Assuming that the gas flows axially symmetrically and stably in the bladeless gap and the radial velocity of the gas is zero, the three-dimensional flow field in the turbomachinery can be artificially decomposed into the radial variation of flow parameters in the bladeless gap and the two-dimensional flow in the cylindrical surface. However, this assumption has obvious approximate properties, especially for turbomachinery with small hub ratio and large meridian expansion angle. In order to make the flow model closer to reality, it is necessary to develop the three-dimensional flow theory of turbomachinery.
1905, H. Lorenz put forward the through-flow theory, that is, the infinite multi-leaf theory. This theory assumes that the number of blades tends to infinity and the thickness of blades tends to infinity. In this way, the shape of each relative flow surface between two adjacent blades is consistent with the central plane of the blades, and the circumferential variation is close to zero. By introducing an imaginary mass force field, the effect of actual blades is considered. In this way, the solution of the airflow on the limit flow surface which coincides with the central plane of the blade can be obtained. In the early 1950s, Wu Zhonghua, a Chinese scientist, perfected the through-flow theory and put forward the general theory of three-dimensional flow in turbomachinery. The concepts of S 1 and S 2 are introduced in this theory, and the basic equations of these two kinds of stream surfaces are derived respectively. Through the proper combination and alternate application of these two flow surfaces, an actual three-dimensional flow problem can be decomposed into two related two-dimensional flow problems along S 1 and S 2 respectively (see Figure [Schematic Diagram of General Theory of Three-dimensional Flow]). In fact, it is usually assumed that the S 1 stream surface is any rotating surface, but only one central stream surface named S 2m is taken in the S 2 stream surface family. In this way, a preliminary approximation of three-dimensional flow can be obtained.
At that time, the general theory of three-dimensional flow in turbomachinery was put forward for pure subsonic flow and pure supersonic flow In fact, transonic flow usually exists in high-speed turbomachinery, that is, there are both subsonic and supersonic regions in the flow field, and there are shock surfaces and sonic surfaces with unknown shapes, numbers and positions. Therefore, transonic flow is one of the research directions of turbomachinery gas dynamics. Given the geometry of the blade channel, directly solving the three-dimensional flow problem is the second research direction. Semi-empirical method is usually used to correct the influence of fluid viscosity in engineering. In addition, the interaction between the boundary layer on the blade profile, the boundary layer on the ring wall of the casing or hub and the mainstream of the boundary layer will produce the so-called "secondary flow" phenomenon, which has a great influence on the performance of turbomachinery. Therefore, the study of viscous flow in turbomachinery is the third direction. Nozzle-a component in a jet engine, which converts high-pressure gas (or air) into kinetic energy, makes the airflow expand and accelerate in it, and ejects at high speed, resulting in reverse thrust, also known as exhaust nozzle, thrust nozzle or tail nozzle. There are many types of nozzles, such as fixed or adjustable convergent nozzle, convergent-divergent nozzle, ejector nozzle and plug nozzle, which are selected according to the performance of the aircraft and the working characteristics of the engine. Most high-speed fighters use adjustable convergent nozzles and adjustable convergent-divergent nozzles or ejector nozzles; Fixed convergent-divergent nozzles are usually used in rocket engines; V/STOL aircraft use reverse nozzles.
The ratio of the total pressure at the nozzle inlet to the static pressure at the nozzle outlet is called nozzle pressure drop ratio, expansion ratio or pressure ratio. The ratio of the outlet area of the nozzle to the critical cross-sectional area (the area at the minimum cross-section) is called the nozzle expansion area ratio, commonly known as the area ratio. When the static pressure at the nozzle outlet is just equal to the external atmospheric pressure, it is called a fully expanded nozzle, and its performance is the best. When the static pressure at the nozzle outlet is greater than the external atmospheric pressure, it is called an incomplete expansion nozzle, and the pressure energy of airflow is not completely converted into kinetic energy. When the static pressure at the nozzle outlet is lower than the external atmospheric pressure, it is called over-expansion nozzle, and then negative pressure thrust will appear.
Nozzle with decreasing cross-sectional area along the flow direction. The convergence half angle is usually 7 ~ 35, which will cause great thrust loss due to incomplete expansion when flying at large Mach number. For example, when Mach number is 1.5, the loss is about14%; When Mach number is 3, the loss exceeds 50%. This nozzle is simple in structure and light in weight, and is used for engines of subsonic or hypersonic aircraft.
Convergent-divergent nozzle A nozzle whose cross-sectional area converges first and then diffuses along the flow direction. It was invented by C.G. Laval of Sweden, so it is also called Laval nozzle. When this nozzle is used in supersonic fighter, the critical area and outlet area need to be adjusted with the flight state; When used in rocket engines, the area ratio can reach 7 ~ 400. The bell nozzle is most commonly used in modern rocket engines, and the outlet half angle is reduced to 2 ~ 8, and the length is short. There are also several short annular nozzles, such as plug nozzle, expansion deflection nozzle, reflux nozzle and advection nozzle. Its * * * feature is that the air flow has a free expansion boundary, which can be automatically adjusted with the external pressure, and it is often in a state of complete expansion, but it is not widely used.
Adjustable nozzle is mainly used for afterburner turbojet engine or afterburner turbofan engine of military aircraft flying at high speed. The nozzle area ratio is easy to adjust, can change with flight conditions, and is often in a fully expanded state. Structural types include balance bar type, folding type, folding petal type, sleeve cone type and so on.
The ejector nozzle consists of an adjustable convergent main nozzle and a fixed or adjustable ejector sleeve. The ejector action of the main stream drives the secondary stream to flow between the main stream gas column and the ejector sleeve, and the secondary stream acts as an air cushion for the main stream to limit its expansion. Adjusting the secondary flow can control the flow area of the main flow, so that it can reach or approach complete expansion. The injector nozzle is light in weight and simple in structure. It can maintain good performance in a wide flight range and has been widely used in many high-performance aircraft.
The exit section of the two-dimensional nozzle is not circular, which is easy to realize the integration of the aircraft afterbody and nozzle, reduce the external resistance and exposed surface of the aircraft, and improve the performance and concealment of the aircraft; It can also realize thrust commutation and reverse, and increase maneuverability.
Nozzle material The choice of nozzle material is closely related to the nozzle structure and cooling mode. Nickel-based superalloy is often used in gas turbine engine nozzle, and stainless steel is used in liquid rocket engine regenerative cooling nozzle. The extension part of the radiation cooling nozzle is made of niobium alloy and other heat-resistant materials; Composite materials are commonly used in solid rocket motors. The parts in contact with airflow are made of high-temperature or corrosion-resistant materials, and the back wall is made of heat insulation materials. The high-temperature resistant layer on the inner side of the throat which is most seriously heated in the nozzle is called throat liner, and high-melting metal such as tungsten and its alloy or sweating material, cermet, graphite, carbon-carbon composite material, etc. can be used. The inlet part is mainly made of graphite phenolic or carbon phenolic materials. High-silica phenolic or carbophenolic materials are usually used in the outlet part.