7.1.7. LASER HEATING Initial period. The laser (abbreviation of the English Light Amplification by Stimulated Emission of Radiation) was created in the second half of the 20th century. and found some application in electrical technology. The idea of ​​the process of stimulated emission was expressed by A. Einstein in 1916. In the 40s, V.A.

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Description

Aerodynamic heating - heating of bodies moving at high speed in air or another gas. Aerodynamic heating is the result of the fact that air (gas) molecules incident on the body slow down near the body. If the flight is from supersonic speed, deceleration occurs primarily in the shock wave that occurs in front of the body. When air molecules decelerate in the boundary layer, directly at the surface of the body, the energy of their chaotic motion increases, which leads to an increase in the gas temperature in this layer and aerodynamic heating of the body. For example, during the flight of a supersonic aircraft at a speed of 1 km/s, the stagnation temperature is about 700 K, and when a spacecraft enters the Earth's atmosphere with the first cosmic velocity (~7.6 km/s), the stagnation temperature reaches 8300 K. If in the first case the temperature of the aircraft skin can be close to the stagnation temperature, then in the second case the surface of the spacecraft will inevitably begin to collapse due to the inability of the materials to withstand such high temperatures.

The maximum temperature to which the gas can be heated in the vicinity of a moving body is close to the so-called stagnation temperature T 0:

,

where is the temperature of the incoming air;

V - body flight speed;

c p is the specific heat capacity of the gas at constant pressure.

As the velocity of the body increases, the temperature of the air behind the shock wave and in the boundary layer increases.

The degree of aerodynamic heating significantly depends on the shape of the body, which is taken into account by introducing the aerodynamic drag coefficient Cx. There are two types of aerodynamic heating: convective and radiation. Convective heating is the transfer of heat from the boundary layer region to the surface of a moving object by conduction and diffusion. Radiative heating is the transfer of heat by radiation of gas molecules. The ratio between convection and radiation heat fluxes depends on the speed of the object. Up to the values ​​of the first cosmic velocity, convective heating prevails, at the second cosmic velocity (~11200 m/s) the convective and radiative fluxes are approximately equal, and at velocities above 13000 m/s, the radiative heat flux becomes predominant.

Characteristics of aerodynamic heating of gases are studied in installations called shock tubes. A shock wave can be created by an explosion, an electrical discharge, etc.

Timing

Initiation time (log to -1 to 2);

Lifetime (log tc 13 to 15);

Degradation time (log td -1 to 2);

Optimal development time (log tk 1 to 2).

Diagram:

Technical realizations of the effect

Technical implementation of the effect

Aerodynamic heating is associated with the "thermal barrier" problem that arises when creating supersonic aircraft and launch vehicles. An important role is played by aerodynamic heating during the return of spacecraft into the Earth's atmosphere, as well as during the entry into the atmosphere of planets with velocities of the order of the second cosmic and higher. To combat aerodynamic heating, special thermal protection systems are used.

Aerodynamic heating usually plays the role of a negative factor. To combat aerodynamic heating, aircraft are equipped with special thermal protection systems. There are active and passive methods of thermal protection. With active methods, a gaseous or liquid coolant is forced to the protected surface. The gaseous coolant, as it were, blocks the surface from the effects of high-temperature external environment, and the liquid coolant forming on the surface protective film, absorbs heat approaching the surface as a result of heating and evaporation of the film, as well as subsequent heating of vapors. With passive methods of thermal protection, the effect of the heat flux is assumed by a specially designed outer shell or a special coating applied to the main structure. The most widespread is thermal protection with the help of collapsing surfaces, in which the heat flux is spent on the processes of melting, evaporation, sublimation and chemical reactions. The materials of such coatings are fiberglass and other plastics on organic and organosilicon binders. Carbon and carbon compositions are also promising.

Aerodynamic calculation is the most important element of the aerodynamic study of an aircraft or its individual parts (hull, wings, empennage, control devices). The results of such a calculation are used in trajectory calculations, in solving problems related to the strength of moving objects, in determining flight performance LA.

When considering aerodynamic characteristics, one can use the principle of dividing the characteristics into separate components for isolated hulls and load-bearing surfaces (wings and plumage), as well as their combinations. In the latter case, the aerodynamic forces and moments are determined as the sum of the corresponding characteristics (for an isolated hull, wings, and empennage) and interference corrections due to interaction effects.

Aerodynamic forces and moments can be determined using aerodynamic coefficients.

According to the representation of the total aerodynamic force and the total aerodynamic moment in projections on the axes of the velocity and associated coordinate systems, respectively, the following names of aerodynamic coefficients are adopted: - aerodynamic coefficients of drag, lifting lateral force; aerodynamic coefficients of roll, yaw and pitch moments.

The given method for determining the aerodynamic characteristics is approximate. The figure shows a diagram of the rocket, here L is the length of the aircraft, dm is the diameter of the aircraft body, is the length of the nose, l is the wing span with the ventral part (Fig. 1).

engine steering rocket flying

lifting force

The lifting force is determined by the formula

where is the velocity head, is the air density, S is the characteristic area, (for example, the area cross section fuselage), is the lift coefficient.

It is customary to determine the coefficient in the velocity coordinate system 0xyz. Along with the coefficient, the normal force coefficient is also considered and is determined in the coupled coordinate system.

These coefficients are related to each other by the relation

We present the aircraft as a combination of the following main parts: body (fuselage), front (I) and rear (II) bearing surfaces. At small angles of attack and deflection angles of the bearing surfaces, the dependences and are close to linear, i.e., can be represented as

here and are the deflection angles of the front and rear bearing surfaces, respectively; and - values ​​and at; , are the partial derivatives of the coefficients and with respect to the angles and taken at.

The values ​​of and for unmanned aircraft in most cases are close to zero, so they are not considered further. Rear bearing surfaces are taken as controls.

Coefficient definition

find the derivative:

At small angles of attack and at, we can put, then equality (2) takes the form. We represent the normal force of the aircraft as the sum of three terms

each of which is expressed in terms of the corresponding coefficient of the normal force:

Dividing equality (3) term by term by and removing the derivative with respect, we obtain at the point 0

Where; - flow deceleration coefficients;

; ; - relative areas of aircraft parts.

Let us consider in more detail the quantities included in the right-hand side of equality (4).

The first term takes into account the normal force of the fuselage, and at low angles of attack it is equal to the normal force of the isolated fuselage (without taking into account the influence of the bearing surfaces)

AERODYNAMIC HEATING- heating of bodies moving at high speed in air or other gas. A. n. inextricably linked with aerodynamic drag, which test bodies during flight in the atmosphere. The energy expended to overcome resistance is partially transferred to the body in the form of A. n. Physical consideration. It is convenient to carry out the processes that determine the A. N. from the point of view of an observer who is on a moving body. In this case, it can be seen that the gas incident on the body is decelerated near the surface of the body. First, braking occurs in shock wave, which is formed in front of the body if the flight occurs at supersonic speed. Further deceleration of the gas occurs, as in the case of subsonic flight speeds, directly at the very surface of the body, where it is caused by the forces of viscosity, forcing the molecules to "stick" to the surface with the formation boundary layer.

When decelerating the flow of gas, its kinetic. energy decreases, which, in accordance with the law of conservation of energy, leads to an increase in ext. gas energy and its temperature. Max. heat content ( enthalpy) of the gas during its deceleration near the body surface is close to the stagnation enthalpy: , where is the enthalpy of the oncoming flow, and is the flight speed. If the flight speed is not too high (1000 m / s), then beats. heat capacity at DC pressure with p can be considered constant and the corresponding gas deceleration rate can be determined from the expression

Where T e- equilibrium temperature-pa (limiting temperature, to which the surface of the body could heat up if there was no energy removal), - coefficient. convective heat transfer, the index marks the parameters on the surface. T e is close to the deceleration temp and can be determined from the expression

Where r-coefficient temperature recovery (for laminar, for turbulent-), T1 And M 1 - temp-pa and mach number to ext. border of the boundary layer, -ratio beats. heat capacities of gas at DC. pressure and volume Pr is the Prandtl number.

The value depends on the speed and altitude of the flight, the shape and size of the body, as well as on some other factors. Similarity theory allows us to represent the laws of heat transfer in the form of relationships between the main dimensionless criteria - Nusselt number , Reynolds number , Prandtl number and temperature factor , taking into account the variability of thermophys. gas properties across the boundary layer. Here and - and the gas velocity, and - coefficient. viscosity and thermal conductivity, L- characteristic body size. Naib. influence on convective A. n. renders the Reynolds number. In the simplest case of a longitudinal flow around a flat plate, the law of convective heat transfer for a laminar boundary layer has the form

where and are calculated at temperature a for a turbulent boundary layer

On the nasal part of the body with blunting spherical. laminar heat transfer is described by the relation:

where r e and m e are calculated at a temperature T e. These formulas can also be generalized to the case of calculating heat transfer in a non-separated flow around bodies of a more complex shape with an arbitrary pressure distribution. In a turbulent flow in the boundary layer, an intensification of the convective A. N. occurs, due to the fact that, in addition to molecular thermal conductivity, beings. turbulent pulsations begin to play a role in the transfer of the energy of the heated gas to the surface of the body.

With the theoretical calculation A. n. For an apparatus flying in dense layers of the atmosphere, the flow near the body can be divided into two regions - inviscid and viscous (boundary layer). From the calculation of the flow of inviscid gas in the external. area is determined by the distribution of pressure over the surface of the body. The flow in a viscous region with a known distribution of pressure along the body can be found by numerically integrating the equations of the boundary layer or, for calculating the A. n. can be used diff. approximate methods.

A. n. plays creatures. role and supersonic flow gas in the channels, primarily in the nozzles of rocket engines. In the boundary layer on the walls of the nozzle, the gas temperature can be close to the temperature in the combustion chamber rocket engine(up to 4000 K). In this case, the same mechanisms of energy transfer to the wall operate as in the boundary layer on a flying body, as a result of which an AE arises. nozzle walls of rocket engines.

To obtain data on A. n., especially for bodies of complex shape, including bodies streamlined with the formation of separation regions, an experiment is carried out. studies on small-scale, geometrically similar models in wind tunnels with reproduction of defining dimensionless parameters (numbers M, Re and temperature factor).

With an increase in the flight speed, the temperature of the gas behind the shock wave and in the boundary layer increases, as a result of which dissociation of the incoming gas molecules also occurs. The resulting atoms, ions and electrons diffuse into a colder region - to the surface of the body. There is a reverse chem. reaction - recombination, going with the release of heat. This gives an addition. contribution to convective A. n. In the case of dissociation and ionization, it is convenient to switch from temperature to enthalpies:

Where - equilibrium enthalpy, and - enthalpy and velocity of the gas at ext. the boundary of the boundary layer, and is the enthalpy of the incoming gas at the surface temperature. In this case, the same critical values ​​can be used to determine. ratio, as for relatively not high speeds flight.

When flying at high altitudes, convective heating can be affected by the non-equilibrium of the physical and chemical. transformations. This phenomenon becomes significant when the characteristic times of dissociation, ionization, and other chem. reactions become equal (in order of magnitude) to the residence time of gas particles in a region with an increased temperature near the body. Influence of physico-chemical. disequilibrium on A. n. manifests itself in the fact that the dissociation and ionization products formed behind the shock wave and in the high-temperature part of the boundary layer do not have time to recombine in the near-wall, relatively cold part of the boundary layer; decreases. In this case, catalytic plays an important role. surface material properties. By using materials or coatings with low catalytic activity with respect to recombination reactions (for example, silicon dioxide), it is possible to significantly reduce the amount of convective A. n.

If a gaseous coolant is supplied ("blowing") into the boundary layer through the permeable surface of the body, then the intensity of convective A. n. decreases. This is happening ch. arr. will add as a result. heat consumption for heating the gases blown into the boundary layer. The effect of reducing the convective heat flux during the injection of foreign gases is the stronger, the lower their molecular weight, since the sp. heat capacity of injected gas. At laminar flow flow in the boundary layer, the blowing effect is more pronounced than in the case of turbulent flow. With moderate beats. blown gas flow rate, the reduction in convective heat flux can be determined by the formula

where is the convective heat flux to the equivalent impermeable surface, G is the sp. mass flow rate of injected gas through the surface, and - coefficient. blowing, which depends on the flow regime in the boundary layer, as well as the properties of the incoming and blown gases. Radiative heating occurs due to the transfer of radiant energy from areas with an increased temperature to the surface of the body. In this case, it plays the greatest role in the UV and visible regions of the spectrum. For the theoretical calculation of radiation heating, it is necessary to solve a system of integro-differential equations of radiation. gas, taking into account own. emission of gas, absorption of radiation by the medium and transfer of radiant energy in all directions in the high-temperature flow region surrounding the body. Integral over the spectrum of radiation. flow q P0 to the body surface can be calculated using Stefan-Boltzmann law of radiation:

where T 2 - gas temp-pa between the shock wave and the body, \u003d 5.67 * 10 -8 W / (m 2 * K 4) - Stefan's constant, - eff. the degree of blackness of the radiating volume of gas, which in the first approximation can be considered as a flat isothermal. layer. The value of e is determined by a combination of elementary processes that cause the emission of gases at high temperatures. It depends on the speed and altitude of the flight, as well as on the distance between the shock wave and the body.

If it relates. the amount of radiation. A. n. great, then creatures. the role begins to play radiats. cooling of the gas behind the shock wave, associated with the removal of energy from the radiating volume into environment and lowering its temperature. In this case, when calculating the radiation. A. n. a correction must be introduced, the value of which is determined by the highlighting parameter:

where is the flight speed, is the density of the atmosphere. When flying in the Earth's atmosphere at speeds below the first cosmic radiation. A. n. small compared to convective. At the second cosmic speeds they are compared in order of magnitude, and at flight speeds of 13-15 km / s, corresponding to the return to Earth after flying to other planets, main. contribution is made by radiative A. n.

A special case of A. n. is the heating of bodies moving upwards. layers of the atmosphere, where the flow regime is free-molecular, i.e., gas molecules are commensurate or even exceed the size of the body. In this case, the formation of a shock wave does not occur even at high flight velocities (of the order of the first cosmic one). a simple formula can be used

where is the angle between the normal to the surface of the body and the velocity vector of the oncoming flow, A- coefficient accommodation, which depends on the properties of the incoming gas and surface material and, as a rule, is close to unity.

With A. n. related to the problem of "thermal barrier", which arises in the creation of supersonic aircraft and launch vehicles. An important role of A. n. plays at the return of space. devices into the Earth's atmosphere, as well as when entering the atmosphere of planets with velocities of the order of the second cosmic and higher. To combat A. n. apply special. systems thermal protection.

Lit.: Radiation properties of gases at high temperatures, M., 1971; Fundamentals of the theory of spacecraft flight, M., 1972; Fundamentals of heat transfer in aviation and rocket and space technology, M., 1975. I. A. Anfimov.

If the heating of projectiles and rockets at low flight speeds is small, then at high speeds it becomes a serious obstacle to the development of aircraft. These devices are heated by the heat radiated by the Sun and the heat released during the operation of the engines and control equipment. In addition, they heat up when moving in the air.

Heating from movement in the air plays the most significant role, especially when returning ballistic missiles to the atmosphere. During the movement of an aircraft in the air, heat arises due to the friction of air on the surface of the rocket and, mainly, the compression of air in front of the flying body.

As you know, a Soviet rocket launched into the Pacific Ocean reached a speed of over 7200 m/sec. If, during its re-entry into the atmosphere, this speed had been maintained and complete deceleration of the air in front of the rocket had been ensured, then, as shown by an elementary calculation based on the energy conservation equation for compressible gases, the temperature of the air in front of the rocket could have increased by almost 26,000 °.

However, let's ask ourselves some questions. First, does the air in front of the rocket actually heat up to the calculated temperature as a result of compression? The answer will be negative. Theoretically, complete deceleration of the air in front of a streamlined body, such as a projectile or rocket, should occur only at one point, namely: in front of the tip of the bow. On the rest of the surface, only partial air deceleration occurs. Therefore, the overall heating of the air near the aircraft is much less. In addition, as the air in front of the rocket heats up and becomes denser, its thermodynamic properties change, in particular, the specific heat capacity increases, and air heating turns out to be less. Finally, the molecules of air, heated to an absolute temperature of 2,500 - 3,000°, begin to "split" into atoms. Atoms turn into ions, that is, they lose electrons. These processes (dissociation and ionization) also take part of the heat, reducing the air temperature.

Secondly, is all the heat that the air possesses transferred to the projectile or rocket as it travels? It turns out not. The heated air gives off a lot of heat to the surrounding air masses through heat transfer and thermal radiation.

Thirdly, if the air in front of the flying body is heated to a certain temperature, does this mean that the rocket is heated to the same degree? Also no. The skin will always have a temperature lower than the air around it.

The aircraft simultaneously with the receipt of heat will give off heat to the surrounding air and cool due to radiation. In general, the apparatus will heat up to a temperature at which some complex thermal balance is established.

In order to estimate the probable heating of a projectile or rocket in flight, one must first of all know at what speed and for how long it will fly through air layers of varying density and temperature. When penetrating the atmosphere upward, the stay of a ballistic missile in a relatively dense atmosphere is very short and is measured in seconds. It develops a high speed, in fact, already at the exit from the atmosphere, i.e., where the air is very rarefied.

All these circumstances, taken together, lead to the fact that the heating intensity of the rocket during the upward flight, although significant, is quite acceptable without taking special constructive measures.

Significantly greater difficulties await the rocket (its warhead) during the return return to the atmosphere. In addition to large aerodynamic loads, the so-called "thermal shock" can occur here, associated with a rapid increase in the temperature of the rocket.

Let us briefly list some of the methods of combating the heating of aircraft, cited in foreign literature * . Firstly, reducing the speed of their forced movement in the atmosphere (for example, during the return of a rocket) by using air brakes, parachutes, brake engines, etc. Secondly, the use of refractory and heat-resistant materials for the construction of the skin. Thirdly, the use of materials or coatings for the shell, which are characterized by high emissivity, i.e., the ability to remove more heat into space. Fourth, careful polishing of the surface, which improves its reflectivity. Fifthly, the thermal insulation of the main components of the structures, i.e., a decrease in the heating rate by applying a layer of a substance with low thermal conductivity to the surface or by creating a layered-porous heat-insulating set between the outer and inner skins.

* ("Airplane" No. 2478.)

And yet, at very high speeds, temperatures develop at which neither metal nor any other materials are suitable without taking measures to forcefully cool the skin. Therefore, the sixth way is to create forced cooling, which can be created in various ways, depending on the purpose of the aircraft.

The head parts of rockets are sometimes covered with so-called burn coatings. The decrease in temperature in this case is achieved by creating such protective skin layers that are designed to melt and burn. Thus, they absorb heat, preventing it from reaching the main structural elements. When the skin layer melts or evaporates, a protective layer is simultaneously formed, which reduces the transfer of heat to the rest of the structure.

The efficiency of aircraft modern level their development is directly related to the solution of the thermal problem. The pinnacle of achievements in this field were flights in a circular orbit with the return to Earth of Soviet cosmonauts Yu. A. Gagarin and G. S. Titov.

Basic data of foreign guided missiles and missiles*

* (The given data is borrowed from foreign press (mainly from "Flight" No. 2602 and 2643). Blank columns mean no published information.)