Turbopump

This article is about the rocket engine component. For turbine devices for producing a high vacuum, see Turbomolecular pump. For the vehicular air pump, see Turbocharger and Supercharger.

A turbopump is a fluid pump with two main components: a liquid pump driven by a gas turbine, usually both mounted on the same shaft, or sometimes geared together^[1]. They were initially developed in Germany in the early 1940s. The most common purpose of a turbopump is to produce a high-pressure fluid for feeding a combustion chamber. While other use cases exist, they are most commonly found in liquid rocket engines. Turbopump fed systems scale much more favorably in large rockets than pressure fed systems, which require increasingly thick and heavy tanks to supply high chamber pressures in the engines.

Part of an axial turbopump designed and built for the M-1 rocket engine

Cutaway of the turbopump used in the F1 rocket engine (Saturn V first stage)

There are two common types of pumps used in turbopumps: a centrifugal pump, where the pumping is done by throwing fluid outward at high speed, or an axial flow pump, where alternating rotating and static blades progressively raise the pressure of a fluid. Axial flow pumps have small diameters but give relatively modest pressure increases. Although multiple compression stages are needed, axial flow pumps work well with low density fluids. Centrifugal pumps are far more powerful for high-density fluids but require large diameters for low density fluids.

A common turbopump arrangement. Note there are many possible, and common, variations to this arrangement including mounting the turbine between the pumps, having a multistage turbine, separating the fuel pump and ox pumps onto dual shafts, etc.

Design Principles

Hydrodynamic Design

The pump side of turbopumps consist of impellers that spin at very high speeds (thousands of RPM) in order to pump liquid propellants^[2]. Impellers are mounted on a central shaft, which also has a turbine mounted to it (or in some cases geared off on a different shaft). The turbine supplies shaft power, which is then consumed by the impellers in order to impart energy to the liquid propellants. Impellers mostly impart this energy by accelerating the liquid to a high velocity. However the ultimate goal is not a fast liquid, but a high pressure one; so surrounding the impeller is either a volute or a diffuser - these are specially shaped housings to decelerate the flow which then consequently dramatically increases its pressure (via Bernoulli's principle). The liquid is then discharged to the rest of the rocket engine, or in some cases to a second impeller and volute/diffuser stage which increases the pressure even further.

Turbopumps on liquid rocket engines virtually always have inducers as well, upstream of the impellers. Inducers are spiral shaped pumping elements that serve to gently raise the pressure of the incoming fluid enough to prevent it cavitating when it reaches the impeller. In many cases the impeller and inducer are manufactured as a single component, with a gradual transition between the axial spiral and the radial blades.

Aerodynamic Design

The turbine side of turbopumps consist of one or more stages, where each stage has a stator and a rotor^[3]. Individual rotor discs in a turbine are more commonly referred to as wheels in the modern day. These turbines are virtually always of the axial type, because of the very high gas flow (volumetrically) needed to supply enough shaft power for a liquid rocket engine. Contrast this with turbochargers, which usually feature radial turbine designs because of their much lower gas flow.

Upstream of the turbine is the turbine manifold, which collects gas from whatever source that rocket engine's cycle has upstream of it, and then disperses it circumferentially along the rim of the turbine. It then flows from the manifold axially downwards to the stages of the turbine. Stators are typically bladed, though it is also quite common (where pressure drop is particularly high, as in gas generator cycles) to forgo blades and drill angled nozzles directly off of the manifold itself to then impinge on the turbine wheel.

Downstream of the turbine varies based on cycle - in closed cycles it leads to the main injector of the engine, where (depending on whether the turbine is fuel-rich or ox-rich), one of the propellants can be injected into the main combustion chamber as a gas which can be very advantageous for promoting propellant atomization and mixing^[4]. In open cycles it is dumped to atmosphere. This can either mean dumped overboard directly off the side of the engine, or it can also lead to a manifold on the rocket engine nozzle which then injects it in the main flowpath, far downstream of the throat where ambient pressure is much lower than the chamber^[5]. The purpose here is to provide extra film cooling to the nozzle, since the hot gas leaving the turbine is nevertheless much cooler than the gas in the main combustion chamber. the latter option is common in vacuum optimized open cycle engines because they have much larger nozzles (with correspondingly large areas that need cooling, often without a regen jacket at its furthest extremes). It is important to note that the dumped gas from the turbine can still provide a non-negligible portion of the engine thrust. For this reason even if it is dumped overboard directly, there will usually still be a housing and a mild converging-diverging nozzle downstream of the turbine to take full advantage of the extra thrust opportunity. There is also an opportunity to extract waste heat from the flow at this point via heat exchangers; useful for heating up repressurizing gas for the tanks, for example.

Cycle Design

Turbomachinery / engine cycle design looks very different in liquid rocket engines compared to air-breathing engines (turbojets) for essentially one main reason:

Turbine materials cannot survive combustion chamber temperatures.

Rocket engine cycles are all various workarounds to this fundamental problem.

The turbine is driven by high pressure gas. The exact source of this gas is the primary differentiator between the various rocket engine cycles. Air-breathing engines (turbojets and similar) mount their turbine downstream of the burner and take direct advantage of the full flow and pressure of the engine. Rocket engines have never been able to do this because their mixture ratios are much closer to stoichiometric (since oxidizer comes at a premium; it must be carried with the rocket) and thus the flame temperature in the combustion chamber is dramatically higher. They are so high that nearly all possible materials would melt, and even the few that do have very little structural strength left at these temperatures.

For this reason, rocket engine cycles are all various schemes to circumvent this and supply hot gas to the turbine that is nevertheless much cooler than the main combustion chamber gas. Gas generator and staged combustion cycles do this by mounting an entirely separate and smaller combustion chamber to the engine, termed the gas generator (whose gas is ultimately dumped overboard) or the preburner (whose "pre-burnt" gas eventually reaches the main combustion chamber after passing through the turbine). These smaller chambers run very far from stoichiometric, either with way too much fuel or way too much oxidizer. Hence you can have "fuel rich" and "ox rich" gas generator and staged combustion cycles. You could also have two preburners, one fuel rich and one ox rich, which is termed "full flow staged combustion".

Beyond these, there are also expander cycles, where liquid propellant is heated (usually fuel) in the regenerative cooling loop of the main combustion chamber, to the point of boiling, and then fed as gas to the turbine. The last major cycle is the tap off cycle, where a portion of the main combustion gas is "tapped off" and routed to the turbine. Because of the aforementioned temperature problem, tap off cycles require large dedicated heat exchangers to rapidly cool the re-routed gas before it reaches the turbine.

Mechanical Design

Simplified mechanical design of a turbopump rotor. The retaining bolts, when torqued, will put net tension in the shaft and net compression in the stackup of components that fully encase it. This preload increase the stiffness of the full rotor assembly.

The collection of all rotating components in a turbopump (i.e. the impellers, inducers, wheels, shaft, parts of the seals, and various spacers) are collectively known as the "rotor". The rotor is spinning at extreme angular velocities: shaft speeds in the tens of thousands of RPM are common. Nominally the only mechanical connection between the rotor and the rest of the turbopump is via the bearings^[6]. Most common by far are ball bearings, with some modern exceptions pivoting to hydrodynamic bearings. The goal of bearing selection is to minimize friction - both because high friction can wear out the bearing, and also because any frictional energy losses are dissipated as heat that must be carried away rapidly to not destroy the bearing. The extra challenge in turbopump design is that the local environment in the pumps is very often at cryogenic temperatures, which virtually all greases and oils normally used to lubricate bearings are not compatible with (they freeze). Therefore turbopump bearings do not use lubricants at all in the traditional sense. Rather they are installed as bare metal, and some amount of cold propellant is intentionally routed through them (i.e. where the balls are) to dissipate the heat generated by their friction. This bearing cooling circuit is a secondary flow that the hydrodynamic designer must also design in addition to the primary flow of the propellant through the inducer/impeller/volute.

Turbopumps can be very sensitive to the exact placement of components and the loads/stresses developed in them. Hydrodynamic considerations typically demand very tight clearances between the impellers / inducers and the pump housings, as well as aerodynamic considerations demanding tight clearances between turbine wheels and stators / manifolds. Furthermore, rotordynamics demands a high stiffness coupling of the rotating components with the shaft, especially when it comes to the turbine wheel.

These considerations and more demand high precision and high stiffness mechanical design. Bolted joints are generally the default method by which to join parts; some turbopumps have welded joints as well but require more careful consideration and analysis because of their generally lower stiffness, potential for thermally induced warpage of the parts during the welding process, as well as increased risk of fatigue over the life of the turbopump. In order for the rotor to act structurally as one rigid object, all of the components are stacked into one long stackup that envelops the entire shaft and then is preloaded onto it from both ends. This moderately loads the ball bearings, which are usually of the angular contact type, which increases their stiffness. Typically the preload is supplied one end by a bolt clamping onto the nose of the inducer and threaded into the end of the shaft below it. Depending on the exact configuration of the turbopump, the other end could be another inducer (for the other propellant), or a turbine wheel which will also have preloaded bolt(s) onto the end of the shaft.

Design of the shaft itself is driven by the need to carry high torque; the more torque it can carry the more power can be transferred from the turbine to the pump(s)^[7]. Shaft power is the product of shaft speed and shaft torque. This high torque requirement drives the designer to maximizing the polar moment of inertia of the shaft. It is not uncommon for shafts to be hollow, as this maximizes this polar moment of inertia for a given weight of material. Shaft also need to transfer torque to the components of the rotor stackup. This can be accomplished via keyways, which carry less torque but are easier to manufacture, splines which generally carry higher torque but more difficult to manufacture and/or shear pins, which are common for components attached to the circular face of the shaft (i.e. turbine wheels).

Seal Design

Turbopumps need to keep fuel and oxidizer apart from each other; otherwise there is high risk of ignition in the turbopump that will cascade into a total failure of the rocket engine. Secondarily they also need to keep propellants out of the turbine cavity; to avoid wastage and also to avoid changing the conditions of the gas flow through the turbine. They especially want to keep oxidizer out of a turbine running fuel-rich, and fuel out of a turbine running ox-rich. This is because leakage in this case would push the oxidizer/fuel (OF) ratio of the working gas closer to stoichiometric, increasing its flame temperature which may be too much for the turbine materials to handle. For these reasons turbopumps always have dynamic seals around their shafts, where one part of the seal is attached to the rotor and corotating with it, while the other part is statically attached to the housing^[8].

The dynamic seals in turbopumps have quite specialized requirements compared to seals in most systems. They must support very high shaft speeds on shaft of significant diameter, meaning rubbing velocities are very high. They usually need to be cryogenically compatible as well, and oxygen-compatible on the seals exposed to the oxidizer side. This eliminates the possibility of elastomer based seals, which will embrittle (and cannot hold up at these speeds anyways). Spring loaded and other compression type seals are also not practical at these speeds.

In practice, turbopumps primarily use three seals: labyrinth seals, face seals, and carbon ring seals^[9]. Labyrinth seals are a non-contact type where the fluid is routed through a circuitous path that minimizes the seal's discharge coefficient, and thus minimizes leakage through it. Labyrinth seals leak the most of the three types, and so are seldom used in isolation. Face seals consist of two metal sealing faces that are lapped to a very smooth finish and are pressed together during assembly. These face seals are typically of the non-contact "lift-off" variety, where they develop a thin microfilm of leakage fluid between them during operation that minimizes friction between the static and rotating face. Carbon ring seals are contact seals that consist of multiple carbon static segments around the shaft. They are pressed tightly around the shaft and during operation will intentionally "wear in" to provide a precision sealing surface with minimal leakage.

In practice all three of these seal types will leak to some extent. A large part of seal design is providing safe flow paths for this leakage. Most imperative is that the interpropellant seal (IPS), which the vast majority of engines have at least one of, does not leak fuel and oxidizer together. This is often accomplished by having a central cavity that is continuously purged with inert gas (e.g. helium, nitrogen) at a higher pressure than the propellants on either side, so that the IPS will leak that inert buffer gas outwards from the cavity instead of propellants inwards to the cavity. The only engines that are able to forgo an IPS entirely are full flow staged combustion cycles, because they have one entirely fuel rich turbopump and one entirely ox rich turbopump that don't interact with each other.

Impellers

A few criteria are used when sizing and designing impellers. The first is specific speed - this is a dimensionless parameter characterizing the impeller discharge, for which certain ranges of values are empirically known to indicate different impeller designs would be optimal.^[10]

${\displaystyle N_{s}={\frac {N_{RPM}\cdot Q_{gpm}^{0.50}}{H_{ft}^{0.75}}}\,\,\,\,\,\,\,\,n_{q}={\frac {N_{RPM}\cdot Q_{m^{3}/s}^{0.50}}{H_{m}^{0.75}}}\,\,\,\,\,\,\,\,\omega _{s}={\frac {\omega _{rad/s}\cdot Q_{m^{3}/s}^{0.50}}{(g\cdot H_{m})^{0.75}}}}$

${\displaystyle \omega _{s}={\frac {N_{s}}{2730}}={\frac {n_{q}}{52.9}}}$

• ${\displaystyle H}$ = head
• ${\displaystyle Q}$ = volumetric flowrate
• ${\displaystyle N}$ = shaft speed

${\displaystyle N_{s}}$ is the imperial version, common in US literature. ${\displaystyle n_{q}}$ is common in European literature. ${\displaystyle \omega _{s}}$ is the dimensionless version, but is not yet commonly seen in pump literature. The second parameter is similar: the suction specific speed. This characterizes the impeller's inlet (suction) conditions, and is used to quantify the required inducer and tank pressures upstream of the impeller.

${\displaystyle N_{ss}={\frac {N\cdot Q^{0.5}}{(NPSH_{R})^{0.75}}}}$

NPSH is net positive suction head; NPSH_R is the amount of head required to be generated in the fluid before it reaches the impeller inlet in order to not excessively cavitate in the impeller. "Excessive" is often defined as the level of cavitation that would degrade the pump's discharge head by 3% – hence it is common to see NPSH_R defined as NPSH_3%.

Another key parameter is the impeller's head coefficient ${\displaystyle \psi }$. This characterizes how effective a given tip speed ${\displaystyle u_{out}}$ is at generating head. Head coefficient is typically selected (for a given specific speed) from empirical curves generated by previous industry experience.

${\displaystyle \psi ={\frac {2gH}{u_{out}^{2}}}}$ in some sources^[10]; ${\displaystyle \psi ={\frac {gH}{u_{out}^{2}}}}$ in others^[11][12].

Centrifugal (Radial) Impellers

Main article: Centrifugal pump

In centrifugal turbopumps a rotating disk throws the fluid to the rim.

Centrifugal impellers are optimal on a range of 500 < ${\displaystyle N_{s}}$ < 2500 (numbers are approximate and vary by source)^[10].

Most turbopumps have centrifugal impellers - the fluid enters the pump along its rotational axis and the impeller accelerates the fluid to high speed. The fluid then passes through a volute (which spirals outwards to the outlet) or a diffuser, which is a ring with multiple diverging channels. This causes a large increase in dynamic pressure as fluid velocity is lost. The volute or diffuser turns the high kinetic energy into high pressures (hundreds of bar is not uncommon), and if the outlet backpressure is not too high, high flow rates can be achieved.

The development of the contours of the impeller blades, especially their inlet and outlet angles, is a major driver of the turbopump's overall hydrodynamic performance. Impeller blade geometry development begins with Euler's pump equation^[12]:

${\displaystyle H=\eta (u_{out}v_{out}-u_{in}v_{in})=\eta \omega (r_{out}v_{out}-r_{in}v_{in})}$

• ${\displaystyle \eta }$ = pump efficiency (unitless). This summarizes all inefficiencies into one term.
• ${\displaystyle u}$ = tangential velocity (m/s)
• ${\displaystyle v}$ = flow velocity (m/s). In a stationary frame of reference.
• ${\displaystyle \omega }$ = angular velocity (rad/s)
• ${\displaystyle r}$ = radial position (m)

Mixed Flow Impellers

Between the specific-speed ranges of radial and axial impellers, nominally lie mixed flow impellers. These are rare in turbopumps. They have increased manufacturing complexity and it is easier to adjust one's specific speed out of this range towards radial or axial designs.

Axial Impellers

Main article: Axial-flow pump

Axial impellers are optimal on a range of 8000 < ${\displaystyle N_{s}}$ < 20,000 (numbers are approximate and vary by source)^[10].

In this case the shaft essentially has (sometimes multiple) rotor wheels and stators along the shaft, and the pump the fluid in a direction parallel with the main axis of the pump^[13]. Compared to centrifugal impellers, axial impellers trade lower head generation for higher volumetric flowrates of propellants. For this reason they are common for pumping liquid hydrogen, because of its significantly lower density than essentially all other propellants which use centrifugal pump designs.

Inducers

Main article: Pump inducer

It is very common for turbopumps to feature inducers as well, upstream of the impellers^[14]. The inducer is an axial, spiral design that raises the fluid pressure enough to prevent cavitation when it reaches the entrance to the impeller. The head pressure that the fluid rises over the length of the inducer is termed the NPSH_A (NPSH available). This must be above the NPSH_R of the impeller: NPSH_margin = NPSH_A / NPSH_R. Turbopumps also require a certain NPSH before it even reaches the inducer, again termed the NPSH_R for the inducer (so the inducer and impeller both have their own individual NPSH_R). This is achieved by pressurizing the propellant tanks to some extent; a few bar is typical. Inducers for cryogenic propellants usually cannot be designed to have zero NPSH_R because a rocket usually fills cryogenic propellants at their saturated state, meaning NPSH_A in the tank is zero. This gives no margin and thus cavitation at the inducer blades becomes likely. This can possibly be overcome with subcooled / densified propellants (e.g. Falcon 9). Regardless, some tank pressure is often desirable for structural stability of the rocket itself and so increases the NPSH_A, reducing the NPSH_R of the inducer (and so probably its axial length) as a side benefit.

Turbines

Main article: Gas turbine

Turbopumps, by definition, are driven by gas turbines. Turbines are typically either of an impulse design (common in gas generator and other open cycles) or of a reaction design (common in staged combustion and other closed cycles). They can consist of one or more stages, where each stage has both a stator, which can be bladed or nozzles, and a wheel (sometimes referred to as a rotor in older papers and aero focused papers)^[3].

Open cycles aim to increase efficiency by minimizing mass flow through the turbine, making up for it by maximizing pressure drop. This is because the mass flow is dumped overboard, a performance hit. Comparatively, maximizing pressure drop is easy to do because it dumps to ambient pressure, which will be significantly lower than the gas generator (GG) chamber pressure. This is true even if GG pressure were the same value as the main chamber pressure, which the pumps have to work hard enough to discharge to anyways. These considerations drive the designer towards impulse designs on the turbine, with gas flow expanded via converging-diverging blades or nozzles to supersonic velocities that then impinge on the turbine wheel.

Closed cycles aim to increase efficiency by minimizing pressure drop across the turbine, making up for it by maximizing mass flow. This is because the downstream pressure must be higher than the main chamber pressure. It is often significantly higher because of injector and regen jacket pressure losses. Consequently, the only method for increasing pressure drop is to increase chamber pressure in the preburner much higher than the main chamber. This puts significantly more load on the pumps which must have a high pressure discharge for the preburner. Comparatively, high mass flow is easy to accomplish because none is being dumped overboard - so it is common to route the entire mass flow of one propellant through the preburner and turbine. Full flow staged combustion cycles take this a step further by routing the entire mass flow (hence 'full flow') of both propellants through preburners and turbines, taking advantage of essentially 100% of possible mass flow through the engine to generate shaft power for the turbopumps. These considerations drive the designer towards reaction type designs on the turbine where gas flow is subsonically expelled from, and reacting against, the wheel blades.

Complexities of centrifugal turbopumps

Turbopumps have a reputation for being difficult to design for optimal performance. Whereas a well engineered and debugged pump can manage 70–90% efficiency, figures less than half that are not uncommon. Low efficiency may be acceptable in some applications, but in rocketry this is a severe problem. Turbopumps in rockets are important and problematic enough that launch vehicles using one have been caustically described as a "turbopump with a rocket attached"–up to 55% of the total cost has been ascribed to this area.^[15]

Common problems include:

1. excessive flow from the high-pressure rim back to the low-pressure inlet along the gap between the casing of the pump and the rotor,
2. excessive recirculation of the fluid at inlet,
3. excessive vortexing of the fluid as it leaves the casing of the pump,
4. damaging cavitation to impeller blade surfaces in low-pressure zones.

In addition, the precise shape of the rotor itself is critical.

History

The V-2 rocket used a circular turbopump to pressurize the propellant.

Early development

High-pressure pumps for larger missiles had been discussed by rocket pioneers such as Hermann Oberth.^[16] In mid-1935 Wernher von Braun initiated a fuel pump project at the southwest German firm Klein, Schanzlin & Becker that was experienced in building large fire-fighting pumps.^[17]: 80 The V-2 rocket design used hydrogen peroxide decomposed through a Walter steam generator to power the uncontrolled turbopump^[17]: 81 produced at the Heinkel plant at Jenbach,^[18] so V-2 turbopumps and combustion chamber were tested and matched to prevent the pump from overpressurizing the chamber.^[17]: 172 The first engine fired successfully in September, and on August 16, 1942, a trial rocket stopped in mid-air and crashed due to a failure in the turbopump.^[17][19] The first successful V-2 launch was on October 3, 1942.^[20]

Starting from the 1938-1940, Robert H. Goddard's team also independently developed small turbopumps.

Development from 1947 to 1949

The principal engineer for turbopump development at Aerojet was George Bosco. During the second half of 1947, Bosco and his group learned about the pump work of others and made preliminary design studies. Aerojet representatives visited Ohio State University where Florant was working on hydrogen pumps, and consulted Dietrich Singelmann, a German pump expert at Wright Field. Bosco subsequently used Singelmann's data in designing Aerojet's first hydrogen pump.^[21]

By mid-1948, Aerojet had selected centrifugal pumps for both liquid hydrogen and liquid oxygen. They obtained some German radial-vane pumps from the Navy and tested them during the second half of the year.^[21]

By the end of 1948, Aerojet had designed, built, and tested a liquid hydrogen pump (15 cm diameter). Initially, it used ball bearings that were run clean and dry, because the low temperature made conventional lubrication impractical. The pump was first operated at low speeds to allow its parts to cool down to operating temperature. When temperature gauges showed that liquid hydrogen had reached the pump, an attempt was made to accelerate from 5000 to 35 000 revolutions per minute. The pump failed and examination of the pieces pointed to a failure of the bearing, as well as the impeller. After some testing, super-precision bearings, lubricated by oil that was atomized and directed by a stream of gaseous nitrogen, were used. On the next run, the bearings worked satisfactorily but the stresses were too great for the brazed impeller and it flew apart. A new one was made by milling from a solid block of aluminum. The next two runs with the new pump were a great disappointment; the instruments showed no significant flow or pressure rise. The problem was traced to the exit diffuser of the pump, which was too small and insufficiently cooled during the cool-down cycle so that it limited the flow. This was corrected by adding vent holes in the pump housing; the vents were opened during cool down and closed when the pump was cold. With this fix, two additional runs were made in March 1949 and both were successful. Flow rate and pressure were found to be in approximate agreement with theoretical predictions. The maximum pressure was 26 atmospheres (26 atm (2.6 MPa; 380 psi)) and the flow was 0.25 kilogram per second.^[21]

After 1949

The Space Shuttle main engine's turbopumps spun at over 30,000 rpm, delivering 150 lb (68 kg) of liquid hydrogen and 896 lb (406 kg) of liquid oxygen to the engine per second.^[22] While not technically a turbopump (in that it lacks a turbine), the Electron Rocket's Rutherford became the first engine to use an electrically-driven pump in flight in 2018.^[23]

Turbopump Examples

Engine	Cycle	Fuel	Oxidizer	Pump Type	Shafts	Shaft Speed, RPM	Outlet Pressure, barA	Turbine Stages	Geared
F-1	Gas Generator	RP-1	LOX	Radial	Single	5488[1]	110 / 128	2	No
RS-25 / SSME	Fuel Rich Staged[i]	Hydrogen	LOX	Axial/Radial	Quad	36000 HPFTP / 16185 LPFTP / 28120 HPOTP / 5150 LPOTP [24]	357 / 585	4 / 6	No
RS-68	Gas Generator	Hydrogen	LOX	Radial	Dual	21000 / 8700[25]		2 / 2	No
J-2	Gas Generator	Hydrogen	LOX	Axial/Radial	Dual	27130 / 8753[1]	77 / 85	2 / 2	No
RL10	Expander (Closed)	Hydrogen	LOX	Radial	Dual	30250 / 12100[1]	41 / 68	2	Yes
RD-107	Catalyst Gas Generator	RP-1	LOX	Radial	Single[ii]			1	Yes[ii]
RD-180	Ox Rich Staged	RP-1	LOX	Radial	Triple[26]			1	No
RD-275	Ox Rich Staged	N2O4	UDMH	Radial	Single				No
YF-100	Ox Rich Staged	RP-1	LOX	Radial	Triple[27]				No
Merlin	Gas Generator	RP-1	LOX	Radial	Single			1	No
Raptor	Full Flow Staged	Methane	LOX	Radial	Dual				No
Archimedes	Ox Rich Staged	Methane	LOX	Radial	Single				No
Rutherford[iii]	Electric[iii]	RP-1	LOX	Radial	Dual				No
Reaver	Tap-Off	RP-1	LOX	Radial	Single				No
Lightning	Tap-Off	RP-1	LOX	Radial	Single			1	No
E-2	Ox Rich Staged	RP-1	LOX	Radial	Single	30000			No
Aeon-R	Gas Generator	Methane	LOX	Radial	Dual			1 / 1	No
Hadley	Ox Rich Staged	RP-1	LOX	Radial	Single				No
Zenith	Full Flow Staged	Methane	LOX	Radial	Dual				No

Where two values are given, fuel side listed first and oxidizer side listed second.

Notes

1. The low pressure fuel turbopump of the RS-25 is technically driven via expander cycle
2. The RD-107 technically has two more secondary shafts to pump hydrogen peroxide and nitrogen
3. Technically Rutherford is not a turbopump because it does not have a turbine. Shown for comparison

See also

• Turboexpander
• Gas-generator cycle
• Staged combustion cycle
• Expander cycle
• Components of jet engines