Propulsion
Propulsion

Jet engine

Rocket

Spacecraft propulsion

Electric propulsion
Jet Engine
A jet engine is a reaction engine that discharges a fast moving jet of fluid to generate
thrust in accordance with Newton's third law of motion. This broad definition of jet
engines includes turbojets, turbofans, rockets, ramjets, pulse jets and pump-jets. In
general, most jet engines are internal combustion engines but non-combusting forms
also exist.
In common usage, the term 'jet engine' generally refers to a gas turbine driven internal
combustion engine, an engine with a rotary compressor powered by a turbine ("Brayton
cycle"), with the leftover power providing thrust. These types of jet engines are
primarily used by jet aircraft for long distance travel. The early jet aircraft used turbojet
engines which were relatively inefficient for subsonic flight. Modern jet aircraft usually
use high-bypass turbofan engines which help give high speeds as well as, over long
distances, giving better fuel efficiency than many other forms of transport.
About 7.2% of the world's oil was ultimately consumed by jet engines in 2004 In 2007,
the cost of jet fuel, while highly variable from one airline to another, averaged 26.5% of
total operating costs, making it the single largest operating expense for most airlines.
History
Jet engines can be dated back to the first century AD, when Hero of Alexandria invented
the aeolipile. This used steam power directed through two jet nozzles so as to cause a
sphere to spin rapidly on its axis. So far as is known, it was little used for supplying
mechanical power, and the potential practical applications of Hero's invention of the jet
engine were not recognized. It was simply considered a curiosity.
Jet propulsion only literally and figuratively took off with the invention of the rocket by
the Chinese in the 11th century. Rocket exhaust was initially used in a modest way for
fireworks but gradually progressed to propel formidable weaponry; and there the
technology stalled for hundreds of years.
In Ottoman Turkey in 1633 Lagari Hasan Çelebi took off with what was described to be a
cone shaped rocket and then glided with wings into a successful landing winning a
position in the Ottoman army. However, this was essentially a stunt.
The problem was that rockets are simply too inefficient at low speeds to be useful for
general aviation. Instead, by the 1930s, the piston engine in its many different forms
(rotary and static radial, aircooled and liquid-cooled inline) was the only type of
powerplant available to aircraft designers. This was acceptable as long as only low
performance aircraft were required, and indeed all that were available.
However, engineers were beginning to realize that the piston engine was self-limiting in
terms of the maximum performance which could be attained; the limit was essentially
one of propeller efficiency. This seemed to peak as blade tips approached the speed of
sound. If engine, and thus aircraft, performance were ever to increase beyond such a
barrier, a way would have to be found to radically improve the design of the piston
engine, or a wholly new type of powerplant would have to be developed. This was the
motivation behind the development of the gas turbine engine, commonly called a "jet"
engine, which would become almost as revolutionary to aviation as the Wright brothers'
first flight.
The earliest attempts at jet engines were hybrid designs in which an external power
source first compressed air, which was then mixed with fuel and burned for jet thrust. In
one such system, called a thermojet by Secondo Campini but more commonly, motorjet,
the air was compressed by a fan driven by a conventional piston engine. Examples of
this type of design were Henri Coandă's Coandă-1910 aircraft, and the much later
Campini Caproni CC.2, and the Japanese Tsu-11 engine intended to power Ohka
kamikaze planes towards the end of World War II. None were entirely successful and
the CC.2 ended up being slower than the same design with a traditional engine and
propeller combination.
The key to a practical jet engine was the gas turbine, used to extract energy from the
engine itself to drive the compressor. The gas turbine was not an idea developed in the
1930s: the patent for a stationary turbine was granted to John Barber in England in
1791. The first gas turbine to successfully run self-sustaining was built in 1903 by
Norwegian engineer Ægidius Elling. The first patents for jet propulsion were issued in
1917. Limitations in design and practical engineering and metallurgy prevented such
engines reaching manufacture. The main problems were safety, reliability, weight and,
especially, sustained operation. In 1923, Edgar Buckingham of the US National Bureau of
Standard published a report expressing scepticism that jet engines would be
economically competitive with prop driven aircraft at low altitude and the airspeeds of
the period: "there does not appear to be, at present, any prospect whatever that jet
propulsion of the sort here considered will ever be of practical value, even for military
purposes."
In 1928, RAF College Cranwell cadet Frank Whittle formally submitted his ideas for a
turbo-jet to his superiors. In October 1929 he developed his ideas further. . On 16
January 1930 in England, Whittle submitted his first patent (granted in 1932). The
patent showed a two-stage axial compressor feeding a single-sided centrifugal
compressor. Practical axial compressors were made possible by ideas from A.A.Griffith
In 1935 Hans von Ohain started work on a similar design in Germany, apparently
unaware of Whittle's work. His first engine was strictly experimental and could only run
under external power, but he was able to demonstrate the basic concept. Ohain was
then introduced to Ernst Heinkel, one of the larger aircraft industrialists of the day, who
immediately saw the promise of the design. Heinkel had recently purchased the Hirth
engine company, and Ohain and his master machinist Max Hahn were set up there as a
new division of the Hirth company. They had their first HeS 1 centrifugal engine running
by September 1937. Unlike Whittle's design, Ohain used hydrogen as fuel, supplied
under external pressure. Their subsequent designs culminated in the gasoline-fuelled
HeS 3 of 1,100 lbf (5 kN), which was fitted to Heinkel's simple and compact He 178
airframe and flown by Erich Warsitz in the early morning of August 27, 1939, from
Marienehe aerodrome, an impressively short time for development. The He 178 was the
world's first jet plane.
Meanwhile, Whittle's engine was starting to look useful, and his Power Jets Ltd. started
receiving Air Ministry money. In 1941 a flyable version of the engine called the W.1,
capable of 1000 lbf (4 kN) of thrust, was fitted to the Gloster E28/39 airframe specially
built for it, and first flew on May 15, 1941 at RAF Cranwell
A British aircraft engine designer, Frank Halford, working from Whittle's ideas developed
a "straight through" version of the centrifugal jet; his design became the de Havilland
Goblin.
One problem with both of these early designs, which are called centrifugal-flow
engines, was that the compressor worked by "throwing" (accelerating) air outward from
the central intake to the outer periphery of the engine, where the air was then
compressed by a divergent duct setup, converting its velocity into pressure. An
advantage of this design was that it was already well understood, having been
implemented in centrifugal superchargers
Austrian Anselm Franz of Junkers' engine division (Junkers Motoren or Jumo) addressed
these problems with the introduction of the axial-flow compressor. Essentially, this is a
turbine in reverse. Air coming in the front of the engine is blown towards the rear of the
engine by a fan stage (convergent ducts), where it is crushed against a set of nonrotating blades called stators (divergent ducts). The process is nowhere near as
powerful as the centrifugal compressor, so a number of these pairs of fans and stators
are placed in series to get the needed compression. Even with all the added complexity,
the resulting engine is much smaller in diameter and thus, more aerodynamic. Jumo was
assigned the next engine number in the RLM numbering sequence, 4, and the result was
the Jumo 004 engine. After many lesser technical difficulties were solved, mass
production of this engine started in 1944 as a powerplant for the world's first jet-fighter
aircraft, the Messerschmitt Me 262 (and later the world's first jet-bomber aircraft, the
Arado Ar 234). A variety of reasons conspired to delay the engine's availability, this
delay caused the fighter to arrive too late to decisively impact Germany's position in
World War II. Nonetheless, it will be remembered as the first use of jet engines in
service.
In the UK, their first axial-flow engine, the Metrovick F.2, ran in 1941 and was first flown
in 1943. Although more powerful than the centrifugal designs at the time, the Ministry
considered its complexity and unreliability a drawback in wartime. The work at
Metrovick led to the Armstrong Siddeley Sapphire engine which would be built in the US
as the J65.
Following the end of the war the German jet aircraft and jet engines were extensively
studied by the victorious allies and contributed to work on early Soviet and US jet
fighters. The legacy of the axial-flow engine is seen in the fact that practically all jet
engines on fixed wing aircraft have had some inspiration from this design.
Centrifugal-flow engines have improved since their introduction. With improvements in
bearing technology the shaft speed of the engine was increased, greatly reducing the
diameter of the centrifugal compressor. The short engine length remains an advantage
of this design, particularly for use in helicopters where overall size is more important
than frontal area. Also, its engine components are robust; axial-flow compressors are
more liable to foreign object damage.
Although German designs were more advanced aerodynamically, the combination of
simplicity and advanced British metallurgy meant that Whittle-derived designs were far
more reliable than their German counterparts. British engines also were licensed widely
in the US (see Tizard Mission),and were sold to the USSR who reverse engineered them
with the Nene going on to power the famous MiG-15. American and Soviet designs,
independent axial-flow types for the most part, would not come fully into their own
until the 1960s, although the General Electric J47 provided excellent service in the F-86
Sabre in the 1950s.
By the 1950s the jet engine was almost universal in combat aircraft, with the exception
of cargo, liaison and other specialty types. By this point some of the British designs were
already cleared for civilian use, and had appeared on early models like the de Havilland
Comet and Canadair Jetliner. By the 1960s all large civilian aircraft were also jet
powered, leaving the piston engine in niche roles here as well.
Relentless improvements in the turboprop pushed the piston engine out of the
mainstream entirely, leaving it serving only the smallest general aviation designs, and
some use in drone aircraft. The ascension of the jet engine to almost universal use in
aircraft took well under twenty years.
However, the story was not quite at an end, for the efficiency of turbojet engines was
still rather worse than piston engines, but by the 1970s with the advent of high bypass
jet engines, an innovation not foreseen by the early commentators like Edgar
Buckingham, at high speeds and high altitudes that seemed absurd to them, only then
did the fuel efficiency finally exceeded that of the best piston and propeller engines, and
the dream of fast, safe, economical travel around the world finally arrived, and their
dour, if well founded for the time, predictions that jet engines would never amount to
much, killed forever.
Types
There are a large number of different types of jet engines, all of which achieve
propulsion from a high speed exhaust jet.
Type
Description
Advantages
Disadvantages
Water jet
Can run in shallow
water, high
acceleration, no risk
of engine overload
(unlike propellers),
less noise and
For propelling
vibration, highly
boats; squirts
manoeuvrable at all
water out the back
boat speeds, high
through a nozzle
speed efficiency, less
vulnerable to damage
from debris, very
reliable, more load
flexibility, less harmful
to wildlife
Can be less efficient
than a propeller at low
speed, more expensive,
higher weight in boat
due to entrained water,
will not perform well if
boat is heavier than the
jet is sized for
Motorjet
Most primitive
airbreathing jet
engine. Essentially
a supercharged
piston engine with
a jet exhaust.
Higher exhaust
velocity than a
propeller, offering
better thrust at high
speed
Heavy, inefficient and
underpowered
Turbojet
Generic term for
simple turbine
engine
Simplicity of design, A basic design, misses
efficient at supersonic many improvements in
speeds (~M2)
efficiency and power for
subsonic flight,
relatively noisy.
Low-bypass
Turbofan
One- or two-stage
fan added in front
bypasses a
proportion of the
air through a
bypass chamber
surrounding the
core. Compared
with its turbojet
ancestor, this
allows for more
efficient operation
with somewhat
less noise. This is
the engine of highspeed military
aircraft, some
smaller private
jets, and older
civilian airliners
such as the Boeing
707, the
McDonnell
Douglas DC-8, and
their derivatives.
High-bypass
Turbofan
First stage
compressor
drastically
enlarged to
provide bypass
airflow around
engine core, and it
provides
significant
amounts of thrust.
Compared to the
low-bypass
turbofan and nobypass turbojet,
the high-bypass
turbfan works on
As with the turbojet,
the design is
aerodynamic, with
only a modest
increase in diameter
over the turbojet
required to
accommodate the
bypass fan and
chamber. It is capable
of supersonic speeds
with minimal thrust
drop-off at high
speeds and altitudes
yet still more efficient
than the turbojet at
subsonic operation.
Noisier and less efficient
than high-bypass
turbofan, with less static
(Mach 0) thrust. Added
complexity to
accommodate dual
shaft designs. More
inefficient than a
turbojet around M2 due
to higher cross-sectional
area.
Quieter due to
greater mass flow and
lower total exhaust
speed, more efficient
for a useful range of
subsonic airspeeds for
same reason, cooler
exhaust temperature.
High bypass variants
exhibit good fuel
economy.
Greater complexity
(additional ducting,
usually multiple shafts)
and the need to contain
heavy blades. Fan
diameter can be
extremely large,
especially in high bypass
turbofans such as the
GE90. More subject to
FOD and ice damage.
Top speed is limited due
to the potential for
shockwaves to damage
engine. Thrust lapse at
higher speeds, which
the principle of
moving a great
deal of air
somewhat faster,
rather than a small
amount extremely
fast. This
translates into less
noise. Most
common form of
jet engine in
civilian use todayused in airliners
like the Boeing
747, most 737s,
and all Airbus
aircraft.
necessitates huge
diameters and
introduces additional
drag.
Rocket
Very few moving
parts, Mach 0 to Mach
25+, efficient at very
high speed (> Mach
10.0 or so),
thrust/weight ratio
over 100, no complex
air inlet, high
compression ratio,
Carries all
very high speed
propellants and
(hypersonic) exhaust,
oxidants on-board, good cost/thrust
emits jet for
ratio, fairly easy to
propulsion
test, works in a
vacuum-indeed works
best exoatmospheric
which is kinder on
vehicle structure at
high speed, fairly
small surface area to
keep cool, and no
turbine in hot exhaust
stream.
Needs lots of
propellant- very low
specific impulse —
typically 100-450
seconds. Extreme
thermal stresses of
combustion chamber
can make reuse harder.
Typically requires
carrying oxidiser onboard which increases
risks. Extraordinarily
noisy.
Ramjet
Intake air is
compressed
entirely by speed
Must have a high initial
speed to function,
inefficient at slow
Very few moving
parts, Mach 0.8 to
Mach 5+, efficient at
of oncoming air
and duct shape
(divergent)
Turboprop
(Turboshaft
similar)
Strictly not a jet at
all — a gas turbine
engine is used as
powerplant to
drive propeller
shaft (or rotor in
the case of a
helicopter)
Turboprop engine
drives one or more
Propfan/Unducted propellers. Similar
Fan
to a turbofan
without the fan
cowling.
Pulsejet
high speed (> Mach
2.0 or so), lightest of
all air-breathing jets
(thrust/weight ratio
up to 30 at optimum
speed), cooling much
easier than turbojets
as no turbine blades
to cool.
High efficiency at
lower subsonic
airspeeds (300 knots
plus), high shaft
power to weight
Limited top speed
(aeroplanes), somewhat
noisy, complex
transmission
Higher fuel efficiency,
potentially less noisy
than turbofans, could
lead to higher-speed
commercial aircraft,
popular in the 1980s
during fuel shortages
Development of propfan
engines has been very
limited, typically more
noisy than turbofans,
complexity
Air is compressed
and combusted
intermittently
Very simple design,
instead of
commonly used on
continuously.
model aircraft
Some designs use
valves.
Similar to a
pulsejet, but
combustion occurs
Pulse detonation as a detonation
Maximum theoretical
engine
instead of a
engine efficiency
deflagration, may
or may not need
valves
Air-augmented
Essentially a
speeds due to poor
compression ratio,
difficult to arrange shaft
power for accessories,
usually limited to a
small range of speeds,
intake flow must be
slowed to subsonic
speeds, noisy, fairly
difficult to test, finicky
to keep lit.
Noisy, inefficient (low
compression ratio),
works poorly on a large
scale, valves on valved
designs wear out quickly
Extremely noisy, parts
subject to extreme
mechanical fatigue, hard
to start detonation, not
practical for current use
Mach 0 to Mach 4.5+ Similar efficiency to
rocket
Scramjet
Turborocket
Precooled jets /
LACE
ramjet where
intake air is
compressed and
burnt with the
exhaust from a
rocket
(can also run
exoatmospheric),
good efficiency at
Mach 2 to 4
rockets at low speed or
exoatmospheric, inlet
difficulties, a relatively
undeveloped and
unexplored type,
cooling difficulties, very
noisy, thrust/weight
ratio is similar to
ramjets.
Still in development
stages, must have a very
high initial speed to
Similar to a ramjet
Few mechanical parts, function (Mach >6),
without a diffuser;
can operate at very
cooling difficulties, very
airflow through
high Mach numbers poor thrust/weight ratio
the entire engine
(Mach 8 to 15) with
(~2), extreme
remains
good efficiencies
aerodynamic
supersonic
complexity, airframe
difficulties, testing
difficulties/expense
A turbojet where
an additional
oxidizer such as
oxygen is added to
the airstream to
increase maximum
altitude
Very close to existing
designs, operates in
very high altitude,
wide range of altitude
and airspeed
Airspeed limited to
same range as turbojet
engine, carrying oxidizer
like LOX can be
dangerous. Much
heavier than simple
rockets.
Easily tested on
ground. Very high
Intake air is chilled thrust/weight ratios
to very low
are possible (~14)
temperatures at
together with good
inlet in a heat
fuel efficiency over a
exchanger before wide range of
passing through a airspeeds, mach 0ramjet or turbojet 5.5+; this combination
engine. Can be
of efficiencies may
combined with a permit launching to
rocket engine for orbit, single stage, or
orbital insertion. very rapid, very long
distance
intercontinental
Exists only at the lab
prototyping stage.
Examples include
RB545, SABRE, ATREX.
Requires liquid
hydrogen fuel which has
very low density and
heavily insulated
tankage.
travel.
All jet engines are reaction engines that generate thrust by emitting a jet of fluid
rearwards at relatively high speed. The forces on the inside of the engine needed to
create this jet give a strong thrust on the engine which pushes the craft forwards.
Jet engines make their jet from propellant from tankage that is attached to the engine
(as in a 'rocket') or from sucking in external fluid (very typically air) and expelling it at
higher speed; or more commonly, a combination of the two sources.
Thrust
The motion impulse of the engine is equal to the fluid mass multiplied by the speed at
which the engine emits this mass:
I=mc
where m is the fluid mass per second and c is the exhaust speed. In other words, a
vehicle gets the same thrust if it outputs a lot of exhaust very slowly, or a little exhaust
very quickly.
However, when an vehicle moves with certain velocity v, the fluid moves towards it,
creating an opposing ram drag at the intake:
mv
Most types of jet engine have an intake, which provides the bulk of the fluid exiting the
exhaust. Conventional rocket motors, however, do not have an intake, the oxidizer and
fuel both being carried within the vehicle. Therefore, rocket motors do not have ram
drag; the gross thrust of the nozzle is the net thrust of the engine. Consequently, the
thrust characteristics of a rocket motor are completely different from that of an air
breathing jet engine.
The jet engine with an intake is only useful if the velocity of the gas from the engine, c, is
greater than the vehicle velocity, v, as the net engine thrust is the same as if the gas
were emitted with the velocity c-v. So the thrust is actually equal to
S = m (c-v)
Energy efficiency
For all jet engines the propulsive efficiency (essentially energy efficiency) is highest when
the engine emits an exhaust jet at a speed that is the same as, or nearly the same as, the
vehicle velocity. The exact formula for air-breathing engines as given in the literature,
Noise
Noise is due to shockwaves that form when the exhaust jet interacts with the external
air.
The intensity of the noise is proportional to the thrust as well as proportional to the
fourth power of the jet velocity.
Generally then, the lower speed exhaust jets emitted from engines such as high bypass
turbofans are the quietest, whereas the fastest jets are the loudest.
Although some variation in jet speed can often be arranged from a jet engine (such as
by throttling back and adjusting the nozzle) it is difficult to vary the jet speed from an
engine over a very wide range. Therefore since engines for supersonic vehicles such as
Concorde, military jets and rockets inherently need to have supersonic exhaust at top
speed, so these vehicles are especially noisy even at low speeds.
Common types
A turbojet engine is a type of internal combustion engine often used to propel aircraft.
Air is drawn into the rotating compressor via the intake and is compressed, through
successive stages, to a higher pressure before entering the combustion chamber. Fuel is
mixed with the compressed air and ignited by flame in the eddy of a flame holder. This
combustion process significantly raises the temperature and volume of the air. Hot
combustion products leaving the combustor expand through a gas turbine, where
power is extracted to drive the compressor. This expansion process reduces both the
gas temperature and pressure but sufficient fuel is burnt so that both parameters are
usually still well above ambient conditions at exit from the turbine. The gas stream is
then expanded to ambient pressure via a propelling nozzle, producing a high velocity jet
as the exhaust. If the jet velocity exceeds the aircraft flight velocity, there is a net
forward thrust upon the airframe.
Under normal circumstances, the pumping action of the compressor prevents any
backflow, thus facilitating the continuous-flow process of the engine. Indeed, the entire
process is similar to a four-stroke cycle, but with induction, compression, ignition,
expansion and exhaust taking place simultaneously, but in different sections of the
engine. The efficiency of a jet engine is strongly dependent upon the overall pressure
ratio (combustor entry pressure/intake delivery pressure) and the turbine inlet
temperature of the cycle.
It is also perhaps instructive to compare turbojet engines with propeller engines.
Turbojet engines take a relatively small mass of air and accelerate it by a large amount,
whereas a propeller takes a large mass of air and accelerates it by a small amount. The
high-speed exhaust of a turbojet engine makes it efficient at high speeds (especially
supersonic speeds) and high altitudes. On slower aircraft and those required to fly short
stages, a gas turbine-powered propeller engine, commonly known as a turboprop, is
more common and much more efficient. Very small aircraft generally use conventional
piston engines to drive a propeller but small turboprops are getting smaller as
engineering technology improves.
The turbojet described above is a single-spool design, in which a single shaft connects
the turbine to the compressor. Two spool designs have two concentric turbinecompressor systems, that spin independently with the turbine and compressors for each
section connected from opposite ends of the engine via concentric shafts. This allows
for a higher compression ratio as well as improved compressor stability during engine
throttle movements. Three spool designs also exist.
Turbofan engines
Most modern jet engines are actually turbofans, where the low pressure compressor
acts as a fan, supplying supercharged air not only to the engine core, but to a bypass
duct. The bypass airflow either passes to a separate 'cold nozzle' or mixes with low
pressure turbine exhaust gases, before expanding through a 'mixed flow nozzle'.
Turbofans are used for airliners because they give an exhaust speed that is better
matched to subsonic airliner's flight speed, conventional turbojet engines generate an
exhaust that ends up travelling very fast backwards, and this wastes energy. By emitting
the exhaust so that it ends up travelling more slowly, better fuel consumption is
achieved. In addition, the lower exhaust speed gives much lower noise.
In the 1960s there was little difference between civil and military jet engines, apart from
the use of afterburning in some (supersonic) applications. Civil turbofans today have a
low exhaust speed (low specific thrust -net thrust divided by airflow) to keep jet noise to
a minimum and to improve fuel efficiency. Consequently the bypass ratio (bypass flow
divided by core flow) is relatively high (ratios from 4:1 up to 8:1 are common). Only a
single fan stage is required, because a low specific thrust implies a low fan pressure
ratio.
Today's military turbofans, however, have a relatively high specific thrust, to maximize
the thrust for a given frontal area, jet noise being of less concern in military uses relative
to civil uses. Multistage fans are normally needed to reach the relatively high fan
pressure ratio needed for high specific thrust. Although high turbine inlet temperatures
are often employed, the bypass ratio tends to be low, usually significantly less than 2.0.
An approximate equation for calculating the net thrust of a jet engine, be it a turbojet or
a mixed turbofan, is:
where:
intake mass flow rate
fully expanded jet velocity (in the exhaust plume)
aircraft flight velocity
While the
term represents the gross thrust of the nozzle, the
represents the ram drag of the intake.
term
Rocket engines
The third most common form of jet engine is the rocket engine.
Rocket engines are used for rockets because their extremely high exhaust velocity and
independence from the atmospheric oxygen permits them to achieve spaceflight.
This is used for launching satellites, space exploration and manned access, and
permitted landing on the moon in 1969.
However, the high exhaust speed and the heavier propellant mass results in less
efficient flight than turbojets, and their use is largely restricted to very high altitudes or
where very high accelerations are needed as rocket engines themselves have a very high
thrust-to-weight ratio.
Major components
The major components of a jet engine are similar across the major different types of
engines, although not all engine types have all components. The major parts include:

Cold Section:
o Air intake (Inlet) — The standard reference frame for a jet engine is the
aircraft itself. For subsonic aircraft, the air intake to a jet engine presents
no special difficulties, and consists essentially of an opening which is
designed to minimise drag, as with any other aircraft component.
However, the air reaching the compressor of a normal jet engine must be
travelling below the speed of sound, even for supersonic aircraft, to
sustain the flow mechanics of the compressor and turbine blades. At
supersonic flight speeds, shockwaves form in the intake system and
reduce the recovered pressure at inlet to the compressor. So some
supersonic intakes use devices, such as a cone or ramp, to increase


pressure recovery, by making more efficient use of the shock wave
system.
o Compressor or Fan — The compressor is made up of stages. Each stage
consists of vanes which rotate, and stators which remain stationary. As
air is drawn deeper through the compressor, its heat and pressure
increases. Energy is derived from the turbine (see below), passed along
the shaft.
o Bypass ducts much of the thrust of essentially all modern jet engines
comes from air from the front compressor that bypasses the combustion
chamber and gas turbine section that leads directly to the nozzle or
afterburner (where fitted).
Common:
o Shaft — The shaft connects the turbine to the compressor, and runs
most of the length of the engine. There may be as many as three
concentric shafts, rotating at independent speeds, with as many sets of
turbines and compressors. Other services, like a bleed of cool air, may
also run down the shaft.
Hot section:
o Combustor or Can or Flameholders or Combustion Chamber — This is a
chamber where fuel is continuously burned in the compressed air.
o Turbine — The turbine is a series of bladed discs that act like a windmill,
gaining energy from the hot gases leaving the combustor. Some of this
energy is used to drive the compressor, and in some turbine engines (ie
turboprop, turboshaft or turbofan engines), energy is extracted by
additional turbine discs and used to drive devices such as propellers,
bypass fans or helicopter rotors. One type, a free turbine, is configured
such that the turbine disc driving the compressor rotates independently
of the discs that power the external components. Relatively cool air, bled
from the compressor, may be used to cool the turbine blades and vanes,
to prevent them from melting.
o Afterburner or reheat (chiefly UK) — (mainly military) Produces extra
thrust by burning extra fuel, usually inefficiently, to significantly raise
Nozzle Entry Temperature at the exhaust. Owing to a larger volume flow
(i.e. lower density) at exit from the afterburner, an increased nozzle flow
area is required, to maintain satisfactory engine matching, when the
afterburner is alight.
o Exhaust or Nozzle — Hot gases leaving the engine exhaust to
atmospheric pressure via a nozzle, the objective being to produce a high
velocity jet. In most cases, the nozzle is convergent and of fixed flow
area.
o Supersonic nozzle — If the Nozzle Pressure Ratio (Nozzle Entry
Pressure/Ambient Pressure) is very high, to maximize thrust it may be
worthwhile, despite the additional weight, to fit a convergent-divergent
(de Laval) nozzle. As the name suggests, initially this type of nozzle is
convergent, but beyond the throat (smallest flow area), the flow area
starts to increase to form the divergent portion. The expansion to
atmospheric pressure and supersonic gas velocity continues downstream
of the throat, whereas in a convergent nozzle the expansion beyond sonic
velocity occurs externally, in the exhaust plume. The former process is
more efficient than the latter.
The various components named above have constraints on how they are put together to
generate the most efficiency or performance. The performance and efficiency of an
engine can never be taken in isolation; for example fuel/distance efficiency of a
supersonic jet engine maximises at about mach 2, whereas the drag for the vehicle
carrying it is increasing as a square law and has much extra drag in the transonic region.
The highest fuel efficiency for the overall vehicle is thus typically at Mach ~0.85.
For the engine optimisation for its intended use, important here is air intake design,
overall size, number of compressor stages (sets of blades), fuel type, number of exhaust
stages, metallurgy of components, amount of bypass air used, where the bypass air is
introduced, and many other factors. For instance, let us consider design of the air
intake.
Air intakes
Pitot intakes are the dominant type for subsonic applications. A subsonic pitot inlet is
little more than a tube with an aerodynamic fairing around it.
At zero airspeed (i.e., rest), air approaches the intake from a multitude of directions:
from directly ahead, radially, or even from behind the plane of the intake lip.
At low airspeeds, the streamtube approaching the lip is larger in cross-section than the
lip flow area, whereas at the intake design flight Mach number the two flow areas are
equal. At high flight speeds the streamtube is smaller, with excess air spilling over the
lip.
Beginning around 0.85 Mach, shock waves can occur as the air accelerates through the
intake throat.
Careful radiusing of the lip region is required to optimize intake pressure recovery (and
distortion) throughout the flight envelope.
Supersonic inlets
Supersonic intakes exploit shock waves to decelerate the airflow to a subsonic condition
at compressor entry.
There are basically two forms of shock waves:
1) Normal shock waves lie perpendicular to the direction of the flow. These form sharp
fronts and shock the flow to subsonic speeds. Microscopically the air molecules smash
into the subsonic crowd of molecules like alpha rays. Normal shock waves tend to cause
a large drop in stagnation pressure. Basically, the higher the supersonic entry Mach
number to a normal shock wave, the lower the subsonic exit Mach number and the
stronger the shock (i.e. the greater the loss in stagnation pressure across the shock
wave).
2) Conical (3-dimensional) and oblique shock waves (2D) are angled rearwards, like the
bow wave on a ship or boat, and radiate from a flow disturbance such as a cone or a
ramp. For a given inlet Mach number, they are weaker than the equivalent normal
shock wave and, although the flow slows down, it remains supersonic throughout.
Conical and oblique shock waves turn the flow, which continues in the new direction,
until another flow disturbance is encountered downstream.
Note: Comments made regarding 3 dimensional conical shock waves, generally also
apply to 2D oblique shock waves.
A sharp-lipped version of the pitot intake, described above for subsonic applications,
performs quite well at moderate supersonic flight speeds. A detached normal shock
wave forms just ahead of the intake lip and 'shocks' the flow down to a subsonic
velocity. However, as flight speed increases, the shock wave becomes stronger, causing
a larger percentage decrease in stagnation pressure (i.e. poorer pressure recovery). An
early US supersonic fighter, the F-100 Super Sabre, used such an intake.
More advanced supersonic intakes, excluding pitots:
a) exploit a combination of conical shock wave/s and a normal shock wave to improve
pressure recovery at high supersonic flight speeds. Conical shock wave/s are used to
reduce the supersonic Mach number at entry to the normal shock wave, thereby
reducing the resultant overall shock losses.
b) have a design shock-on-lip flight Mach number, where the conical/oblique shock
wave/s intercept the cowl lip, thus enabling the streamtube capture area to equal the
intake lip area. However, below the shock-on-lip flight Mach number, the shock wave
angle/s are less oblique, causing the streamline approaching the lip to be deflected by
the presence of the cone/ramp. Consequently, the intake capture area is less than the
intake lip area, which reduces the intake airflow. Depending on the airflow
characteristics of the engine, it may be desirable to lower the ramp angle or move the
cone rearwards to refocus the shockwaves onto the cowl lip to maximise intake airflow.
c) are designed to have a normal shock in the ducting downstream of intake lip, so that
the flow at compressor/fan entry is always subsonic. However, if the engine is throttled
back, there is a reduction in the corrected airflow of the LP compressor/fan, but (at
supersonic conditions) the corrected airflow at the intake lip remains constant, because
it is determined by the flight Mach number and intake incidence/yaw. This discontinuity
is overcome by the normal shock moving to a lower cross-sectional area in the ducting,
to decrease the Mach number at entry to the shockwave. This weakens the shockwave,
improving the overall intake pressure recovery. So, the absolute airflow stays constant,
whilst the corrected airflow at compressor entry falls (because of a higher entry
pressure). Excess intake airflow may also be dumped overboard or into the exhaust
system, to prevent the conical/oblique shock waves being disturbed by the normal
shock being forced too far forward by engine throttling.
Many second generation supersonic fighter aircraft featured an inlet cone, which was
used to form the conical shock wave. This type of inlet cone is clearly seen at the very
front of the English Electric Lightning and MiG-21 aircraft, for example.
The same approach can be used for air intakes mounted at the side of the fuselage,
where a half cone serves the same purpose with a semicircular air intake, as seen on the
F-104 Starfighter and BAC TSR-2.
Some intakes are biconic; that is they feature two conical surfaces: the first cone is
supplemented by a second, less oblique, conical surface, which generates an extra
conical shockwave, radiating from the junction between the two cones. A biconic intake
is usually more efficient than the equivalent conical intake, because the entry Mach
number to the normal shock is reduced by the presence of the second conical shock
wave.
A very sophisticated conical intake was featured on the SR-71's Pratt & Whitney J58s
that could move a conical spike fore and aft within the engine nacelle, preventing the
shockwave formed on the spike from entering the engine and stalling the engine, while
keeping it close enough to give good compression. Movable cones are uncommon.
A more sophisticated design than cones is to angle the intake so that one of its edges
forms a ramp. An oblique shockwave will form at the start of the ramp. The Century
Series of US jets featured several variants of this approach, usually with the ramp at the
outer vertical edge of the intake, which was then angled back inward towards the
fuselage. Typical examples include the Republic F-105 Thunderchief and F-4 Phantom.
Later this evolved so that the ramp was at the top horizontal edge rather than the outer
vertical edge, with a pronounced angle downwards and rearwards. This design
simplified the construction of intakes and allowed use of variable ramps to control
airflow into the engine. Most designs since the early 1960s now feature this style of
intake, for example the F-14 Tomcat, Panavia Tornado and Concorde.
From another point of view, like in a supersonic nozzle the corrected (or nondimensional) flow has to be the same at the intake lip, at the intake throat and at the
turbine. One of this three can be fixed. For inlets the throat is made variable and some
air is bypassed around the turbine and directly fed into the afterburner. Unlike in a
nozzle the inlet is either unstable or inefficient, because a normal shock wave in the
throat will suddenly move to the lip, thereby increasing the pressure at the lip, leading
to drag and reducing the pressure recovery, leading to turbine surge and the loss of one
SR-71.
Compressors
Axial compressors rely on spinning blades that have aerofoil sections, similar to
aeroplane wings. As with aeroplane wings in some conditions the blades can stall. If this
happens, the airflow around the stalled compressor can reverse direction violently. Each
design of a compressor has an associated operating map of airflow versus rotational
speed for characteristics peculiar to that type (see compressor map).
At a given throttle condition, the compressor operates somewhere along the steady
state running line. Unfortunately, this operating line is displaced during transients. Many
compressors are fitted with anti-stall systems in the form of bleed bands or variable
geometry stators to decrease the likelihood of surge. Another method is to split the
compressor into two or more units, operating on separate concentric shafts.
Another design consideration is the average stage loading. This can be kept at a sensible
level either by increasing the number of compression stages (more weight/cost) or the
mean blade speed (more blade/disc stress).
Although large flow compressors are usually all-axial, the rear stages on smaller units
are too small to be robust. Consequently, these stages are often replaced by a single
centrifugal unit. Very small flow compressors often employ two centrifugal
compressors, connected in series. Although in isolation centrifugal compressors are
capable of running at quite high pressure ratios (e.g. 10:1), impeller stress
considerations limit the pressure ratio that can be employed in high overall pressure
ratio engine cycles.
Increasing overall pressure ratio implies raising the high pressure compressor exit
temperature. This implies a higher high pressure shaft speed, to maintain the datum
blade tip Mach number on the rear compressor stage. Stress considerations, however,
may limit the shaft speed increase, causing the original compressor to throttle-back
aerodynamically to a lower pressure ratio than datum.
Combustors
Great care must be taken to keep the flame burning in a moderately fast moving
airstream, at all throttle conditions, as efficiently as possible. Since the turbine cannot
withstand stoichiometric temperatures (a mixture ratio of around 15:1), some of the
compressor air is used to quench the exit temperature of the combustor to an
acceptable level (an overall mixture ratio of between 45:1 and 130:1 is used). Air used
for combustion is considered to be primary airflow, while excess air used for cooling is
called secondary airflow. Combustor configurations include can, annular, and canannular.
Turbines
Because a turbine expands from high to low pressure, there is no such thing as turbine
surge or stall. The turbine needs fewer stages than the compressor, mainly because the
higher inlet temperature reduces the deltaT/T (and thereby the pressure ratio) of the
expansion process. The blades have more curvature and the gas stream velocities are
higher.
Designers must, however, prevent the turbine blades and vanes from melting in a very
high temperature and stress environment. Consequently bleed air extracted from the
compression system is often used to cool the turbine blades/vanes internally. Other
solutions are improved materials and/or special insulating coatings. The discs must be
specially shaped to withstand the huge stresses imposed by the rotating blades. They
take the form of impulse, reaction, or combination impulse-reaction shapes. Improved
materials help to keep disc weight down.
Turbopumps
Turbopumps are centrifugal pumps which are spun by gas turbines and are used to raise
the propellant pressure above the pressure in the combustion chamber so that it can be
injected and burnt. Turbopumps are very commonly used with rockets, but ramjets and
turbojets also have been known to use them.
Due to temperature limitations with the gas turbines, jet engines do not consume all the
oxygen in the air ('run stoichiometric'). Afterburners burn the remaining oxygen after
exiting the turbines, but usually do so inefficiently due to the low pressures typically
found at this part of the jet engine; however this gains significant thrust, which can be
useful. Engines intended for extended use with afterburners often have variable nozzles
and other details.
Nozzles
The primary objective of a nozzle is to expand the exhaust stream to atmospheric
pressure, and form it into a high speed jet to propel the vehicle. For airbreathing
engines, if the fully expanded jet has a higher speed than the aircraft's airspeed, then
there is a net rearward momentum gain to the air and there will be a forward thrust on
the airframe.
Simple convergent nozzles are used on many jet engines. If the nozzle pressure ratio is
above the critical value (about 1.8:1) a convergent nozzle will choke, resulting in some
of the expansion to atmospheric pressure taking place downstream of the throat (i.e.
smallest flow area), in the jet wake. Although much of the gross thrust produced will still
be from the jet momentum, additional (pressure) thrust will come from the imbalance
between the throat static pressure and atmospheric pressure.
Many military combat engines incorporate an afterburner (or reheat) in the engine
exhaust system. When the system is lit, the nozzle throat area must be increased, to
accommodate the extra exhaust volume flow, so that the turbomachinery is unaware
that the afterburner is lit. A variable throat area is achieved by moving a series of
overlapping petals, which approximate the circular nozzle cross-section.
At high nozzle pressure ratios, the exit pressure is often above ambient and much of the
expansion will take place downstream of a convergent nozzle, which is inefficient.
Consequently, some jet engines (notably rockets) incorporate a convergent-divergent
nozzle, to allow most of the expansion to take place against the inside of a nozzle to
maximise thrust. However, unlike the fixed con-di nozzle used on a conventional rocket
motor, when such a device is used on a turbojet engine it has to be a complex variable
geometry device, to cope with the wide variation in nozzle pressure ratio encountered
in flight and engine throttling. This further increases the weight and cost of such an
installation.
The simpler of the two is the ejector nozzle, which creates an effective nozzle through a
secondary airflow and spring-loaded petals. At subsonic speeds, the airflow constricts
the exhaust to a convergent shape. As the aircraft speeds up, the two nozzles dilate,
which allows the exhaust to form a convergent-divergent shape, speeding the exhaust
gasses past Mach 1. More complex engines can actually use a tertiary airflow to reduce
exit area at very low speeds. Advantages of the ejector nozzle are relative simplicity and
reliability. Disadvantages are average performance (compared to the other nozzle type)
and relatively high drag due to the secondary airflow. Notable aircraft to have utilized
this type of nozzle include the SR-71, Concorde, F-111, and Saab Viggen
For higher performance, it is necessary to use an iris nozzle. This type uses overlapping,
hydraulically adjustable "petals". Although more complex than the ejector nozzle, it has
significantly higher performance and smoother airflow. As such, it is employed primarily
on high-performance fighters such as the F-14, F-15, F-16, though is also used in highspeed bombers such as the B-1B. Some modern iris nozzles additionally have the ability
to change the angle of the thrust (see thrust vectoring).
Rocket motors also employ convergent-divergent nozzles, but these are usually of fixed
geometry, to minimize weight. Because of the much higher nozzle pressure ratios
experienced, rocket motor con-di nozzles have a much greater area ratio (exit/throat)
than those fitted to jet engines. The Convair F-106 Delta Dart has used such a nozzle
design, as part of its overall design specification as a aerospace interceptor for highaltitude bomber interception, where conventional nozzle design would prove
ineffective.
At the other extreme, some high bypass ratio civil turbofans use an extremely low area
ratio (less than 1.01 area ratio), convergent-divergent, nozzle on the bypass (or mixed
exhaust) stream, to control the fan working line. The nozzle acts as if it has variable
geometry. At low flight speeds the nozzle is unchoked (less than a Mach number of
unity), so the exhaust gas speeds up as it approaches the throat and then slows down
slightly as it reaches the divergent section. Consequently, the nozzle exit area controls
the fan match and, being larger than the throat, pulls the fan working line slightly away
from surge. At higher flight speeds, the ram rise in the intake increases nozzle pressure
ratio to the point where the throat becomes choked (M=1.0). Under these
circumstances, the throat area dictates the fan match and being smaller than the exit
pushes the fan working line slightly towards surge. This is not a problem, since fan surge
margin is much better at high flight speeds.
Thrust reversers
These either consist of cups that swing across the end of the nozzle and deflect the jet
thrust forwards (as in the DC-9), or they are two panels behind the cowling that slide
backward and reverse only the fan thrust (the fan produces the majority of the thrust).
This is the case on many large aircraft such as the 747, C-17, KC-135, etc.
Cooling systems
All jet engines require high temperature gas for good efficiency, typically achieved by
combusting hydrocarbon or hydrogen fuel. Combustion temperatures can be as high as
3500K (5841F) in rockets, far above the melting point of most materials, but normal
airbreathing jet engines use rather lower temperatures.
Cooling systems are employed to keep the temperature of the solid parts below the
failure temperature.
Air systems
A complex around combustor and is injected into the rim of the rotating turbine disc.
The cooling air then passes through complex passages within the turbine blades. After
removing heat from the blade material, the air (now fairly hot) is vented, via cooling
holes, into the main gas stream. Cooling air for the turbine vanes undergoes a similar
process.
Cooling the leading edge of the blade can be difficult, because the pressure of the
cooling air just inside the cooling hole may not be much different from that of the
oncoming gas stream. One solution is to incorporate a cover plate on the disc. This acts
as a centrifugal compressor to pressurize the cooling air before it enters the blade.
Another solution is to use an ultra-efficient turbine rim seal to pressurize the area where
the cooling air passes across to the rotating disc.
Seals are used to prevent oil leakage, control air for cooling and prevent stray air flows
into turbine cavities.
A series of (e.g. labyrinth) seals allow a small flow of bleed air to wash the turbine disc
to extract heat and, at the same time, pressurize the turbine rim seal, to prevent hot
gases entering the inner part of the engine. Other types of seals are hydraulic, brush,
carbon etc.
Small quantities of compressor bleed air are also used to cool the shaft, turbine shrouds,
etc. Some air is also used to keep the temperature of the combustion chamber walls
below critical. This is done using primary and secondary airholes which allow a thin layer
of air to cover the inner walls of the chamber preventing excessive heating.
Exit temperature is dependent on the turbine upper temperature limit depending on
the material. Reducing the temperature will also prevent thermal fatigue and hence
failure. Accessories may also need their own cooling systems using air from the
compressor or outside air.
Air from compressor stages is also used for heating of the fan, airframe anti-icing and
for cabin heat. Which stage is bled from depends on the atmospheric conditions at that
altitude.
Fuel system
Apart from providing fuel to the engine, the fuel system is also used to control propeller
speeds, compressor airflow and cool lubrication oil. Fuel is usually introduced by an
atomized spray, the amount of which is controlled automatically depending on the rate
of airflow.
So the sequence of events for increasing thrust is, the throttle opens and fuel spray
pressure is increased, increasing the amount of fuel being burned. This means that
exhaust gases are hotter and so are ejected at higher acceleration, which means they
exert higher forces and therefore increase the engine thrust directly. It also increases
the energy extracted by the turbine which drives the compressor even faster and so
there is an increase in air flowing into the engine as well.
Obviously, it is the rate of the mass of the airflow that matters since it is the change in
momentum (mass x velocity) that produces the force. However, density varies with
altitude and hence inflow of mass will also vary with altitude, temperature etc. which
means that throttle values will vary according to all these parameters without changing
them manually.
This is why fuel flow is controlled automatically. Usually there are 2 systems, one to
control the pressure and the other to control the flow. The inputs are usually from
pressure and temperature probes from the intake and at various points through the
engine. Also throttle inputs, engine speed etc. are required. These affect the high
pressure fuel pump.
Fuel control unit (FCU)
This element is something like a mechanical computer. It determines the output of the
fuel pump by a system of valves which can change the pressure used to cause the pump
stroke, thereby varying the amount of flow.
Take the possibility of increased altitude where there will be reduced air intake
pressure. In this case, the chamber within the FCU will expand which causes the spill
valve to bleed more fuel. This causes the pump to deliver less fuel until the opposing
chamber pressure is equivalent to the air pressure and the spill valve goes back to its
position.
When the throttle is opened, it releases i.e. lessens the pressure which lets the throttle
valve fall. The pressure is transmitted (because of a back-pressure valve i.e. no air gaps
in fuel flow) which closes the FCU spill valves (as they are commonly called) which then
increases the pressure and causes a higher flow rate.
The engine speed governor is used to prevent the engine from over-speeding. It has the
capability of disregarding the FCU control. It does this by use of a diaphragm which
senses the engine speed in terms of the centrifugal pressure caused by the rotating
rotor of the pump. At a critical value, this diaphragm causes another spill valve to open
and bleed away the fuel flow.
There are other ways of controlling fuel flow for example with the dash-pot throttle
lever. The throttle has a gear which meshes with the control valve (like a rack and
pinion) causing it to slide along a cylinder which has ports at various positions. Moving
the throttle and hence sliding the valve along the cylinder, opens and closes these ports
as designed. There are actually 2 valves viz. the throttle and the control valve. The
control valve is used to control pressure on one side of the throttle valve such that it
gives the right opposition to the throttle control pressure. It does this by controlling the
fuel outlet from within the cylinder.
So for example, if the throttle valve is moved up to let more fuel in, it will mean that the
throttle valve has moved into a position which allows more fuel to flow through and on
the other side, the required pressure ports are opened to keep the pressure balance so
that the throttle lever stays where it is.
At initial acceleration, more fuel is required and the unit is adapted to allow more fuel
to flow by opening other ports at a particular throttle position. Changes in pressure of
outside air i.e. altitude, speed of aircraft etc are sensed by an air capsule.
Fuel pump
Fuel pumps are used to raise the fuel pressure above the pressure in the combustion
chamber so that the fuel can be injected. Fuel pumps are usually driven by the main
shaft, via gearing.
Turbopumps are very commonly used with liquid-fuelled rockets and rely on the
expansion of an onboard gas through a turbine.
Ramjet turbopumps use ram air expanding through a turbine.
Engine starting system
The fuel system as explained above, is one of the 2 systems required for starting the
engine. The other is the actual ignition of the air/fuel mixture in the chamber. Usually,
an auxiliary power unit is used to start the engines. It has a starter motor which has a
high torque transmitted to the compressor unit. When the optimum speed is reached,
i.e. the flow of gas through the turbine is sufficient, the turbines take over. There are a
number of different starting methods such as electric, hydraulic, pneumatic etc.
The electric starter works with gears and clutch plate linking the motor and the engine.
The clutch is used to disengage when optimum speed is achieved. This is usually done
automatically. The electric supply is used to start the motor as well as for ignition. The
voltage is usually built up slowly as starter gains speed.
Some military aircraft need to be started quicker than the electric method permits and
hence they use other methods such as a turbine starter. This is an impulse turbine
impacted by burning gases from a cartridge. It is geared to rotate the engine and also
connected to an automatic disconnect system. The cartridge is set alight electrically and
used to turn the turbine.
Another turbine starter system is almost exactly like a little engine. Again the turbine is
connected to the engine via gears. However, the turbine is turned by burning gases usually the fuel is isopropyl nitrate stored in a tank and sprayed into a combustion
chamber. Again, it is ignited with a spark plug. Everything is electrically controlled, such
as speed etc.
Most Commercial aircraft and large Military Transport airplanes usually use what is
called an auxiliary power unit or APU. It is normally a small gas turbine. Thus, one could
say that using such an APU is using a small gas turbine to start a larger one. High
pressure air from the compressor section of the APU is bled off through a system of
pipes to the engines where it is directed into the starting system. This "bleed air" is
directed into a mechanism to start the engine turning and begin pulling in air. When the
rotating speed of the engine is sufficient to pull in enough air to support combustion,
fuel is introduced and ignited. Once the engine ignites and reaches idle speed, the bleed
air is shut off.
The APUs on aircraft such as the Boeing 737 and Airbus A320 can be seen at the
extreme rear of the aircraft. This is the typical location for an APU on most commercial
airliners although some may be within the wing root (Boeing 727) or the aft fuselage
(DC-9/MD80) as examples and some military transports carry their APU's in one of the
main landing gear pods (C-141).
The APUs also provide enough power to keep the cabin lights, pressure and other
systems on while the engines are off. The valves used to control the airflow are usually
electrically controlled. They automatically close at a pre-determined speed. As part of
the starting sequence on some engines fuel is combined with the supplied air and
burned instead of using just air. This usually produces more power per unit weight.
Usually an APU is started by its own electric starter motor which is switched off at the
proper speed automatically. When the main engine starts up and reaches the right
conditions, this auxiliary unit is then switched off and disengages slowly.
Hydraulic pumps can also be used to start some engines through gears. The pumps are
electrically controlled on the ground.
A variation of this is the APU installed in a Boeing F/A-18 Hornet; it is started by a
hydraulic motor, which itself receives energy stored in an accumulator. This
accumulator is recharged after the right engine is started and develops hydraulic
pressure, or by a hand pump in the right hand main landing gear well.
Ignition
Usually there are 2 igniter plugs in different positions in the combustion system. A high
voltage spark is used to ignite the gases. The voltage is stored up from a low voltage
supply provided by the starter system. It builds up to the right value and is then released
as a high energy spark. Depending on various conditions, the igniter continues to
provide sparks to prevent combustion from failing if the flame inside goes out. Of
course, in the event that the flame does go out, there must be provision to relight.
There is a limit of altitude and air speed at which an engine can obtain a satisfactory
relight.
For example, the General Electric F404-400 uses one ignitor for the combustor and one
for the afterburner; the ignition system for the A/B incorporates an ultraviolet flame
sensor to activate the ignitor.
It should be noted that most modern ignition systems provide enough energy to be a
lethal hazard should a person be in contact with the electrical lead when the system is
activated, so team communication is vital when working on these systems.
Lubrication system
A lubrication system serves to ensure lubrication of the bearings and to maintain
sufficiently cool temperatures, mostly by eliminating friction.
The lubrication system as a whole should be able to prevent foreign material from
entering the plane, and reaching the bearings, gears, and other moving parts. The
lubricant must be able to flow easily at relatively low temperatures and not disintegrate
or break down at very high temperatures.
Usually the lubrication system has subsystems that deal individually with the pressure of
an engine, scavenging, and a breather.
The pressure system components are an oil tank and de-aerator, main oil pump, main
oil filter/filter bypass valve, pressure regulating valve (PRV), oil cooler/by pass valve and
tubing/jets.
Usually the flow is from the tank to the pump inlet and PRV, pumped to main oil filter or
its bypass valve and oil cooler, then through some more filters to jets in the bearings.
Using the PRV method of control, means that the pressure of the feed oil must be below
a critical value (usually controlled by other valves which can leak out excess oil back to
tank if it exceeds the critical value). The valve opens at a certain pressure and oil is kept
moving at a constant rate into the bearing chamber.
If the engine speed increases, the pressure within the bearing chamber also increases,
which means the pressure difference between the lubricant feed and the chamber
reduces which could reduce slow rate of oil when it is needed even more. As a result,
some PRVs can adjust their spring force values using this pressure change in the bearing
chamber proportionally to keep the lubricant flow constant.
Advanced designs
J-58 combined ramjet/turbojet
The SR-71's Pratt & Whitney J58 engines were rather unusual. They could convert in
flight from being largely a turbojet to being largely a compressor-assisted ramjet. At
high speeds (above Mach 2.4), the engine used variable geometry vanes to direct excess
air through 6 bypass pipes from downstream of the fourth compressor stage into the
afterburner. 80% of the SR-71's thrust at high speed was generated in this way, giving
much higher thrust, improving specific impulse by 10-15%, and permitting continuous
operation at Mach 3.2. The name coined for this setup is turbo-ramjet.
Hydrogen fuelled jet engines
Jet engines can be run on almost any fuel. Hydrogen is a highly desirable fuel, as,
although the energy per mole is not unusually high, the molecule is very much lighter
than other molecules. It turns out that the energy per kg of hydrogen is twice that of
more common fuels and this gives twice the specific impulse. In addition jet engines
running on hydrogen are quite easy to build- the first ever turbojet was run on
hydrogen.
However, in almost every other way, hydrogen is problematic. The downside of
hydrogen is its density, in gaseous form the tanks are impractical for flight, but even in
liquid form it has a density one fourteenth that of water. It is also deeply cryogenic and
requires very significant insulation that precludes it being stored in wings. The overall
vehicle ends up very large, and they would be difficult for most airports to
accommodate. Finally, pure hydrogen is not found in nature, and must be manufactured
either via steam reforming or expensive electrolysis. Both are relatively inefficient
processes.
Precooled jet engines
An idea originated by Robert P. Carmichael in 1955 is that hydrogen fuelled engines
could theoretically have much higher performance than hydrocarbon fuelled engines if a
heat exchanger were used to cool the incoming air. The low temperature allows lighter
materials to be used, a higher mass-flow through the engines, and permits combustors
to inject more fuel without overheating the engine.
This idea leads to plausible designs like SABRE, that might permit single-stage-to-orbit,
and ATREX, that might permit jet engines to be used up to hypersonic speeds and high
altitudes for boosters for launch vehicles. The idea is also being researched by the EU for
a concept to achieve non-stop antipodal supersonic passenger travel at Mach 5
(Reaction Engines A2).
Nuclear-powered ramjet
Project Pluto was a nuclear-powered ramjet, intended for use in a cruise missile. Rather
than combusting fuel as in regular jet engines, air was heated using a high-temperature,
unshielded nuclear reactor. This dramatically increased the engine burn time, and the
ramjet was predicted to be able to cover any required distance at supersonic speeds
(Mach 3 at tree-top height).
However, there was no obvious way to stop it once it had taken off, which would be a
great disadvantage in any non-disposable application. Also, because the reactor was
unshielded, it was dangerous to be in or around the flight path of the vehicle (although
the exhaust itself wasn't radioactive). These disadvantages limit the application to
warhead delivery system for all-out nuclear war, which it was being designed for.
Scramjets
Scramjets are an evolution of ramjets that are able to operate at much higher speeds
than any other kind of airbreathing engine. They share a similar structure with ramjets,
being a specially-shaped tube that compresses air with no moving parts through ram-air
compression. Scramjets, however, operate with supersonic airflow through the entire
engine. Thus, scramjets do not have the diffuser required by ramjets to slow the
incoming airflow to subsonic speeds.
Scramjets start working at speeds of at least Mach 4, and have a maximum useful speed
of approximately Mach 17. Due to aerodynamic heating at these high speeds, cooling
poses a challenge to engineers.
Environmental considerations
Jet engines are usually run on fossil fuel propellant, and in that case, are a net source of
carbon to the atmosphere.
Some scientists believe that jet engines are also a source of global dimming due to the
water vapour in the exhaust causing cloud formations.
Nitrogen compounds are also formed from the combustion process from atmospheric
nitrogen. At low altitudes this is not thought to be especially harmful, but for supersonic
aircraft that fly in the stratosphere some destruction of ozone may occur.
Sulphates are also emitted if the fuel contains sulphur.
Safety and reliability
Jet engines are usually very reliable and have a very good safety record. However
failures do sometimes occur.
One class of failures that has caused accidents in particular is uncontained failures,
where rotary parts of the engine break off and exit through the case. These can cut fuel
or control lines, and can penetrate the cabin. Although fuel and control lines are usually
duplicated for reliability the United Airlines Flight 232 was caused when all control lines
were simultaneously severed.
The most likely failure is compressor blade failure, and modern jet engines are designed
with structures that can catch these blades and keep them contained them within the
engine casing. Verification of a jet engine design involves testing that this system works
correctly.
Bird strike
Bird strike is an aviation term for a collision between a bird and an aircraft. It is a
common threat to aircraft safety and has caused a number of fatal accidents. In 1988 an
Ethiopian Airlines Boeing 737 sucked pigeons into both engines during take-off and then
crashed in an attempt to return to the Bahir Dar airport; of the 104 people aboard, 35
died and 21 were injured. In another incident in 1995, a Dassault Falcon 20 crashed at a
Paris airport during an emergency landing attempt after sucking lapwings into an
engine, which caused an engine failure and a fire in the airplane fuselage; all 10 people
on board were killed.
Modern jet engines have the capability of surviving an ingestion of a bird. Small fast
planes, such as military jet fighters, are at higher risk than big heavy multi-engine ones.
This is due to the fact that the fan of a high-bypass turbofan engine, typical on transport
aircraft, acts as a centrifugal separator to force ingested materials (birds, ice, etc.) to the
outside of the fan's disc. As a result, such materials go through the relatively
unobstructed bypass duct, rather than through the core of the engine, which contains
the smaller and more delicate compressor blades. Military aircraft designed for highspeed flight typically have pure turbojet, or low-bypass turbofan engines, increasing the
risk that ingested materials will get into the core of the engine to cause damage.
The highest risk of the bird strike is during the takeoff and landing, in low altitudes,
which is in the vicinity of the airports.
Rocket
A rocket or rocket vehicle is a missile, aircraft or other vehicle which obtains thrust by
the reaction of the rocket to the ejection of fast moving fluid from a rocket engine.
Chemical rockets work by the action of hot gas produced by the combustion of the
propellant against the inside of combustion chambers and expansion nozzles. This
generates forces that accelerate the gas to extremely high speed and exerts a large
thrust on the rocket (since every action has an equal and opposite reaction).
The history of rockets goes back to at least the 13th century. By the 20th century, they
have enabled human spaceflight to the Moon. In the 21st century, they have made
commercial space tourism possible.
Rockets are used for fireworks and weaponry, as launch vehicles for artificial satellites,
human spaceflight and exploration of other planets. While inefficient for low speed use,
they are, compared to other propulsion systems, very lightweight and powerful, capable
of attaining extremely high speeds with reasonable efficiency.
Chemical rockets store a large amount of energy in an easily-released form, and can be
very dangerous. However, careful design, testing, construction, and use minimizes the
risks.
In antiquity
According to the writings of the Roman Aulus Gellius, in c. 400 BC, a Greek Pythagorean
named Archytas propelled a wooden bird using steam. However, the only knowledge
that exists of it is in Aulus's writings, which dates from 5 centuries later. No diagrams
survive, and whether it was truly propelled by rocket power is unknown.
The availability of black powder (gunpowder) to propel projectiles was a precursor to
the development of the first solid rocket. Ninth Century Chinese Taoist alchemists
discovered black powder while searching for the Elixir of life; this accidental discovery
led to experiments in the form of weapons like bombs, cannon, incendiary fire arrows
and rocket-propelled fire arrows.
Exactly when the first flights of rockets occurred is contested. Some say that the first
recorded use of a rocket in battle was by the Chinese in 1232 against the Mongol
hordes. There were reports of fire arrows and 'iron pots' that could be heard for 5
leagues (15 miles) when they exploded upon impact, causing devastation for a radius of
2,000 feet, apparently due to shrapnel. The lowering of the iron pots may have been a
way for a besieged army to blow up invaders. The fire arrows were either arrows with
explosives attached, or arrows propelled by gunpowder, such as the Korean Hwacha.
Less controversially, one of the earliest devices recorded that used internal-combustion
rocket propulsion was the 'ground-rat,' a type of firework, recorded in 1264 as having
frightened the Empress-Mother Kung Sheng at a feast held in her honor by her son the
Emperor Lizong.
Subsequently, one of the earliest texts to mention the use of rockets was the
Huolongjing, written by the Chinese artillery officer Jiao Yu in the mid-14th century. This
text also mentioned the use of the first known multistage rocket, the 'fire-dragon issuing
from the water' (huo long chu shui), used mostly by the Chinese navy. Frank H. Winter
proposed in The Proceedings of the Twentieth and Twenty-First History Symposia of the
International Academy of Astronautics that southern China and the Laotian community
rocket festivals might have been key in the subsequent spread of rocketry in the Orient.
Spread of rocket technology
Rocket technology first became known to Europeans following their use by the Mongols
Genghis Khan and Ögedei Khan when they conquered parts of Russia, Eastern, and
Central Europe. The Mongolians had acquired the Chinese technology by conquest of
the northern part of China and also by the subsequent employment of Chinese rocketry
experts as mercenaries for the Mongol military. Reports of the Battle of Sejo in the year
1241 describe the use of rocket-like weapons by the Mongols against the Magyars.
Rocket technology also spread to Korea, with the 15th century wheeled hwacha that
would launch singijeon rockets. These first Korean rockets had an amazingly long range
at the time, and were designed and built by Byun Eee-Joong. They were just like arrows
but had small explosives attached to the back, and were fired in swarms.
Additionally, the spread of rockets into Europe was also influenced by the Ottomans at
the siege of Constantinople in 1453, although it is very likely that the Ottomans
themselves were influenced by the Mongol invasions of the previous few centuries.
They appear in literature describing the capture of Baghdad in 1258 by the Mongols.
In their history of rockets published on the Internet, NASA says “the Arabs adopted the
rocket into their own arms inventory and, during the Seventh Crusade, used them
against the French Army of King Louis IX in 1268.".
The name Rocket comes from the Italian Rocchetta (i.e. little fuse), a name of a small
firecracker created by the Italian artificer Muratori in 1379.
"Artis Magnae Artilleriae pars prima" ("Great Art of Artillery, the First Part", also known
as "The Complete Art of Artillery"), first printed in Amsterdam in 1650, was translated to
French in 1651, German in 1676, English and Dutch in 1729 and Polish in 1963. For over
two centuries, this work of Polish-Lithuanian Commonwealth nobleman Kazimierz
Siemienowicz was used in Europe as a basic artillery manual. The book provided the
standard designs for creating rockets, fireballs, and other pyrotechnic devices. It
contained a large chapter on caliber, construction, production and properties of rockets
(for both military and civil purposes), including multi-stage rockets, batteries of rockets,
and rockets with delta wing stabilizers (instead of the common guiding rods).
In 1792, iron-cased rockets were successfully used militarily by Tipu Sultan, Ruler of the
Kingdom of Mysore in India against the larger British East India Company forces during
the Anglo-Mysore Wars. The British then took an active interest in the technology and
developed it further during the 19th century. The major figure in the field at this time
was William Congreve. From there, the use of military rockets spread throughout
Europe. At the Battle of Baltimore in 1814, the rockets fired on Fort McHenry by the
rocket vessel HMS Erebus were the source of the rockets' red glare described by Francis
Scott Key in The Star-Spangled Banner. Rockets were also used in the Battle of
Waterloo.
Accuracy of early rockets
Early rockets were very inaccurate. Without the use of spinning or any gimballing of the
thrust, they had a strong tendency to veer sharply off course. The early British Congreve
rockets reduced this somewhat by attaching a long stick to the end of a rocket (similar
to modern bottle rockets) to make it harder for the rocket to change course. The largest
of the Congreve rockets was the 32-pound (14.5 kg) Carcass, which had a 15-foot (4.6
m) stick. Originally, sticks were mounted on the side, but this was later changed to
mounting in the center of the rocket, reducing drag and enabling the rocket to be more
accurately fired from a segment of pipe.
The British were greatly impressed by the Mysorean Rocket artillery made from iron
tubes used by the armies of Tipu Sultan and his father, Haidar Ali. Tipu Sultan
championed the use of mass attacks with rocket brigades in the army. The effect of
these weapons on the British during the Second, Third and Fourth Mysore Wars was
sufficiently impressive to inspire William Congreve to develop his own rocket designs.
Several Mysore rockets were sent to England, and after thoroughly examining the Indian
specimens, from 1801, William Congreve, son of the Comptroller of the Royal Arsenal,
Woolwich, London, set on a vigorous research and development programme at the
Arsenal's laboratory. Congreve prepared a new propellant mixture, and developed a
rocket motor with a strong iron tube with conical nose, weighing about 32 pounds (14.5
kilograms). The Royal Arsenal's first demonstration of solid fuel rockets was in 1805. The
rockets were effectively used during the Napoleonic Wars and the War of 1812.
Congreve published three books on rocketry.
In 1815, Alexander Dmitrievich Zasyadko began his work on creating military gunpowder
rockets. He constructed rocket-launching platforms, which allowed to fire in salvos (6
rockets at a time), and gun-laying devices. Zasyadko elaborated a tactic for military use
of rocket weaponry. In 1820, Zasyadko was appointed head of the Petersburg Armory,
Okhtensky Powder Factory, pyrotechnic laboratory and the first Highest Artillery School
in Russia. He organized rocket production in a special rocket workshop and created the
first rocket sub-unit in the Russian army.
The accuracy problem was mostly solved in 1844 when William Hale modified the rocket
design so that thrust was slightly vectored, causing the rocket to spin along its axis of
travel like a bullet. The Hale rocket removed the need for a rocket stick, travelled
further due to reduced air resistance, and was far more accurate.
Early manned rocketry
According to legend, a manned rocket sled with 47 gunpowder-filled rockets was
attempted in China by Wan Hu in the 16th Century. The alleged flight is said to have
been interrupted by an explosion at the start, and the pilot did not seem to have
survived (he was never found). There are no known Chinese sources for this event, and
the earliest known account is an unsourced reference in a book by an American, Herbert
S. Zim in 1945.
In Ottoman Turkey in 1633, Lagari Hasan Çelebi took off with what was described as a
cone-shaped rocket, glided with wings through Bosporus from Topkap Palace, and made
a successful landing, winning him a position in the Ottoman army. The flight was
accomplished as a part of celebrations performed for the birth of Ottoman Emperor
Murat IV's daughter and was rewarded by the sultan. The device was composed of a
large winged cage with a conical top with 7 rockets filled with 70 kg of gunpowder. The
flight was estimated to have lasted about 200 seconds and the maximum height
reached around 300 metres.
Theories of interplanetary rocketry
In 1903, high school mathematics teacher Konstantin Tsiolkovsky (1857-1935) published
Исследование мировых пространств реактивными приборами (The Exploration of
Cosmic Space by Means of Reaction Devices), the first serious scientific work on space
travel. The Tsiolkovsky rocket equation—the principle that governs rocket propulsion—
is named in his honor (although it had been discovered previously). His work was
essentially unknown outside the Soviet Union, where it inspired further research,
experimentation and the formation of the Cosmonautics Society.
In 1920, Robert Goddard published A Method of Reaching Extreme Altitudes, the first
serious work on using rockets in space travel after Tsiolkovsky. The work attracted
worldwide attention and was both praised and ridiculed, particularly because of its
suggestion that a rocket theoretically could reach the Moon. A New York Times editorial
famously expressed disbelief that it was possible at all as it stated that: "after the rocket
quits our air and really starts on its longer journey it will neither be accelerated nor
maintained by the explosion of the charges it then might have left" and suggested that
Professor Goddard actually: "does not know of the relation of action to reaction, and the
need to have something better than a vacuum against which to react" and talked of
"such things as intentional mistakes or oversights."
Goddard, the Times declared, apparently suggesting bad faith, "only seems to lack the
knowledge ladled out daily in high schools."
After these and other scathing criticisms, Goddard began working in isolation, and
avoided publicity.
Nevertheless, in Russia, Tsiolkovsky's work was republished in the 1920s in response to
Russian interest raised by the work of Robert Goddard. Among other ideas, Tsiolkovsky
accurately proposed to use liquid oxygen and liquid hydrogen as a nearly optimal
propellant pair and determined that building staged and clustered rockets to increase
the overall mass efficiency would dramatically increase range.
In 1923, Hermann Oberth (1894-1989) published Die Rakete zu den Planetenräumen
("The Rocket into Planetary Space"), a version of his doctoral thesis, after the University
of Munich rejected it.
Modern rocketry
Modern rockets were born when Goddard attached a supersonic (de Laval) nozzle to a
liquid fuelled rocket engine's combustion chamber. These nozzles turn the hot gas from
the combustion chamber into a cooler, hypersonic, highly directed jet of gas, more than
doubling the thrust and raising the engine efficiency from 2% to 64%. Early rockets had
been grossly inefficient because of the thermal energy that was wasted in the exhaust
gases. In 1926, Robert Goddard launched the world's first liquid-fueled rocket in
Auburn, Massachusetts.
During the 1920s, a number of rocket research organizations appeared in America,
Austria, Britain, Czechoslovakia, France, Italy, Germany, and Russia. In the mid-1920s,
German scientists had begun experimenting with rockets which used liquid propellants
capable of reaching relatively high altitudes and distances. 1927 the German car
manufacturer Opel began to research with rockets together with Mark Valier and the
rocket builder Friedrich Wilhelm Sander. In 1928, Fritz von Opel drove with a rocket car,
the Opel RAK1 on the Opel raceway in Rüsselsheim, Germany. In 1929 von Opel started
at the Frankfurt-Rebstock airport with the Opel-Sander RAK 1-airplane. This was maybe
the first flight with a manned rocket-aircraft. In 1927 and also in Germany, a team of
amateur rocket engineers had formed the Verein für Raumschiffahrt (German Rocket
Society, or VfR), and in 1931 launched a liquid propellant rocket (using oxygen and
gasoline).
From 1931 to 1937, the most extensive scientific work on rocket engine design occurred
in Leningrad, at the Gas Dynamics Laboratory. Well-funded and staffed, over 100
experimental engines were built under the direction of Valentin Glushko. The work
included regenerative cooling, hypergolic propellant ignition, and fuel injector designs
that included swirling and bi-propellant mixing injectors. However, the work was
curtailed by Glushko's arrest during Stalinist purges in 1938. Similar work was also done
by the Austrian professor Eugen Sänger who worked on rocket powered spaceplanes
such as Silbervogel (sometimes called the 'antipodal' bomber.)
On November 12, 1932 at a farm in Stockton NJ, the American Interplanetary Society's
attempt to static fire their first rocket (based on German Rocket Society designs) fails in
a fire.
In 1932, the Reichswehr (which in 1935 became the Wehrmacht) began to take an
interest in rocketry. Artillery restrictions imposed by the Treaty of Versailles limited
Germany's access to long distance weaponry. Seeing the possibility of using rockets as
long-range artillery fire, the Wehrmacht initially funded the VfR team, but seeing that
their focus was strictly scientific, created its own research team, with Hermann Oberth
as a senior member. At the behest of military leaders, Wernher von Braun, at the time a
young aspiring rocket scientist, joined the military (followed by two former VfR
members) and developed long-range weapons for use in World War II by Nazi Germany,
notably the A-series of rockets, which led to the infamous V-2 rocket (initially called A4).
World War II
In 1943, production of the V-2 rocket began. The V-2 had an operational range of 300
km (185 miles) and carried a 1000 kg (2204 lb) warhead, with an amatol explosive
charge. Highest point of altitude of its flight trajectory is 90 km. The vehicle was only
different in details from most modern rockets, with turbopumps, inertial guidance and
many other features. Thousands were fired at various Allied nations, mainly England, as
well as Belgium and France. While they could not be intercepted, their guidance system
design and single conventional warhead meant that the V-2 was insufficiently accurate
against military targets. The later versions however, were more accurate, sometimes
within metres, and could be devastating. 2,754 people in England were killed, and 6,523
were wounded before the launch campaign was terminated. While the V-2 did not
significantly affect the course of the war, it provided a lethal demonstration of the
potential for guided rockets as weapons.
Under Projekt Amerika Nazi Germany also tried to develop and use the first submarinelaunched ballistic missile (SLBMs) and the first intercontinental ballistic missiles (ICBMs)
A9/A10 Amerika-Raketen to bomb New York and other American cities. The tests of
SLBM-variants of the A4 rocket was achieved with U-boat submarines towing launch
platforms. The second stage of the A9/A10 rocket was tested a few times in January,
February and March 1945.
In parallel with the guided missile programme in Nazi Germany, rockets were also being
used for aircraft, either for rapid horizontal take-off (JATO) or for powering the aircraft
(Me 163,etc) and for vertical take-off (Bachem Ba 349 "Natter").
Post World War II
At the end of World War II, competing Russian, British, and U.S. military and scientific
crews raced to capture technology and trained personnel from the German rocket
program at Peenemünde. Russia and Britain had some success, but the United States
benefited the most. The US captured a large number of German rocket scientists (many
of whom were members of the Nazi Party, including von Braun) and brought them to
the United States as part of Operation Paperclip. In America, the same rockets that were
designed to rain down on Britain were used instead by scientists as research vehicles for
developing the new technology further. The V-2 evolved into the American Redstone
rocket, used in the early space program.
After the war, rockets were used to study high-altitude conditions, by radio telemetry of
temperature and pressure of the atmosphere, detection of cosmic rays, and further
research; notably for the Bell X-1 to break the sound barrier. This continued in the U.S.
under von Braun and the others, who were destined to become part of the U.S.
scientific complex.
Independently, research continued in the Soviet Union under the leadership of the chief
designer Sergei Korolev. With the help of German technicians, the V-2 was duplicated
and improved as the R-1, R-2 and R-5 missiles. German designs were abandoned in the
late 1940s, and the foreign workers were sent home. A new series of engines built by
Glushko and based on inventions of Aleksei Mihailovich Isaev formed the basis of the
first ICBM, the R-7. The R-7 launched the first satellite, and Yuri Gagarin, the first man
into space and the first lunar and planetary probes, and is still in use today. These
events attracted the attention of top politicians, along with more money for further
research.
Rockets became extremely important militarily in the form of modern intercontinental
ballistic missiles (ICBMs) when it was realised that nuclear weapons carried on a rocket
vehicle were essentially not defensible against once launched, and ICBM/Launch
vehicles such as the R-7, Atlas and Titan became the delivery platform of choice for
these weapons.
Fueled partly by the Cold War, the 1960s became the decade of rapid development of
rocket technology particularly in the Soviet Union (Vostok, Soyuz, Proton) and in the
United States (e.g. the X-15 and X-20 Dyna-Soar aircraft). There was also significant
research in other countries, such as Britain, Japan, Australia, etc. and their growing use
for Space exploration, with pictures returned from the far side of the Moon and
unmanned flights for Mars exploration.
In America the manned programmes, Project Mercury, Project Gemini and later the
Apollo programme culminated in 1969 with the first manned landing on the moon via
the Saturn V, causing the New York Times to retract their earlier editorial implying that
spaceflight couldn't work:
"Further investigation and experimentation have confirmed the findings of Isaac Newton
in the 17th century and it is now definitely established that a rocket can function in a
vacuum as well as in an atmosphere. The Times regrets the error."
In the 1970s America made further lunar landings, before abandoning the Apollo launch
vehicle. The replacement vehicle, the partially reusable 'Space Shuttle' was intended to
be cheaper, but this large reduction in costs was largely not achieved. Meanwhile in
1973, the expendable Ariane programme was begun, a launcher that by the year 2000
would capture much of the geosat market.
Current day
Rockets remain a popular military weapon. The use of large battlefield rockets of the V-2
type has given way to guided missiles. However rockets are often used by helicopters
and light aircraft for ground attack, being more powerful than machine guns, but
without the recoil of a heavy cannon. In the 1950s there was a brief vogue for air-to-air
rockets, ending with the AIR-2 'Genie' nuclear rocket, but by the early 1960s these had
largely been abandoned in favor of air-to-air missiles.
Economically, rocketry is the enabler of all space technologies particularly satellites,
many of which impact people's everyday lives in almost countless ways, satellite
navigation, communications satellites and even things as simple as weather satellites.
Scientifically, rocketry has opened a window on our universe, allowing the launch of
space probes to explore our solar system, satellites to view the Earth itself, and spacebased telescopes to obtain a clearer view of the rest of the universe.
However, in the minds of much of the public, the most important use of rockets is
perhaps manned spaceflight. Vehicles such as the Space Shuttle for scientific research,
the Soyuz for orbital tourism and SpaceShipOne for suborbital tourism may show a
trend towards greater commercialisation of manned rocketry, away from government
funding, and towards more widespread access to space.
Types
There are many different types of rockets, and a comprehensive list of the basic engine
types can be found in rocket engine — the vehicles themselves range in size from tiny
models such as water rockets or small solid rockets that can be purchased at a hobby
store, to the enormous Saturn V used for the Apollo program, and in many different
vehicle types such as rocket cars and rocket planes.
Most current rockets are chemically powered rockets (usually internal combustion
engines, but some employ a decomposing monopropellant) that emit a hot exhaust gas.
A chemical rocket engine can use gas propellant, solid propellant, liquid propellant, or a
hybrid mixture of both solid and liquid. With combustive propellants a chemical reaction
is initiated between the fuel and the oxidizer in the combustion chamber, and the
resultant hot gases accelerate out of a nozzle (or nozzles) at the rearward-facing end of
the rocket. The acceleration of these gases through the engine exerts force ("thrust") on
the combustion chamber and nozzle, propelling the vehicle (in accordance with
Newton's Third Law). See rocket engine for details.
Rockets in which the heat is supplied from a source other than a propellant, such as
solar thermal rockets, can be classed as external combustion engines. Other examples of
external combustion rocket engines include most designs for nuclear powered rocket
engines. Use of hydrogen as the propellant for such engines gives very high exhaust
velocities (around 6-10 km/s).
Steam rockets, are another example of non chemical rockets. These rockets release very
hot water through a nozzle where, due to the lower pressure there, it instantly flashes
to high velocity steam, propelling the rocket. The efficiency of steam as a rocket
propellant is relatively low, but it is simple and reasonably safe, and the propellant is
cheap and widely available. Most steam rockets have been used for propelling landbased vehicles but a small steam rocket was tested in 2004 on board the UK-DMC
satellite, as an alternative, with higher performance, to cold gas thrusters for attitude
jets. There are even proposals to use steam rockets for interplanetary transport using
either nuclear or solar heating as the power source to vaporize water collected from
around the solar system, at system costs that are claimed to be greatly lower than
hydrogen-based systems.
Uses
Rockets or other similar reaction devices carrying their own propellant must be used
when there is no other substance (land, water, or air) or force (gravity, magnetism, light)
that a vehicle may usefully employ for propulsion, such as in space. In these
circumstances, it is necessary to carry all the propellant to be used.
However, they are also useful in other situations:
Weaponry
In many military weapons, rockets are used to propel payloads to their targets. A rocket
and its payload together are generally referred to as a missile, especially when the
weapon has a guidance system.
Science
Sounding rockets are commonly used to carry instruments that take readings from 50
kilometers (30 mi) to 1,500 kilometers (930 mi) above the surface of the Earth, the
altitudes between those reachable by weather balloons and satellites.
Launch
Due to their high exhaust velocity (Mach ~10+), rockets are particularly useful when
very high speeds are required, such as orbital speed (Mach 25+). Indeed, rockets remain
the only way to launch spacecraft into orbit. They are also used to rapidly accelerate
spacecraft when they change orbits or de-orbit for landing. Also, a rocket may be used
to soften a hard parachute landing immediately before touchdown (see Soyuz
spacecraft). Spacecraft delivered into orbital trajectories become artificial satellites.
Hobby, sport and entertainment
Hobbyists build and fly Model rockets of various types and rockets are used to launch
both commercially available fireworks and professional fireworks displays.
Hydrogen peroxide rockets are used to power jet packs, and have been used to power
cars and a rocket car holds the all time drag racing record.
Components of a rocket
Rockets at minimum have a place to put propellant (such as a propellant tank), one or
more rocket engines and nozzle, directional stabilization device(s) (such as fins, attitude
jets or engine gimbals) and a structure (typically monocoque) to hold these components
together. Rockets intended for high speed atmospheric use also have an aerodynamic
fairing such as a nose cone.
As well as these components, rockets can have any number of other components, such
as wings (rocketplanes), wheels (rocket cars), even, in a sense, a person (rocket belt).
Noise
For all but the very smallest sizes, rocket exhaust compared to other engines is generally
very noisy. As the hypersonic exhaust mixes with the ambient air, shock waves are
formed. The sound intensity from these shock waves depends on the size of the rocket.
The sound intensity of large rockets could potentially kill at close range.
The Space Shuttle generates over 200 dB(A) of noise around its base. A Saturn V launch
was detectable on seismometers a considerable distance from the launch site.
Generally speaking, noise is most intense when a rocket is close to the ground, since the
noise from the engines radiates up away from the plume, as well as reflecting off the
ground. This noise can be reduced somewhat by flame trenches with roofs, by water
injection around the plume and by deflecting the plume at an angle.
For manned rockets various methods are used to reduce the sound intensity for the
passengers as much as possible, and typically the placement of the astronauts far away
from the rocket engines helps significantly. For the passengers and crew, when a vehicle
goes supersonic the sound cuts off as the sound waves are no longer able to keep up
with the vehicle.
Physics
Operation
In all rockets, the exhaust is formed from propellants carried within the rocket prior to
use. Rocket thrust is due to the rocket engine, which propels the rocket forwards by
exhausting the propellant rearwards at extreme high speed.
In a closed chamber, the pressures are equal in each direction and no acceleration
occurs. If an opening is provided at the bottom of the chamber then the pressure is no
longer acting on that side. The remaining pressures give a resultant thrust on the side
opposite the opening; as well as permitting exhaust to escape. Using a nozzle increases
the forces further, in fact multiplies the thrust as a function of the area ratio of the
nozzle, since the pressures also act on the nozzle. As a side effect the pressures act on
the exhaust in the opposite direction and accelerate this to very high speeds (in
accordance with Newton's Third Law).
If propellant gas is continuously added to the chamber then this disequilibrium of
pressures can be maintained for as long as propellant remains.
It turns out (from conservation of momentum) that the speed of the exhaust of a rocket
determines how much momentum increase is created for a given amount of propellant,
and this is termed a rocket's specific impulse.
As the remaining propellant decreases, the vehicle's becomes lighter and acceleration
tends to increase until eventually it runs out of propellant, and this means that much of
the speed change occurs towards the end of the burn when the vehicle is much lighter.
Forces on a rocket in flight
The general study of the forces on a rocket or other spacecraft is called astrodynamics.
Flying rockets are primarily affected by the following:

Thrust from the engine(s)



Gravity from celestial bodies
Drag if moving in the atmosphere
Lift; usually relatively small effect except for rocket-powered aircraft
In addition, the inertia/centrifugal pseudo-force can be significant due to the path of the
rocket around the center of a celestial body; when high enough speeds in the right
direction and altitude are achieved a stable orbit or escape velocity is obtained.
During a rocket launch, there is a point of maximum aerodynamic drag called Max Q.
This determines the minimum aerodynamic strength of the vehicle.
These forces, with a stabilizing tail present will, unless deliberate control efforts are
made, to naturally cause the vehicle to follow a trajectory termed a gravity turn, and
this trajectory is often used at least during the initial part of a launch. This means that
the vehicle can maintain low or even zero angle of attack. This minimizes transverse
stress on the launch vehicle; allowing for a weaker, and thus lighter, launch vehicle.
Net thrust
The thrust of a rocket is often deliberately varied over a flight, to provide a way to
control the airspeed of the vehicle so as to minimize aerodynamic losses but also so as
to limit g-forces that would otherwise occur during the flight as the propellant mass
decreases, which could damage the vehicle, crew or payload.
Below is an approximate equation for calculating the gross thrust of a rocket:
where:
propellant flow (kg/s or lb/s)
jet velocity at nozzle exit plane (m/s or s)
flow area at nozzle exit plane (m2 or ft2)
static pressure at nozzle exit plane (Pa or lb/ft2)
ambient (or atmospheric) pressure (Pa or lb/ft2)
Since, unlike a jet engine, a conventional rocket motor lacks an air intake, there is no
'ram drag' to deduct from the gross thrust. Consequently the net thrust of a rocket
motor is equal to the gross thrust.
The
term represents the momentum thrust, which remains constant at a given
throttle setting, whereas the
term represents the pressure thrust
term. At full throttle, the net thrust of a rocket motor improves slightly with increasing
altitude, because the reducing atmospheric pressure increases the pressure thrust term.
Specific impulse
As can be seen from the thrust equation the effective speed of the exhaust, Ve, has a
large impact on the amount of thrust produced from a particular quantity of fuel burnt
per second. The thrust-seconds (impulse) per unit of propellant is called Specific Impulse
(Isp) or effective exhaust velocity and this is one of the most important figures that
describes a rocket's performance.
Vacuum Isp
Due to the specific impulse varying with pressure, a quantity that is easy to compare
and calculate with is useful. Because rockets choke at the throat, and because the
supersonic exhaust prevents external pressure influences travelling upstream, it turns
out that the pressure at the exit is ideally exactly proportional to the propellant flow ,
provided the mixture ratios and combustion efficiencies are maintained. It is thus quite
usual to rearrange the above equation slightly:
and so define the vacuum Isp to be:
Vevac = Cfc *
Where:
the speed of sound constant at the throat
the thrust coefficient constant of the nozzle (typically between 0.8 and
1.9)
And hence:
Delta-v (rocket equation)
The delta-v capacity of a rocket is the theoretical total change in velocity that a rocket
can achieve without any external interference (without air drag or gravity or other
forces).
The delta-v that a rocket vehicle can provide can be calculated from the Tsiolkovsky
rocket equation:
where:
m0 is the initial total mass, including propellant, in kg (or lb)
m1 is the final total mass in kg (or lb)
ve is the effective exhaust velocity in m/s or (ft/s) or
is the delta-v in m/s (or ft/s)
Delta-v can also be calculated for a particular manoeuvre; for example the delta-v to
launch from the surface of the Earth to Low earth orbit is about 9.7 km/s, which leaves
the vehicle with a sideways speed of about 7.8 km/s at an altitude of around 200 km. In
this manoeuvre about 1.9 km/s is lost in air drag, gravity drag and gaining altitude.
Mass ratios
Mass ratio is the ratio between the initial fuelled mass and the mass after the 'burn'.
Everything else being equal, a high mass ratio is desirable for good performance, since it
indicates that the rocket is lightweight and hence performs better, for essentially the
same reasons that low weight is desirable in sports cars.
Rockets as a group have the highest thrust-to-weight ratio of any type of engine; and
this helps vehicles achieve high mass ratios, which improves the performance of flights.
The higher this ratio, the less engine mass is needed to be carried and permits the
carrying of even more propellant, this enormously improves performance.
Achievable mass ratios are highly dependent on many factors such as propellant type,
the design of engine the vehicle uses, structural safety margins and construction
techniques.
Vehicle
Takeoff Mass
Ariane 5 (vehicle +
payload)
746,000 kg
Titan 23G first stage
258,000 lb
(117,020 kg)
Saturn V
3,038,500 kg
Space Shuttle (vehicle +
payload)
Saturn 1B (stage only)
2,040,000 kg
448,648 kg
Final Mass
2,700 kg + 16,000
kg
10,500 lb (4,760
kg)
13,300 kg +
118,000 kg
104,000 kg +
28,800 kg
41,594 kg
Mass
ratio
Mass
fraction
39.9
0.975
24.6
0.959
23.1
0.957
15.4
0.935
10.7
0.907
V2
X-15
Concorde
747
12.8 ton (13000 kg)
34,000 lb (15,420 14,600 lb (6,620
kg)
kg)
400,000 lb
800,000 lb
3.85
0.74
2.3
0.57
2
2
0.5
0.5
Staging
Often, the required velocity (delta-v) for a mission is unattainable by any single rocket
because the propellant, tankage, structure, guidance, valves and engines and so on, take
a particular minimum percentage of take-off mass.
The mass ratios that can be achieved with a single set of fixed rocket engines and
tankage varies depends on acceleration required, construction materials, tank layout,
engine type and propellants used, but for example the first stage of the Saturn V,
carrying the weight of the upper stages, was able to achieve a mass ratio of about 10.
This problem is frequently solved by staging — the rocket sheds excess weight (usually
empty tankage and associated engines) during launch to reduce its weight and
effectively increase its mass ratio. Staging is either serial where the rockets light after
the previous stage has fallen away, or parallel, where rockets are burning together and
then detach when they burn out.
Typically, the acceleration of a rocket increases with time (if the thrust stays the same)
as the weight of the rocket decreases as propellant is burned. Discontinuities in
acceleration will occur when stages burn out, often starting at a lower acceleration with
each new stage firing.
Energy efficiency
Rocket launch vehicles take-off with a great deal of flames, noise and drama, and it
might seem obvious that they are grievously inefficient. However while they are far
from perfect, their energy efficiency is not as bad as might be supposed.
The energy density of rocket propellant is around 1/3 that of conventional hydrocarbon
fuels; the bulk of the mass is in the form of (often relatively inexpensive) oxidiser.
Nevertheless, at take-off the rocket has a great deal of energy in the form of fuel and
oxidiser stored within the vehicle, and it is of course desirable that as much of the
energy stored in the propellant ends up as kinetic or potential energy of the body of the
rocket as possible.
Energy from the fuel is lost in air drag and gravity drag and is used to gain altitude.
However, much of the lost energy ends up in the exhaust.
100% efficiency within the engine (ηc) would mean that all of the heat energy of the
combustion products is converted into kinetic energy of the jet. This is not possible, but
the high expansion ratio nozzles that can be used with rockets come surprisingly close:
when the nozzle expands the gas, the gas is cooled and accelerated, and an energy
efficiency of up to 70% can be achieved. Most of the rest is heat energy in the exhaust
that is not recovered. This compares very well with other engine designs. The high
efficiency is a consequence of the fact that rocket combustion can be performed at very
high temperatures and the gas is finally released at much lower temperatures, and so
giving good Carnot efficiency.
However, engine efficiency is not the whole story. In common with many jet-based
engines, but particularly in rockets due to their high and typically fixed exhaust speeds,
rocket vehicles are extremely inefficient at low speeds irrespective of the engine
efficiency. The problem is that at low speeds, the exhaust carries away a huge amount
of kinetic energy rearward. This phenomenon is termed propulsive efficiency (ηp).
However, as speeds rise, the resultant exhaust speed goes down, and the overall vehicle
energetic efficiency rises, reaching a peak of around 100% of the engine efficiency when
the vehicle is travelling exactly at the same speed that the exhaust is emitted. In this
case the exhaust would ideally stop dead in space behind the moving vehicle, taking
away zero energy, and from conservation of energy, all the energy would end up in the
vehicle. The efficiency then drops off again at even higher speeds as the exhaust ends
up travelling forwards behind the vehicle.
From these principles it can be shown that the propulsive efficiency ηp for a rocket
moving at speed u with an exhaust velocity c is:
And the overall energy efficiency η is:
η = ηpηc
Since the energy ultimately comes from fuel, these joint considerations mean that
rockets are mainly useful when a very high speed is required, such as ICBMs or orbital
launch, and they are rarely if ever used for general aviation. For example, from the
equation, with an ηc of 0.7, a rocket flying at Mach 0.85 (which most aircraft cruise at)
with an exhaust velocity of Mach 10, would have a predicted overall energy efficiency of
5.9%, whereas a conventional, modern, air breathing jet engine achieves closer to 30%
or more efficiency. Thus a rocket would need about 5x more energy; and allowing for
the ~3x lower specific energy of rocket propellant than conventional air fuel, roughly
15x more mass of propellant would need to be carried for the same journey.
Thus jet engines which have a better match between speed and jet exhaust speed such
as turbofans (in spite of their worse ηc) dominate for subsonic and supersonic
atmospheric use while rockets work best at hypersonic speeds. On the other hand
rockets do also see many short-range relatively low speed military applications where
their low-speed inefficiency is outweighed by their extremely high thrust and hence high
accelerations.
Safety, reliability and accidents
Rockets are not inherently highly dangerous. In military usage quite adequate reliability
is obtained.
Because of the enormous chemical energy in all useful rocket propellants (greater
energy per weight than explosives, but lower than gasoline), accidents can and have
happened. The number of people injured or killed is usually small because of the great
care typically taken, but this record is not perfect.
Spacecraft propulsion
Spacecraft propulsion is any method used to change the velocity of spacecraft and
artificial satellites. There are many different methods. Each method has drawbacks and
advantages, and spacecraft propulsion is an active area of research. However, most
spacecraft today are propelled by exhausting a gas from the back/rear of the vehicle at
very high speed through a supersonic de Laval nozzle. This sort of engine is called a
rocket engine.
All current spacecraft use chemical rockets (bipropellant or solid-fuel) for launch,
though some (such as the Pegasus rocket and SpaceShipOne) have used air-breathing
engines on their first stage. Most satellites have simple reliable chemical thrusters
(often monopropellant rockets) or resistojet rockets for orbital station-keeping and
some use momentum wheels for attitude control. Soviet bloc satellites have used
electric propulsion for decades, and newer Western geo-orbiting spacecraft are starting
to use them for north-south stationkeeping. Interplanetary vehicles mostly use chemical
rockets as well, although a few have experimentally used ion thrusters (a form of
electric propulsion) to great success.
The necessity for propulsion system
Artificial satellites must be launched into orbit, and once there they must be placed in
their nominal orbit. Once in the desired orbit, they often need some form of attitude
control so that they are correctly pointed with respect to the Earth, the Sun, and
possibly some astronomical object of interest. They are also subject to drag from the
thin atmosphere, so that to stay in orbit for a long period of time some form of
propulsion is occasionally necessary to make small corrections (orbital stationkeeping).
Many satellites need to be moved from one orbit to another from time to time, and this
also requires propulsion. When a satellite has exhausted its ability to adjust its orbit, its
useful life is over.
Spacecraft designed to travel further also need propulsion methods. They need to be
launched out of the Earth's atmosphere just as satellites do. Once there, they need to
leave orbit and move around.
For interplanetary travel, a spacecraft must use its engines to leave Earth orbit. Once it
has done so, it must somehow make its way to its destination. Current interplanetary
spacecraft do this with a series of short-term trajectory adjustments. In between these
adjustments, the spacecraft simply falls freely along its orbit. The simplest fuel-efficient
means to move from one circular orbit to another is with a Hohmann transfer orbit: the
spacecraft begins in a roughly circular orbit around the Sun. A short period of thrust in
the direction of motion accelerates or decelerates the spacecraft into an elliptical orbit
around the Sun which is tangential to its previous orbit and also to the orbit of its
destination. The spacecraft falls freely along this elliptical orbit until it reaches its
destination, where another short period of thrust accelerates or decelerates it to match
the orbit of its destination. Special methods such as aerobraking are sometimes used for
this final orbital adjustment.
Some spacecraft propulsion methods such as solar sails provide very low but
inexhaustible thrust; an interplanetary vehicle using one of these methods would follow
a rather different trajectory, either constantly thrusting against its direction of motion in
order to decrease its distance from the Sun or constantly thrusting along its direction of
motion to increase its distance from the Sun.
Spacecraft for interstellar travel also need propulsion methods. No such spacecraft has
yet been built, but many designs have been discussed. Since interstellar distances are
very great, a tremendous velocity is needed to get a spacecraft to its destination in a
reasonable amount of time. Acquiring such a velocity on launch and getting rid of it on
arrival will be a formidable challenge for spacecraft designers.
Effectiveness of propulsion systems
When in space, the purpose of a propulsion system is to change the velocity, or v, of a
spacecraft. Since this is more difficult for more massive spacecraft, designers generally
discuss momentum, mv. The amount of change in momentum is called impulse. So the
goal of a propulsion method in space is to create an impulse.
When launching a spacecraft from the Earth, a propulsion method must overcome a
higher gravitational pull to provide a net positive acceleration. In orbit, any additional
impulse, even very tiny, will result in a change in the orbit path.
The rate of change of velocity is called acceleration, and the rate of change of
momentum is called force. To reach a given velocity, one can apply a small acceleration
over a long period of time, or one can apply a large acceleration over a short time.
Similarly, one can achieve a given impulse with a large force over a short time or a small
force over a long time. This means that for maneuvering in space, a propulsion method
that produces tiny accelerations but runs for a long time can produce the same impulse
as a propulsion method that produces large accelerations for a short time. When
launching from a planet, tiny accelerations cannot overcome the planet's gravitational
pull and so cannot be used.
The Earth's surface is situated fairly deep in a gravity well and it takes a velocity of 11.2
kilometers/second (escape velocity) or more to escape from it. As human beings
evolved in a gravitational field of 1g (9.8 m/s²), an ideal propulsion system would be one
that provides a continuous acceleration of 1g (though human bodies can tolerate much
larger accelerations over short periods). The occupants of a rocket or spaceship having
such a propulsion system would be free from all the ill effects of free fall, such as
nausea, muscular weakness, reduced sense of taste, or leaching of calcium from their
bones.
The law of conservation of momentum means that in order for a propulsion method to
change the momentum of a space craft it must change the momentum of something
else as well. A few designs take advantage of things like magnetic fields or light pressure
in order to change the spacecraft's momentum, but in free space the rocket must bring
along some mass to accelerate away in order to push itself forward. Such mass is called
reaction mass.
In order for a rocket to work, it needs two things: reaction mass and energy. The
impulse provided by launching a particle of reaction mass having mass m at velocity v is
mv. But this particle has kinetic energy mv²/2, which must come from somewhere. In a
conventional solid, liquid, or hybrid rocket, the fuel is burned, providing the energy, and
the reaction products are allowed to flow out the back, providing the reaction mass. In
an ion thruster, electricity is used to accelerate ions out the back. Here some other
source must provide the electrical energy (perhaps a solar panel or a nuclear reactor),
while the ions provide the reaction mass.
When discussing the efficiency of a propulsion system, designers often focus on
effectively using the reaction mass. Reaction mass must be carried along with the rocket
and is irretrievably consumed when used. One way of measuring the amount of impulse
that can be obtained from a fixed amount of reaction mass is the specific impulse, the
impulse per unit weight-on-Earth (typically designated by Isp). The unit for this value is
seconds. Since the weight on Earth of the reaction mass is often unimportant when
discussing vehicles in space, specific impulse can also be discussed in terms of impulse
per unit mass. This alternate form of specific impulse uses the same units as velocity
(e.g. m/s), and in fact it is equal to the effective exhaust velocity of the engine (typically
designated ve). Confusingly, both values are sometimes called specific impulse. The two
values differ by a factor of gn, the standard acceleration due to gravity 9.80665 m/s²
(Ispgn = ve).
A rocket with a high exhaust velocity can achieve the same impulse with less reaction
mass. However, the energy required for that impulse is proportional to the square of
the exhaust velocity, so that more mass-efficient engines require much more energy,
and are typically less energy efficient. This is a problem if the engine is to provide a large
amount of thrust. To generate a large amount of impulse per second, it must use a large
amount of energy per second. So highly (mass) efficient engines require enormous
amounts of energy per second to produce high thrusts. As a result, most high-efficiency
engine designs also provide very low thrust.
Delta-v and propellant use
Burning the entire usable propellant of a spacecraft through the engines in a straight
line in free space would produce a net velocity change to the vehicle; this number is
termed 'delta-v' (Δv).
If the exhaust velocity is constant then the total Δv of a vehicle can be calculated using
the rocket equation, where M is the mass of propellant, P is the mass of the payload
(including the rocket structure), and ve is the velocity of the rocket exhaust. This is
known as the Tsiolkovsky rocket equation:
For historical reasons, as discussed above, ve is sometimes written as
ve = Ispgo
where Isp is the specific impulse of the rocket, measured in seconds, and go is the
gravitational acceleration at sea level.
For a high delta-v mission, the majority of the spacecraft's mass needs to be reaction
mass. Since a rocket must carry all of its reaction mass, most of the initially-expended
reaction mass goes towards accelerating reaction mass rather than payload. If the
rocket has a payload of mass P, the spacecraft needs to change its velocity by Δv, and
the rocket engine has exhaust velocity ve, then the mass M of reaction mass which is
needed can be calculated using the rocket equation and the formula for Isp:
For Δv much smaller than ve, this equation is roughly linear, and little reaction mass is
needed. If Δv is comparable to ve, then there needs to be about twice as much fuel as
combined payload and structure (which includes engines, fuel tanks, and so on). Beyond
this, the growth is exponential; speeds much higher than the exhaust velocity require
very high ratios of fuel mass to payload and structural mass.
For a mission, for example, when launching from or landing on a planet, the effects of
gravitational attraction and any atmospheric drag must be overcome by using fuel. It is
typical to combine the effects of these and other effects into an effective mission deltav. For example a launch mission to low Earth orbit requires about 9.3-10 km/s delta-v.
These mission delta-vs are typically numerically integrated on a computer.
Power use and propulsive efficiency
Although solar power and nuclear power are virtually unlimited sources of energy, the
maximum power they can supply is substantially proportional to the mass of the
powerplant. For fixed power, with a large ve which is desirable to save propellant mass,
it turns out that the maximum acceleration is inversely proportional to ve. Hence the
time to reach a required delta-v is proportional to ve. Thus the latter should not be too
large. It might be thought that adding power generation is helpful, however this takes
mass away from payload, and ultimately reaches a limit as the payload fraction tends to
zero.
For all reaction engines (such as rockets and ion drives) some energy must go into
accelerating the reaction mass. Every engine will waste some energy, but even assuming
100% efficiency, to accelerate a particular mass of exhaust the engine will need energy
amounting to
which is simply the energy needed to accelerate the exhaust. This energy is not
necessarily lost- some of it usually ends up as kinetic energy of the vehicle, and the rest
is wasted in residual motion of the exhaust.
Comparing the rocket equation (which shows how much energy ends up in the final
vehicle) and the above equation (which shows the total energy required) shows that
even with 100% engine efficiency, certainly not all energy supplied ends up in the
vehicle - some of it, indeed usually most of it, ends up as kinetic energy of the exhaust.
The exact amount depends on the design of the vehicle, and the mission. However there
are some useful fixed points:

if the Isp is fixed, for a mission delta-v, there is a particular Isp that minimises the
overall energy used by the rocket. This comes to an exhaust velocity of about ⅔
of the mission delta-v (see the energy computed from the rocket equation).
Drives with a specific impulse that is both high and fixed such as Ion thrusters
have exhaust velocities that can be enormously higher than this ideal for many
missions.

if the exhaust velocity can be made to vary so that at each instant it is equal and
opposite to the vehicle velocity then the absolute minimum energy usage is
achieved. When this is achieved, the exhaust stops in space ^ and has no kinetic
energy; and the propulsive efficiency is 100%- all the energy ends up in the
vehicle (in principle such a drive would be 100% efficient, in practice there would
be thermal losses from within the drive system and residual heat in the exhaust).
However in most cases this uses an impractical quantity of propellant, but is a
useful theoretical consideration. Another complication is that unless the vehicle
is moving initially, it cannot accelerate, as the exhaust velocity is zero at zero
speed.
Some drives (such as VASIMR or Electrodeless plasma thruster ) actually can significantly
vary their exhaust velocity. This can help reduce propellant usage or improve
acceleration at different stages of the flight. However the best energetic performance
and acceleration is still obtained when the exhaust velocity is close to the vehicle speed.
Proposed ion and plasma drives usually have exhaust velocities enormously higher than
that ideal (in the case of VASIMR the lowest quoted speed is around 15000 m/s
compared to a mission delta-v from high Earth orbit to Mars of about 4000m/s).
Example
Suppose we want to send a 10,000 kg space probe to Mars. The required Δv from LEO is
approximately 3000 m/s, using a Hohmann transfer orbit. (A manned craft would need
to take a faster route and use more fuel). For the sake of argument, let us say that the
following thrusters may be used:
Engine
Effectiv
e
Specific Fuel
Exhaust impulse mass
Velocity
(s)
(kg)
(km/s)
Energy
Energy
Power
per kg
minimum
require
generator
of
power/thrus
d
mass/thrust
propellan
t
(GJ)
*
t
Solid rocket 1
Bipropellan
5
t rocket
Ion thruster 50
Advance
electrically
1,000
powered
drive

100
190,00
95
0
500 kJ
0.5 kW/N
N/A
500
8,200
103
12.6 MJ
2.5 kW/N
N/A
5,000
620
775
1.25 GJ
25 kW/N
25 kg/N
15,000
500 GJ
500 kW/N
500 kg/N
100,00
30
0
- assumes a specific power of 1kW
Observe that the more fuel-efficient engines can use far less fuel; its mass is almost
negligible (relative to the mass of the payload and the engine itself) for some of the
engines. However, note also that these require a large total amount of energy. For Earth
launch, engines require a thrust to weight ratio of more than unity. To do this they
would have to be supplied with Gigawatts of power — equivalent to a major
metropolitan generating station. From the table it can be seen that this is clearly
impractical with current power sources.
Instead, a much smaller, less powerful generator may be included which will take much
longer to generate the total energy needed. This lower power is only sufficient to
accelerate a tiny amount of fuel per second, and would be insufficient for launching
from the Earth but in orbit, where there is no friction, over long periods the velocity will
be finally achieved. For example. it took the Smart 1 more than a year to reach the
Moon, while with a chemical rocket it takes a few days. Because the ion drive needs
much less fuel, the total launched mass is usually lower, which typically results in a
lower overall cost.
Mission planning frequently involves adjusting and choosing the propulsion system
according to the mission delta-v needs, so as to minimise the total cost of the project,
including trading off greater or lesser use of fuel and launch costs of the complete
vehicle.
Space propulsion methods
Propulsion methods can be classified based on their means of accelerating the reaction
mass. There are also some special methods for launches, planetary arrivals, and
landings.
Reaction engines
Rocket engines
Most rocket engines are internal combustion heat engines (although non combusting
forms exist). Rocket engines generally produce a high temperature reaction mass, as a
hot gas. This is achieved by combusting a solid, liquid or gaseous fuel with an oxidiser
within a combustion chamber. The extremely hot gas is then allowed to escape through
a high-expansion ratio nozzle. This bell-shaped nozzle is what gives a rocket engine its
characteristic shape. The effect of the nozzle is to dramatically accelerate the mass,
converting most of the thermal energy into kinetic energy. Exhaust speeds as high as 10
times the speed of sound at sea level are common.
Ion propulsion rockets can heat a plasma or charged gas inside a magnetic bottle and
release it via a magnetic nozzle, so that no solid matter need come in contact with the
plasma. Of course, the machinery to do this is complex, but research into nuclear fusion
has developed methods, some of which have been proposed to be used in propulsion
systems, and some have been tested in a lab.
Electromagnetic acceleration of reaction mass
Rather than relying on high temperature and fluid dynamics to accelerate the reaction
mass to high speeds, there are a variety of methods that use electrostatic or
electromagnetic forces to accelerate the reaction mass directly. Usually the reaction
mass is a stream of ions. Such an engine very typically uses electric power, first to ionise
atoms, and then uses a voltage gradient to accelerate the ions to high exhaust
velocities.
For these drives, at the highest exhaust speeds, energetic efficiency and thrust are all
inversely proportional to exhaust velocity. Their very high exhaust velocity means they
require huge amounts of energy and thus with practical power sources provide low
thrust, but use hardly any fuel.
For some missions, particularly reasonably close to the Sun, solar energy may be
sufficient, and has very often been used, but for others further out or at higher power,
nuclear energy is necessary; engines drawing their power from a nuclear source are
called nuclear electric rockets.
With any current source of electrical power, chemical, nuclear or solar, the maximum
amount of power that can be generated limits the amount of thrust that can be
produced to a small value. Power generation adds significant mass to the spacecraft,
and ultimately the weight of the power source limits the performance of the vehicle.
Current nuclear power generators are approximately half the weight of solar panels per
watt of energy supplied, at terrestrial distances from the Sun. Chemical power
generators are not used due to the far lower total available energy. Beamed power to
the spacecraft shows some potential. However, the dissipation of waste heat from any
power plant may make any propulsion system requiring a separate power source
infeasible for interstellar travel.
Some electromagnetic methods:



Ion thrusters (accelerate ions first and later neutralize the ion beam with an
electron stream emitted from a cathode called a neutralizer)
o Electrostatic ion thruster
o Field Emission Electric Propulsion
o Hall effect thruster
o Colloid thruster
Plasma thrusters (where both ions and electrons are accelerated simultaneously,
no neutralizer is required)
o Magnetoplasmadynamic thruster
o Helicon Double Layer Thruster
o Electrodeless plasma thruster
o Pulsed plasma thruster
o Pulsed inductive thruster
o Variable specific impulse magnetoplasma rocket (VASIMR)
Mass drivers (for propulsion)
Systems without reaction mass carried within the spacecraft
The law of conservation of momentum states that any engine which uses no reaction
mass cannot move the center of mass of a spaceship (changing orientation, on the other
hand, is possible). But space is not empty, especially space inside the Solar System;
there are gravitation fields, magnetic fields, solar wind and solar radiation. Various
propulsion methods try to take advantage of these. However, since these phenomena
are diffuse in nature, corresponding propulsion structures need to be proportionately
large.
There are several different space drives that need little or no reaction mass to function.
A tether propulsion system employs a long cable with a high tensile strength to change a
spacecraft's orbit, such as by interaction with a planet's magnetic field or through
momentum exchange with another object. Solar sails rely on radiation pressure from
electromagnetic energy, but they require a large collection surface to function
effectively. The magnetic sail deflects charged particles from the solar wind with a
magnetic field, thereby imparting momentum to the spacecraft. A variant is the minimagnetospheric plasma propulsion system, which uses a small cloud of plasma held in a
magnetic field to deflect the Sun's charged particles.
For changing the orientation of a satellite or other space vehicle, conservation of
angular momentum does not pose a similar constraint. Thus many satellites use
momentum wheels to control their orientations. These cannot be the only system for
controlling satellite orientation, as the angular momentum built up due to torques from
external forces such as solar, magnetic, or tidal forces eventually needs to be "bled off"
using a secondary system.
Gravitational slingshots can also be used to carry a probe onward to other destinations.
Planetary and atmospheric spacecraft propulsion
Launch mechanisms
High thrust is of vital importance for Earth launch, thrust has to be greater than weight
Many of the propulsion methods above give a thrust/weight ratio of much less than 1,
and so cannot be used for launch.
All current spacecraft use chemical rocket engines (bipropellant or solid-fuel) for launch.
Other power sources such as nuclear have been proposed, and tested, but safety,
environmental and political considerations have so far curtailed their use.
One advantage that spacecraft have in launch is the availability of infrastructure on the
ground to assist them. Proposed non-rocket spacelaunch ground-assisted launch
mechanisms include:








Space elevator (a geostationary tether to orbit)
Launch loop (a very fast rotating loop about 80km tall)
Space fountain (a very tall building held up by a stream of masses fired from
base)
Orbital ring (a ring around the Earth with spokes hanging down off bearings)
Hypersonic skyhook (a fast spinning orbital tether)
Electromagnetic catapult (railgun, coilgun) (an electric gun)
Space gun (Project HARP, ram accelerator) (a chemically powered gun)
Laser propulsion (Lightcraft) (rockets powered from ground-based lasers)
Airbreathing engines for orbital launch
Studies generally show that conventional air-breathing engines, such as ramjets or
turbojets are basically too heavy (have too low a thrust/weight ratio) to give any
significant performance improvement when installed on a launch vehicle itself.
However, launch vehicles can be air launched from separate lift vehicles (e.g. B-29,
Pegasus Rocket and White Knight) which do use such propulsion systems.
On the other hand, very lightweight or very high speed engines have been proposed
that take advantage of the air during ascent:
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SABRE - a lightweight hydrogen fuelled turbojet with precooler
ATREX - a lightweight hydrogen fuelled turbojet with precooler
Liquid air cycle engine - a hydrogen fuelled jet engine that liquifies the air before
burning it in a rocket engine
Scramjet - jet engines that use supersonic combustion
Normal rocket launch vehicles fly almost vertically before rolling over at an altitude of
some tens of kilometers before burning sideways for orbit; this initial vertical climb
wastes propellant but is optimal as it greatly reduces airdrag. Airbreathing engines burn
propellant much more efficiently and this would permit a far flatter launch trajectory,
the vehicles would typically fly approximately tangentially to the earth surface until
leaving the atmosphere then perform a rocket burn to bridge the final delta-v to orbital
velocity.
Planetary arrival and landing
When a vehicle is to enter orbit around its destination planet, or when it is to land, it
must adjust its velocity. This can be done using all the methods listed above (provided
they can generate a high enough thrust), but there are a few methods that can take
advantage of planetary atmospheres and/or surfaces.
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Aerobraking allows a spacecraft to reduce the high point of an elliptical orbit by
repeated brushes with the atmosphere at the low point of the orbit. This can
save a considerable amount of fuel since it takes much less delta-V to enter an
elliptical orbit compared to a low circular orbit. Since the braking is done over
the course of many orbits, heating is comparatively minor, and a heat shield is
not required. This has been done on several Mars missions such as Mars Global
Surveyor, Mars Odyssey and Mars Reconnaissance Orbiter, and at least one
Venus mission, Magellan.
Aerocapture is a much more aggressive manoeuver, converting an incoming
hyperbolic orbit to an elliptical orbit in one pass. This requires a heat shield and
much trickier navigation, since it must be completed in one pass through the
atmosphere, and unlike aerobraking no preview of the atmosphere is possible. If
the intent is to remain in orbit, then at least one more propulsive maneuver is
required after aerocapture—otherwise the low point of the resulting orbit will
remain in the atmosphere, resulting in eventual re-entry. Aerocapture has not
yet been tried on a planetary mission, but the re-entry skip by Zond 6 and Zond 7
upon lunar return were aerocapture maneuvers, since they turned a hyperbolic
orbit into an elliptical orbit. On these missions, since there was no attempt to
raise the perigee after the aerocapture, the resulting orbit still intersected the
atmosphere, and re-entry occurred at the next perigee.
Parachutes can land a probe on a planet with an atmosphere, usually after the
atmosphere has scrubbed off most of the velocity, using a heat shield.
Airbags can soften the final landing.
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Lithobraking, or stopping by simply smashing into the target, is usually done by
accident. However, it may be done deliberately with the probe expected to
survive (see, for example, Deep Space 2). Very sturdy probes and low approach
velocities are required.
Proposed spacecraft methods that may violate the laws of physics
In addition, a variety of hypothetical propulsion techniques have been considered that
would require entirely new principles of physics to realize and that may not actually be
possible. To date, such methods are highly speculative and include:
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Diametric drive
Pitch drive
Bias drive
Disjunction drive
Alcubierre drive (a form of Warp drive)
Differential sail
Wormholes - theoretically possible, but impossible in practice with current
technology
Antigravity - requires the concept of antigravity; theoretically impossible
Reactionless drives - breaks the law of conservation of momentum; theoretically
impossible
EmDrive - tries to circumvent the law of conservation of momentum; may be
theoretically impossible
A "hyperspace" drive based upon Heim theory
Table of methods and their specific impulse
Below is a summary of some of the more popular, proven technologies, followed by
increasingly speculative methods.
Four numbers are shown. The first is the effective exhaust velocity: the equivalent
speed that the propellant leaves the vehicle. This is not necessarily the most important
characteristic of the propulsion method, thrust and power consumption and other
factors can be, however:
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if the delta-v is much more than the exhaust velocity, then exorbitant amounts
of fuel are necessary (see the section on calculations, above)
if it is much more than the delta-v, then, proportionally more energy is needed;
if the power is limited, as with solar energy, this means that the journey takes a
proportionally longer time
The second and third are the typical amounts of thrust and the typical burn times of the
method. Outside a gravitational potential small amounts of thrust applied over a long
period will give the same effect as large amounts of thrust over a short period. (This
result does not apply when the object is significantly influenced by gravity.)
The fourth is the maximum delta-v this technique can give (without staging). For rocketlike propulsion systems this is a function of mass fraction and exhaust velocity. Mass
fraction for rocket-like systems is usually limited by propulsion system weight and
tankage weight. For a system to achieve this limit, typically the payload may need to be
a negligible percentage of the vehicle, and so the practical limit on some systems can be
much lower.
Propulsion methods
Effective
Exhaust
Thrust
Velocity
(N)
(km/s)
Method
Propulsion methods in current
use
Solid rocket
Hybrid rocket
Monopropellant rocket
Bipropellant rocket
Tripropellant rocket
Resistojet rocket
Arcjet rocket
Hall effect thruster (HET)
Electrostatic ion thruster
Field Emission Electric
Propulsion (FEEP)
Pulsed plasma thruster (PPT)
Pulsed inductive thruster (PIT)
Nuclear electric rocket
Currently feasible propulsion
methods
Solar sails
Tether propulsion
Mass drivers (for propulsion)
Launch loop
Firing Duration
1 - 4.7
2.5 - 5.3
2-6
4 - 16
8 - 50
15 - 80
103 - 107 minutes
<0.1 - 107 minutes
milliseconds 0.1 - 100
minutes
0.1 - 107 minutes
minutes
10-2 - 10 minutes
10-2 - 10 minutes
10-3 - 10 months/years
10-3 - 10 months/years
100 - 130
10-6 - 10-3 months/years
1-4
1.5 - 4.2
1-3
Maximum
Delta-v
(km/s)
~7
>3
~3
~9
~9
> 100
> 100
~ 2,000 - ~
10,000 hours
50
20
months
As electric propulsion method used
~ 20
N/A
N/A
30 - ?
N/A
~ 0.1
9 per km²
Indefinite
(at 1 AU)
1 - 1012 minutes
104 - 108 months
~104
minutes
> 40
~7
>> 11
Orion Project (Near term
nuclear pulse propulsion)
Magnetic field oscillating
amplified thruster
Variable specific impulse
magnetoplasma rocket
(VASIMR)
Magnetoplasmadynamic
thruster (MPD)
Nuclear thermal rocket
Solar thermal rocket
Radioisotope rocket
20 - 100
109 - 1012 several days
~30-60
10 - 130
0,1 - 1
days - months
> 100
10 - 300
40 1,200
days - months
> 100
20 - 100
100
weeks
9
7 - 12
7-8
105
1 - 100
Air-augmented rocket
5-6
0.1 - 107
Liquid air cycle engine
4.5
minutes
weeks
months
secondsminutes
secondsminutes
minutes
1000 107
0.1 - 107
SABRE
30/4.5
Dual mode propulsion rocket
Technologies requiring further
research
Magnetic sails
N/A
Indefinite Indefinite
Mini-magnetospheric plasma
200
~1 N/kW months
propulsion
Nuclear pulse propulsion
20 - 1,000 109 - 1012 years
(Project Daedalus' drive)
Gas core reactor rocket
10 - 20
10³ - 106
Nuclear salt-water rocket
100
10³ - 107 half hour
As propulsion method powered by
Beam-powered propulsion
beam
Fission sail
Fission-fragment rocket
1,000
Nuclear photonic rocket 300,000 10-5 - 1 years-decades
Fusion rocket 100 - 1,000
Space Elevator
N/A
N/A
Indefinite
Significantly beyond current
engineering
Antimatter catalyzed nuclear
200 - 4,000
days-weeks
pulse propulsion
> ~ 20
> ~ 20
> 7?
?
9.4
~15,000
> 12
10,000 100,000
Bussard ramjet 2.2 - 20,000
Gravitoelectromagnetic toroidal
launchers
Antimatter rocket
indefinite
~30,000
<300,000
Testing spacecraft propulsion
Spacecraft propulsion systems are often first statically tested on the Earth's surface,
within the atmosphere but many systems require a vacuum chamber to test fully.
Rockets are usually tested at a rocket engine test facility well away from habitation and
other buildings for safety reasons. Ion drives are far less dangerous and require much
less stringent safety, usually only a large-ish vacuum chamber is needed.
Famous static test locations can be found at Rocket Ground Test Facilities
Some systems cannot be adequately tested on the ground and test launches may be
employed at a Rocket Launch Site.
Electric propulsion
Electric propulsion is a form of spacecraft propulsion used in outer space. This type of
rocket-like reaction engine utilize electric energy to obtain thrust from propellant
carried with the vehicle. Unlike rocket engines these kinds of engines do not necessarily
have rocket nozzles, and thus many types are not considered true rockets.
While electric thrusters typically offer much higher specific impulse, due to electrical
power constraints thrust is weaker compared to chemical thrusters by several orders of
magnitude.
The idea of electric propulsion dates back to 1906, when Robert Goddard considered
the possibility in his personal notebook. Konstantin Tsiolkovsky published the idea in
1911.
Types of electric propulsion
The various technologies of electric propulsion for spacecraft are usually grouped in
three families based on the type of force used to accelerate the ions of the plasma.
Electric propulsion systems can also be characterized in terms of their operation in
either steady (continuous firing for a prescribed duration) or unsteady (pulsed firings
accumulating to a desired thrust bit).
Electrostatic
If the acceleration is caused mainly by the Coulomb Force (i.e application of a static
electric field in the direction of the acceleration) the device is considered electrostatic.
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Ion thruster
Hall effect thruster
Field Emission Electric Propulsion
Colloid thruster
Electrothermal
The electrothermal category groups the devices where electromagnetic fields are used
to generate a plasma to increase the heat of the bulk propellant. The thermal energy
imparted to the propellant gas is then converted into kinetic energy by a nozzle of either
physical material construction or by magnetic means. Low molecular weight gases are
preferred propellants (e.g. hydrogen, helium, ammonia) for this kind of system.
Performance of electrothermal systems in terms of specific impulse (Isp) is somewhat
modest (500 to ~1000 seconds), but exceeds that of cold gas thrusters, monopropellant
thrusters, and even most bi-propellant thrusters. In the USSR electrothermal engines
were used since 1971, Soviet "Meteor-3", "Meteor-Priroda", "Resurs-O" satellite series
and Russian "Elektro" satellite are equipped with them. Electrothermal systems by
Aerojet (MR-510) are currently used on Lockheed-Martin A2100 satellites using
hydrazine as a propellant.
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DC arcjet
microwave arcjet
VASIMR
Electromagnetic
If the ions are accelerated either by the Lorentz Force or by the effect of an
electromagnetic fields where the electric field is not in the direction of the acceleration,
the device is considered electromagnetic.
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MPD thruster
Electrodeless plasma thruster
Pulsed inductive thruster