Learn Military and Aerospace Propulsion Explained Simply – Turbofans (Brayton Cycle, Bypass Ratio, Low/High-Bypass), Turbojets, Turboprops, Rocket Engines (Propellants, Nozzles, Cooling), Ramjets & Scramjets, Naval Propulsion (Gas Turbine, Nuclear, Diesel Electric) and Land Vehicle Propulsion

Introduction:

Military propulsion systems form the engineering foundation of modern defense capabilities across aerospace, naval, and land domains. Each platform whether a supersonic fighter, a nuclear submarine, or a main battle tank demands a propulsion solution uniquely optimized for its operational environment, speed requirements, endurance needs, and mission profile. Military propulsion systems encompass an extraordinarily diverse range of technologies each optimized for specific operational requirements. Airbreathing engines dominate aerospace applications with turbofans providing the best balance of efficiency and performance for most aircraft, while turbojets, ramjets and scramjets serve specialized high speed roles from subsonic cruise to hypersonic flight. Rockets enable space access and ballistic missile capabilities with their unique ability to operate in vacuum. Naval propulsion has evolved from direct mechanical systems to sophisticated electrical architectures incorporating diesel electric and integrated electric propulsion, while nuclear power provides unmatched endurance for submarines and aircraft carriers. Land vehicle propulsion remains dominated by internal combustion engines though hybrid and turbine technologies continue to find niche applications in main battle tanks and specialized vehicles.

The ongoing development across all these domains reflects the fundamental importance of propulsion to military capability. Each new generation of engines pushes the boundaries of efficiency, power and operational flexibility, driven by advances in materials science, thermodynamics, aerodynamics and control systems. From the high temperature superalloys enabling hotter turbine operation to the sophisticated cooling techniques that allow rocket engines to survive extreme combustion temperatures, every component and system represents a triumph of engineering applied to the demanding requirements of military service. This article provides a detailed examination of the major propulsion technologies used across military domains, emphasizing the underlying engineering principles, thermodynamic cycles, and design challenges, while reducing historical narrative to focus on how these systems actually work.

Turbofan Engines:

The turbofan, also known as a fanjet, is the dominant engine type in modern aviation and represents a sophisticated evolution of the basic gas turbine. Its fundamental purpose is to improve fuel efficiency by moving a larger mass of air at a lower velocity rather than accelerating a smaller mass to extremely high speeds.

Principles of Operation and the Brayton Cycle:

Turbofans operate on the Brayton thermodynamic cycle which consists of four processes: isentropic compression, constant pressure heat addition, isentropic expansion and constant pressure heat rejection. In the turbofan, incoming air is divided into two paths. A portion enters the engine core passing through the compressor, combustion chamber and turbine. The remainder bypasses the core and is accelerated solely by the fan. The ratio of bypass airflow to core airflow is the bypass ratio, a metric that fundamentally determines engine characteristics.

High-bypass turbofans, where fan thrust dominates, are standard in commercial and military transport aviation. These engines achieve remarkable fuel efficiency and relatively low noise signatures, making them ideal for sustained subsonic cruise. Low-bypass turbofans, producing more thrust from the jet exhaust relative to the fan, are typically found in military fighter aircraft where supersonic capability and afterburning take priority over fuel economy.

Thermodynamic Efficiency and the Role of Bypass:

The turbofan was conceived to address inherent inefficiencies in the turbojet. For aircraft speeds below approximately 500 miles per hour, a turbojet suffers from two major penalties. First, the energy required for a given thrust increases as exhaust velocity rises so efficiency improves when larger quantities of air are propelled at lower speeds. Second, attempts to improve thermal efficiency by raising pressure ratios and turbine temperatures produce higher exhaust gas speeds, which actually worsen propulsive efficiency. The turbofan overcomes these limitations by using an additional turbine to drive a ducted fan, transferring energy from the core to the bypass stream. This allows the engine to produce the same thrust with reduced specific fuel consumption.

The working substance of the thermodynamic cycle is the mass accelerated to produce thrust. For subsonic flight, the speed of the propelling jet must be reduced because there is a penalty in producing thrust. The energy required to accelerate the gas inside the engine is expended in two ways: producing a change in momentum and producing a wake, an unavoidable consequence of thrust generation by an airbreathing engine. The wake velocity and fuel burned to produce it can be reduced while maintaining thrust by increasing the mass accelerated. A turbofan achieves this by transferring energy from the gas generator to a ducted fan which produces an additional mass of accelerated air.

The transfer of energy from the core to bypass air results in lower pressure and temperature gas entering the core nozzle, producing lower exhaust velocity and fan produced higher pressure and temperature bypass air entering the fan nozzle. The amount of energy transferred depends on the fan pressure ratio. The best energy exchange, yielding the lowest fuel consumption depends on losses in the fan turbine and fan.

The fan flow has lower exhaust velocity, providing much more thrust per unit energy lower specific thrust. Both airstreams contribute to gross thrust. The additional air for the bypass stream increases ram drag in the intake stream tube but there remains a significant increase in net thrust. The overall effective exhaust velocity of the two jets can be made closer to a normal subsonic aircraft's flight speed, approaching the ideal Froude efficiency.

Modern turbofans may feature either a large single stage fan or a smaller fan with multiple stages. An early configuration combined a low pressure turbine and fan in a single rear mounted unit. The low pressure turbine driving the fan is more robust than that of a comparable turbojet, often containing additional stages to handle the increased workload. Some advanced designs incorporate a gearbox between the fan and the low pressure turbine allowing each to operate at its optimal rotational speed.

Low Bypass Turbofans:

Low bypass turbofans typically employ a multi stage fan with inlet guide vanes, developing a relatively high pressure ratio and yielding high exhaust velocity. These engines require sufficient core airflow to generate the power needed to drive the fan. Achieving a smaller core flow and higher bypass ratio requires raising the inlet temperature of the high pressure turbine. Advances in turbine cooling and materials technology enable higher turbine inlet temperatures which are necessary to compensate for increased cooling air temperatures resulting from higher overall pressure ratios.

The resulting turbofan assuming reasonable efficiencies and duct losses, typically operates at a higher nozzle pressure ratio than a comparable turbojet though with lower exhaust temperature to maintain net thrust. Since the overall temperature rise through the engine is lower, fuel flow at dry power settings is reduced, yielding better specific fuel consumption.

Since the 1970s, most jet fighter engines have been low or medium bypass turbofans equipped with afterburners and variable area exhaust nozzles. The afterburner is a combustor located after the turbine blades and bypass duct in the jetpipe. It has its own fuel system with fuel nozzle rings, flame holders, igniters, and fuel control that ensures the correct nozzle area for the amount of fuel being burned. When activated, it raises turbine exhaust gas temperature, increasing exhaust velocity and thrust. The exhaust nozzle must open wider to accommodate the larger exhaust volume and prevent excessive back pressure. Afterburning provides a significant thrust boost for takeoff, transonic acceleration and combat maneuvers, but consumes fuel very rapidly.

In terms of global examples, the United States fields the Pratt & Whitney F119 for the F 22 and F135 for the F 35, alongside the General Electric F110. Russia relies on the Saturn AL 31 for its Su 27 and Su-30 series and the AL 41 for the Su 57. The United Kingdom contributes the Eurojet EJ200 for the multinational Eurofighter. France employs the Snecma M88 for the Rafale. China has developed the WS 10 for its J 10 and J 20 fighters, along with the WS 15 and WS 18 for transports. India’s GTRE GTX 35VS Kaveri remains under development, while Japan has produced the IHI XF9 1 for its future F 3 fighter and the experimental XF5 1.

High Bypass Turbofans:

High bypass turbofans evolved from low bypass designs used in military aircraft during the 1960s. By replacing the multi stage fan with a single stage unit and scaling it appropriately, engineers achieve lower specific thrust and improved fuel economy. Unlike many military engines, modern civil turbofans lack stationary inlet guide vanes in front of the fan rotor. The core engine must generate sufficient power to drive the fan at its rated mass flow and pressure ratio. Improvements in turbine cooling and materials allow higher turbine rotor inlet temperatures, enabling smaller and lighter cores that improve thermal efficiency. Reducing core mass flow tends to increase load on the low pressure turbine, potentially requiring additional stages to maintain efficiency. Reducing core flow also increases bypass ratio, with ratios exceeding 5:1 becoming increasingly common. The Pratt & Whitney PW1000G, which entered commercial service in 2016 attains a bypass ratio of 12.5:1.

Further improvements in core thermal efficiency come from raising overall pressure ratio. Advances in blade aerodynamics can reduce the number of required compressor stages while variable geometry stators enable high pressure ratio compressors to operate without surge across all throttle settings.

The first experimental high bypass turbofan was the AVCO Lycoming PLF1A-2, derived from the Honeywell T55 turboshaft engine, first run in February 1962. It featured a 40-inch diameter geared fan stage, produced 4,320 pounds of static thrust, and had a bypass ratio of 6:1. The General Electric TF39 became the first production model, designed for the Lockheed C-5 Galaxy military transport. The civil General Electric CF6 used a derived design. Other high-bypass turbofans include the Pratt & Whitney JT9D, the three shaft Rolls Royce RB211, CFM International CFM56, and the smaller TF34. More recent large high bypass turbofans include the Pratt & Whitney PW4000, the three shaft Rolls Royce Trent, General Electric GE90 and GEnx, and the GP7000 produced jointly by GE and P&W.

The global turbofan market is dominated by General Electric, Rolls Royce and Pratt & Whitney. General Electric and Safran of France operate a joint venture called CFM International, while Pratt & Whitney partners with Japanese Aero Engine Corporation and MTU Aero Engines of Germany in International Aero Engines. Pratt & Whitney and General Electric also collaborate through Engine Alliance.

The lower the specific thrust of a turbofan, the lower the mean jet outlet velocity, which translates to a high thrust lapse rate, decreasing thrust with increasing flight speed. Consequently, an engine sized for high subsonic cruise generates relatively high thrust at low speeds, enhancing runway performance. High bypass turbofans on twin engined transport aircraft produce enough take off thrust to continue take-off on one engine after a critical point in the take off run.

Turbojet Engines:

The turbojet is the original gas turbine jet engine design taking in air compressing it, mixing it with fuel, burning the mixture and expelling hot exhaust through a nozzle to create thrust. Unlike turbofans, 100 percent of the air entering the intake passes directly through the engine core.

Components and Their Functions:

The main components of a turbojet are the inlet, compressor, combustion chamber, turbine and nozzle. The inlet scoops in air as the aircraft moves forward. The compressor uses fan blades to squeeze the air to high pressure. In the combustion chamber, fuel mixes with pressurized air and ignites, expanding the gas. The turbine extracts power from the hot rushing gas to spin the compressor via a central shaft. The nozzle constricts and accelerates the gas to high speed generating thrust.

The compressor may be axial, centrifugal or mixed flow with early designs often using centrifugal compressors while modern designs favor axial for smaller diameter and higher efficiency. Compressor types used in turbojets typically had low pressure ratios up to about 5:1. Aerodynamic improvements including splitting the compressor into two separately rotating parts, incorporating variable blade angles for entry guide vanes and stators and bleeding air from the compressor enabled later turbojets to have overall pressure ratios of 15:1 or more.

The combustion chamber is where the air fuel mixture burns in a continuous flowing process with no pressure build up. Instead, a small pressure loss occurs in the combustor. The fuel air mixture can only burn in slow moving air, so an area of reverse flow is maintained by the fuel nozzles for approximately stoichiometric burning in the primary zone. Further compressed air is introduced which completes the combustion process and reduces the temperature of the combustion products to a level the turbine can accept. Less than 25 percent of the air is typically used for combustion, as an overall lean mixture is required to stay within turbine temperature limits.

The turbine is where hot gases leaving the combustor expand. Typical materials for turbines include Inconel and Nimonic. The hottest turbine vanes and blades have internal cooling passages with air from the compressor passed through these to keep metal temperature within limits. In the first stage, the turbine is largely an impulse turbine that rotates because of the impact of the hot gas stream. Later stages are convergent ducts that accelerate the gas. Energy is transferred into the shaft through momentum exchange in the opposite way to energy transfer in the compressor. The power developed by the turbine drives the compressor and accessories.

The propelling nozzle after the turbine expands the gases producing a high velocity jet. In a convergent nozzle, the ducting narrows progressively to a throat. The nozzle pressure ratio on a turbojet is high enough at higher thrust settings to cause the nozzle to choke. If a convergent divergent de Laval nozzle is fitted, the divergent section allows the gases to reach supersonic velocity, generating additional thrust.

Turbojets supply bleed air from the compressor to the aircraft for environmental control systems, anti icing and fuel tank pressurization. The engine itself needs air at various pressures and flow rates to keep it running, preventing turbine overheating, lubricating oil leakage and ice formation. Bleed air decreases engine efficiency since it has been compressed but does not contribute to producing thrust.

Turbojets have poor efficiency at low vehicle speeds, limiting their usefulness outside aircraft applications. They were widely used for early supersonic fighters through the third generation. Since most fighters spend little time traveling supersonically, fourth generation fighters and subsequent designs are powered by more efficient low bypass turbofans with afterburners. Turbojets remain common in medium range cruise missiles due to their high exhaust speed, small frontal area, and relative simplicity.

In terms of global examples, the United States produced the Pratt & Whitney J57 for the B-52, J79 for the F-4, and GE J85 for the F-5. Russia developed the Tumansky R-15 for the MiG-25 and Klimov RD-9 for the MiG-19. The United Kingdom produced the Rolls-Royce Avon for the Hawker Hunter and jointly developed the Olympus 593 for Concorde with France. France’s SNECMA Atar powered the Mirage III and IV. China developed the WP-7 and WP-13 for its fighters. India has delivered its first fully indigenous 350 kg thrust class expendable turbojet engine, designed by the Gas Turbine Research Establishment under the DRDO, for use in tactical cruise missiles and long-range unmanned drones. Japan produced the Ishikawajima Ne-20 during World War II, its first turbojet.

Turboprop Engines:

A turboprop engine uses a gas turbine to drive a propeller. Instead of using hot exhaust gases directly for thrust, the turbine extracts most of the energy to spin a reduction gearbox that drives the propeller which creates nearly all of the forward thrust.

Operating Principles and Gearbox Design:

The operating principle follows the same basic gas turbine cycle: air enters through an intake, is compressed, mixed with fuel and burned and the expanding gases spin turbine wheels. The turbine powers both the compressor and a reduction gearbox that slows the speed for efficient propeller operation. The exhaust jet produces only about 10 percent of the total thrust with the propeller providing the remainder.

Turboprops have bypass ratios of 50 to 100, though the propulsion airflow is less clearly defined for propellers than for fans. The propeller is coupled to the turbine through a reduction gear that converts high RPM and low torque output to low RPM and high torque. Two primary designs exist: free-turbine and fixed. A free turbine design, found on the Pratt & Whitney Canada PT6, has only the turbine connected to the propeller while the gas generator spins independently. This allows propeller damage to occur without harming the gas generator and reduces stress during engine ground starts. A fixed shaft engine connects the gearbox, turbine and gas generator on a shared shaft, as on the Honeywell TPE331.

The propeller itself is normally a constant speed, variable pitch type. Due to the turbine engine's slow response to power inputs, particularly at low speeds, the propeller has a greater range of selected travel to make rapid thrust changes for taxi, reverse and ground operations. The propeller has two modes: Alpha for all flight operations including takeoff and Beta for all ground operations aside from takeoff. Beta mode ranges from zero to negative thrust, used for taxi and backing operations. Feathering of these propellers is performed by the propeller control lever.

The constant speed propeller is distinguished from the reciprocating engine constant speed propeller by the control system. The turboprop system consists of three propeller governors: a governor, an overspeed governor and a fuel topping governor. To make the engine more compact, reverse airflow can be used, where the compressor intake is at the aft of the engine and the exhaust is situated forward, reducing the distance between the turbine and the propeller.

Efficiency and Applications:

Unlike the small diameter fans of turbofans, the propeller has a large diameter that accelerates a large volume of air. Since it is more efficient at low speeds to accelerate a large amount of air by a small degree than a small amount by a large degree, low disc loading increases energy efficiency and reduces fuel consumption. Propellers work efficiently until blade tip speeds approach the speed of sound, beyond which efficiency drops dramatically. For this reason, turboprops are not commonly used on aircraft flying faster than Mach 0.6 to 0.7, with some exceptions such as the Tupolev Tu-95.

Modern turboprop airliners operate at nearly the same speed as small regional jets but burn two  thirds of the fuel per passenger. Compared to piston engines, their greater power to weight ratio and reliability offset higher initial cost, maintenance and fuel consumption. Since jet fuel is more readily available than avgas in remote areas, turboprop aircraft are used as bush airplanes.

Examples of turboprop engines include the US Allison T56 powering the C-130 and Honeywell TPE331 in various applications. Russia produces the Kuznetsov NK-12 for the Tu-95 and the TV3-117 for helicopters. The UK developed the Rolls-Royce Dart and Tyne. France offers the Turbomeca Ardiden for helicopters and light aircraft. China has the WJ-6 for its Y-8 and Y-9 transports and the WJ-9 for turboprop trainers. India has experimental HTSE-1200 and proposed HPT-700 engines, while Japan relies mostly on imported turboshafts for this category.

Rocket Engines:

Rocket engines differ fundamentally from airbreathing engines in that they carry both fuel and oxidizer on board. This allows them to operate in the vacuum of space and achieve very high speeds. A rocket engine produces thrust by ejecting reaction mass rearward, typically a high speed jet of high temperature gas produced by combustion of rocket propellant stored inside the vehicle.

Propellant Types and Combustion:

Chemical rockets are classified by propellant state. Solid-fuel rockets use propellant in solid state, stored within the combustion chamber itself. Liquid propellant rockets use liquid propellants fed from tanks by pumps. Hybrid rockets use solid propellant in the combustion chamber with a liquid or gaseous oxidizer added. Monopropellant rockets use a single propellant decomposed by a catalyst with hydrazine and hydrogen peroxide being common examples.

Liquid fueled rockets force separate fuel and oxidizer into the combustion chamber where they mix and burn. Both liquid and hybrid rockets use injectors to introduce propellant, often with jets that cause the propellants to collide, breaking up the flow into smaller droplets that burn more easily. In order for fuel and oxidizer to flow into the chamber, the pressure of the propellants entering the combustion chamber must exceed the pressure inside the combustion chamber itself. This may be accomplished by turbopumps or in simpler engines, via sufficient tank pressure.

For chemical rockets, the combustion chamber is typically cylindrical. The dimensions are such that propellant combusts thoroughly with different propellants requiring different chamber sizes. Temperatures and pressures are extreme compared to non afterburning airbreathing engines. No atmospheric nitrogen dilutes and cools the combustion, so propellant mixtures can reach true stoichiometric ratios. Heat conduction through chamber walls is very high. To achieve practical thermal efficiency, rocket combustion chambers reach temperatures that can approach 3,300°C, substantially exceeding the melting points of nozzle and combustion chamber materials.

Nozzle Design and Expansion:

The most commonly used nozzle is the de Laval nozzle, a fixed geometry design with a high expansion ratio. The large bell or cone shape beyond the throat gives rocket engines their characteristic appearance. Rocket thrust comes partly from unbalanced pressures inside the combustion chamber and partly from pressures acting against the inside of the nozzle. As gas expands, pressure against the nozzle walls forces the engine one direction while accelerating gas the other.

For optimal performance, exhaust pressure at the nozzle exit should equal ambient pressure. If exit pressure is lower than ambient, the vehicle is slowed by the pressure difference. If exit pressure is higher, energy is wasted. To maintain ideal equality as altitude changes, nozzle diameter would need to increase with altitude, which is difficult to arrange in a lightweight fashion. Exotic designs like plug nozzles, stepped nozzles, expanding nozzles and aerospikes have been proposed to adapt to changing ambient pressure.

The four expansion regimes of a de Laval nozzle are under expanded, perfectly expanded, over expanded, and grossly over expanded. In practice, perfect expansion is only achievable with a variable exit area nozzle. If the nozzle is not perfectly expanded, loss of efficiency occurs. Grossly over expanded nozzles lose less efficiency but can cause mechanical problems.

Propellant efficiency depends on heating the propellant to the highest possible temperature, using low density gases, and using propellants that form simple molecules with few degrees of freedom. Exhaust speed is an excellent measure of engine propellant efficiency. For aerodynamic reasons, flow goes sonic at the nozzle throat. Since speed of sound in gases increases with temperature, hot exhaust gas greatly improves performance. Speed of sound in rocket exhaust can exceed 1,700 meters per second, compared to 340 meters per second at room temperature in air. Rocket propellants are chosen for low molecular mass, further increasing velocity.

Expansion in the nozzle multiplies speed by 1.5 to 2 times, producing a highly collimated hypersonic exhaust jet. The speed increase is determined by area expansion ratio and gas properties. Larger ratio nozzles are more massive but extract more heat from combustion gases, increasing exhaust velocity.

Thrust vectoring is achieved through several methods: mounting the entire engine on a gimbal, gimbaling only the combustion chamber and nozzle using multiple throttled engines canted at slight angles, or using high temperature vanes that protrude into the exhaust.

Cooling Methods:

Rocket engines use various cooling methods. Ablative cooling lines the combustion chamber with material that carries heat away as it vaporizes. Radiative cooling uses refractory materials that glow red or white hot, radiating heat away. Dump cooling passes cryogenic propellant around the nozzle and discards it, wasting propellant. Regenerative cooling routes fuel around the nozzle before injection, the most widely applied method. Film cooling injects additional propellant through orifices in the chamber wall, cooling as it evaporates. Transpiration cooling, a more efficient subtype, passes propellant through a porous inner wall.

Regenerative cooling creates a boundary layer in the coolant channels. This boundary layer acts as insulation, so its thickness must be minimized by maximizing coolant velocity. Liquid fueled engines are often run fuel rich to lower combustion temperatures reducing heat loads and allowing lower cost materials and simplified cooling systems. This can also increase performance by lowering exhaust molecular weight and improving conversion efficiency.

Global rocket engine examples include the US RS-25 for the Space Shuttle, F-1 for Saturn V, Raptor for Starship, and RL10. Russia produces the RD-180 for Atlas, NK-33 for Soyuz, and RD-107 for Soyuz. France developed the Vulcain for Ariane 5 and HM7B for Ariane upper stages. China has the YF-100 and YF-77 for its Long March rockets. India has the CE-20 for GSLV and Vikas for PSLV. Japan developed the LE-7 for H-II and LE-9 for H-III.

Ramjets and Scramjets:

Ramjet Principles:

A ramjet is an airbreathing jet engine that requires forward motion to provide air for combustion. It uses the vehicle's speed to compress incoming air, eliminating the need for the rotating compressor found in turbojets. Ramjets work most efficiently at supersonic speeds around Mach 3 and can operate up to Mach 6.

The first part of a ramjet is its diffuser, where forward motion raises air pressure as required for combustion. Air is compressed, heated by combustion, and expanded in the Brayton cycle before passing through a nozzle to accelerate to supersonic speeds and generate thrust. The diffuser converts high velocity incoming air into high static pressure required for combustion. High combustion pressures minimize entropy rise during heat addition, reducing wasted thermal energy. Subsonic and low supersonic ramjets use a pitot type opening followed by a widening internal passage. At higher supersonic speeds, a protruding spike or cone produces oblique shock waves in front of a final normal shock at the inlet entrance.

The combustor raises air temperature by burning fuel with a small pressure loss. Air velocity entering the combustor must be low enough for continuous combustion in sheltered zones provided by flame holders. A ramjet combustor can safely operate at stoichiometric fuel air ratios producing combustion exit temperatures around 2,400 K for kerosene. The combustor must operate over a wide range of throttle settings matching flight speeds and altitudes.

The propelling nozzle accelerates exhaust flow to produce thrust. Subsonic ramjets use a convergent nozzle, while supersonic flight typically requires a convergent-divergent nozzle.

Ramjets give little thrust below about Mach 0.5 and are highly inefficient due to low pressure ratios. Above this speed, given sufficient initial velocity, a ramjet is self sustaining. Unless vehicle drag is extremely high, the engine tends to accelerate to higher speeds increasing air intake temperature. The fuel control system must reduce flow to stabilize speed and temperature.

Ramjets can be classified by fuel type: liquid or solid. Liquid fuel ramjets inject hydrocarbon fuel into the combustor ahead of a flame holder, requiring a means of pressurizing and supplying fuel. Solid fuel ramjets cast fuel along the outer wall of the combustor, with ablation providing fuel injection. Integrated boosters package booster propellant inside the combustor, with separate nozzles ejected after burnout.

Ramjets are particularly appropriate for applications requiring a compact high speed mechanism, such as missiles. Weapons designers are investigating ramjet technology for artillery shells to increase range; a 120mm ramjet assisted mortar shell is thought to be able to travel 35 kilometers.

Global ramjet examples include the US SM-6 Standard Missile, Russia’s Kh-31 and 3M-54 Kalibr variants, the UK’s retired Bristol Thor for the Bloodhound SAM, France’s ASMP nuclear missile and the joint UK-France Meteor air-to-air missile, China’s YJ-12 anti-ship missile, India’s joint Russia BrahMos, and Japan’s ASM-3 anti-ship missile.

Scramjet Principles:

The scramjet supersonic combustion ramjet is an advanced airbreathing engine that operates at hypersonic speeds starting around Mach 5. Like a ramjet, it uses forward motion to compress air, but unlike a ramjet, it maintains supersonic airflow through the combustor for fuel combustion.

Scramjets have no moving parts. They lack fans, turbines or rotating compressors found in normal jet engines. Air compression is achieved by the extreme forward speed forcing and compressing incoming atmospheric air through the engine inlet. Fuel, often hydrogen is injected and ignited while airflow remains supersonic. Expanding hot exhaust gases push backward through a diverging nozzle to generate immense forward thrust.

The scramjet consists of three basic components: a converging inlet where incoming air is compressed, a combustor where fuel burns with atmospheric oxygen and a diverging nozzle where heated air accelerates to produce thrust. No rotating components are needed. However, scramjet operation is limited to near hypersonic velocities requiring a vehicle to be accelerated to about Mach 4 by some other means of propulsion before the scramjet can function.

The scramjet was developed to overcome the performance limitations of ramjets at very high speeds. In a typical ramjet, supersonic inflow is decelerated to subsonic speeds at the inlet and then reaccelerated through the nozzle. This deceleration, produced by a normal shock creates total pressure loss that limits the upper operating point. By keeping the flow supersonic through the combustor, the scramjet avoids this pressure loss and can operate efficiently at much higher speeds.

Scramjets are designed to operate in the hypersonic flight regime, beyond the reach of turbojet engines. Turbomachinery based engines become increasingly inefficient at transonic speeds because compressor rotors require subsonic speeds to operate. Around Mach 3 to 4, turbomachinery is no longer useful, and ram style compression becomes preferred.

The technical challenges are considerable. The minimum Mach number at which a scramjet can operate is limited by the need for compressed flow to be hot enough to burn fuel and have high enough pressure for the reaction to finish before exiting the engine. The compressed flow must also remain supersonic after combustion. If gas within a scramjet goes below Mach 1, the engine chokes, transitioning to subsonic combustion. Additionally, heating from combustion causes the speed of sound to increase and Mach number to decrease even though gas speed remains constant, a phenomenon called thermal choking. An isolator between the inlet and combustion chamber is often included to improve flow homogeneity and extend the operating range.

Global scramjet examples include the US X-43A and X-51A Waverider, Russia’s GLL-8 experimental tests, the UK’s Reaction Engines SABRE precooled concept, France’s ONERA and MBDA experimental tests, China’s Star-1 and other experimental flights, India’s 2016 ISRO scramjet test on an RH-560 sounding rocket, and Japan’s JAXA scramjet experiments.

Gas Turbine Engines:

Gas turbine engines are continuous flow internal combustion machines that draw in ambient air, compress it, mix it with fuel and ignite it to produce high pressure, high temperature gas. This expanding gas spins turbine blades to create rotational mechanical power or high speed jet thrust. They operate on the Brayton cycle and are used in aircraft propulsion, electric power generation, marine propulsion, industrial machinery, pumps, gas compressors and some land vehicles.

In the ideal Brayton cycle, the working fluid undergoes isentropic compression, constant pressure heat addition, isentropic expansion, and constant pressure heat rejection. In a real gas turbine, shaft work supplied to the compressor raises pressure and temperature of incoming air. Heat is added in the combustor by burning fuel. The resulting high temperature gas expands through the turbine, producing work to drive the compressor and additional shaft power or jet thrust. Real compressors, combustors and turbines depart from the ideal cycle due to friction, turbulence, pressure losses, heat transfer and other irreversible effects.

The compressor, combustor and compressor driving turbine together form the core of the engine, also known as the gas generator. If the engine has an additional turbine to drive an industrial generator, helicopter rotor, marine propeller, fan or other load, a larger share of flow energy is extracted as shaft work. In some designs, this turbine is mechanically independent of the gas generator and is called a free power turbine.

The smaller the engine, the higher the rotation rate of the shaft must be to attain required blade tip speed. Blade tip speed affects pressure ratios obtainable by turbine and compressor, along with blade aerodynamics, stage count, material limits and component efficiency. Large jet engines operate around 10,000 to 25,000 RPM, while micro turbines may spin as fast as 500,000 RPM.

Creep and Materials Technology:

A major challenge facing turbine design is reducing the creep induced by high temperatures and stresses during operation. Higher operating temperatures are sought to increase efficiency but come at the cost of higher creep rates. Several methods have been employed to achieve optimal performance while limiting creep, with the most successful being high performance coatings and single crystal superalloys. These technologies limit deformation through mechanisms that can be broadly classified as dislocation glide, dislocation climb, and diffusional flow.

Protective coatings provide thermal insulation of the blade and offer oxidation and corrosion resistance. Thermal barrier coatings are often stabilized zirconium dioxide based ceramics and oxidation resistant bond coats typically consist of aluminides or MCrAlY alloys. Using thermal barrier coatings limits the temperature exposure of the superalloy substrate, decreasing diffusivity and reducing creep. A coating of 1 to 200 micrometers can decrease blade temperatures by up to 200°C.

Nickel based superalloys boast improved strength and creep resistance due to their composition and resultant microstructure. The gamma FCC nickel is alloyed with aluminum and titanium to precipitate a uniform dispersion of coherent Ni3(Al,Ti) gamma-prime phases. These finely dispersed precipitates impede dislocation motion and introduce a threshold stress, increasing the stress required for the onset of creep. Further refractory elements such as rhenium and ruthenium can be added in solid solution to improve creep strength. The development of single crystal superalloys has led to significant improvements in creep resistance, as single crystals eliminate Coble creep and decrease the creep rate.

Marine Applications:

In naval applications, gas turbines are valued for high power to weight ratio and resulting acceleration and ability to get underway quickly. The first gas-turbine-powered naval vessel was the Royal Navy's motor gunboat MGB 2009, converted in 1947. The first large scale, partially gas turbine powered ships were the Royal Navy's Type 81 frigates with combined steam and gas powerplants. The Soviet Navy commissioned the first of 25 Kashin class destroyers in 1962 with four gas turbines in combined gas and gas propulsion, the first large ships in the world powered solely by gas turbines.

Global marine gas turbine examples include the US GE LM2500 for Arleigh Burke destroyers and LM6000 for FREMM frigates, Russia’s M8E for Kashin class vessels, the UK’s Rolls Royce Olympus for Type 42 and MT30 for Queen Elizabeth carriers, France’s limited use of imported LM2500, China’s QC 280 for Type 052D destroyers, India’s LM2500 for Kolkata class destroyers, and Japan’s LM2500 for Kongo and Maya-class destroyers.

Nuclear Marine Propulsion:

Nuclear marine propulsion uses heat from a nuclear reactor to create steam, which drives turbines that turn a ship's propeller or power an electric motor. This technology gives warships and icebreakers near limitless range without needing frequent refueling.

Pressurized Water Reactor Principles:

Most naval nuclear reactors are pressurized water types. The primary water circuit transfers heat from nuclear fission in the fuel to a steam generator. This water is kept under pressure so it does not boil operating at temperatures around 250 to 300°C. Any radioactive contamination in the primary water is confined. Water is circulated by pumps at lower power levels, reactors designed for submarines may rely on natural circulation to reduce pump noise.

The hot water from the reactor heats a separate water circuit in the steam generator. That water is converted to steam and passes through steam driers to the steam turbine. Spent steam at low pressure runs through a condenser cooled by seawater and returns to liquid form. The water is pumped back to the steam generator continuing the cycle.

In the turbine, steam expands and reduces pressure as it imparts energy to rotating blades. There may be many stages of rotating blades and fixed guide vanes. The output shaft may connect to a gearbox to reduce rotation speed, then to the vessel's propellers. In another form, the turbine turns an electrical generator and the electric power feeds drive motors for the propellers.

Design Differences from Land Reactors:

Marine type reactors differ from land based commercial power reactors in several respects. While land based reactors produce up to around 1600 megawatts net electrical power, a typical marine propulsion reactor produces no more than a few hundred megawatts. Tight space constraints dictate that a marine reactor must be physically small, generating higher power per unit of space. Components are subject to greater stresses, and mechanical systems must operate flawlessly under adverse sea conditions including vibration and pitching and rolling. Reactor shutdown mechanisms cannot rely on gravity to drop control rods into place as in a land based reactor. Salt water corrosion complicates maintenance.

As the core of a seagoing reactor is much smaller than a power reactor, the probability of a neutron intersecting with a fissionable nucleus before escaping is much lower. Fuel is typically more highly enriched containing a higher concentration of uranium 235 than used in land based plants. Some marine reactors run on relatively low enriched uranium, requiring more frequent refueling. Others run on highly enriched uranium, from 20% to over 96% uranium 235 in US submarines, resulting in a smaller, quieter core. Using more highly enriched fuel increases power density and extends fuel life but is more expensive and a greater proliferation risk.

A marine nuclear propulsion plant must be highly reliable and self sufficient, requiring minimal maintenance. Fuel elements must withstand large amounts of radiation damage. Marine reactors use metal zirconium alloy rather than ceramic uranium dioxide. They are designed for long core life, enabled by relatively high enrichment and incorporating burnable poison in fuel elements that is slowly depleted as fuel ages. Gradual dissipation of nuclear poison increases reactivity to compensate for aging fuel extending usable life. The compact reactor pressure vessel includes an internal neutron shield to reduce steel damage from constant neutron bombardment.

Nuclear power is particularly suitable for vessels needing to remain at sea for long periods without refueling or for powerful submarine propulsion. Over 160 ships are powered by more than 200 small nuclear reactors. The US Navy has accumulated over 6200 reactor years of accident free experience over more than 50 years. It operated 81 nuclear powered ships with 92 reactors in 2017. Russia built 248 nuclear submarines and five naval surface vessels plus nine icebreakers.

Global nuclear marine propulsion examples include the US A4W for Nimitz carriers, A1B for Ford carriers, S6G for Los Angeles submarines, and S9G for Virginia submarines. Russia uses OK 900 for icebreakers, VM 5 for submarines and RITM 200 for LK 60 icebreakers. The UK operates PWR1, PWR2 for Vanguard and Astute, and PWR3 for Dreadnought. France has K15 for Charles de Gaulle and Triomphant, and CAS48 for Rubis. China has unknown designations for Type 94 Jin and Type 95 Shang class submarines. India has an indigenous Arihant class reactor. Japan has no nuclear ships or submarines.

Diesel Electric and Integrated Electric Propulsion:

Diesel electric transmission is a system powered by diesel engines that generate electricity to power electric motors. A major advantage is that it avoids the need for a gearbox by converting mechanical force of the diesel engine into electrical energy through an alternator and using that energy to drive traction motors.

In marine applications, diesel electric propulsion emerged early in the 1900s but was confined to specialist niches. A much wider application is now enjoyed thanks to developments in AC drive technology and increasing interest in low emissions and propulsion plant redundancy. The first diesel motorship was also the first diesel electric ship, the Russian tanker Vandal from Branobel launched in 1903.

Most early submarines used a direct mechanical connection between the combustion engine and propeller. In a true diesel electric transmission arrangement, the propeller is always driven by electric motors, while diesel generators provide electric energy for charging batteries and driving motors. This mechanically isolates the noisy engine compartment from the outer pressure hull protecting the submarine from detection by reducing its acoustic signature.

Global diesel electric examples include the US Virginia class using backup diesels, Russia’s Kilo class submarines, the UK’s Astute class backup diesels, France’s Scorpène class for export, China’s Yuan class Type 039, India’s Kalvari class Scorpène and Japan’s Soryu and Taigei class with AIP.

Integrated electric propulsion also known as full electric propulsion is an arrangement where gas turbines or diesel generators generate electricity that powers electric motors turning propellers or waterjet impellors. It eliminates the need for clutches and reduces or eliminates the need for gearboxes. Eliminating the mechanical connection provides increased freedom in engine placement, acoustical decoupling making the ship less noisy and reduction in weight and volume. Reducing acoustic signature is particularly important for naval vessels.

A typical integrated electric propulsion arrangement on larger vessels includes both diesel generators and gas turbines. The advantages of gas turbines include much lower weight and smaller size than diesels of similar power with much less noise and vibration. However, gas turbines are efficient only at or near maximum power. Diesel generators have high efficiency over a wide range of power levels. Using both types allows for full range operational efficiency.

Global IEP examples include the US DDG 1000 Zumwalt, Russia’s future icebreakers, the UK’s Type 45 Daring class and Queen Elizabeth class carriers, France’s Mistral class with some electric drive, China’s future frigates and research vessels, India’s future planned vessels and Japan’s future support vessels.

Land Vehicle Propulsion:

Internal combustion engines remain the primary propulsion for most military land vehicles. Diesel engines are preferred for their fuel efficiency, durability, and the lower flammability of diesel fuel compared to petrol. Global examples include the US Detroit Diesel and Cummins in HMMWVs, Russia’s V84 and V92 in T72 and T90 tanks, the UK’s Perkins CV12 in Challenger 2, France’s hyperbar diesel in Leclerc, China’s Type 99 and Type 15 diesels, India’s MTU diesel in Arjun MBT and Japan’s Type 10 diesel.

Gas turbines have been used in some tanks, with the most prominent examples being the US M1 Abrams powered by the Honeywell AGT1500 and the Russian T 80 series powered by GTD 1000 and GTD 1250 turbines. Gas turbines in tanks are lighter and smaller than diesel engines at the same sustained power output, but models installed to date are less fuel efficient, especially at idle. Successive models of M1 have addressed this with battery packs or secondary generators. A turbine is theoretically more reliable and easier to maintain than a piston engine due to simpler construction with fewer moving parts, but turbine parts experience higher wear rates due to higher working speeds and are sensitive to dust and sand.

In railway applications, diesel electric technology first saw limited use in switcher locomotives in the 1920s and became popular because they greatly simplified power transmission to wheels and had greatly reduced maintenance requirements. Global examples include the US GE Evolution and EMD SD series locomotives, Russia’s TEP series, the UK’s Class 66 and 68, France’s SNCF locomotives, China’s HXN series, India’s WDP and WDG series and Japan’s DE series.

Hybrid electric vehicles couple a conventional internal combustion engine with one or more electric engines into a combined propulsion system. Modern HEVs use energy recovery technologies such as motor generator units and regenerative braking. Some varieties use an internal combustion engine to directly drive an electrical generator, which recharges batteries or directly powers electric traction motors. Many HEVs reduce idle emissions by shutting down the combustion engine at idle and restarting when needed. Some 70 years after Porsche's pioneering hybrid-drivetrain efforts, the United States Army's Future Combat System all used a hybrid electric drive, though these were put on hold in the 2010 Department of Defense budget.

Global HEV examples include the US Ford Escape Hybrid and Chevy Volt, Russia’s limited Yandex taxi hybrids, the UK’s Vauxhall and JLR hybrids, France’s Renault Captur and Peugeot 3008 hybrids, China’s BYD Tang and Qin, India’s Toyota Camry Hybrid and Maruti Suzuki Grand Vitara and Japan’s global leader Toyota Prius, Honda Insight and Nissan Note e-Power.

Specialized and experimental propulsion technologies continue to be researched globally. The US has explored tip jets for helicopter rotors and pulse detonation engines through AFRL. Russia, the UK, France, China, India and Japan all have active research programs in rotating detonation engines, aerospike engines, and air-augmented rockets. Microturbines are produced by Capstone in the US with domestic designs in Russia, China and Japan. Auxiliary power units are manufactured globally, with US Honeywell and Hamilton Sunstrand, Russia’s TA series, the UK’s Rolls Royce, France’s Safran Microturbo, China’s domestic APUs, India’s HAL and Japan’s IHI designs.