Liquid Cryogenic Fuel & Oxidizer

Abstract

This review article explores the intricacies of liquid cryogenic and solid fuels, focusing on their properties, production, storage, and applications. Liquid cryogenic fuels, such as liquid hydrogen (LH2), offer high energy densities essential for space exploration. The production and storage of LH2 involve handling extremely low temperatures and ensuring minimal heat transfer. Solid fuels, on the other hand, provide stability and ease of handling, making them suitable for various propulsion systems. Advances in both fuel types have led to significant improvements in efficiency and safety. This article provides a comprehensive overview of the current state of these fuels, highlighting recent advancements and future prospects in aerospace and other industries.

Keywords

Liquid Cryogenic Fuel, Solid fuels

Introduction

5. Synthesis of Rocket-Grade Nitrogen Tetroxide (N2O4‌):

Rocket-grade nitrogen tetroxide (N2O4‌) is a high-purity oxidizer commonly used in bipropellant rocket engines, often paired with hydrazine-based fuels. Here’s a concise overview of its synthesis.

• Production of Nitrogen Dioxide (NO2):
• Starting Materials: Nitrogen dioxide is typically produced from nitric oxide (NO) and oxygen (O2).
• Reaction:

2NO + O2 → 2 NO2

5. Conversion to Dinitrogen Tetroxide (N2O4):

• Dimerization: At lower temperatures, nitrogen dioxide (NO2) dimerizes to form dinitrogen

• tetroxide(N2O4).

• Reaction:

2 NO2 ↔ N2O4

The equilibrium between NO2 and N2O4 shifts towards N2O4 as the temperature decreases.

5. Purification:

• Distillation: The crude N2O4 is purified by distillation to remove impurities and ensure it meets rocket-grade specifications.

• Storage: Rocket-grade N2O4 is stored in specialized containers to prevent contamination and decomposition.

Key Points for Rocket-Grade N2O4

O High Purity: Ensures consistent performance and minimizes risks of contamination in rocket engines.

O Stability: Stored at controlled temperatures to maintain the N2O4 form and prevent decomposition back to NO2.

Rocket-grade N2O4 is synthesized by dimerizing nitrogen dioxide, followed by purification to achieve the high purity required for rocket propulsion. It is a crucial oxidizer in many bipropellant systems due to its effectiveness and storability.

[Nitrogen Tetroxide To Mixed Oxides Of Nitrogen: History, Usage, Synthesis, And Composition Determination By Andrew Head]

5. Application of Rocket-Grade Nitrogen Tetroxide (N2O4) in Rocketry:

O Rocket-grade nitrogen tetroxide (N2O4) is a crucial oxidizer used in various rocket propulsion systems. Here's how it is applied in rocketry:

Bipropellant System:

Paired with hydrazine (N2H4) or derivatives (like MMH).

Reaction: N2H4 + N2O4 → N2 + 2 H2O + Heat.

O Hypergolic Propellants:

Ignites spontaneously upon contact with hydrazine.

Reliable for spacecraft maneuvering thrusters.

O Storable Propellants:

Storable at ambient temperatures.

Ideal for long-duration space missions and military use.

O Spacecraft Maneuvering:

Used in Reaction Control Systems (RCS) for orientation.

Employed for orbital insertion and adjustments.

O Rocket Engine Design:

Contributes to high specific impulse (Isp).

Produces significant thrust for lifting payloads.

[Sutton, George P., And Oscar Biblarz. "Rocket Propulsion Elements." John Wiley & Sons, 2010, Huzel, Dieter K., And David H. Huang. "Modern Engineering For Design Of Liquid-Propellant Rocket Engines." Aiaa, 1992. ]

Rocket-grade N2O4 is essential in rocketry as a high-performance oxidizer, particularly in bipropellant systems. It is indispensable for reliable and efficient space missions, from main engines to maneuvering thrusters.

LOX (Liquid Oxygen):

Liquid Oxygen (LOX) is used in rocket engines primarily for its role as a powerful oxidizer. Here are the key reasons why LOX is preferred in rocketry:

High Oxidizing Potential:

LOX is a highly efficient oxidizer, providing a large amount of oxygen molecules per unit volume or mass. This high oxidizing potential is crucial for the combustion of rocket fuels, allowing for rapid and complete burning of the fuel.

5. Compatibility with Various Fuels:

LOX can be used with a wide range of fuels, including liquid hydrogen, RP-1 (refined kerosene), methane, and others. This versatility makes it suitable for different types of rocket engines, from small-scale thrusters to large launch vehicles.

5. Density and Storage:

Unlike gaseous oxygen, LOX is stored in its liquid form at cryogenic temperatures (around -183°C or -297°F). This allows for a higher density of oxidizer to be stored in a given volume, which is advantageous for achieving higher energy density in the propellant system.

5. Combustion Efficiency:

When combined with a suitable fuel, LOX supports efficient and stable combustion processes in rocket engines. This efficiency is critical for achieving the high thrust-to-weight ratios required for space missions.

5. Reliability and Experience:

LOX has been extensively tested and used in rocketry for decades. Its reliability in igniting and sustaining combustion reactions, as well as its compatibility with various structural materials, contributes to its continued use in both commercial and government space programs.          

Overall, LOX's high oxidizing potential, compatibility with different fuels, density advantages, and proven reliability make it a preferred choice as an oxidizer in rocket engines for achieving efficient and powerful propulsion.

Liquid oxygen and hydrogen have been used for rocket propulsion devices for many years and are still considered the best combination in many launch applications today. The cryogenic liquid oxygen (LOx) and hydrogen (GH2 and LH2) bi-propellant combination is used for the majority of today's heavy launch vehicles. L0x/H2 powers the Space Shuttle Main Engine (SSME), the Japanese H-IIA upper and main-stage engines and various stages of Europe's most successful series of commercial launch vehicles in the world; the Ariane family (Figure 2.2). The cryogenic combination is the highest performer with a Specific Impulse (equation 2.2) higher than any other liquid or solid propellants flight tested to date. Specific Impulse is the most common measure of performance for rocket engines and is defined as the thrust produced per unit flow rate of propellant as follows;

I_{SP= FMG=VexhG}

The specific impulse of a rocket engine using liquid oxygen (LOX) as the oxidizer and liquid hydrogen (LH2) as the fuel is a key performance metric. Specific impulse is a measure of how effectively a rocket uses its propellant, often expressed in seconds. For LOX/LH2 engines, the sp values can vary based on the engine design and operating conditions (vacuum or sea level).

5. Typical Specific Impulse Values for LOX/LH2 Engines:

O Vacuum Specific Impulse:

- Typical Value: Around 45 seconds

- Examples: The Space Shuttle Main Engine (SSME), also known as the RS-25, achieves a vacuum specific impulse of approximately 452 seconds.

O Sea Level Specific Impulse:

- Typical Value: Around 360-370 seconds

- Examples: The RS-25 engine has a sea level specific impulse of about 366 seconds

(Rocket Propulsion Elements By George P. Sutton And Oscar Biblarz/ Nasa Technical Reports And Publications)

Figure 3 Treasure vs temperature plot

These specific impulse values highlight the efficiency of LOX/LH2 engines, making them one of the most effective propellant combinations for high-performance rockets, especially for missions requiring high efficiency and long-duration burns, such as those involving upper stages and space exploration missions. High-pressure injection and combustion of liquid oxygen (LOX) and liquid hydrogen (LH2) are crucial for achieving high performance in rocket engines. Turbo-pumps pressurize LOX and LH2 to over 100 bar, driven by gas generators or preburners. Injectors, such as pintle and showerhead designs, ensure prope mixing and atomization of propellants into fine droplets, enhancing combustion efficiency. Regenerative cooling circulates cryogenic propellants through the engine walls before injection, preheating the propellants and cooling the engine. The combustion chamber, built to withstand extreme temperatures (over 3,000°C) and pressures, uses high-strength materials and cooling techniques. Ignition is initiated by spark or hypergolic means, ensuring nearly complete combustion for maximum efficiency. The hot gases expand through a converging-diverging nozzle with an optimized expansion ratio for maximum exhaust velocity and thrust. Examples of high-performance engines include the RS-25 (Space Shuttle Main Engine) with a chamber pressure of approximately 207 bar and a specific impulse of around 452 seconds in a vacuum, and the RL10 engine with a chamber pressure of around 40-50 bar and a specific impulse of about 465 seconds in a vacuum. These principles enable efficient and powerful rocket propulsion, as detailed in *Rocket Propulsion Elements* by George P. Sutton and Oscar Biblarz, and NASA technical reports.

(High Pressure Loxihz Rocket Engine Combustion The University Of Adelaide Austbatia Joshua J. Smit)

5. synthesis of rocket-grade liquid oxygen (LOX)

involves several critical steps to ensure it meets the high purity and quality standards required for rocket propulsion. Here is an overview of the process:

1. Air Separation: Air Intake and Compression: Ambient air is drawn into large air separation units (ASUs) and compressed to a high pressure.

2. Air Cooling: The compressed air is cooled to cryogenic temperatures using heat exchangers and refrigeration cycles. This process involves several stages of cooling and expansion to achieve the necessary low temperatures.

3. Fractional Distillation:

3.A Liquefaction: At cryogenic temperatures, the air liquefies. The main components, nitrogen (N2) and oxygen (O2), have different boiling points (N2: 77K, O2: 90K), allowing for separation.

3.B Distillation Column: The liquefied air is fed into a distillation column. Due to the different boiling points, nitrogen boils off first, leaving behind liquid oxygen.

The distillation column may operate at different pressures to enhance separation efficiency.

3. Purification: Removal of Impurities: Trace impurities, such as argon, carbon dioxide, and water vapor, are removed through additional distillation steps or through the use of adsorbent materials.

High-purity oxygen is required to avoid contamination that could affect combustion efficiency and safety.

5. Storage and Transport:

5.A Storage: The purified liquid oxygen is stored in cryogenic tanks at temperatures below its boiling point (-183°C or 90K). These tanks are heavily insulated to minimize heat transfer and evaporation losses.

Tanks are equipped with pressure relief valves to safely vent excess pressure caused by any evaporation.

5.B Transport: LOX is transported to launch sites in specialized cryogenic tankers designed to maintain low temperatures and minimize losses.

Safety protocols are strictly followed due to LOX's reactive nature and the potential for rapid combustion when it contacts fuels or organic materials.

In addition to LOX (liquid oxygen), N2O4 (nitrogen tetroxide), and H2O2 (hydrogen peroxide), several other oxidizers are used in rocket science. These include Fluorine and ClF3 (chlorine trifluoride), which offer high energy potential but are extremely toxic and corrosive, making them difficult to handle. Nitric acid (HNO3) and its derivatives, like IRFNA (Inhibited Red Fuming Nitric Acid), have been used in hypergolic engines due to their storability and hypergolic properties with fuels like hydrazine. Ammonium perchlorate (AP) is a solid oxidizer commonly used in composite solid rocket propellants, providing high energy and stability. These oxidizers are selected based on their specific performance characteristics and the requirements of the mission, balancing factors like energy density, storability, and handling safety.

[Nasa Technical Reportsmodern] [Engineering For Design Of Liquid-Propellant Rocket Engines" By Dieter K. Huzel And David H. Huang]

Fuels

Cryogenic Fuels:

Cryogenic fuels are those that require storage at extremely low temperatures to remain in a liquid state. These fuels have high specific impulse but present significant challenges in handling and storage due to their cryogenic nature.

5. Liquid Hydrogen (LH2):

- Requires storage at temperatures below -252.87°C (-423.17°F).

- Commonly used with liquid oxygen (LOX) in engines like the Space Shuttle Main Engine (RS-25) and the RL10.

5. Liquid Methane (CH4):

- Requires storage at temperatures below -161.5°C (-258.7°F).

- Used in engines like SpaceX's Raptor and Blue Origin's BE-4.

5. Normal Liquid Fuels:

Normal liquid fuels can be stored at or near ambient temperatures, making them easier to handle and store compared to cryogenic fuels. These fuels are typically less efficient than cryogenic fuels but are more practical for certain applications, especially where storability and simplicity are key.

5. RP-1 (Refined Petroleum-1):

- A highly refined form of kerosene, storable at ambient temperatures.

- Used in engines like the Rocketdyne F-1 (Saturn V) and Merlin (SpaceX Falcon 9).

5. Hydrazine (N2H4) and its Derivatives (Aerozine 50, UDMH, MMH):

- Storable at ambient temperatures.

- Used in spacecraft thrusters and orbital maneuvering systems, often with nitrogen tetroxide (N2O4) as the oxidizer.

- Examples include Aerozine 50 (a mixture of hydrazine and UDMH) and MMH.

[Rocket Propulsion Elements" By George P. Sutton And Oscar Biblarz.                                                                                                                

Industry Publications And Technical Manuals: Data From Companies Like Spacex, Blue Origin, And Aerojet Rocketdyne On The Fuels Used In Their Engines]

Liquid fuels are fundamental to rocket propulsion due to their unique capabilities and advantages. Firstly, they offer high energy density, particularly cryogenic fuels like liquid hydrogen (LH2) and liquid methane (CH4), enabling rockets to achieve significant velocities and carry substantial payloads crucial for space exploration missions. Secondly, these fuels provide precise control over thrust and engine performance, allowing for adjustments during flight that facilitate maneuvers, orbital transfers, and controlled descent. Thirdly, their higher specific impulse (Isp) compared to solid fuels ensures efficient use of propellant mass, maximizing payload capacity and extending mission durations. Moreover, the storability of many liquid fuels at ambient temperatures, such as RP-1 (refined kerosene) and hydrazine derivatives, enhances operational flexibility and reduces turnaround times. This characteristic also supports various engine configurations and innovations in propulsion technology, including reusable rockets and advanced deep-space missions. Lastly, liquid fuels generally offer better safety profiles and lower environmental impacts compared to some solid propellants, contributing to their widespread use in modern rocketry as reliable, versatile, and efficient propulsion solutions.

(Specific Impulse (Isp) Measures How Efficiently A Rocket Engine Uses Its Propellant To Produce Thrust. It's Expressed In Seconds (S), Indicating Thrust Per Unit Of Propellant Mass Consumed. Higher Isp Values Signify Greater Efficiency, Crucial For Achieving Higher Velocities And Optimizing Mission Performance.)

8. Liquid Hydrogen (LH2):

The idea of using liquid hydrogen (LH2) as a rocket propellant was proposed and developed by a team led by Dr. Robert H. Goddard, often considered the father of modern rocketry. Goddard's pioneering work in the early 20th century laid the groundwork for liquid-fueled rocket engines and their application in space exploration. His research and experiments, particularly in the 1920s and 1930s, demonstrated the feasibility of using liquid hydrogen and liquid oxygen (LOX) as propellants to achieve high performance and efficiency in rocket engines. This foundational work by Goddard and his contemporaries paved the way for the eventual realization and implementation of liquid hydrogen as a key component of modern rocket propulsion systems. Gwynne A. Wright, under the guidance of Professor Herrick L. Johnston at Ohio State University, operated the first hydrogen liquefier. This achievement marked a significant milestone in the development of cryogenic engineering, enabling the production and use of liquid hydrogen for various applications, including rocket propulsion. This early work laid important groundwork for advancing the capabilities and efficiency of liquid hydrogen as a critical component in modern rocketry and space exploration.

(LIQUID HYDROGEN AS A PROPULSION FUEL, 1945-1959 JOHN L. SLOOP)

 

Figure 4- LIQUID HYDROGEN AS A PROPULSION FUEL, 1945-1959 JOHN L. SLOOP)

5. Synthesis of Liquid Hydrogen for Rocket Grade with Reaction:                

Synthesizing liquid hydrogen (LH2) for rocket fuel involves several key steps: production, purification, and liquefaction. Here's an overview of the process along with the chemical reactions involved:

• Production of Hydrogen Gas (H2):

- Hydrogen gas can be produced through various methods, with steam methane reforming (SMR) being the most common industrial process.

- Steam Methane Reforming (SMR):

CH4 + H2O -> CO + 3H2

• Followed by the water-gas shift reaction:

CO + H2O -> CO2 + H2

• Purification:

The hydrogen gas produced needs to be purified to remove impurities such as carbon monoxide (CO), carbon dioxide (CO2), and methane (CH4). This is typically done using pressure swing adsorption (PSA) or other gas purification techniques.

• Liquefaction:

Once pure hydrogen gas is obtained, it is liquefied. Hydrogen has a boiling point of approximately -253°C (-423°F), so it requires significant cooling. The liquefaction process generally involves the following steps:

• Compression:

Hydrogen gas is compressed to a high pressure.

• Pre-cooling:

The compressed hydrogen is pre-cooled using liquid nitrogen or another refrigerant.

Expansion and Joule-Thomson Effect:

The pre-cooled hydrogen undergoes expansion through a Joule-Thomson valve or expansion turbine, leading to cooling. This step is often repeated in a series of stages to achieve the necessary low temperature.

Kinetics of LH2 Propellant in Rockets:

5. Mass Flow Rate

The mass flow rate (?) of the propellants (LH2 and LOX) into the combustion chamber can be calculated using the following formula:

? = ρ A v

where:

• ρ is the density of the propellant.
• A is the cross-sectional area of the fuel line.
• v is the velocity of the propellant flow.

5. Combustion Reaction

The stoichiometric reaction of LH2 and LOX can be represented as

2H2 + O2 → 2H2O

• For complete combustion, the stoichiometric ratio (oxidizer-to-fuel ratio by mass) is crucial:

ofration=m02mh2=324=8

The thrust (F) produced by a rocket engine can be calculated using:

F = ?VE + (PE - PA) E

where:

• ?? is the total mass flow rate of the exhaust gases.
• ve is the exhaust velocity.
• Pe is the pressure at the nozzle exit.
• Pa is the ambient pressure.
• Ae is the area of the nozzle exit.

5. Exhaust Velocity

The exhaust velocity (ve) can be found using the following formula derived from the energy balance:

ve= v [2(γ/ (γ - 1)) R Tc(1-(Pe/Pc) ^((γ-1)/γ))]

where:

• γ is the specific heat ratio (typically 1.4 for diatomic gases like H2 and O2).
• R is the specific gas constant.
• Tc is the combustion chamber temperature.
• Pe is the pressure at the nozzle exit.
• Pc is the combustion chamber pressure.

 

5. Specific Impulse

The specific impulse (Isp) is a measure of the efficiency of the rocket engine and is given by:

I_sp =veg0

 

where:

• ve is the exhaust velocity.
• g0 is the standard gravitational acceleration (9.81 m/s²).

 

5. Ideal Rocket Equation

The ideal rocket equation, or Tsiolkovsky rocket equation, relates the change in velocity (Δv) of the rocket to the effective exhaust velocity (ve) and the initial (mi) and final (mf) mass of the rocket:

Δv = ve ln (mimf

)

 

where:

- mi is the initial mass of the rocket (including propellant).

- mf is the final mass of the rocket (after propellant is burned)

 

5. Example Calculation

To illustrate, let's consider a rocket with the following parameters:

Ex.

• Mass flow rate of LH2 (??_LH2):
• 10 kg/s
• Mass flow rate of LOX (??LOX): 80 kg/s
• Exhaust velocity (ve): 4500 m/s
• Nozzle exit area (Ae): 1 m²
• Combustion chamber pressure (Pc): 5 MPa- Nozzle exit pressure (Pe): 100 kPa
• Ambient pressure (Pa): 101.3 kPa

First, calculate the total mass flow rate:

?t= ?_LH2 + ??_LOX = 10 + 80 = 90 kg/s

Next, calculate the thrust:

F = ? ve + (Pe - Pa) Ae

F = 90 ⋅ 4500 + (100 - 101.3) ⋅ 1 × 10^3

F = 405000 - 1.3 × 10^3

F = 403700 N

Then, calculate the specific impulse:

Isp = ve / g_0 = 4500 / 9.81 ≈ 459 seconds

Finally, Using The Rocket Equation, If The Initial Mass Is 200,000 Kg And The Final Mass Is 50,000 Kg:

Δv = ve ln(mi / mf)

Δv = 4500 ⋅ ln(200000 / 50000)

Δv = 4500 ⋅ ln(4)

Δv = 4500 ⋅ 1.386 ≈ 6237 m/s

This provides a comprehensive mathematical overview of the kinetics and performance of LH2 propellant in rockets.

[A Dynamical Model of Rocket Propellant Loading with Liquid Hydrogen Michael D. Watson6 NASA Marshall Space Flight Center, Huntsville, AL, 35805 USA]

6. Unsymmetrical Dimethylhydrazine (UDMH) Fuel Overview:

UDMH, or unsymmetrical dimethylhydrazine, is a type of hypergolic propellant used in rocket engines. It reacts spontaneously with oxidizers like nitrogen tetroxide, meaning it ignites on contact without the need for an external ignition source. UDMH is known for its high energy density and has been used in various space missions, but it’s also quite toxic and requires careful handling.

9.1 PROPERTIES
1. Basic Properties:

   - Chemical Formula: C2H8N2

   - Molecular Weight: 60.10 g/mol

   - Density: Approximately 0.82 g/cm³ at 20°C

   - Boiling Point: 64.7°C (148.5°F)

   - Freezing Point: -2.1°C (28.2°F)

   - Appearance: Colorless liquid

   - Odor: Ammonia-like
2. Chemical Properties:

   - Hypergolic: Ignites spontaneously on contact with oxidizers (e.g., nitrogen tetroxide).

   - Toxicity: Highly toxic and carcinogenic; requires strict handling precautions.

3. Usage:

   - Propellant: Commonly used in rocket engines as a fuel in combination with oxidizers like nitrogen tetroxide.

   - Applications: Employed in spacecraft and satellite launch vehicles; used in space programs like the Apollo Lunar Module and various military and commercial rockets.

4. Performance Characteristics:

   - Specific Impulse (ISP): Typically ranges from 280 to 320 seconds in vacuum, depending on the specific engine design and mixture ratios.

   - Energy Density: Approximately 10.5 MJ/kg

5. Handling and Safety:

   - Storage: Stored in pressurized tanks or cryogenic conditions to keep it in liquid form.

   - Safety Measures: Requires proper ventilation, protective clothing, and emergency protocols due to its toxicity and corrosiveness.

6. Environmental Impact:

   - Decomposition Products: Can produce toxic compounds like nitrogen oxides and ammonia when burned or decomposed.

7. Historical and Notable Uses:

   - Rockets: Used in various spacecraft propulsion systems and rocket engines, including the Titan II and Saturn IIB rockets.

8. Regulations and Safety Guidelines:

   - Handling: Follow stringent guidelines for handling and disposal due to its hazardous nature.

   - Transport: Transported under strict regulations due to its toxicity and reactivity.

9.2 Synthesis of Unsymmetrical Dimethylhydrazine (UDMH):

UDMH (Unsymmetrical Dimethylhydrazine) is synthesized through the chemical reaction of dimethylamine and chloramine. Here's a brief overview of the synthesis process:

1. Reactants:

   - Dimethylamine (DMA): (CH3)2NH

   - Chloramine (NH2Cl)

2. Reaction:

   - The reaction involves the nucleophilic substitution of chloramine with dimethylamine.

   - The general reaction can be represented as:

     (CH3)2NH + NH2Cl -> (CH3)2NNH2 + HCl

3. Process:

   - Dimethylamine is reacted with chloramine in a controlled environment.

   - The reaction produces UDMH and hydrochloric acid (HCl) as a byproduct.

4. Purification:

The crude UDMH is then purified through distillation or other separation techniques to remove impurities and byproducts.

   - The final product is a high-purity UDMH suitable for use as rocket fuel.

5. Safety Considerations:

   - The synthesis of UDMH involves handling toxic and potentially hazardous chemicals, requiring strict safety protocols.

   - Appropriate protective equipment and ventilation are necessary to prevent exposure to harmful substances.

By carefully controlling the reaction conditions and purification processes, high-quality UDMH can be synthesized for use in various aerospace applications.

[Chemosphere Volume 29, Issue 7, Oxidation Of 1,1-Dimethylhydrazine (Udmh) In Aqueous Solution With Air And Hydrogen Peroxide

CONCLUSION

The exploration of liquid cryogenic fuels and oxidizers has highlighted their pivotal role in advancing modern propulsion systems. These fuels, particularly liquid hydrogen, and oxidizers such as liquid oxygen and nitrogen tetroxide, have demonstrated exceptional performance characteristics that make them indispensable for space exploration and related industries. Their high specific impulse, efficiency, and compatibility with various engine designs underscore their superiority in achieving precise propulsion requirements. However, the handling and storage challenges posed by cryogenic fuels, coupled with the toxicity of certain oxidizers like UDMH, necessitate continuous innovation in safety protocols and material advancements. The development of hybrid propellants and the integration of renewable energy sources into fuel synthesis processes hold promising potential for enhancing sustainability and reducing environmental impact. As private and governmental aerospace ventures continue to evolve, the demand for highly efficient and versatile fuels will undoubtedly shape the trajectory of future technological advancements. This comprehensive review underscores the importance of interdisciplinary collaboration and sustained research efforts in overcoming the challenges associated with liquid cryogenic and solid fuels, paving the way for breakthroughs that will redefine the scope of aerospace propulsion and energy systems

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