In pursuit of specific impulse. Conversations about rocket engines How to convert specific impulse into ordinary impulse

The development of a project for a working rocket model is closely related to the issue of the engine. Which engine is better to put on the model? Which of its characteristics are the main ones? What is their essence? The modeler needs to understand these issues.

This chapter talks as simply as possible about the characteristics of the engine, that is, those factors that determine its features. A clear understanding of the value of engine thrust, its operating time, total and specific impulse and their influence on the flight quality of the rocket model will help the model-student designer choose the right engine for the rocket model, and therefore ensure success in competitions.

The main characteristics of the rocket engine are:

• 1. Engine thrust P (kg)
• 2. Operating time t (sec)
• 3. Specific thrust R ud (kg sec/kg)
• 4. Total (total) impulse J ∑ (10 n sec ≈ 1 kg sec)
• 5. Fuel weight G T (kg)
• 6. Secondary fuel consumption ω (kg)
• 7. Gas flow rate W (m/sec)
• 8. Engine weight G dv (kg)
• 9. Engine dimensions l, d (mm)

1. Engine thrust

In the case of an open nozzle (Fig. 11, b), the gases located in the combustion chamber under pressure rush at high speed through the nozzle. In this case, the part of the chamber opposite the nozzle turns out to be unbalanced. The pressure forces acting on that part of the chamber bottom area that is located opposite the nozzle opening are also unbalanced, resulting in thrust.

If we consider only the translational movement of gases along the combustion chamber and nozzle, then the distribution of gas velocity along this path can be characterized by a curve (Fig. 12, a). The pressure on the surface elements of the chamber and nozzle is distributed as shown in Fig. 12, b.

The size of the uncompensated area of ​​the combustion chamber bottom is equal to the area of ​​the smallest cross-section of the nozzle. Obviously, the larger the area of ​​this section, the large quantity gases will be able to leave the combustion chamber per unit time.

Thus, we can conclude: the engine thrust depends on the amount of gases leaving the combustion chamber per unit time as a result of the uncompensated area and gas flow rate caused by pressure imbalance.

To obtain a quantitative relationship, consider the change in the momentum of gases as they flow out of the combustion chamber. Let us assume that during time t a certain amount of gas leaves the combustion chamber of the engine, the mass of which will be denoted by m. If we assume that the translational velocity of the gases in the combustion chamber is zero, and at the exit from the nozzle reaches the value W m/sec, then the change in gas velocity will be equals W m/sec. In this case, the change in the momentum of the mentioned mass of gas will be written as an equality:

However, a change in the momentum of gases can only occur if a certain force P acts on the gas for some time t, then

Replacing the value ΔQ in formula (1) with one equal to J ∑ =P·t, we obtain:

In accordance with the laws of mechanics, the force with which the walls of the chamber and nozzle act on the gas is equal in magnitude to the force P, with which the gas, in turn, acts on the walls of the chamber and nozzle. This force P is the engine thrust.

Substituting into formula (5) instead of the gas mass m its similar value from formula (6), we obtain:

Thus, the final formula for rocket engine thrust is:

To calculate thrust, you can use a formula that has a similar meaning to formula (5), with P=const:

For a constant thrust value, the formula is often used

In cases of variable thrust, for example, when determining the initial, maximum, average thrust and thrust at any time during engine operation, it is necessary to enter into this formula true values U and F of a given engine.

So, thrust is the product of the effective gas flow rate W and the mass per second fuel consumption ω/g.

Problem 1. Determine the thrust of a rocket engine of the DB-Z-SM-10 type, having the following data: P stroke = 45.5 kg sec/kg; G T =0.022 kg; t=4 sec.

Solution. Effective flow rate of gases from the nozzle:

Problem 2. Determine the thrust of a rocket engine of type DB-Z-SM-10, having the following data: 1 kg·sec; G T =0.022 kg; t=4 sec.

Solution. We use formula (11):

2. Gas flow rate

The rate of gas flow depends on various factors. The most important parameter characterizing the state of gases in the combustion chamber is temperature (T°K). The flow rate is directly proportional to the square root of the temperature of the gases in the chamber. The temperature, in turn, depends on the amount of heat released during fuel combustion. Thus, the exhaust rate depends primarily on the quality of the fuel and its energy resource.

3. Specific thrust and specific impulse

Modern model solid propellant rocket engines have low specific thrust values: from 28 to 50 kg sec/kg. There are also new engines with a specific thrust of 160 kg sec/kg and higher, with a lower pressure limit of no higher than 3 kg/cm 2 and a relatively high specific gravity of the fuel - more than 2 g/cm 3 .

Specific thrust shows the efficiency of using one kilogram of fuel in a given engine. The higher the specific thrust of the engine, the less fuel is consumed to obtain the same total engine impulse. This means that with the same fuel weight and engine size, the one with a higher specific thrust will be preferable.

Problem 3. Determine the weight of the fuel in each of the four engines with a total impulse of 1 kg sec, but with different specific thrusts: a) P ud = 28 kg sec/kg; b) P ud =45.5 kg sec/kg; c) P ud =70 kg sec/kg; d) P ud =160 kg sec/kg.

Solution. The weight of the fuel is determined from the formula:

The specific impulse Jsp is understood as the ratio of the total thrust impulse during the time t of engine operation to the weight of the fuel consumed during this time G T .

At constant thrust, i.e. at constant pressure in the combustion chamber and engine operation on the ground, J beat = P beat.

4. Calculation of engine characteristics DB-1-SM-6

Problem 4. Calculate the main characteristics of the DB-1-SM-6 engine, whose body is a 12-gauge paper hunting cartridge case. The fuel is mixture No. 1 (potassium nitrate - 75, sulfur - 12 and charcoal - 26 parts). Compaction density (fuel specific gravity) γ = 1.3-1.35 g/cm 2 , P = 30 kg sec/kg, K = 100. We set the maximum pressure in the combustion chamber to be within 8 kg/cm 2 . The burning rate of a given fuel depending on pressure at normal temperature environment presented on the graph in Fig. 13.

Solution. First of all, it is necessary to draw the engine housing, i.e., a 12-gauge sleeve (Zhevelo), which makes it possible to visually follow the progress of the calculations (Fig. 14). The engine housing (sleeve) has a ready-made nozzle (hole for the Zhevelo piston). The hole diameter is 5.5 mm, the sleeve length is 70 mm, its internal diameter is 18.5 mm, its external diameter is 20.5 mm, the nozzle length is 9 mm. The engine fuel block must have free space - a longitudinal channel, thanks to which it is possible to increase the fuel combustion area in the engine to its maximum value. The shape of the channel is a truncated cone, the lower base of which corresponds to the size of the hole in the sleeve (5.5 mm), and during calibration can be equal to 6 mm. The diameter of the upper base is 4 mm. The upper base is made somewhat smaller due to technological considerations and safety precautions when removing the metal cone from the powder mass. To determine the length of the cone (rod), initial data is required, which is obtained in the following order.

Using formula (15), the possible maximum combustion area is determined:

Channel length l=h 1 -r=4.27-0.7=3.57 cm. We will immediately plot the resulting dimensions on the drawing (Fig. 15). Length of the rod for pressing: 3.57 + 0.7 = 4.27 cm (0.7 cm - height of the sleeve base).

Let's move on to determining the height of the sustainer part of the fuel bomb. This part of the fuel block does not have a channel, i.e. it is pressed in completely. Its purpose is to obtain a cruising section, preferably with constant thrust, after achieving the maximum thrust value. The height of the marching part of the checker must be strictly defined. The combustion of the propellant part of the rocket fuel occurs in the engine with a slight pressure of 0.07-0.02 kg/cm 2 . Based on this, according to the graph in Fig. 13 we determine the burning rate of the propulsion part of the fuel: U=0.9 cm/sec.

The height of the main part h 2 for the burning time t=1.58 sec. will make up.

The content of the article

ROCKET, an aircraft moving due to the ejection of high-speed hot gases created by a jet (rocket) engine. In most cases, the energy to propel a rocket is obtained from the combustion of two or more chemical components (fuel and oxidizer, which together form rocket fuel) or from the decomposition of one high-energy chemical. Most rockets are one of two types - solid propellant or liquid propellant. These terms refer to the form in which the propellant is stored before it is burned in the rocket engine chamber. The rocket consists of a propulsion system (engine and fuel compartment), control and guidance systems, payload and some auxiliary systems.

THEORY OF MOTION

Two familiar examples explain the principle of rocket motion. When a gun is fired, the powder gases, expanding in the barrel, push the bullet forward and the gun back. The bullet flies to the target, and the shooter (or the carriage of the artillery gun) absorbs the recoil energy due to the force of friction with the surface of the earth. If the shooter were standing on skates on ice, the recoil would cause him to roll backwards (and only stop due to friction with the air and ice).

Another example is an inflated balloon. While the ball hole is closed, the internal air pressure is balanced by the elastic forces of the ball shell. If you open the hole, air will escape from the ball, and its unbalanced pressure on the shell will push the ball forward. Note that the ball is driven by a force acting only on the area of ​​the hole. All other forces acting on the shell are balanced and do not affect the movement of the ball, which is chaotic due to the continuous change in the shape of the ball and the flexibility of its neck.

A rocket engine works in a similar way, except that combustion or chemical decomposition reactions produce a steady stream of hot gases that are expelled out through a nozzle. There are other methods for producing a jet gas stream ( see below), however, none of them has become as widespread as the chemical one.

All the above examples of the movement of a shooter and a bullet, an inflated ball and a rocket are described by Newton's third law of motion, which states that every action has an opposite and equal reaction. Mathematically, this law is expressed as the equality of quantities of motion MV = mv. It is important to note that the total change in momentum (momentum) in the system is zero. If two masses M And m are equal, then their speeds V And v are also equal. If the mass of one of the interacting bodies is greater than the mass of the other, then its speed will be correspondingly lower. In the example with the shooter, the impulse mv, communicated to the bullet, is exactly the same as the impulse MV, reported to the shooter, however, due to the low mass of the bullet, its speed is much greater than the shooter's speed. In the case of a rocket, the release of gases in one direction (action) causes the rocket to move in the opposite direction (reaction).

ROCKET ENGINE

Inside a working rocket engine, an intense process of rapid, controlled combustion occurs. To carry out a combustion reaction (the release of energy during the reaction of two chemical substances, which results in the formation of products with less latent energy), the presence of an oxidizing agent (oxidizer) and a reducing agent (fuel) is necessary. During combustion, energy is released in the form of heat, i.e. internal movement of atoms and molecules as a result of increased temperature.

Design.

A rocket engine consists of two main parts: the combustion chamber and the nozzle. The chamber must have sufficient volume for complete mixing, evaporation and combustion of fuel components. The chamber itself and the fuel supply system must be designed in such a way that the gas speed in the chamber is below the speed of sound, otherwise combustion will be ineffective. As in the case of an inflatable balloon, gas molecules collide with the walls of the chamber and exit through a narrow opening (the neck of the nozzle). When the gas flow is constrained in the tapering part of the nozzle, its speed increases to the speed of sound in the neck, and in the diverging part of the nozzle the gas flow becomes supersonic. A nozzle of this design was proposed by Carl de Laval, a Swedish engineer working in the field steam turbines, in the 1890s.

The contour of the expanding part of the nozzle and the degree of its expansion (the ratio of the areas at the outlet and in the neck) are selected based on the speed of the gas jet and the ambient pressure, so that the pressure of the exhaust gases on the walls of the supersonic part of the nozzle increases the thrust force created by the gas pressure on the front part combustion chambers. Since external (atmospheric) pressure decreases with increasing altitude, and the flared portion of the nozzle profile can only be optimized for one altitude, the expansion ratio is chosen to provide acceptable efficiency for all altitudes. The engine for low altitudes should have a short nozzle with a small expansion ratio. Nozzles designed for adjustable expansion ratio. However, in practice they turn out to be too complex and expensive and are therefore rarely used.

Thrust and specific impulse of thrust.

Engine thrust F is equal to the product of the pressure created by the exhaust gases and the exit cross-sectional area of ​​the nozzle, minus the force of environmental pressure on the same area. The efficiency of an engine is measured by its specific impulse Isp, which has several different units of measurement. One of the units is thrust divided by total second fuel consumption ( w), i.e. I sp = F/w. The other is the effective exhaust velocity C, divided by the acceleration due to gravity g, in this case I sp = C/g. Specific impulse is usually expressed in seconds (SI Isp measured in LF s/kg or m/s), and in this case its value is equal to the number of kilograms of thrust obtained from the combustion of one kilogram of fuel. Magnitude Isp depends on a number of factors, mainly the energy released during combustion of the fuel and the efficiency of using this energy in the engine (for example, a short conical nozzle in a vacuum will be less efficient than a long and carefully shaped one).

Relative initial mass and characteristic speed of the rocket.

These quantities are the main characteristics of the rocket as aircraft. Relative initial mass is the ratio of the initial mass of the rocket W to its final mass after fuel burnout w. Magnitude Isp depends on the design perfection of the rocket and the efficiency of its engine; these parameters determine the final speed that the rocket develops. The characteristic final speed of a rocket is determined by the Tsiolkovsky formula

V b 0 = (gIsp ln[ W/w]) – (V Lg + V Ld + V Lt),

Where V Lg, V Ld And V Lt– speed losses (determined from additional equations) associated with gravity, atmospheric resistance and lower traction in the atmosphere.

As can be seen from this formula, to increase the final speed of the rocket it is necessary: ​​1) increase the relative initial mass ( W/w) due to lightweight design; 2) increase specific impulse through the use of higher energy fuel; 3) reduce drag by improving flow around and reducing the size of the rocket. However, due to the fact that the flight mission of a rocket (especially a space rocket) changes from flight to flight, and during the flight, external conditions continuously change, compromises have to be made when designing a rocket.

The charge geometry can be neutral, progressive or regressive depending on how the engine's thrust is to be varied. A charge of neutral geometry is a solid cast cylindrical rod that burns at one end (end combustion charge). Special protective coatings prevent fuel from burning from the edges. The progressive geometry charge is usually cast in the form of a tube; combustion occurs on the inside (channel combustion charge). As such a charge burns out, the combustion surface and, accordingly, the thrust increase. By giving the channel a star shape, it is possible to ensure that the burnout rate and draft decrease over time; The conical channel allows you to smoothly adjust the traction.

By giving the charge a special shape or combining several simple shapes, you can obtain the desired law for changing the rocket's thrust in flight. For an air-to-air projectile, for example, a progressive geometry charge can be used to obtain the high accelerations necessary to intercept a target. In spaceflight launch vehicles, on the other hand, a combination of progressive and regressive charge geometries is more useful in order to obtain more thrust at launch, when the rocket is at maximum mass and atmospheric drag is high, and less thrust in the upper atmosphere, when the rocket mass is low. and the accelerations are high.

Composition and production technology.

The solid fuel mixture most commonly used in the United States is ammonium perchlorate as an oxidizer and aluminum powder as a fuel with a polymer binder, nitrile butadiene rubber (Russian designation SKN - synthetic nitrile rubber). Iron oxide powder is added to control the burning rate. Mixtures of these components in various proportions are used for space launch vehicles, ballistic and tactical missiles. These fuels have a specific impulse of 280 to 300 s depending on the mixture composition. The combustion products of such solid propellant engines contain hydrogen chloride and aluminum oxide particles.

The fuel described above is obtained by grinding individual components into a fine powder and then mixing them with elastic SKN in special mixers, similar in design to conventional industrial dough mixers. Once the mixture is sufficiently mixed, it is poured into the engine housing. A special mold is inserted into the engine to obtain the desired charge configuration (the process is similar to making a sponge cake). The charge is then polymerized at a carefully controlled temperature. After the polymerization process is completed, the insert is removed, and the nozzle, ignition device and other elements necessary for starting the engine and flying the rocket are attached to the body.

The manufacture of even the simplest solid fuel motor is highly hazardous and requires careful controls, such as protection against static electricity, the use of non-sparking materials, and good ventilation of fumes and dust to ensure worker safety. Industrial premises for equipment, solid propellant motors are usually separated by thick walls and have weak roofs so that the blast wave in the event of an accident goes up and does not cause much damage.

The body of a solid propellant motor is usually made by welding high-quality metal alloys or composite materials, wound around a mandrel that follows the outer contours of the propellant charge. The body must have very high strength to withstand internal combustion pressure, especially at the end of flight. Once the housing is complete, it is cleaned and insulated to prevent burnout. For better contact between insulation and charge, a binder is often used.

One of the last stages of manufacturing a solid fuel engine is checking it for defects and foreign inclusions. Cracks in the charge serve as additional combustion surfaces, which can lead to an increase in thrust and a change in the flight path. In the worst case, the pressure in the combustion chamber can become so high that the engine is destroyed. The process of equipping the engine is completed by installing the starting igniter on its front bottom and the nozzle on the rear. The pilot igniter is usually a small rocket engine containing a fast-burning propellant that emits a flame and ignites the propellant charge.

Some military applications require accelerations that SKN-based engines cannot provide; then metallized mixed fuels based on nitroglycerin or other powerful explosives are used. In these cases, a controlled explosion process occurs in the engine. To control the explosion process, special chemical reaction retardants are added. Other military needs required the development of smokeless-burning tactical missiles so that it would be impossible to trace where the missile was launched from.

Tests.

Solid propellant motors are usually tested on firing stands, where the engine is mounted motionless in a horizontal or vertical position and the operation of all its systems is checked. During engine operation, sensors installed on it measure thrust, pressure and temperature of combustion products, loads on the body, etc. During fire tests, all possible operating modes are checked, including off-design ones, which should not exist during normal flight.

Advantages and disadvantages.

Solid fuel engines are used in cases where the main requirements are simplicity, ease of maintenance, quick starting and high power in a small volume. The first American ballistic missiles used liquid fuel, but starting in the 1960s there was a transition to solid fuel, which was due to the improvement of its production technology. Solid propellant rocket engines have always been used in small military projectiles and rockets, ejection devices on jet planes and for separating rocket stages.

The main disadvantage of solid fuel engines is the practical impossibility of regulating thrust during flight, as well as the difficulty of turning off the engine. In some solid propellant rocket engines, thrust cut-off is accomplished by opening holes in the front of the engine. When the holes are opened (usually this happens with the help of special squibs), the pressure inside the engine drops and, accordingly, the combustion intensity decreases. In addition, reverse thrust occurs, opposite to the normal thrust of the main nozzle, and the acceleration of the rocket stops. Since the thrust of a solid propellant rocket engine is determined by the geometry and chemical composition of the charge, changing engine parameters to obtain a different thrust versus time relationship may require a full cycle of testing of the new engine.

LIQUID ROCKET STAGES

The most efficient rockets use liquid fuel because the chemical energy of liquid components is greater than that of solid components, and their combustion products have a lower molecular weight.

Cryogenic and self-igniting fuels.

Liquid fuels that have a high calorific value include some cryogenic substances - gases that turn into liquid at very low temperatures, such as liquid oxygen (at temperatures below - 183 ° C) and liquid hydrogen (below - 253 ° C). On the other hand, the use of cryogenic components has a number of disadvantages, which include the need to maintain large industrial installations for liquefying gases, a long time for refueling a rocket (several hours) and the need for thermal insulation of fuel tanks. Therefore, America's first cryogenically fueled intercontinental ballistic missiles, the Atlas and Titan I, were vulnerable to surprise attack with only a few minutes to retaliate.

Liquid rocket engines (LPREs), which use self-igniting liquid fuel that can be stored at normal temperatures for long periods of time and ignite when components come into contact with each other, were created in the 1950s to meet military needs for easier operation and reduced preparation time. to launch ballistic missiles. In such engines, nitrogen tetroxide (N 2 O 4) was used as an oxidizer, and hydrazine (N 2 H 4) or unsymmetrical dimethylhydrazine (NH 2 - N 2) as a fuel - a combination that gives a specific impulse of about 340 s. Self-igniting fuel components are extremely toxic and quite corrosive, so they require extreme care in handling and periodic replacement of structural elements that contain them or are in contact with them. And although liquid-propellant ballistic missiles with self-igniting fuel were subsequently replaced by solid propellant, this fuel is still indispensable in attitude control and correction engines.

Two-component rocket engines.

In the liquid-propellant rocket engines described above, the fuel and oxidizer are stored in separate tanks and, by displacement or using pumps, are fed into the combustion chamber, where they ignite and burn, creating a high-speed gas jet. Liquid oxygen is often used as an oxidizing agent due to the ease of its production from atmospheric air. Although compared to many others chemicals Liquid oxygen is relatively safe; only very clean containers should be used to store it, because oxygen reacts chemically even with grease stains left by fingerprints, which can lead to a fire.

Heavy hydrocarbons or liquid hydrogen are most often used as fuel paired with oxygen. The heat of combustion of hydrocarbon fuel per unit volume, for example, refined kerosene or alcohol, is higher than that of hydrogen. Hydrocarbon fuel burns with a bright orange flame. The main combustion products of the oxygen/hydrocarbon mixture are carbon dioxide and water vapor. The specific impulse of such fuel can reach 350 s.

Liquid hydrogen requires deeper cooling than liquid oxygen, but its calorific value per unit mass is higher than that of hydrocarbon fuels. Hydrogen burns with an almost invisible blue flame. The main product of combustion of the oxygen-hydrogen mixture is superheated water vapor. The specific impulse of engines using this fuel can reach from 450 to 480 s, depending on the engine design. (Engines using liquid hydrogen usually operate in excess fuel mode, which reduces fuel mass consumption and improves efficiency.)

Over the years, many other fuel and oxidizer combinations have been tested, but most have had to be abandoned due to their toxicity. For example, fluorine is a more effective oxidizer than oxygen, but it is extremely toxic and aggressive both in its initial state and in combustion products. Various mixtures nitric acid with nitrogen oxides were previously used as an oxidizer, but their advantages were outweighed by the dangers of storing and operating such engines and rockets.

It is not always easy to make a choice between hydrocarbon fuel and liquid hydrogen. Typically, the first stages of rockets use liquid hydrocarbon (or mixed solid) fuel to pass through dense layers of the atmosphere in the first minutes of flight. Of course, liquid hydrogen is a very efficient fuel, but due to its low density, the first stage would require large fuel tanks, which would increase the weight of the structure and the drag of the rocket. At high altitudes and in space, hydrogen engines are more often used, where their advantages are fully demonstrated.

Three-component rocket engines.

Since the early 1970s, the concept of three-component engines, which would combine the advantages of minimum volume and minimum weight in a single engine, has been studied in Russia and the United States. When starting, such an engine would run on oxygen and kerosene, and at high altitudes it would switch to using liquid oxygen and hydrogen. This approach could possibly make it possible to create a single-stage rocket, but the design of the engine would be significantly more complicated.

Single-component rocket engines.

Such engines use single-component liquid fuel, which, when interacting with a catalyst, decomposes to form hot gas. Although single-component liquid-propellant rocket engines develop a small specific impulse (in the range from 150 to 255 s) and are much inferior in efficiency to two-component ones, their advantage is the simplicity of their design. Fuel such as hydrazine or hydrogen peroxide is stored in a single container. Under the action of displacement pressure, the liquid enters the combustion chamber through the valve, in which a catalyst, for example, iron oxide, causes its decomposition (hydrazine into ammonia and hydrogen, and hydrogen peroxide into water vapor and oxygen). Single-component liquid-propellant rocket engines are usually used as low-thrust engines (sometimes their thrust is only a few Newtons) in orientation and stabilization systems for spacecraft and tactical missiles, for which simplicity and reliability of design and low mass are the determining criteria. There is a remarkable example of the use of a hydrazine thruster on board the first American communications satellite, TDRS-1; this engine ran for several weeks to propel the satellite into geostationary orbit after the booster failed and the satellite was stranded in a much lower orbit.

The simplest single-component engine is powered by a bottle of compressed cold gas (such as nitrogen) released through a valve. Such jet engines are used where the thermal and chemical effects of the exhaust jet of gas or combustion products are unacceptable and where the main requirement is simplicity of design. These requirements are met, for example, by individual cosmonaut maneuvering devices (UMDs), located in the backpack behind the back and intended for movement when working outside the spacecraft. The MCUs operate from two cylinders of compressed nitrogen, which is supplied through solenoid valves to a propulsion system consisting of 16 engines.

Propulsion system.

The greater power, controllability, and high specific impulse of liquid-propellant rocket engines come at the cost of design complexity. Special systems must ensure the supply of fuel and oxidizer in strictly defined quantities from fuel tanks to the combustion chamber. The supply of fuel components is carried out using pumps or by displacing them with gas pressure. In displacement systems, typically used in small propulsion systems, fuel is supplied by pressurizing the tanks; in this case, the pressure in the tank must be greater than in the combustion chamber.

The pumping system uses mechanical pumps to deliver fuel, although some tank pressurization is also used (to prevent pump cavitation). The most commonly used are turbopump units (TPU), where the turbine is powered by gas from its own propulsion system. Sometimes the turbine is powered by gas obtained from the evaporation of liquid oxygen as it passes through the engine cooling circuit. In other cases, a special gas generator is used in which a small amount of basic fuel or special single-component fuel is burned.

The Shuttle's propellant engine, with its pump fuel supply system, is one of the most advanced engines ever flown into space. Each engine has two pumps - booster (low-pressure) and main (high-pressure). Fuel and oxidizer have the same supply systems. The booster pump, driven by the expanding gas, increases the pressure of the working fluid before it enters the main pump, in which the pressure increases even more. Most of the liquid oxygen passes through the combustion chamber cooling path and nozzles (and in some designs, the fuel pump) before it is introduced into the combustion chamber. Part of the liquid oxygen is supplied to the gas generators of the main fuel pumps, where it reacts with hydrogen; This produces hydrogen-rich steam, which expands in the turbine, drives the pumps, and is then fed into the combustion chamber, where it burns with the remaining oxygen. Although small amounts of oxygen and hydrogen are consumed to drive the booster pumps and pressurize the oxygen and hydrogen tanks, they also eventually pass through the main combustion chamber and contribute to the generation of thrust. This process provides overall engine efficiency of up to 98%.

Production.

The production of liquid rocket engines is more complex and requires greater precision than the production of solid rocket engines, since they contain high-speed rotating parts (up to 38,000 rpm in the main fuel pumps of the Shuttle propulsion engine). The slightest inaccuracy in the manufacture of rotating parts can lead to vibration and destruction.

Even when the blades, wheels and shafts of engine turbines and pumps are properly balanced, other problems can arise. Experience with the J-2 oxygen-hydrogen engine used in the second and third stages of the Saturn 5 rocket has shown that such engines often suffer from high-frequency instability. Even if the engine is properly balanced, the interaction of the fuel pump with the combustion process can cause vibration at a frequency close to the rotational speed of the hydrogen pump. Engine vibrations occur in specific directions rather than randomly. With this instability, the level of vibration can become so great that it requires shutting down the engine to avoid engine damage. Combustion chambers are usually a welded or stamped thin-walled metal structure with a cooling path and a mixing head for supplying fuel.

Tests.

A necessary stage in the development of a liquid-propellant rocket engine and its units is testing them on hydraulic and firing stands. During fire tests, the engine operates at pressures and rotation speeds of the pump, which exceed normal operating values, so that the permissible maximum loads on individual units and the structure as a whole can be assessed. Flight engine samples must undergo acceptance tests, which include short-term and control-selective fire tests simulating the main stages of the flight. The total time of testing and operation of the engine in flight should not exceed its total service life.

Shutdown, restart and traction control.

The main advantage of the liquid-propellant rocket engine is the ability to turn off, restart and regulate thrust. The Shuttle propulsion engine, for example, can operate steadily in the range from 65 to 104% of rated thrust. The crew of the lunar module of the Apollo spacecraft, maneuvering during landing, could adjust the thrust of the engines to 10% of the nominal value. On the contrary, the thrust of the engines providing the launch of the module from the Moon was not regulated, which made it possible to increase their efficiency and reliability.

The possibility of restarting a liquid-propellant rocket engine in space is a problem, since the fuel, like any objects in zero gravity, is chaotically located inside the tanks and will not enter the engine power system in the absence of acceleration. The simplest way to solve the problem is to use special low-thrust engines, which create a slight acceleration sufficient for fuel to flow into the pipelines. Starting of these engines is ensured either by small elastic bags of fuel attached to pipelines, or by means of special grids, on which, due to surface tension forces, enough fuel is retained to start the engine. Elastic fuel containers and liquid collection devices are also used for direct launch of space rocket engines.

CONTROL AND GUIDANCE SYSTEMS

Important integral part Missiles are control and guidance systems. The guidance system determines the position and course of the missile and provides the control system with the necessary data to control its flight. The rocket's flight is controlled by small steering motors or by changing the direction of the main engine's thrust vector.

In large solid propellant engines, the body-nozzle connection may be made of many thin layers of steel and heat-resistant rubber, allowing the nozzle to rotate several degrees in any direction. With the help of one or two hydraulic drives, the nozzle is deflected, changing the direction of the thrust vector. The drives use the energy of a small turbopump unit running on hydrazine decomposition products. In some solid propellant engines, hot gas (from a small auxiliary engine) is supplied through several valves located circumferentially in the diverging part of the nozzle. When one or more valves are closed, the direction of the main jet and, accordingly, the thrust vector change. The liquid-propellant rocket engine is installed in rotating axles or in a gimbal suspension, which allows the entire engine to be rotated.

HISTORICAL REFERENCE

Antiquity and the Middle Ages.

Although rocketry developed in connection with modern military needs and space exploration, the history of rockets goes back to ancient Greece. In the steam engine named after him, Heron demonstrated the principle of jet propulsion. A small metal vessel, shaped like a bird and filled with water, was suspended over the fire. When the water boiled, a stream of steam was released from the bird's tail, pushing the vessel forward. This device was not found practical application, and the principle itself was subsequently forgotten.

In China around 960 AD. For the first time, black gunpowder was used - a mixture of saltpeter (an oxidizing agent) and charcoal with sulfur (fuel) - for throwing projectiles, and in the 11th century. a throwing range of about 300 m was achieved for such projectiles. These “missiles” were bamboo tubes filled with gunpowder and were not particularly accurate in flight. Their main purpose in battle was to cause panic in people and horses. In the 13th century Together with the Mongol conquerors, rockets came to Europe, and in 1248 the English philosopher and naturalist Roger Bacon published a work on their use. The period of use of such unguided rockets for military purposes was short-lived, since they were soon replaced by artillery pieces.

Tsiolkovsky, Oberth and Goddard.

Modern rocket technology owes its development mainly to the works and research of three outstanding scientists: Konstantin Tsiolkovsky (1857–1935) from Russia, Hermann Oberth (1894–1989) from Romania, and Robert Goddard (1882–1945) from the USA. Although these ascetics worked independently of each other and their ideas were often ignored at the time, they laid the theoretical and practical foundations of rocketry and astronautics. Their works inspired generations of dreamers and, most importantly, a few enthusiasts who gave life to their works. see also GODDARD, ROBERT HUTCHINGS ; OBERT, HERMAN; TSIOLKOVSKY, KONSTANTIN EDUARDOVICH.

Tsiolkovsky, a schoolteacher, first wrote about liquid-propellant rockets and artificial satellites in 1883 and 1885. In his work Exploration of world spaces using jet instruments(1903) he outlined the principles of interplanetary flights. Tsiolkovsky argued that the most efficient fuel for rockets would be a combination of liquid oxygen and hydrogen (although even laboratory quantities of these substances were quite expensive at the time), and proposed using a bunch of small engines instead of one large one. He also proposed using multi-stage rockets instead of one large one to facilitate interplanetary travel. Tsiolkovsky developed the basic ideas of crew life support systems and some other aspects of space travel.

In my books Rocket into interplanetary space (Die Rakete zu den Planetenraumen,1923) and Ways to carry out space flights (Wege zur Raumschiffahrt, 1929) G. Oberth outlined the principles of interplanetary flight and performed preliminary calculations of the mass and energy necessary for flights to the planets. His strong point there was a mathematical theory, but in practical activities it did not progress beyond bench testing of rocket engines.

The gap between theory and practice was filled by R. Goddard. As a young man, he was captivated by the idea of ​​interplanetary flight. His first research was in the field of solid propellant rockets, for which he received his first patent in 1914. By the end of World War I, Goddard was well advanced in developing barrel-launched rockets, which were not used by the US Army due to the advent of peace; During World War II, however, his developments led to the creation of the legendary bazooka, the first effective anti-tank missile. The Smithsonian Institution awarded Goddard a research grant in 1917, which resulted in his classic monograph Method for reaching extreme heights (A Method of Reaching Extreme Altitudes,1919). Goddard began work on the rocket engine in 1923, and a working prototype was created by the end of 1925. On March 16, 1926, he launched the first liquid-propellant rocket, using gasoline and liquid oxygen as fuel, in Auburn, Massachusetts. During World War II, Goddard worked on launch boosters for naval aviation.

The work of Tsiolkovsky, Oberth and Goddard was continued by groups of rocketry enthusiasts in the USA, USSR, Germany and Great Britain. IN THE USSR research papers conducted by the Jet Propulsion Research Group (Moscow) and the Gas Dynamics Laboratory (Leningrad). Members of the British Interplanetary Society BIS, limited in their testing by the British Fireworks Act dating back to the Gunpowder Plot (1605) to blow up Parliament, focused their efforts on developing a "manned lunar spacecraft" based on the technology available at the time.

The German Society for Interplanetary Communications VfR in 1930 was able to create a primitive installation in Berlin, and on March 14, 1931, VfR member Johannes Winkler carried out the first successful launch of a liquid-propellant rocket in Europe.

Nazi Germany.

The German army saw missiles as weapons that it could use without fear of international sanctions, since the Treaty of Versailles (which concluded the First World War) and subsequent military treaties made no mention of missiles. After Hitler came to power, the German military department was allocated additional funds for the development of rocket weapons, and in the spring of 1936 a program was approved to build a rocket center at Peenemünde (von Braun was appointed its technical director) on the northern tip of the island of Usedom off the Baltic coast of Germany.

The next rocket, the A-3, had a 15 kN thrust engine with a liquid nitrogen pressurization system and a steam generator, a gyroscopic control and guidance system, a flight parameters control system, electromagnetic servo valves for supplying fuel components and gas rudders. Although all four A-3 rockets exploded on or shortly after launch from the Peenemünde test site in December 1937, the technical experience gained from these launches was used to develop the 250 kN thrust engine for the A-4 rocket, the first successful launch of which was October 3, 1942.

After two years of design tests, pre-production and troop training, the A-4 rocket, renamed V-2 ("Weapon of Vengeance 2") by Hitler, was deployed starting in September 1944 against targets in England, France and Belgium.

Post-war period.

The A-4 missile demonstrated the enormous capabilities of rocketry, and the most powerful post-war powers—the United States and the Soviet Union—soon became embroiled in the development of ballistic guided missiles capable of delivering nuclear weapons. Advances in rocket technology also made it possible to create tactical missiles that radically changed the nature of warfare.

While the military departments of both countries were improving combat missiles, many scientists (S.P. Korolev in the USSR, W. von Braun in the USA) sought to use the capabilities of rocket technology to deliver scientific instruments and, ultimately, people into space. Since the launch of the first satellite in 1957 and the first cosmonaut Yu. Gagarin in 1961, rocket and space technology has come a long way.

ADVANCED MISSILE SYSTEMS

Until the end of the 20th century. fuel combustion remained the main source of energy for jet propulsion. Although many promising technical concepts have been proposed since the 1920s, most of them have not been implemented in practice.

Hybrid engines.

A tempting alternative to solid propellant rocket engines and liquid propellant engines is the idea of ​​a hybrid engine, which combines best qualities both. A hybrid engine uses a solid fuel and a liquid oxidizer, such as liquid oxygen or nitrogen tetroxide. This approach makes it possible to simplify the fuel supply system by half while maintaining the inherent compactness of the solid propellant rocket engine. Since the oxidizer and fuel are stored separately, cracks in the solid propellant charge are less dangerous than in a traditional solid propellant rocket, which simplifies its manufacture. However, despite significant research efforts, especially in the 1980s, this idea has never been found. wide application. The main problem was the insufficiently stable and efficient combustion process.

Electric rocket engine.

Electricity can be used to heat the working fluid. An example of such an engine is an ion engine, which uses a high-voltage arc to ionize a working fluid, such as argon or mercury vapor, and an electric field to accelerate the flow of ions. The fundamental advantage of such an engine is a very high specific impulse (up to 5000 s, depending on the design of the engine and the working fluid used). The thrust of ion engines is very small and is usually in the range from 0.02 to 0.03 N. Ion engines are intended for long-term space flights, when a significant total increase in speed is obtained over months of operation in zero gravity conditions. Ion thrusters have also found use on geostationary satellites, where they provide a constant, small pulse sufficient to control attitude and maintain orbit. Other electric propulsion schemes use high-energy plasma and the magnetohydrodynamic effect.

Nuclear rocket engines.

Another reactive system that has almost achieved practical implementation is nuclear. In the United States, as part of the NERVA program to create a nuclear rocket engine (NRE), a graphite reactor was developed, cooled by liquid hydrogen, which was evaporated, heated and ejected through a rocket nozzle. Graphite was chosen for its high temperature resistance. According to the NERVA project, the YARD was supposed to develop a thrust of 1100 kN for one hour and have a specific impulse of 800 s, which is almost twice the corresponding figure for chemical engines. The NERVA program was canceled in 1972 because the manned mission to Mars for which it had been developed was postponed indefinitely.

A variation of the fission nuclear engine is the gas-phase nuclear engine, in which a slow-moving gas stream of fissile plutonium is surrounded by a faster flow of cooling hydrogen. This idea, however, did not go beyond the preliminary research stage.

An interesting idea of ​​​​creating an engine that uses the reaction of annihilation of matter and antimatter was studied as part of the US Strategic Defense Initiative (SDI) program. Antimatter in the form of atoms is stored in an electromagnetic trap and, through a magnetic field, is fed into the engine chamber, where it interacts with ordinary matter, turning into gamma radiation, which heats working fluid and creates a jet stream. Although magnetic traps are used in high-energy physics, obtaining the few grams of antimatter needed for flight requires great amount energy.

External energy sources.

The SDI and National Aeronautics and Space Administration (NASA) programs also studied a rocket system with a high-power laser that heats the working fluid on board the rocket. The rocket itself has a small mass, since the bulk of the system is the laser, which can be located on Earth. Such a system requires extremely precise aiming of the laser beam at the target, so as not to burn the rocket instead of heating the working fluid. The idea of ​​using large mirrors to focus the sun's rays onto the engine was also considered.

Using the energy of an atomic explosion.

In the 1960s, NASA and the US Atomic Energy Commission explored one rather exotic method of generating thrust as part of the Orion project. In this method, rocket acceleration to high speed, necessary for flight to other planets, was supposed to be carried out through successive explosions of small atomic charges ejected behind the rocket. Special dampers were supposed to smooth out the effects of explosions. However, the Orion project was canceled in accordance with international treaties on the use of outer space and the limitation of nuclear weapons.

Photon engines.

The possibility of using light to generate thrust in space was also studied. Particles of light - photons - create a very small reactive impulse when exposed to a surface. The simplest engine of this kind is a huge plastic mirror that reflects the sun's rays and pushes the spacecraft away from the Sun ( sunny wind creates an additional impulse). In a real photon engine, due to the annihilation of ordinary matter and antimatter, a flux of gamma radiation should be created, providing jet thrust for the movement of the spacecraft.

Everyone probably knows that space mainly consists of vacuum. And in this vacuum there is practically nothing from which to push off, just as we push off from the floor to walk. And if so, then in order to change our movement in the way we need, we need to throw something out of ourselves. And finally, everyone knows that a vehicle that can do this is called a rocket.
Rockets were invented a very, very long time ago, more than one and a half thousand years ago. But they were able to seriously understand the theory of jet propulsion only towards the very end of the 19th century. In particular, it was then that the great Russian scientist Konstantin Eduardovich Tsiolkovsky derived his famous formula:

here V is the final speed of the rocket, I is the specific impulse, M is the mass of the fueled rocket, and m is the mass of the rocket without fuel (or other working fluid).

Specific impulse is the ratio of engine thrust to fuel consumption or other working fluid. In the SI system we measure flow in kg/s and thrust in newtons. Newton, in turn, is equal to kg*m/s 2. As a result, we find that specific impulse is measured, like speed, in meters per second. In essence, it is speed - the effective speed of the working fluid jet escaping from the engine nozzle.
There is another definition of specific impulse: the time during which, with the help of 1 kg of fuel (or other working fluid), the engine can create a thrust of 1 kgf (kilogram-force). Then it is measured in seconds.
The specific impulse from the first definition must be substituted into the Tsiolkovsky formula, but the second definition is often more convenient in calculations. If we want to convert one variant of specific impulse to another, then we can use a simple formula: 1 m/s = 9.81 s. Although most often it is further simplified to: 1 m/s = 10 s. I will use the latter here. Of course, both formulas are applicable only for specific impulse; it is not worth converting the time of milk “escaping” into the speed of the cook’s running along them necessary to save the stove :-)

What's so interesting about this formula? Quite obvious things: the faster the gas stream and the more fuel in the rocket, the faster it will fly.
And the interesting thing about it is the logarithm. This function increases very slowly with the mass ratio underneath it. For the logarithm to equal 1, it must be 2.72. Those. In order for a rocket with a “dry” mass of 10 tons to accelerate to the speed of the working fluid ejected by it, it needs more than 17 tons of this same working fluid. To accelerate this rocket to two speeds of the working fluid, 64 tons of fuel are needed. For three - 191 tons. Finally, for four speeds of the working fluid, 534 tons of working fluid will be required. It is obvious that placing 534 tons of working fluid in a rocket weighing 10 tons, i.e. more than fifty times its own mass - this is a very difficult task. Four jet outflow speeds are the approximate technical limit for the rocket's speed.

Of course, gravity is not taken into account here. It greatly slows down the rocket when moving away from the Earth or from the Sun, but it accelerates the rocket when approaching the Earth and the Sun, as well as when flying past planets along certain trajectories (flying along other trajectories can slow down). As a result, after the launch vehicle engines are turned off, their speed less than that, which can be calculated using this formula, but the maximum speed ever achieved by a spacecraft is several times greater than what a modern rocket can achieve. But now it doesn't matter to us.

Well, why am I doing all this? And how important specific impulse is.
Let's say we need to reach a speed of 18 km/s. This is approximately what is needed to fly beyond the solar system (the exact speed required for such a flight depends on the direction in which we start).
Let the specific impulse of our rocket engine be 450 s or 4500 m/s. This is in line with the best liquid-propellant rocket engines and close to the theoretical limit for chemical engines (if you don't use too toxic components like fluorine).
In this case, to accelerate a rocket weighing 10 tons, exactly those 534 tons of fuel and oxidizer (in this case, liquid oxygen and hydrogen) will be required. A fueled rocket will weigh 544 tons at launch and only 10 will accelerate to the speed we need...
And if you make the specific impulse only twice as large: 900 s or 9000 m/s? Then, to accelerate a rocket weighing 10 tons, only 64 tons of working fluid will be required! Those. the rocket at launch will weigh only 74 tons! If at launch the rocket weighs the same 544 tons, then more than 73 tons will accelerate to 18 km/s!
Thus, a doubling of the specific impulse makes it possible to accelerate more than seven times more cargo, spending less working fluid.
What if we have a specific impulse of 1350 s or 13,500 m/s? We get 28 tons of working fluid per 10 tons of rocket mass, i.e. 38 tons of starting mass. Or the ability to accelerate 143 tons of 544 tons of starting mass to 18 km/s.
Finally, let's dream about 3600 s or 36,000 m/s... 6.5 tons of working fluid to accelerate 10 tons, i.e. 16.5 t starting mass. Or acceleration of 330 tons from 544 starting ones.
An increase in specific impulse by 2 times improves our rocket (reduces the launch mass or increases the accelerated mass) by 7.3 times, an increase by 3 times - by 14.3 times, and an increase by 8 times - an improvement by 33 times!

But how can we achieve such a specific impulse?..
Surely many have heard about plasma and ion engines, and maybe about electric rocket engines in general. In such engines, to accelerate the working fluid, it is not energy contained in the working fluid itself, but energy supplied from outside that is used. Due to this, such engines fundamentally have no specific impulse limitation. At least 1,000,000 m/s! There's just one BUT...
With a specific impulse of 450 s, we will spend approximately 541 MJ of energy to accelerate 1 kg to that same 18 km/s. At 900 s - 259 MJ. At 1350 s - 255 MJ. So far so good. But then things get worse... At 3600 s - 421 MJ. A further increase in specific impulse will lead to an even greater increase in energy costs, because the mass of the working fluid will no longer decrease as quickly as the square of its speed will increase. This energy will be minimal at a specific impulse equal to approximately 0.63 of the final speed. In our case it is 1130 s or 11,300 m/s.
“So what?” the reader will rightly ask, “After all, now we spend 541 MJ, and at 3600 s we will spend only 421!”
And the fact is that now all these 541 MJ are contained in the working fluid itself, and in the case of electric rocket engines we need to supply them from the outside...
Chemical current sources, obviously, do not make sense here: rather than convert hydrogen and oxygen into water in a fuel cell (which is by no means light) in order to power an ion engine from it, which will accelerate some xenon, it is much easier and more efficient to immediately burn hydrogen in the combustion chamber of a conventional rocket engine. Solar panels potentially have an unlimited supply of energy, but their power is very small, so the thrust of the engine will be low. Moreover, these batteries weigh a lot. So they are only suitable for powering satellite orbit correction engines. If we want to send humans to other planets, we will need something different...
A nuclear reactor is a great solution. It contains a lot of energy, can have great power and at the same time relatively small mass. Now there is already a project for a powerful plasma engine powered by a nuclear reactor, which is planned to be used for a flight to Mars (VASIMR). But, alas, this system is far from ideal... Still, even a nuclear reactor does not have such a high power-to-mass ratio that it would be advisable to make an ion engine with a very high specific impulse. If we increase the impulse, we will slightly reduce the mass of the working fluid, but we will greatly increase the mass of the reactor... And still, such a system will provide an acceleration of no more than 0.1 m/s 2 . The acceleration will be long, and there is no talk of starting from the surface of the Earth.

So what to do?.. It’s simple: you need to throw out the extra links in the chain of energy transfer from the reactor to the working fluid! Ideally - to zero. The working fluid must receive energy from the reactor directly. And such systems were created. Soviet and American nuclear rocket engines, which were actually created “in metal” during testing, completely achieved a specific impulse of around 900 seconds! In them, liquid hydrogen passed through a reactor core heated to thousands of degrees (but still solid), where it evaporated and heated, after which it was ejected through a nozzle.
Calculations show that if you make a reactor designed to melt the core, then 1350 seconds is by no means the specific impulse limit. And such reactors can be created with the current level of technology.
Finally, there are projects for gas-phase nuclear rocket engines... In them, uranium will evaporate, and the specific impulse will be the same 3600 seconds or even higher - up to 4500 seconds.
At the same time, nuclear rocket engines not only hypothetically can, but actually work in the atmosphere, and their thrust can be several times greater than their weight, making it possible to launch directly from Earth.
It’s a pity that work on such engines has not received proper funding for a long time... I think it’s already quite obvious what enormous advantages even a 2-3-fold increase in specific impulse provides, not to mention its increase by 8 or even 10 times .

But is 4500 seconds the limit for the specific impulse of sufficiently powerful (capable of providing a rocket acceleration of more than 0.1 m/s2) engines or not?.. Theoretically, no.
During thermonuclear reactions, the reaction products fly apart at a speed of more than 10,000,000 m/s, i.e. the specific impulse of a hypothetical fusion rocket engine could be 1,000,000 or even 1,500,000 seconds. And, best of all, the energy for accelerating the working fluid is again contained in the working fluid itself! By the way, the technical speed limit for a rocket with such an engine can reach 20% of the speed of light...
Alas, thermonuclear research has not yet gone far enough to create a thermonuclear rocket engine. On the other hand, there is every reason to believe that it will be even easier to create than a thermonuclear power plant. When launching from orbit (and, alas, such engines will not work in the atmosphere), we will not have problems creating and maintaining a vacuum, the engine does not need to work continuously for months, like power plant reactors, and finally, we do not need it to provide us with electricity ! To power the ship itself, you can use a separate nuclear reactor, and let the thermonuclear reactor power only itself.
With a specific impulse of even just 450,000 seconds, a rocket with a launch mass of 11 tons, of which only 1 ton will be thermonuclear fuel, will accelerate to almost 430 km/s. If we want to accelerate the ship, decelerate, then accelerate again and decelerate again without refueling, then the same ratio (11 tons at start, of which 1 ton is fuel) is enough for flight at speeds of more than 100 km/s. If we take a launch mass of 12 tons, of which 2 tons are thermonuclear fuel, then the speed of such a flight (round trip) will already be 200 km/s. So in a month you can fly to Mars, work there for a couple of weeks, and return home...

So, dear readers, the exploration of the solar system is already closer than on the horizon :-)

This article is about the characteristics of jet engines. For a concept from explosion engineering, see Explosion impulse.

Specific impulse- an indicator of the efficiency of a jet engine. Sometimes the synonym “specific thrust” is used for jet engines (the term has other meanings), while specific thrust usually used in internal ballistics, while specific impulse- in external ballistics. The dimension of specific impulse is the dimension of speed; in SI units it is meters per second.

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Definitions

Specific impulse- characteristic of a jet engine, equal to the ratio of the impulse (amount of motion) it creates to fuel consumption (usually mass, but can also be related, for example, to the weight or volume of fuel). The greater the specific impulse, the less fuel must be spent to obtain a certain amount of movement. Theoretically, the specific impulse is equal to exhaust speed combustion products may actually differ from it. Therefore, specific impulse is also called effective (or equivalent) exhaust velocity combustion products.

Specific thrust- characteristic of a jet engine, equal to the ratio of the thrust it creates to the mass fuel consumption. It is measured in meters per second (m/s = N s/kg = kgf s/t.e.m.) and means, in this dimension, how many seconds a given engine can create a thrust of 1 N, while expending 1 kg fuel (or thrust of 1 kgf, having consumed 1 t.e.m. of fuel). With another interpretation, the specific thrust is equal to the ratio of thrust to weight fuel consumption; in this case it is measured in seconds (s = N s/N = kgf s/kgf) - this value can be considered as the time during which the engine can develop a thrust of 1 kgf using a mass of fuel of 1 kg (i.e. weighing 1 kgf). To convert weight specific thrust into mass thrust, it must be multiplied by the acceleration of free fall (taken equal to 9.80665 m/s²).

The approximate formula for calculating the specific impulse (exhaust velocity) for jet engines using chemical fuel looks like [ clarify (no comment specified) ]

Where T k is the gas temperature in the combustion (decomposition) chamber; p k and p a is the gas pressure in the combustion chamber and at the nozzle exit, respectively; M- molecular mass of gas in the combustion chamber; u- coefficient characterizing the thermophysical properties of the gas in the chamber (usually u≈ 15 ). As can be seen from the formula to a first approximation, the higher the temperature of the gas, the lower its molecular weight and the higher the ratio of pressures in the RD chamber to the surrounding space, the higher the specific impulse.

Comparison of efficiency of different types of engines

Specific impulse is an important engine parameter that characterizes its efficiency. This value is not directly related to the energy efficiency of the fuel and the thrust of the engine; for example, ion engines have very little thrust, but due to their high specific impulse they are used as maneuvering engines in space technology.

It operates in the mode of short-term periodic switching on (pulses), the total number of which is usually many thousands. Characteristic is the pulse modulation mode with thrust pulses of constant amplitude and variable duration (width) and frequency (from several tens of pulses per second to 1 every few days). According to the value of the total thrust impulse developed over a certain time, pulse rocket engine equivalent to a taxiway operating continuously at lower thrust. However, the advantage is the ability to quickly and with great accuracy obtain different values ​​of the total thrust impulse by changing the engine operating mode, which is not feasible when using a taxiway operating continuously. TO pulse rocket engine there are requirements for speed, stability of characteristics, output of a minimum value of a single thrust impulse, and low power consumption by control valves. Ideal pulse rocket engine must produce rectangular thrust pulses that coincide in time with electrical commands. In real pulse rocket engine thrust impulses are trapezoidal or bell-shaped; they are wider than command impulses and lag behind them. Wasteful consumption of rocket fuel during multiple start-up and shutdown modes reduces the resulting specific impulse of the rocket launcher. develop low thrust, most of them are rocket-propelled micromotors. used in individual rocket propulsion systems and are the main type of rocket propulsion systems for spacecraft control systems. The speed of operation ensures flight control with low flow rate of the working fluid. When performing maneuvers involving relatively large energy expenditures, pulse rocket engines operate continuously (up to several hours when the location of synchronous satellites changes).

Pulse rocket engines operate both on two-component self-igniting fuel and on single-component fuel. Example pulse rocket engine the R-4D, designed for reactive control systems Apollo spacecraft. Hydrazine is widely used as a single-component fuel. In particular, a typical reactive control system of a connected satellite stabilized by rotation (usually with a frequency of ~ 1 s -1) contains several pairs of hydrazine pulse rocket engines thrust ~ 20 N each. Disadvantages of hydrazine pulse rocket engines are the destruction and loss of quality of the catalyst with a large number of “cold” inclusions. Increase in resource pulse rocket engines is achieved by maintaining the catalyst at an elevated temperature (for example, 600 K) by electrical heating of the remote control. Hydrazine compounds have been created pulse rocket engines with more than 1 million inclusions.

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