (N₂O₄ / Hydrazine derivatives) are acutely toxic, volatile, and carcinogenic. Ground operations require hazardous fueling protocols, extensive personal protective equipment (SCAPE suits), and dedicated infrastructure. Silicon and oxygen are completely non-toxic and inherently safe to handle prior to launch.In-Situ Resource Utilization (ISRU):
Lunar regolith is roughly 45% Oxygen and 21% Silicon by mass. A lunar lander or hopping vehicle utilizing Si/O₂ could synthesize both its fuel and oxidizer directly from lunar soil via molten salt electrolysis, eliminating the need to transport propellant from Earth.Nozzle Erosion & Slag Accumulation:
The condensation of liquid SiO₂ during burn cycles causes slag accumulation inside the combustion chamber and severe abrasive erosion on nozzle throats, complicating multi-start or long-duration burn profiles.
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In this post I asked Gemini to think about GW Johnson's space tug application. It appears to me the key failing of this propulsion method is accumulation of slag in the engine. The low ISP is due to the high molecular weight of Silicon Dioxide.
'
Executive Summary: Si–O₂ Propulsion for an Earth–Mars Space Tug Industry
An Earth–Mars space tug architecture relies heavily on three core factors: propellant availability, thrust-to-weight ratio, and long-term operating costs.
Evaluating a Silicon–Oxygen (Si–O₂) propulsion system sourced from the Moon against competing architectures (such as Lunar LOX/LH₂, Lunar LOX/Al, nuclear thermal, or high-power electric) reveals clear operational trade-offs for a commercial transit service.
--------------------------------------------------
Strategic Value: The Lunar Sourcing Advantage
The primary economic argument for Si–O₂ is In-Situ Resource Utilization (ISRU) abundance:
Regolith Composition: Lunar regolith is roughly 45% Oxygen and 21% Silicon by weight. Both elements are universally available across almost the entire lunar surface, unlike water ice which is concentrated exclusively in permanently shadowed polar craters.
Simplified Extraction: Molten salt electrolysis (such as the FFC Cambridge process) can extract pure elemental silicon and gaseous oxygen directly from bulk dry regolith in a single processing loop.
Gravitational Energy Savings: Launching propellant mass out of the Moon's gravity well requires a fraction of the velocity increment (Δv ≈ 2.4 km/s) compared to launching from Earth (Δv ≈ 9.3 – 10.0 km/s).
--------------------------------------------------
Technical Feasibility & System Trade-Offs
While the propellant is inexpensive to source on the Moon, using it for high-energy interplanetary transits introduces critical engineering challenges.
1. The Specific Impulse (Isp) Penalty for Transits
An Earth–Mars transfer (Hohmann or fast-transit) requires significant velocity maneuvers at Trans-Mars Injection (TMI) and Mars Orbit Insertion (MOI):
Δv_TMI ≈ 3.6 to 4.5 km/s (from Low Earth Orbit)
Because final vehicle mass scales exponentially with Δv / (g_0 * Isp), the modest performance of Si–O₂ (Isp ≈ 210 – 250 s) requires a massive propellant mass fraction compared to hydrolox or methalox:
Silicon / Oxygen (Si/O₂):
• Lunar Feasibility: Universal (Dry regolith everywhere)
• Vacuum Isp: 230 s
• Propellant Mass Ratio for 4.0 km/s Δv: 5.88 to 1 (Requires 5.88x dry mass in propellant)Aluminum / Oxygen (Al/O₂):
• Lunar Feasibility: Moderate (Requires processing Anorthite)
• Vacuum Isp: 260 s
• Propellant Mass Ratio for 4.0 km/s Δv: 4.61 to 1Hydrogen / Oxygen (LOX/LH₂):
• Lunar Feasibility: Restricted (Polar water ice only)
• Vacuum Isp: 450 s
• Propellant Mass Ratio for 4.0 km/s Δv: 2.48 to 1Nuclear Thermal (NTP - H₂):
• Lunar Feasibility: Earth-supplied reactant
• Vacuum Isp: 900 s
• Propellant Mass Ratio for 4.0 km/s Δv: 1.57 to 1While the propellant mass originates on the Moon rather than Earth, a Si–O₂ tug must carry over twice as much propellant mass per maneuver as a hydrogen-fueled tug.
2. Operational Modes for the Space Tug Architecture
A space tug system using Si–O₂ can be structured in two primary operational ways:
High-Impulse "Shove" Engine (Cislunar / TMI Booster):
• Role: The tug stays strictly within the Earth–Moon system. It docks with customer vessels in LEO or High Earth Orbit (HEO), provides a high-thrust TMI "shove" toward Mars, detaches, and uses lunar flybys or aerocapture to return to a lunar staging depot.
• Verdict: Viable. Since the tug does not make the full trip to Mars, the low Isp penalty is offset by zero-cost propellant sourcing from the Moon.Full Earth–Mars Cycler / Transit Tug:
• Role: The tug accompanies the payload all the way to Mars, performs capture, and returns.
• Verdict: Challenging. Carrying heavy silicon combustion products and cryogenic oxygen over interplanetary distances with a 230 s Isp results in severe dry-mass scale-up.3. Engine Reuse & Maintenance Issues
For a commercial space tug company, engine lifespan and turnaround time are primary financial metrics:
Slag Accumulation: Silicon dioxide (SiO₂) condenses into liquid glass/slag at chamber pressures and temperatures. Slag pooling in the chamber or throat alters nozzle geometry, degrading performance over multi-minute burns.
Refueling Logistics: Refinement of solid silicon grains requires physical mechanical handling (inserting structured porous silicon blocks or slurries into hybrid fuel casings), whereas liquid-liquid systems (LOX/LH₂ or LOX/CH₄) allow rapid fluid line quick-disconnect refueling.
--------------------------------------------------
Verdict for Decision Makers
Strategic Trade-Off: Si–O₂ trades chemical efficiency (Isp) for unrestricted supply-chain availability.
Where it wins: As a Cislunar Booster / Earth-Departure Shove Engine. If lunar polar ice proves too scarce or expensive to mine, dry-regolith Si–O₂ offers an infinite, globally accessible propellant supply on the Moon.
Where it loses: As a long-haul interplanetary cruise motor. For full transits, high-efficiency systems (such as high-power solar/nuclear electric propulsion or nuclear thermal) remain vastly superior in payload mass fraction.
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In this post, I asked Gemini to think about the slag problem. Apparently considerable research has been done by NASA and others, to see if the problem of slag might be addressed. One detail I found interesting is that high thrust burns such as would be performed by a space tug apparently sustain operating conditions that prevent formation of slag, and addition of a flow of oxygen in the border may prevent solidification in the expansion bell.
Summary of Slag Research & Mitigation for Si–O₂ Engines
The formation of liquid glass (SiO₂) slag during combustion is recognized in aerospace literature as a primary hurdle for reusable metal/silicon propulsion. Research indicates three primary methods to manage or prevent engine degradation:
Vortex Gas Boundaries: Using swirling oxygen injection to keep liquid droplets centered in the flow stream, preventing contact with chamber walls.
Refractory Nozzle Liners: Utilizing Silicon Carbide (SiC) or Rhenium coatings to resist abrasive erosion and chemical reactions with hot silica.
Sub-Micron Pore Control: Micro-structuring the solid silicon matrix so that exhaust particles remain tiny aerosols suspended in the gas stream rather than coalescing into heavy liquid pools.
Commercial Bottom Line: While slag prevents infinite engine reuse without maintenance, active fluid-dynamic shielding and refractory materials allow multi-burn operation suitable for cislunar boost missions.
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Gemini offered to prepare the full slag mitigation strategy for us:
Technical Deep-Dive: Slag Mitigation & Nozzle Erosion in Si–O₂ Rocket Engines
In high-temperature Silicon–Oxygen (Si–O₂) combustion, the primary reaction product is silicon dioxide (SiO₂). At chamber temperatures exceeding 2,800 K, SiO₂ exists as a mix of gas, vaporized silicon monoxide (SiO), and liquid droplets.
As these products expand through the nozzle, the rapid temperature drop causes liquid glass droplets to coalesce, impinge on nozzle walls, and freeze. Understanding and mitigating this behavior is critical for engine efficiency and reusability.
--------------------------------------------------
1. Fluid-Dynamic & Thermal Mechanisms of Slagging
Phase Transition During Expansion:
Inside the combustion chamber, temperatures exceed the boiling point of SiO₂ (~2,500 °C / 2,773 K). However, as gas expands past the nozzle throat, static temperature drops sharply below the melting point (~1,713 °C / 1,986 K).Droplet Coalescence & Surface Tension:
In microgravity, liquid SiO₂ micro-droplets do not fall out due to gravity; instead, surface tension causes them to collide and merge into larger liquid glass pools along the chamber walls.Two-Phase Flow Velocity Lag:
Condensed liquid and solid particulates cannot expand like gas. The gas accelerates faster than the heavy droplets, creating a velocity lag:Lag Ratio = V_liquid / V_gas < 1.0
This momentum lag reduces the effective exhaust velocity, resulting in a 5% to 15% reduction in overall Isp compared to theoretical gas-only calculations.
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2. Nozzle Erosion & Material Degradation
When molten silica contacts nozzle throat materials at high velocities, three distinct degradation mechanisms occur:
Abrasive Hydro-Erosion: High-velocity liquid glass droplets act as a severe abrasive liquid jet, mechanically scouring the throat material during multi-minute burns.
Carbothermic Chemical Attack:
Standard graphite or carbon-composite nozzles suffer severe chemical degradation. Hot silicon dioxide reacts chemically with carbon at high temperatures:SiO₂ (liquid) + 3 C (solid) → SiC (solid) + 2 CO (gas)
This reaction converts the structural carbon nozzle wall into gas and weak porous silicon carbide, causing rapid throat expansion and loss of chamber pressure.
Thermal Shock & Glass Stressing:
During engine shutdown, residual liquid glass cools into an amorphous silica layer. Because silica and metals/graphite have radically different Coefficients of Thermal Expansion (CTE), repeated thermal cycles cause surface cracking and spalling.--------------------------------------------------
3. Engineered Slag Mitigation Strategies
To prevent throat destruction and maintain steady thrust profiles, propulsion researchers utilize four primary engineering defenses:
Vortex / Swirl Gas Barrier (Boundary Layer Shielding):
Gaseous oxygen is injected tangentially near the injector face to create a high-speed, swirling gas layer along the combustion chamber wall. This centrifugal boundary layer forces heavy liquid SiO₂ droplets into the core stream, keeping them away from the throat walls.Sub-Micron Porous Silicon Grain Fabrication:
By structuring the solid silicon grain at nanoscale porosity (pores < 100 nm), combustion occurs at ultra-fine surface boundaries. This ensures that exhaust particles remain tiny aerosol droplets (sub-micron scale) that stay entrained in the gas stream rather than merging into large liquid slag pools.Refractory & Chemical Coating Selection:
To survive chemical attack and abrasive scouring, nozzles for Si–O₂ engines require specialized non-carbon refractory materials:
• Silicon Carbide (SiC): Chemically compatible with silica; resists oxidation.
• Rhenium-Coated Carbon-Carbon: Provides an impervious metallic barrier against carbothermic reduction up to 2,400 °C.
• Tungsten / Tantalum-Hafnium Carbide alloys: Extremely durable for long-duration throat inserts.Hydrogen / Fluorine Additive Stripping:
Introducing trace amounts of hydrogen (or fluorinated compounds) into the oxidizer feed alters reaction thermodynamics. Hydrogen reacts to form gaseous water vapor and silicon monoxide (SiO), suppressing liquid phase formation:Si + H₂O → SiO (gas) + H₂ (gas)
This keeps a larger fraction of the exhaust in the gaseous phase throughout expansion.
--------------------------------------------------
4. Summary for Operational Space Tug Design
Engineering Takeaway: Slag accumulation and nozzle erosion are manageable engineering problems rather than fatal design flaws. By employing swirl oxygen wall-shielding and Rhenium or SiC throat liners, a Si–O₂ engine can achieve multi-start capability and controlled burn profiles for cislunar boost operations.
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In this post Gemini took up a possible space tug business scenario. It added a detail about shutdown of the engine that looks helpful to me.
Operational Profile: Sustained High-Thrust Burns & Dual-Mode Tug Architecture
Using a dual-propulsion strategy—combining a High-Thrust Si–O₂ core for major velocity changes with a clean secondary system for docking—directly addresses the physical limitations of Silicon–Oxygen combustion.
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1. Dynamics of Sustained Heavy Burns
During continuous, maximum-thrust maneuvers (e.g., Trans-Mars Injection or high-energy capture):
Vapor-Phase Retention: Peak chamber temperatures (>2,900 K) and high chamber pressures keep a larger proportion of silicon exhaust in gaseous form (SiO and SiO₂ gas).
High-Velocity Dynamic Sweep: High mass flow rates create intense boundary-layer shear forces, sweeping micro-droplets cleanly out of the nozzle bell before they can coalesce into heavy liquid pools.
Post-Burn Purge Protocol: To prevent residual silica from freezing on the nozzle walls during shutdown cooling, the engine executes a brief 2-second gaseous oxygen (GOX) purge cycle immediately at main-engine cut-off (MECO).
--------------------------------------------------
2. Dual-Mode Architecture for Proximity Operations
Maneuvering & Docking (Secondary System):
Using clean monopropellants, compressed gas, or electric thrusters for rendezvous and proximity operations (RPO) avoids low-temperature Si–O₂ pulsing. This prevents slag formation during fine adjustments and protects customer spacecraft optics and solar panels from particle contamination.Main Impulse Shove (Primary Si–O₂ Core):
Engaged only after structural hard-docking is confirmed and the combined vehicle stack is aligned along the target vector, maximizing the economic benefit of lunar-sourced propellant for high-mass acceleration.--------------------------------------------------
Conclusion: Reserving Si–O₂ strictly for long, continuous, high-thrust burns while utilizing clean secondary thrusters for docking creates a robust, highly practical operational framework for a commercial space tug fleet.
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In this post, I asked Gemini to think about the three major active engine scenarios.
It appears that landing will be a problem, but the vessel will be as light as it will ever be while landing, so the maneuvering engines may be able to handle the task.
Evaluation of Silicon–Oxygen Burns: LEO Breakout, Lunar Ascent, and Landing
Evaluating the three primary delta-v maneuvers required for a lunar-supplied space tug reveals that burn intensity directly dictates slag risk.
--------------------------------------------------
Burn-by-Burn Thermal & Slag Breakdown
1. LEO Breakout / Trans-Mars Injection (Full Thrust):
• Slag Risk: Very Low
• Physics: Continuous 100% thrust maintains maximum chamber pressure and flame temperatures (>2,900 K). Intense gas shear forces sweep micro-particulates out of the nozzle before pooling can occur.
• Protocol: Standard 2-second gaseous oxygen (GOX) purge at main-engine cut-off (MECO).2. Fully Loaded Lunar Liftoff (Ascent Burn):
• Slag Risk: Very Low
• Physics: Liftoff with a maximum payload and propellant load requires full mass flow rate. Chamber conditions remain well above the SiO₂ condensation threshold throughout ascent.
• Protocol: Standard GOX purge upon reaching orbit.3. Powered Descent & Lunar Landing (Throttled Burn):
• Slag Risk: High / Critical
• Physics: Pinpoint landing requires throttling down to 20%–40% thrust. Reduced chamber pressure lowers core flame temperatures, while decreased mass flow velocity permits liquid glass to pool on chamber walls and the nozzle throat.--------------------------------------------------
Engineered Solutions for the Landing Phase
To prevent throat degradation during the throttled landing maneuver, three operational approaches can be implemented:
Auxiliary System Terminal Touchdown:
The main Si–O₂ core executes the high-energy Powered Descent Initiation burn at 100% thrust. The engine is shut down and purged at ~500 m altitude, allowing the clean secondary maneuvering thrusters to handle final hover and touchdown without producing slag or kicking back glass particles.Oxidizer-Biased Wall Throttling:
Throttling is achieved by reducing solid silicon consumption while maintaining high-speed GOX injection along the chamber boundary layer. This preserves high wall velocity to sweep particulates out even at lower net thrust.Ground Depot Servicing:
Because the lander touches down directly at the lunar processing facility, the engine undergoes automated mechanical inspection, cleaning, and solid silicon core reloading prior to the next ascent.--------------------------------------------------
Summary: Heavy breakout burns and fully loaded liftoffs operate at ideal full-thrust conditions to prevent slagging. The throttled landing burn represents the primary technical bottleneck, which is cleanly resolved by executing terminal touchdown on secondary clean thrusters or utilizing wall-biased GOX throttling.
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The AI pointed out exactly the same difficulty that I pointed out in my email to you. The silicon oxide is significantly a condensed species that cannot expand in the nozzle, thus not producing any thrust. The AI used the term "two-phase losses" pointing this out. It's not an either-or thing, but it is a significant loss, which is why the Isp estimated for it is actually lower than solids with sea level nozzles, even AN-oxidized solids.
If you plan on accepting an Isp that low, why nor just use the solid instead, and avoid all that silica particle contamination it leaves behind? About the worst offender in a solid's plume might be carbon soot plus some HCl (hydrochloric acid).
Most solids today are AP-oxidized, might or might not include aluminum, and are bound with HTPB or CTPB (not PBAN, it cracks upon soaking out cold to around 0 F). Isp with SL nozzles is near 250 s, and with a vacuum nozzle, can approach 280. And if you stay away from PBAN binder, they can easily be qualified to Mil Std 210 hot and cold soak, which is 145 F down to -65 F. The older 210 standard when I first entered the industry had -65 F, but the hot soak was higher at 165 F.
Propellant soak temperature affects burn rate (higher hotter, slower colder) which affects both chamber pressure and thrust, and burn time. It does not affect total impulse hardly at all. The higher pressure fully hot is what you size the case to take. In tactical sizes that is usually in the vicinity of 2000 psia. In shuttle booster sizes, it was nearer 1000 psia, because steel strength does not scale up with size.
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Last edited by GW Johnson (Yesterday 10:01:21)
GW Johnson
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"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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