Combustion product gasses enter the nozzle through a throat.
Exhaust exits the rocket.
A liquid-propellant rocket or liquid rocket uses a rocket engine burning liquid propellants. (Alternate approaches use gaseous or solid propellants.) Liquids are desirable propellants because they have reasonably high density and their combustion products have high specific impulse (Isp). This allows the volume of the propellant tanks to be relatively low.
Most designs of liquid rocket engines are throttleable
for variable thrust operation. Some allow control of the propellant
mixture ratio (ratio at which oxidizer and fuel are mixed). Some can be
shut down and, with a suitable ignition system or self-igniting
propellant, restarted.
Hybrid rockets apply a liquid or gaseous oxidizer to a solid fuel.
Advantages and disadvantages
The use of liquid propellants has a number of advantages:
A liquid rocket engine can be tested prior to use, whereas for a solid rocket motor a rigorous quality management must be applied during manufacturing to ensure high reliability.
Liquid systems enable higher specific impulse than solids and hybrid rocket motors and can provide very high tankage efficiency.
A liquid rocket engine can also usually be reused for several flights, as in the Space Shuttle and Falcon 9 series rockets, although reuse of solid rocket motors was also effectively demonstrated during the Shuttle program.
The flow of propellant into the combustion chamber can be throttled,
which allows for control over the magnitude of the thrust throughout
the flight. This enables real-time error correction during the flight
along with efficiency gains.
Shutdown and restart capabilities allow for multiple burn cycles throughout a flight.
In the case of an emergency, liquid propelled rockets can be
shutdown in a controlled manner, which provides an extra level of safety
and mission abort capability.
Bipropellant
liquid rockets are simple in concept but due to high temperatures and
high speed moving parts, very complex in practice.
Use of liquid propellants can also be associated with a number of issues:
Because the propellant is a very large proportion of the mass of the vehicle, the center of mass
shifts significantly rearward as the propellant is used; one will
typically lose control of the vehicle if its center mass gets too close
to the center of drag/pressure.
When operated within an atmosphere, pressurization of the typically very thin-walled propellant tanks must guarantee positive gauge pressure at all times to avoid catastrophic collapse of the tank.
Liquid propellants are subject to slosh,
which has frequently led to loss of control of the vehicle. This can be
controlled with slosh baffles in the tanks as well as judicious control
laws in the guidance system.
They can suffer from pogo oscillation where the rocket suffers from uncommanded cycles of acceleration.
Liquid propellants often need ullage motors
in zero-gravity or during staging to avoid sucking gas into engines at
start up. They are also subject to vortexing within the tank,
particularly towards the end of the burn, which can also result in gas
being sucked into the engine or pump.
Liquid propellants can leak, especially hydrogen, possibly leading to the formation of an explosive mixture.
Turbopumps
to pump liquid propellants are complex to design, and can suffer
serious failure modes, such as overspeeding if they run dry or shedding
fragments at high speed if metal particles from the manufacturing
process enter the pump.
Cryogenic propellants,
such as liquid oxygen, freeze atmospheric water vapor into ice. This
can damage or block seals and valves and can cause leaks and other
failures. Avoiding this problem often requires lengthy chilldown
procedures which attempt to remove as much of the vapor from the system
as possible. Ice can also form on the outside of the tank, and later
fall and damage the vehicle. External foam insulation can cause issues
as shown by the Space Shuttle Columbia disaster. Non-cryogenic propellants do not cause such problems.
Non-storable liquid rockets require considerable preparation immediately before launch. This makes them less practical than solid rockets for most weapon systems.
Principle of operation
Liquid
rocket engines have tankage and pipes to store and transfer propellant,
an injector system and one or more combustion chambers with associated nozzles.
Typical liquid propellants have densities roughly similar to water, approximately 0.7 to 1.4 g/cm3 (0.025 to 0.051 lb/cu in). An exception is liquid hydrogen which has a much lower density, while requiring only relatively modest pressure to prevent vaporization.
The density and low pressure of liquid propellants permit lightweight
tankage: approximately 1% of the contents for dense propellants and
around 10% for liquid hydrogen. The increased tank mass is due to liquid
hydrogen's low density and the mass of the required insulation.
For injection into the combustion chamber, the propellant
pressure at the injectors needs to be greater than the chamber pressure.
This is often achieved with a pump. Suitable pumps usually use
centrifugal turbopumps due to their high power and light weight, although reciprocating pumps
have been employed in the past. Turbopumps are usually lightweight and
can give excellent performance; with an on-Earth weight well under 1% of
the thrust. Indeed, overall thrust to weight ratios including a turbopump have been as high as 155:1 with the SpaceX Merlin 1D rocket engine and up to 180:1 with the vacuum version. Instead of a pump, some designs use a tank of a high-pressure inert gas
such as helium to pressurize the propellants. These rockets often
provide lower delta-v
because the mass of the pressurant tankage reduces performance. In some
designs for high altitude or vacuum use the tankage mass can be
acceptable.
Liquid propellants are often pumped into the combustion chamber with a lightweight centrifugal turbopump.
Recently, some aerospace companies have used electric pumps with
batteries. In simpler, small engines, an inert gas stored in a tank at a
high pressure is sometimes used instead of pumps to force propellants
into the combustion chamber. These engines may have a higher mass ratio,
but are usually more reliable, and are therefore used widely in
satellites for orbit maintenance.
One of the most efficient mixtures, oxygen and hydrogen,
suffers from the extremely low temperatures required for storing liquid
hydrogen (around 20 K or −253.2 °C or −423.7 °F) and very low fuel
density (70 kg/m3 or 4.4 lb/cu ft, compared to RP-1 at 820 kg/m3 or 51 lb/cu ft), necessitating large tanks that must also be lightweight and insulating. Lightweight foam insulation on the Space Shuttle external tank led to the Space ShuttleColumbia's destruction, as a piece broke loose, damaged its wing and caused it to break up on atmospheric reentry.
Liquid methane/LNG has several advantages over LH2. Its performance (max. specific impulse) is lower than that of LH2
but higher than that of RP1 (kerosene) and solid propellants, and its
higher density, similarly to other hydrocarbon fuels, provides higher
thrust to volume ratios than LH2, although its density is not as high as that of RP1. This makes it specially attractive for reusable launch systems because higher density allows for smaller motors, propellant tanks and associated systems. LNG also burns with less or no soot (less or no coking) than RP1, which
eases reusability when compared with it, and LNG and RP1 burn cooler
than LH2 so LNG and RP1 do not deform the interior structures
of the engine as much. This means that engines that burn LNG can be
reused more than those that burn RP1 or LH2. Unlike engines that burn LH2,
both RP1 and LNG engines can be designed with a shared shaft with a
single turbine and two turbopumps, one each for LOX and LNG/RP1. In space, LNG does not need heaters to keep it liquid, unlike RP1. LNG is less expensive, being readily available in large quantities. It
can be stored for more prolonged periods of time, and is less explosive
than LH2.
Liquid oxygen (LOX) and carbon monoxide (CO) – proposed for a Mars hopper vehicle (with a specific impulse of approximately 250s), principally because carbon monoxide and oxygen can be straightforwardly produced by Zirconia electrolysis from the Martian atmosphere without requiring use of any of the Martian water resources to obtain Hydrogen.
Many non-cryogenic bipropellants are hypergolic (self igniting).
T-Stoff (80% hydrogen peroxide, H2O2 as the oxidizer) and C-Stoff (methanol, CH3OH, and hydrazine hydrate, N2H4·n(H2O) as the fuel) – used for the Hellmuth-Walter-Werke HWK 109-509A, -B and -C engine family used on the Messerschmitt Me 163B Komet, an operational rocket fighter plane of World War II, and Ba 349 Natter crewed VTO interceptor prototypes.
For storableICBMs
and most spacecraft, including crewed vehicles, planetary probes, and
satellites, storing cryogenic propellants over extended periods is
unfeasible. Because of this, mixtures of hydrazine or its derivatives in combination with nitrogen oxides are generally used for such applications, but are toxic and carcinogenic. Consequently, to improve handling, some crew vehicles such as Dream Chaser and Space Ship Two plan to use hybrid rockets with non-toxic fuel and oxidizer combinations.
Injectors
Rocketdyne F-1 rocket engine fuel injector
The injector implementation in liquid rockets determines the percentage of the theoretical performance of the nozzle that can be achieved. A poor injector performance causes unburnt propellant to leave the engine, giving poor efficiency.
Additionally, injectors are also usually key in reducing thermal
loads on the nozzle; by increasing the proportion of fuel around the
edge of the chamber, this gives much lower temperatures on the walls of
the nozzle.
Types of injectors
Injectors
can be as simple as a number of small diameter holes arranged in
carefully constructed patterns through which the fuel and oxidizer
travel. The speed of the flow is determined by the square root of the
pressure drop across the injectors, the shape of the hole and other
details such as the density of the propellant.
The first injectors used on the V-2 created parallel jets of fuel
and oxidizer which then combusted in the chamber. This gave quite poor
efficiency.
Injectors today classically consist of a number of small holes
which aim jets of fuel and oxidizer so that they collide at a point in
space a short distance away from the injector plate. This helps to break
the flow up into small droplets that burn more easily.
The RS-25 engine designed for the Space Shuttle
uses a system of fluted posts, which use heated hydrogen from the
preburner to vaporize the liquid oxygen flowing through the center of
the posts and this improves the rate and stability of the combustion process; previous engines such as the F-1 used for the Apollo program
had significant issues with oscillations that led to destruction of the
engines, but this was not a problem in the RS-25 due to this design
detail.
Valentin Glushko
invented the centripetal injector in the early 1930s, and it has been
almost universally used in Russian engines. Rotational motion is applied
to the liquid (and sometimes the two propellants are mixed), then it is
expelled through a small hole, where it forms a cone-shaped sheet that
rapidly atomizes. Goddard's first liquid engine used a single impinging
injector. German scientists in WWII experimented with impinging
injectors on flat plates, used successfully in the Wasserfall missile.
Combustion stability
To avoid instabilities such as chugging,
which is a relatively low speed oscillation, the engine must be
designed with enough pressure drop across the injectors to render the
flow largely independent of the chamber pressure. This pressure drop is
normally achieved by using at least 20% of the chamber pressure across
the injectors.
Nevertheless, particularly in larger engines, a high speed
combustion oscillation is easily triggered, and these are not well
understood. These high speed oscillations tend to disrupt the gas side
boundary layer of the engine, and this can cause the cooling system to
rapidly fail, destroying the engine. These kinds of oscillations are
much more common on large engines, and plagued the development of the Saturn V, but were finally overcome.
Some combustion chambers, such as those of the RS-25 engine, use Helmholtz resonators as damping mechanisms to stop particular resonant frequencies from growing.
To prevent these issues the RS-25 injector design instead went to
a lot of effort to vaporize the propellant prior to injection into the
combustion chamber. Although many other features were used to ensure
that instabilities could not occur, later research showed that these
other features were unnecessary, and the gas phase combustion worked
reliably.
Testing for stability often involves the use of small explosives.
These are detonated within the chamber during operation, and causes an
impulsive excitation. By examining the pressure trace of the chamber to
determine how quickly the effects of the disturbance die away, it is
possible to estimate the stability and redesign features of the chamber
if required.
Engine cycles
For
liquid-propellant rockets, four different ways of powering the
injection of the propellant into the chamber are in common use.
Fuel and oxidizer must be pumped into the combustion chamber
against the pressure of the hot gasses being burned, and engine power is
limited by the rate at which propellant can be pumped into the
combustion chamber. For atmospheric or launcher use, high pressure, and
thus high power, engine cycles are desirable to minimize gravity drag. For orbital use, lower power cycles are usually fine.
The propellants are forced in from pressurised (relatively heavy)
tanks. The heavy tanks mean that a relatively low pressure is optimal,
limiting engine power, but all the fuel is burned, allowing high
efficiency. The pressurant used is frequently helium due to its lack of
reactivity and low density. Examples: AJ-10, used in the Space Shuttle OMS, Apollo SPS, and the second stage of the Delta II.
An electric motor, generally a brushless DC electric motor, drives the pumps. The electric motor is powered by a battery pack. It is relatively simple to implement and reduces the complexity of the turbomachinery design, but at the expense of the extra dry mass of the battery pack. Example engine is the Rutherford designed and used by Rocket Lab.
A small percentage of the propellants are burnt in a preburner to
power a turbopump and then exhausted through a separate nozzle, or low
down on the main one. This results in a reduction in efficiency since
the exhaust contributes little or no thrust, but the pump turbines can
be very large, allowing for high power engines. Examples: Saturn V's F-1 and J-2, Delta IV's RS-68, Ariane 5's HM7B, Falcon 9's Merlin.
Takes hot gases from the main combustion chamber of the rocket engine and routes them through engine turbopump
turbines to pump propellant, then is exhausted. Since not all
propellant flows through the main combustion chamber, the tap-off cycle
is considered an open-cycle engine. Examples include the J-2S and BE-3.
Cryogenic fuel (hydrogen, or methane) is used to cool the walls of
the combustion chamber and nozzle. Absorbed heat vaporizes and expands
the fuel which is then used to drive the turbopumps before it enters the
combustion chamber, allowing for high efficiency, or is bled overboard,
allowing for higher power turbopumps. The limited heat available to
vaporize the fuel constrains engine power. Examples: RL10 for Atlas V and Delta IV second stages (closed cycle), H-II's LE-5 (bleed cycle).
A fuel- or oxidizer-rich mixture is burned in a preburner and then
drives turbopumps, and this high-pressure exhaust is fed directly into
the main chamber where the remainder of the fuel or oxidizer undergoes
combustion, permitting very high pressures and efficiency. Examples: SSME, RD-191, LE-7.
Fuel- and oxidizer-rich mixtures are burned in separate preburners
and drive the turbopumps, then both high-pressure exhausts, one oxygen
rich and the other fuel rich, are fed directly into the main chamber
where they combine and combust, permitting very high pressures and high
efficiency. Example: SpaceX Raptor.
Engine cycle tradeoffs
Selecting
an engine cycle is one of the earlier steps to rocket engine design. A
number of tradeoffs arise from this selection, some of which include:
Tradeoff comparison among popular engine cycles
Cycle type
Gas generator
Expander cycle
Staged-combustion
Pressure-fed
Advantages
Simple; low dry mass; allows for high power turbopumps for high thrust
High specific impulse; fairly low complexity
High specific impulse; high combustion chamber pressures allowing for high thrust
Simple; no turbopumps; low dry mass; high specific impulse
Disadvantages
Lower specific impulse
Must use cryogenic fuel; heat transfer to the fuel limits available power to the turbine and thus engine thrust
Greatly increased complexity &, therefore, mass (more-so for full-flow)
Tank pressure limits combustion chamber pressure and thrust; heavy tanks and associated pressurization hardware
Injectors are commonly laid out so that a fuel-rich layer is created
at the combustion chamber wall. This reduces the temperature there, and
downstream to the throat and even into the nozzle and permits the
combustion chamber to be run at higher pressure, which permits a higher
expansion ratio nozzle to be used which gives a higher ISP and better system performance. A liquid rocket engine often employs regenerative cooling, which uses the fuel or less commonly the oxidizer to cool the chamber and nozzle.
Ignition
Ignition
can be performed in many ways, but perhaps more so with liquid
propellants than other rockets a consistent and significant ignitions
source is required; a delay of ignition (in some cases as small as a few
tens of milliseconds) can cause overpressure of the chamber due to
excess propellant. A hard start can even cause an engine to explode.
Generally, ignition systems try to apply flames across the
injector surface, with a mass flow of approximately 1% of the full mass
flow of the chamber.
Safety interlocks are sometimes used to ensure the presence of an
ignition source before the main valves open; however reliability of the
interlocks can in some cases be lower than the ignition system. Thus it
depends on whether the system must fail safe, or whether overall
mission success is more important. Interlocks are rarely used for upper,
uncrewed stages where failure of the interlock would cause loss of
mission, but are present on the RS-25 engine, to shut the engines down
prior to liftoff of the Space Shuttle. In addition, detection of
successful ignition of the igniter is surprisingly difficult, some
systems use thin wires that are cut by the flames, pressure sensors have
also seen some use.
Methods of ignition include pyrotechnic, electrical (spark or hot wire), and chemical. Hypergolic
propellants have the advantage of self igniting, reliably and with less
chance of hard starts. In the 1940s, the Russians began to start
engines with hypergols, to then switch over to the primary propellants
after ignition. This was also used on the American F-1 rocket engine on the Apollo program.
Ignition with a pyrophoric agent: Triethylaluminium
ignites on contact with air and will ignite and/or decompose on contact
with water, and with any other oxidizer—it is one of the few substances
sufficiently pyrophoric to ignite on contact with cryogenic liquid oxygen. The enthalpy of combustion, ΔcH°, is −5,105.70 ± 2.90 kJ/mol (−1,220.29 ± 0.69 kcal/mol). Its easy ignition makes it particularly desirable as a rocket engineignitor. May be used in conjunction with triethylborane to create triethylaluminum-triethylborane, better known as TEA-TEB.
Rocket 09 (left) and 10 (GIRD-09 and GIRD-X). Museum of Cosmonautics and Rocket Technology; St. Petersburg.
The idea of a liquid-fueled rocket as understood in the modern context first appeared in 1903 in the book Exploration of the Universe with Rocket-Propelled Vehicles by the Russian rocket scientist Konstantin Tsiolkovsky. The magnitude of his contribution to astronautics is astounding, including the Tsiolkovsky rocket equation, multi-staged rockets, and using liquid oxygen and liquid hydrogen in liquid propellant rockets. Tsiolkovsky influenced later rocket scientists throughout Europe, like Wernher von Braun. Soviet search teams at Peenemünde
found a German translation of a book by Tsiolkovsky of which "almost
every page...was embellished by von Braun's comments and notes." Leading Soviet rocket-engine designer Valentin Glushko and rocket designer Sergey Korolev studied Tsiolkovsky's works as youths and both sought to turn Tsiolkovsky's theories into reality.
From 1929 to 1930 in Leningrad Glushko pursued rocket research at the Gas Dynamics Laboratory (GDL), where a new research section was set up for the study of liquid-propellant and electric rocket engines. This resulted in the creation of ORM (from "Experimental Rocket Motor" in Russian) engines ORM-1 [ru] to ORM-52 [ru]. A total of 100 bench tests of liquid-propellant rockets were conducted
using various types of fuel, both low and high-boiling and thrust up to
300 kg was achieved.[19][18]
During this period in Moscow, Fredrich Tsander
– a scientist and inventor – was designing and building liquid rocket
engines which ran on compressed air and gasoline. Tsander investigated
high-energy fuels including powdered metals mixed with gasoline. In
September 1931 Tsander formed the Moscow based 'Group for the Study of Reactive Motion', better known by its Russian acronym "GIRD". In May 1932, Sergey Korolev replaced Tsander as the head of GIRD. On 17 August 1933, Mikhail Tikhonravov
launched the first Soviet liquid-propelled rocket (the GIRD-9), fueled
by liquid oxygen and jellied gasoline. It reached an altitude of 400
metres (1,300 ft). In January 1933 Tsander began development of the GIRD-X rocket. This
design burned liquid oxygen and gasoline and was one of the first
engines to be regeneratively cooled by the liquid oxygen, which flowed
around the inner wall of the combustion chamber before entering it.
Problems with burn-through during testing prompted a switch from
gasoline to less energetic alcohol. The final missile, 2.2 metres
(7.2 ft) long by 140 millimetres (5.5 in) in diameter, had a mass of 30
kilograms (66 lb), and it was anticipated that it could carry a 2
kilograms (4.4 lb) payload to an altitude of 5.5 kilometres (3.4 mi). The GIRD X rocket was launched on 25 November 1933 and flew to a height of 80 meters.
In 1933 GDL and GIRD merged and became the Reactive Scientific Research Institute (RNII). At RNII Gushko continued the development of liquid propellant rocket engines ОРМ-53 to ОРМ-102, with ORM-65 [ru] powering the RP-318 rocket-powered aircraft. In 1938 Leonid Dushkin replaced Glushko and continued development of the ORM engines, including the engine for the rocket powered interceptor, the Bereznyak-Isayev BI-1. At RNII Tikhonravov worked on developing oxygen/alcohol liquid-propellant rocket engines. Ultimately liquid propellant rocket engines were given a low priority
during the late 1930s at RNII, however the research was productive and
very important for later achievements of the Soviet rocket program.
Peruvian Pedro Paulet, who had experimented with rockets throughout his life in Peru, wrote a letter to El Comercio in Lima in 1927, claiming he had experimented with a liquid rocket engine while he was a student in Paris three decades earlier.Historians of early rocketry experiments, among them Max Valier, Willy Ley, and John D. Clark,
have given differing amounts of credence to Paulet's report. Valier
applauded Paulet's liquid-propelled rocket design in the Verein für
Raumschiffahrt publication Die Rakete, saying the engine had "amazing power" and that his plans were necessary for future rocket development. Hermann Oberth would name Paulet as a pioneer in rocketry in 1965. Wernher von Braun would also describe Paulet as "the pioneer of the liquid fuel propulsion motor" and stated that "Paulet helped man reach the Moon". Paulet was later approached by Nazi Germany, being invited to join the Astronomische Gesellschaft
to help develop rocket technology, though he refused to assist after
discovering that the project was destined for weaponization and never
shared the formula for his propellant. According to filmmaker and researcher Álvaro Mejía, Frederick I. Ordway III would later attempt to discredit Paulet's discoveries in the context of the Cold War and in an effort to shift the public image of von Braun away from his history with Nazi Germany.
United States
Robert H. Goddard, bundled against the cold New England weather of March 16, 1926, holds the launching frame of his most notable invention — the first liquid rocket.
The first flight of a liquid-propellant rocket took place on March 16, 1926 at Auburn, Massachusetts, when American professor Dr. Robert H. Goddard launched a vehicle using liquid oxygen and gasoline as propellants. The rocket, which was dubbed "Nell", rose just 41 feet during a
2.5-second flight that ended in a cabbage field, but it was an important
demonstration that rockets using liquid propulsion were possible.
Goddard proposed liquid propellants about fifteen years earlier and
began to seriously experiment with them in 1921. The German-Romanian Hermann Oberth published a book in 1923 suggesting the use of liquid propellants.
Germany
In
Germany, engineers and scientists became enthralled with liquid
propulsion, building and testing them in the late 1920s within Opel RAK, the world's first rocket program, in Rüsselsheim. According to Max Valier's account, Opel RAK rocket designer, Friedrich Wilhelm Sander launched two liquid-fuel rockets at Opel Rennbahn in Rüsselsheim
on April 10 and April 12, 1929. These Opel RAK rockets have been the
first European, and after Goddard the world's second, liquid-fuel
rockets in history. In his book "Raketenfahrt" Valier describes the size
of the rockets as of 21 cm in diameter and with a length of 74 cm,
weighing 7 kg empty and 16 kg with fuel. The maximum thrust was 45 to 50
kp, with a total burning time of 132 seconds. These properties indicate
a gas pressure pumping. The main purpose of these tests was to develop
the liquid rocket-propulsion system for a Gebrüder-Müller-Griessheim
aircraft under construction for a planned flight across the English channel. Also spaceflight historian Frank H. Winter,
curator at National Air and Space Museum in Washington, DC, confirms
the Opel group was working, in addition to their solid-fuel rockets used
for land-speed records and the world's first crewed rocket-plane
flights with the Opel RAK.1, on liquid-fuel rockets. By May 1929, the engine produced a thrust of 200 kg (440 lb.) "for
longer than fifteen minutes and in July 1929, the Opel RAK collaborators
were able to attain powered phases of more than thirty minutes for
thrusts of 300 kg (660-lb.) at Opel's works in Rüsselsheim," again
according to Max Valier's account. The Great Depression brought an end
to the Opel RAK activities. After working for the German military in the
early 1930s, Sander was arrested by Gestapo in 1935, when private
rocket-engineering became forbidden in Germany. He was convicted of
treason to 5 years in prison and forced to sell his company, he died in
1938. Max Valier's (via Arthur Rudolph
and Heylandt), who died while experimenting in 1930, and Friedrich
Sander's work on liquid-fuel rockets was confiscated by the German
military, the Heereswaffenamt and integrated into the activities under General Walter Dornberger in the early and mid-1930s in a field near Berlin. Max Valier was a co-founder of an amateur research group, the VfR,
working on liquid rockets in the early 1930s, and many of whose members
eventually became important rocket technology pioneers, including Wernher von Braun. Von Braun served as head of the army research station that designed the V-2 rocket weapon for the Nazis.
Drawing of the He 176 V1 prototype rocket aircraft
By the late 1930s, use of rocket propulsion for crewed flight began to be seriously experimented with, as Germany's Heinkel He 176 made the first crewed rocket-powered flight using a liquid rocket engine, designed by German aeronautics engineer Hellmuth Walter on June 20, 1939. The only production rocket-powered combat aircraft ever to see military service, the Me 163Komet in 1944-45, also used a Walter-designed liquid rocket engine, the Walter HWK 109-509, which produced up to 1,700 kgf (16.7 kN) thrust at full power.
Post World War II
After
World War II the American government and military finally seriously
considered liquid-propellant rockets as weapons and began to fund work
on them. The Soviet Union did likewise, and thus began the Space Race.
