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The Tsiolkovsky rocket equation, or ideal rocket equation, describes the motion of vehicles that follow the basic principle of a rocket: a device that can apply acceleration to itself (a thrust) by expelling part of its mass with high speed and move due to the conservation of momentum. The equation relates the delta-v (the maximum change of speed of the rocket if no other external forces act) with the effective exhaust velocity and the initial and final mass of a rocket (or other reaction engine).
For any such maneuver (or journey involving a number of such maneuvers):
(The equation can also be written using the specific impulse instead of the effective exhaust velocity by applying the formula where is the specific impulse expressed as a time period and is Standard Gravity.)
The equation is named after Russian scientist Konstantin Tsiolkovsky who independently derived it and published it in his 1903 work. The equation had been derived earlier by the British mathematician William Moore in 1813.
This equation was independently derived by Konstantin Tsiolkovsky towards the end of the 19th century and is sometimes known under his name, but more often simply referred to as 'the rocket equation' (or sometimes the 'ideal rocket equation'). However, a recently discovered pamphlet "A Treatise on the Motion of Rockets" by William Moore shows that the earliest known derivation of this kind of equation was in fact at the Royal Military Academy at Woolwich in England in 1813, and was used for weapons research.
While the derivation of the rocket equation is a straightforward calculus exercise, Tsiolkovsky is honored as being the first to apply it to the question of whether rockets could achieve speeds necessary for space travel.
In the following derivation, "the rocket" is taken to mean "the rocket and all of its unburned propellant".
Newton's second law of motion relates external forces () to the change in linear momentum of the whole system (including rocket and exhaust) as follows:
where is the momentum of the rocket at time t=0:
and is the momentum of the rocket and exhausted mass at time :
and where, with respect to the observer:
|is the velocity of the rocket at time t=0|
|is the velocity of the rocket at time|
|is the velocity of the mass added to the exhaust (and lost by the rocket) during time|
|is the mass of the rocket at time t=0|
|is the mass of the rocket at time|
The velocity of the exhaust in the observer frame is related to the velocity of the exhaust in the rocket frame by (since exhaust velocity is in the negative direction)
and, using , since ejecting a positive results in a decrease in mass,
If there are no external forces then (conservation of linear momentum) and
Assuming is constant, this may be integrated to yield:
where is the initial total mass including propellant, the final total mass, and the velocity of the rocket exhaust with respect to the rocket (the specific impulse, or, if measured in time, that multiplied by gravity-on-Earth acceleration).
The value is the total mass of propellant expended, and hence:
(delta v) is the integration over time of the magnitude of the acceleration produced by using the rocket engine (what would be the actual acceleration if external forces were absent). In free space, for the case of acceleration in the direction of the velocity, this is the increase of the speed. In the case of an acceleration in opposite direction (deceleration) it is the decrease of the speed. Of course gravity and drag also accelerate the vehicle, and they can add or subtract to the change in velocity experienced by the vehicle. Hence delta-v is not usually the actual change in speed or velocity of the vehicle.
If special relativity is taken into account, the following equation can be derived for a relativistic rocket, with again standing for the rocket's final velocity (after burning off all its fuel and being reduced to a rest mass of ) in the inertial frame of reference where the rocket started at rest (with the rest mass including fuel being initially), and standing for the speed of light in a vacuum:
Writing as , a little algebra allows this equation to be rearranged as
Then, using the identity (here "exp" denotes the exponential function; see also Natural logarithm as well as the "power" identity at Logarithmic identities) and the identity (see Hyperbolic function), this is equivalent to
The rocket equation captures the essentials of rocket flight physics in a single short equation. It also holds true for rocket-like reaction vehicles whenever the effective exhaust velocity is constant; and can be summed or integrated when the effective exhaust velocity varies. It takes only the propulsive force of the engine into account, neglecting aerodynamic or gravitational forces on the vehicle. As such, it cannot be used by itself to accurately calculate the propellant requirement for launch from (or powered descent to) a planet with an atmosphere, and does not apply to non-rocket systems such as aerobraking, gun launches, space elevators, launch loops, or tether propulsion.
Also, the equation strictly applies only to a theoretical impulsive maneuver, in which the propellant is discharged and delta-v applied instantaneously. Orbital maneuvers involving significantly large delta-v (such as translunar injection) still are under the influence of gravity for the duration of the propellant discharge, which influences the vehicle's velocity. The equation is most accurately applied to relatively small delta-v maneuvers such as those involved in fine-tuning space rendezvous, or mid-course corrections in translunar or interplanetary flights where the gravity field is relatively weak.
Nevertheless, the equation is useful for estimating the propellant requirement to perform a given orbital maneuver, assuming a required delta-v. To achieve a large delta-v, either must be huge (growing exponentially as delta-v rises), or must be tiny, or must be very high, or some combination of all of these. In practice, very high delta-v has been achieved by a combination of
Assume an exhaust velocity of 4,500 meters per second (15,000 ft/s) and a of 9,700 meters per second (32,000 ft/s) (Earth to LEO, including to overcome gravity and aerodynamic drag).
In the case of sequentially thrusting rocket stages, the equation applies for each stage, where for each stage the initial mass in the equation is the total mass of the rocket after discarding the previous stage, and the final mass in the equation is the total mass of the rocket just before discarding the stage concerned. For each stage the specific impulse may be different.
For example, if 80% of the mass of a rocket is the fuel of the first stage, and 10% is the dry mass of the first stage, and 10% is the remaining rocket, then
With three similar, subsequently smaller stages with the same for each stage, we have
and the payload is 10%*10%*10% = 0.1% of the initial mass.
A comparable SSTO rocket, also with a 0.1% payload, could have a mass of 11.1% for fuel tanks and engines, and 88.8% for fuel. This would give
If the motor of a new stage is ignited before the previous stage has been discarded and the simultaneously working motors have a different specific impulse (as is often the case with solid rocket boosters and a liquid-fuel stage), the situation is more complicated.
|This article includes a list of references, but its sources remain unclear because it has insufficient inline citations. (February 2009)|