Diadia_Sidor>yeah, to hell with it (politics). I'm better for you that I'll tell you a thread from the life of TNA

warban>come on. it's interesting and worthy

Ask, I love you, he also has a place in the yard

About pumps - are screw centrifugal pumps used by everyone or is it a local flavor?

All of them, at first, were centrifugal, then they began to put an auger in front of the pump, and then, with the development of technology, they merged into a single whole. but still the most popular is the auger separately, the wheel separately

And I designed a screw centrifugal pump for ... hot water at the factory. On one wheel

It also depends a lot on what to download.
it’s one thing to gain the necessary pressure on water and another on hydrogen, not only you can’t get by with a centrifugal screw, you need a hefty wheel, but not one step

By the way, in TNA centrifugal pumps without competition? I have not met others

And you won't meet

What about turbines? Are there more axial ones? And how many steps?

Turbines are used different, axial single and two-stage, on smelly motors - radial in open circuits make partial turbines

But this did not occur - what kind of partial?

They must have met. in a partial turbine, gas is supplied not over the entire area of ​​\u200b\u200bthe entrance to the turbine grate, but in some sector

What for? so the wheel is not fully used

With an open circuit, the flow rate on the turbine is not very large, and if it is smeared over the entire inlet, blades are too short, and the efficiency of the turbine drops sharply. and partiality losses are quite tolerable

I recalled that I had a drawing of a TNA of some kind of RD-350 (or I’m mistaken), there is such a story, a single-stage turbine and with a gas supply through the window

This is a common thing v-2 rd-107/108

Does the V-2 run on hydrogen peroxide?

Yeah, there the nozzle apparatus is funny, several inclined nozzles sawn by hand

Turbine nozzle?

What about bearings? go at 20,000 rpm even with fau

The next day:

warban>greetings!

Diadia_Sidor>and you!

Why are we suddenly on you? Or are you talking to my computer flock?

And shtosh I can’t turn to a good person?

You can, you can;) let's resume the interview on tna, but first and more systematically

With pleasure, only I can have more pauses, the work is low, and today I plan to dump it earlier

And I may have pauses - one computer dies
In light of the non-existent chapter in the development of RD TNA, there should be a special place - if there weren’t TNA (a purely hypotonic case;) and LRE would have lost the competition with solid propellant rocket engines. It was TNA that made it possible to create a liquid propellant rocket engine, much easier than solid propellant rocket engines similar in thrust / impulse.

Not similar, but superior

Yes, superior By the way, an interesting digression into the past:
why Zander stubbornly developed a device for powering solid propellant rocket motors with charges. After all, he's not a fool! Only in this way it is possible to bring solid propellant rocket engines closer to pumping rocket engines. After all fundamental The advantage of an LRE with a TNA over a solid propellant rocket engine is that the LRE chamber is hundreds and thousands of times smaller than the tanks (or the solid propellant rocket engine chamber is the same tank). Make the solid propellant rocket motor power supply with checkers, and this will allow you to reduce the chamber (it’s a completely different matter that this motor will combine not so much the advantages of liquid propellant rocket engines and solid propellant rocket engines as disadvantages)

Shuki jokes, but at NIITP they worked on a solid propellant rocket engine with a continuous supply system, don’t ask for details - I don’t know, and I forgot what I knew

And I've heard that - but I won't go into it either. Purely for ethical reasons - the development is not over. I expressed all this to the developer, and he seems to understand. By the way, he does not feed a saber into the camera, but pushes the camera onto the saber - the camera is smaller!

Hey, let's get to the pumps

How did it all start? Let's not talk about ideas - there were enough of them. Who would you put as the first TNA-schik from real designers?

And then there’s nothing to think about the first normal tna - V-2, we had interesting developments at Dushkin, they even tested it, but how bench models still work, but they NEVER FLYED, but FAU FLYED

And for the main rivals, everything went from the V-2 - both for the Ams and for you?

Note: Just don’t think that I’m sliding down to the primitive level of a journalist - I have a good idea of ​​​​both the Soviet and American contributions to rocket technology - it is so great that recognition of the influence of the V-2 cannot reduce it in any way - it will simply be unfair not to mention.

That's exactly how it is. Moreover, fundamentally, nothing has changed in TNA since then. Of course, technologies have changed, other components have appeared, but everything was already in TNA A.4

Is it possible to say this - others (Ams, Russians, French) did not have TNA, because they did not need it? After all, they made medium-sized rockets, that is, they worked in an area where the displacement supply of components is not much or does not lose at all to the pumping one.

It is debatable here, the same Dushkin made TNA for an aircraft engine (Bi-1, I-301) a small thing, but with TNA completely different characteristics. on the other hand, they didn’t make real rockets, and they did fine without pumps

Then on the contrary - A-4 could not appear without TNA - try to make strong tanks of such sizes.
note: Although the tanks were plug-in, not load-bearing - a fact, and perhaps a mistake. Even then, they could have made the tanks carriers, in any case, one of the first improvements to the A-4 in the Union is exactly that.

It's connected
there cannot be a large rocket without a TNA, just as a TNA is not needed by itself, it is needed for a large rocket

And Germany's successes in TNA are not related to its work on turbojet engines - after all, turbojet aircraft and turbopump rockets appeared at the same time?

Naturally, this speaks of the level of technology, and the technology of blade machines, in particular, in Germany at that time. The largest gas speaker of that time, Theodor von Karman, worked there, and they are still learning from it

For completeness, let's list the German companies that made LRE with TNA. As far as I know, they were made then (and today) by the same company.

The TNA for the FAU was serially produced at the FAU serial factories, you need to look at the exact cooperation at home, but BMW and Jumo specialists were also involved in the development, for sure

Of course, manufacturers / developers and the first turbojet engines.
But back to the TNA of the A-4 engine. He had everything - a turbine and two pumps - alcohol and oxygen - were sitting on one shaft. The partial turbine (as they found out yesterday) worked in an open circuit, on hydrogen peroxide, which decomposed in the gas generator (by the way, there is a noticeable connection with the torpedoes of the German fleet - they tried peroxide there too).
As you said, everything was fundamentally in this TNA.
And how did the TNA of the first "space" engines - 107/108 differ from the A-4 TNA?

Structurally, few other materials were used, more technologically advanced. the layout changed to the cantilever fastening of the turbine, accordingly, seals appeared between the pumps of the components, the wheels began to be cast and not milled, a heat exchanger was installed behind the turbine for nitrogen gasification, and the tanks were pressurized. augers, in my opinion, have not yet appeared

Cantilever mounting of the turbine: in the sense were there bearings between the oxygen and kerosene pumps and between the kerosene pump and the turbine?
Gasification of what nitrogen?
And the screw-centrifugal stage was at the oxygen pump.

About the oxygen pump - however, I don’t have a picture before my eyes, and the FAU and 107/108 have only two bearings
at the fau on the shaft: bearing, pump, turbine, pump, bearing
107/108: bearing, pump, pump, bearing, turbine
nitrogen is gasified, which is used to pressurize tanks

Clear about nitrogen.
Note (of course, not to you) I remind you that supercharging achieves a lot of things:
-tanks become stiffer - they will resist aerodynamic loads well;
-beast-pump, don't be pressurized, crush the tank like an empty cigarette pack
- the most important thing, in my opinion, is that it provides support to the first stage of the pump and thereby contributes to the cavitation-free operation of the aggregate
Any other thoughts on the usefulness of boost?

I laid out everything
an instructive thing, the first atlases had very thin steel tanks, so they were inflated right at the factory, otherwise the rocket folded under its own weight, not even refueled

So it is the same with Russian ones - when they transported Energy tanks to Myasishchevo, they had to pressurize them.
Now let's talk about the action - we'll talk about the cavitation breakdown of the TNA

Eco got you
Let's go
but one more story. My wife did not immediately enter the institute and worked for a year as a typist at a rocket engine department. Somehow, someone Kozlov (the largest specialist in tank processes) slipped her an article to reprint. And the woman, surprised at the rare illiteracy of the professor, sent the pressurization to the pressurization of the tanks in the entire text.

))
Did cavitation breakdown take place in real tests?
Note: If the centrifugal pump is driven by an (a)synchronous motor, cavitation stall is not so bad. Another thing is if the engine has the characteristics of a turbine - when the load drops, the speed increases proportionally. And cavitation breakdown can lead to the separation of the unit.

This is exactly what happens, but there is another unpleasant phenomenon that leads to a complete kirdyk. when a cavitation cavity appears at the inlet to the pump, and it is carried to the exit, under the influence of increased pressure it collapses; several such collapses and the wall - 3.1415926zdets!

And in conclusion, it must be said that the cavitation breakdown is so fleeting that the automatic control of the turbine will not track the increase in speed. The best thing is to cut off the TNA and the engine, obviously. And since marching rocket engines are not yet able to launch in flight, even stinkers, at best, non-induction. At worst, the destruction of the TNA and a fire on board (just under the tanks with components) By the way, any accident in the TNA leads to this

If a cavitation breakdown develops - kirdyk, definitely

And on the stands, what do the rocket engines feed? If the regular TNA has not yet been worked out?

Displacement, naturally, usually with nitrogen

I thought so. With the first Soviet space TNA sorted out. What about competitors? I heard that there were emergency launches of ams at the beginning of the space race, mainly due to the lack of control of the remote control. After a few setbacks, Brown was promoted to lead, and before that it was a mess.

Yes, it's a funny story too. as for combat missiles (Redstone, Thor, Atlas), they did not fantasize and honestly used German developments (Brown commanded the construction of Redstone from the very beginning), and they wanted to go into space on a small Vanguard, which was built just for this. That's where they screwed up. TNA made with a gear train. And in space, anyway, Redstone was the first to fly

Exactly! I heard about this, there were other jokes like:
On the Amov rocket, the GTZA failed (the main turbo-gear unit is an indispensable accessory for turbine ships) - probably the coal for the boilers was of poor quality. With such information, Kulagin diversified our lectures on ballistics;)
Do you know why the Ams went the way of TZA? Did you make teapots with lamers? Or did the teeth break during the development of the direct drive?

There is logic in this. the oxidizer and fuel have different densities, respectively, the maximum efficiency of each wheel will be achieved at different speeds, so they suffered
and by the way, the RL-10 engine, the world's first hydrogen generator, still flies with its turbo-gear unit

E, hydrogen and oxygen have significantly different densities - 0.011 and 1.6. So there is logic - the pump for hydrogen is direct-drive, and for oxygen - the speed has been reduced. By the way, are there separate TNAs on the SSME from there?
And for the Avant-garde kerosene stove, the density is two times different, taking into account the ratio - it will be the most, only the pump wheels are of different diameters. It's still not clear. It is necessary to dig up links to the Vanguard in the internet and figure it out, although maybe. and not worth it. The avant-garde did not give offspring.

The diameter, of course, is different, but with an increase in diameter, the flow area at the outlet (blade height) decreases and the efficiency drops sharply, in general, instead of looking for a compromise, the Americans decided to put toothy wheels, but in vain
about CCME, besides this, there is still a reason for their separation, if a leak occurs along the shaft and the components meet somewhere, you can safely turn off the light,
and we solved this problem and built, constructively, a simpler motor

That's right, design is always a compromise.
> if there is a leak along the shaft and the components meet somewhere, you can safely turn off the light
Of course, a typical intracameral process will take place;)
And did those dictate the layout of the TNA A-4 - is there a turbine - in the middle?

About A.4 - no, the components are not self-igniting, it’s just that this is the simplest (reliable) layout, but if oxygen with hydrogen or UDMH with AT converge on the shaft, then the fun will begin

Summary so far:
The first period - fau and fau-like,
the second - the initial space one - improving the layout, using the best materials, developing the theory of TNA and bringing it to the engineering level (so far all open circuits)
and the third is a high-boiling poison or is it still hydrogen?
And where did the revolution on the way from V to SSME / RD-0120 take place?

I'm leaning towards gradual evolution
the fundamental difference between smelly closed TNA from the previous ones is a high-flow turbine
multi-stage pumps appeared on hydrogen HPs, respectively, the problem of axial force compensation
I would not divide them by generation

Did it happen that TNA was changed on serial engines?

No - the motor and THA are a single whole

Clearly, if you remake TNA, then the camera is a couple of trifles?
And slightly off-topic And what explains the amazing longevity of the RD-107/108? After all, the first space, bourgeois peers have long been in museums only!
I run away to eat, otherwise the unfortunate journalist is hungry to eat! You write

About the camera - I would not say that this is a couple of trifles, because everything works at the limit of the possible, the engine is a single whole
and 107/108 fulfills its task, the technology has been worked out to a flashlight, and God bless him

With the camera, the question was somewhat provocative;),
but this is a separate topic of chatter
Let's touch on sex itself - closed circuits
The fundamental difference is that the fuel for the turbine is then burned in the chamber.
Learn about sweet and sour gases.
But how has THA changed compared to open circuit THA?

Tna has changed for the better.
gas consumption has increased
it became possible to tie with partial turbines (to increase efficiency),
make normally high blades (increase efficiency)
the turbine has become, unambiguously cantilever,
and diverting gas from the radial turbine into the chamber is just a pleasure

However, something has become more difficult - the pressure in the open turbines is less than in the chamber, and hence in the pumps - that is, it is necessary to deal with the leakage of the liquid component
To turbine, seal is COLD. Does TNA have a closed circuit on the contrary?

The pressure is not much more. mainly on the difference in the turbine, and the difference, on such turbines, is less than in an open circuit, there is enough flow. leaks cause a problem, not because the pressure is higher (it is slightly higher), but because the components are still the same

Yes, I also came up with this:
The pressure in the centrifugal pump at the periphery is high. And in the center of the impeller, it is smaller - at my plant there were cases of air leakage or, what is worse, water - in conc. nitrogen.
And Amov's SSMEs also have turbopumps. What kind of animal is this?

I’m talking about funny boosters (pre-pumps that help fight cavitation in the main pumps) are similar in design. little or little flow is taken from the high pressure line and put into a hydraulic turbine, which turns a low-pressure auger-centrifugal wheel, and the spent low flow of high pressure simply enters the main stream
on many engines, the same ejectors were installed and are installed at the inlet

Yes, but what about Soviet practice? I practically don’t know a single engine with a turbopump - everything is a screw-centrifugal first stage, and then centrifuges

120th motor
170/180th with normal boosters
and a bunch of stinky engines with ejector boosters

And what, stinkers have better ejectors? Unclear

Simply put, no moving parts

And the efficiency is lower

For the booster, this is not important, his task is to organize any thread of pressure at the entrance to the main wheel

Yes exactly. And the efficiency screws I love are no different. But nothing works even in the mode of developed cavitation.
Details about THA SSME:
The liquid oxidizer pump has a capacity of 400 kg/s and the hydrogen pump has an exceptionally high power-to-weight ratio of 16.2 kW/kg.
Power TNA - 56 MW, all this together with the pump has a mass of only 345 kg!
I'm crazy about these numbers.
Unfortunately, I have no data on THA 0120 and 170 - there should be even worse!

Exactly
in terms of energy intensity, the 170th has no equal

Because they divided TNA?

At 170 - there are boosters in the same place - count a couple more TNA

I understood
it makes sense to separate the booster, since this is a completely different unit, low pressures, low revs, simple design

And so am I. And if you optimize the pumps, then it’s more convenient not to think about all sorts of attachments like an auger - there is a backwater and that’s it.

And it's easier to work

I'll write down the negotiations in a file.
While I'm cooking, tell me something instructive as an ending and I'll send it with Zeus.

I saw how they burn at the TNA stand
a fireball just flashes and after a couple of seconds a melted shaft remains with lumps of everything else on it

Well... a chill ran down my back.
although instructive - it means everything is openwork and light, the shaft is only massive
Did they burn it on purpose or did it work abnormally?

On purpose, looking for the cause of a series of accidents

More details? And then it is not always possible to spread about participation in emergency commissions.

Yes, it was all torn on the stands, so it's okay

No, I'm talking about a series of accidents

Here on the stands the series was

Now it's clear

Okay, I'll send it, but in a separate topic - TNA-talk.

With an increase in the volumetric performance of the pump, an increase in the pump power is observed, Fig. 74, a:

The efficiency of the pump is determined by the formula:

With an increase in volumetric productivity, volumetric efficiency g| about also increases, because the share of leaks in relation to the flow rate of the liquid pumped by the pump decreases, Fig. 74, b.

Hydraulic efficiency decreases with the growth of volumetric productivity, tk. the fluid velocity increases, which means friction and impact losses, Fig. 74, b.

With an increase in the volumetric performance of the pump, the proportion of mechanical losses, compared with an increase in power, decreases, therefore, increases, Fig. 74, b.

Dependence of power and efficiency pump from its volumetric performance.

Turbine THA

One of the main elements of TNA is a gas turbine. In the turbine, the potential energy of the combustion products from the gas generator or coolant vapor is converted into the mechanical work of the turbine. The turbine is designed to drive the HP pumps. The turbine consists of a nozzle apparatus 1, an impeller 2 with two rows of rotor blades 3 and 4, a guide vane 5 and a turbine housing 6 with an outlet pipe 7, Fig.75.

The first stage of the turbine is a set of nozzle apparatus 1 and impeller blades 3, the second is formed by fixed blades of the guide vane 5 and the second row of rotor blades 4.

The conversion of the enthalpy of the gas flow into the mechanical energy of the shaft rotation is carried out in two stages: the enthalpy of the gas flow - into the kinetic energy of the jet (in the nozzle apparatus); kinetic energy of the jet - into the mechanical energy of the rotation of the shaft (on the impeller).

THA turbine design

Shafts of turbopump units (TPU) operate at high loads and high speeds. To lighten the weight, they are made hollow. The greatest alternating stresses in the metal of the shaft occur on its outer surface. In this case, any kind of sharp transitions, traces from the cutting tool and other surface defects are stress concentrators. In these places, cracks may form during operation, which will lead to breakage of the shaft. Therefore, special attention is paid to the cleanliness of the surface finish of the shaft with the introduction in some cases of hardening operations. Not only places for bearings, seals, landings, but also all other parts of the shaft that are not mated with other parts are subjected to finishing.

Large number of revolutions (10000-20000 rpm and more) force the designer to assign very tight tolerances for the alignment of the necks and seats, the accuracy of the location of the axial hole, the difference in wall thickness and other dimensions. The slightest geometric errors lead to an uneven distribution of the rotating masses of metal, which causes vibrations and shaking of the TPU.

Fig. 76 shows the two most characteristic types of shafts: with a flange (a) and without a flange (b).

The most critical shafts are made of high quality alloy steel with a tensile strength after appropriate heat treatment of 1000-1200 Mn/m 2 (100-120 kg/mm 2 ). Steels 2X13, 18HNVA, 40HNMA, 12HNZA and some others are used.

For less critical shafts, steels of the 38XA type or steel 45 are used.

Turbine disks of TNA operate at high speeds, as a result of which high stresses arise in the metal from the action of centrifugal forces. In addition, thermal stresses arise from the uneven heating of the disk metal.

Typical shaft types

Turbine disks are made from high-alloy steels and alloys with high strength and heat resistance: steels EI415, EI481, EI395, Kh18N9T, alloys EI437B, EI617 (KhN70VMTYu) and others.

The shape of the disks is determined from the condition of equal strength, i.e., approximately equal loading of the metal in all sections of the disk.

Figure 77 shows several typical designs of turbine disks. The disk consists of a hub for connection with the shaft, a rim for fastening the blades and a middle part connecting the hub with the rim. The load from centrifugal forces increases as it approaches the hub, which makes it necessary to make the middle part with a gradual thickening towards the hub. Profiles A and B the middle parts are complex, which makes it difficult to process the disc. Although the end surfaces A And B do not mate with other parts, they must be made accurately, with a high surface finish. All machining defects in the form of scratches (traces from the cutter) or transitions are stress concentrators and reduce the mechanical strength of the disk. Very great importance has a uniform distribution of the metal mass over the disk.

Even small one-sided thickenings lead to an uneven mass distribution, which leads to imbalance. With the rapid rotation of unbalanced disks, unacceptable vibrations of the turbine appear.

bins that could cause an accident. Therefore, when designing disks, tight tolerances are specified for all disk sizes.

Turbine disk design

Particularly high requirements for processing accuracy are imposed on the mating dimensions - the mounting hole in the hub or mounting belts and the grooves for attaching the blades. Landing belts and holes in the hub are usually made according to the 2nd accuracy class. Tolerances for the dimensions of the groove for fastening the blades - 0.01-0.03 mm. Permissible runout of the outer surfaces of the seats - 0.03-0.06 mm.

The transmission of torque from the disk to the shaft is carried out by bolts or pins inserted into the holes G (see Fig. 77, a) or splines E(see fig.77, b). Sometimes the shaft is machined along with the flange, and the turbine disk is welded to the shaft flange, as shown in Fig. 77, V. With this design of the disk, savings are achieved on expensive heat-resistant alloys, since the shaft is made of cheaper steels.

When designing turbine disks, very great attention is given to a rational way of attaching the blades, taking into account the structural strength and manufacturability of the design.

The greatest structural strength with a minimum weight of the disk is achieved when the blades are made in one piece with the disk. With such discs, the rim is the lightest. However, the technology of their manufacture is complex and labor-intensive. In addition, the quality of the processing of the profile of the blades is higher if the blades are manufactured separately from the rotor. Increased roughness or discrepancy between the blade profile and the calculated one reduces the efficiency of the turbine. All these factors are analyzed in detail and the most rational solution is found in each specific HP design.

Despite the apparent advantages of obtaining blanks for turbine disks in one piece with the blades, in real conditions it is sometimes more expedient to manufacture the blades separately and then connect them to the disk using locks or welding.

The blade of a gas turbine consists of two main structural elements - pen And root part with a lock. The feather is the working element of the blade, and the root part, or lock, serves to connect the feather to the turbine disk. The blade airfoil has a complex shape determined by gas-dynamic calculation. The concave side of the pen is called the trough, and the convex side is called the back. The profiles of the trough and the back are connected, forming the edges of the feather: the front, or inlet, edge from the side of the gas inlet to the blade and the rear, or outlet, edge. In practice, three characteristic types of gas turbine blades of THA are widely used:

a blade made separately and connected to the turbine disk by welding or a lock;

blades of an open type, made in one piece with the turbine disk;

blades made in one piece with the turbine disk, connected from above by a shroud ring.

Each of these types of blades has its own advantages and disadvantages, both operational and technological.

The blades of the first type are made separately from the disk and can be made more accurately and with better surface finish than the other types of blades.

A large number of blades are used for each turbine, which makes it possible to organize in-line production of blades using special equipment and high-performance tooling even in small-scale production of HPP. However, the need to fasten separately made blades to the disk using locks complicates technological process and makes the turbine disk heavier. This drawback is largely eliminated by welding the blades to the disk.

Blades of the second type are the most rational in terms of design, since they do not require fastening. However, such blades cannot be manufactured by conventional machining. To select the metal between the blades, it is necessary to use electroerosive, ultrasonic or other methods, which are significantly inferior in performance to conventional machining. In addition, the manufacture of this type of blades requires very precise adherence to the technological process, since the presence of one rejected blade leads to the rejection of the entire turbine disk. Blades of the second and third types cannot be made of a metal or an alloy other than the metal of the disk (since they are one with the disk), which is not always rational, and sometimes even unacceptable.

The blades of the third type are just as rational from a constructive point of view as the blades of the second type. The presence of a bandage made in one piece with the blades even improves their characteristics, but the manufacturing technology of such blades does not allow obtaining the exact geometric dimensions of the blade profile. Lost wax casting gives significant errors, and the processing of closed blade profiles is difficult.

The technological process of manufacturing each of the three types of blades has its own characteristics. The material of the blades also has a great influence on the technological process.

Gas turbine blades operate under harsh conditions - at high temperatures and high stresses from centrifugal forces. The material of the blades must have good heat resistance and, at the same time, be satisfactorily processed by cutting and pressure. The material for cast blades must have high casting properties. The material of the welded blades must be well welded to the material of the disk. The following steels and alloys are used for the manufacture of turbine blades: 1Kh18N9T, ZOHGSA, EI69, VL7-20 and others.

For short-term operation at not very high temperatures, aluminum-based alloys of the AK4 type can be used.

Housing parts of turbopump units can be divided into the following main groups:

Pump housings.

Turbine casings.

Exhaust pipes and manifolds.

Fig.78 TPU body parts Most of THA body parts, fig.78, have a complex shape formed by curved, flat and cylindrical surfaces. Curvilinear surfaces forming volutes, cavities, recesses are not subjected to machining, but are cleaned to remove surface irregularities. Some of these surfaces are marked with the letter Y.

To install bearings, seals and other parts adjacent to the shafts of turbines and pumps, bores, grooves, and landing belts are made in the housings. These seats are machined with high precision - according to the 2nd or 1st class. The mutual runout of the seating surfaces is allowed within 0.03-0.05 mm, and the non-parallelism of the ends is 0.03-0.08 mm. With the same high accuracy, the joints of body parts with each other along the parting planes are processed. P. Particularly stringent requirements for seats and docking places are imposed in HPP designs that have a common shaft of the turbine and pumps.

The combination in one part of raw surfaces with relatively coarse tolerances, with surfaces machined with high precision, is one of the characteristic features of body parts.

The material for the housings is selected based on the conditions of their work, possibly the minimum weight and manufacturability of the design. Pump casings are most often made of cast aluminum alloys of the AL4 type, which have high casting properties with sufficient strength.

Turbine housings are also preferably made of AL4 type alloys, if this is allowed by temperature conditions. At high gas temperatures, turbine housings are made of heat-resistant stainless steels of the 1X18H9T type. Pump casings for pumping aggressive liquids are made of titanium alloys with high corrosion resistance. Sometimes, due to the conditions of minimum weight and design considerations, body parts are made by stamping from a sheet, followed by welding. For welded stamped bodies, alloys EI606, EI654, steel 1Kh18N9T and others are used.

Welded housings made of sheet materials are usually cheaper and lighter than cast ones, so they are widely used.

Welded turbine housing:

1-flange; 2 - collector; 3-ring

Figure 79 shows an example of fabrication of a welded turbine housing with an exhaust manifold.

The body is divided into three elementary parts. Middle part - collector 2 is made by stamping from a thin sheet, and the flange 1 and the seat ring 3 obtained by turning. Elementary parts are connected by two circumferential welds C. Welding is carried out in a special device, the parts are rotated by a welding manipulator.

The invention relates to rocket technology, specifically to liquid rocket engines operating on a cryogenic oxidizer and hydrocarbon fuel. The turbopump unit (TPU) of a liquid-propellant rocket engine contains the impeller of the oxidizer pump, the impeller of the fuel pump and the impeller of the turbine installed on the shaft of the rotor of the turbopump unit, the impeller of the additional fuel pump with the shaft and the impeller of the additional fuel pump, according to the invention, between the impeller the turbine and the impeller of the oxidizer pump are equipped with a magnetic coupling and a multiplier. A magnetic coupling and a multiplier can be installed between the oxidizer pump and the fuel pump. A magnetic coupling and a multiplier can be installed between the fuel pump and the additional fuel pump. The invention provides an increase in the reliability of TNA. 2 w.p. f-ly, 3 ill.

The invention relates to rocket technology, specifically to liquid rocket engines LRE, operating on a cryogenic oxidizer and hydrocarbon fuel.

A liquid rocket engine is known according to the RF patent for invention No. 2095607, intended for use as part of space upper stages, stages of launch vehicles and as a sustainer engine of spacecraft, includes a combustion chamber with a regenerative cooling path, a turbopump unit - TNA. TNA contains pumps for supplying components - fuel and oxidizer with a turbine on one shaft, into which a capacitor is inserted. The outlet of the condenser through the refrigerant line is connected to the inlet to the combustion chamber and to the inlet to the regenerative cooling path of the combustion chamber. The outlet from the condenser through the coolant line is connected to the inlet to the pump of one of the components. The pump outlet of the same component is connected to the condenser inlet through the refrigerant line. The second inlet of the condenser is connected to the outlet of the turbine. The pump outlet of the other component is in communication with the combustion chamber inlet.

The disadvantage of the THA engine is the deterioration of the cavitation properties of the pump when the condensate is bypassed. Such a property of the pump inevitably leads to a decrease in the flow rate of one of the fuel components through the HP, a drop in rocket thrust by several times and a disruption of the rocket flight program or to a catastrophe.

Known method of operation of the LRE and liquid rocket engine according to the RF patent for the invention No. 2187684. The method of operation of a liquid-propellant rocket engine consists in supplying fuel components to the combustion chamber of the engine, gasifying one of the components in the combustion chamber cooling path, supplying it to the turbine of the turbopump unit, and then discharging it into the nozzle head of the combustion chamber. Part of the consumption of one of the fuel components is sent to the combustion chamber, and the remaining part is gasified and sent to the turbines of the turbopump units. The gaseous component exhausted on the turbines is mixed with the liquid component entering the engine at a pressure exceeding the saturated vapor pressure of the resulting mixture. The liquid rocket engine contains a combustion chamber with a regenerative cooling path, pumps for supplying fuel components and a turbine. Pumps and turbines are arranged in two HPs: main and booster. The engine comprises a supply pump of one of the fuel components of the main turbopump unit, a booster turbopump unit pump and a mixer installed in series before the pump. The output of the pump of the main turbopump unit is connected both to the injector head of the combustion chamber and to the path of the regenerative cooling of the combustion chamber. The regenerative cooling path, in turn, is connected to the turbines of the main and booster turbopump units, the outputs of which are connected to the mixer.

The disadvantage of this scheme is that the thermal energy removed during the cooling of the combustion chamber may not be enough to drive the turbopump unit of a very high power engine.

Known LRE under the RF patent for the invention No. 2190114, IPC 7 F02K 9/48, publ. 09/27/2002 This rocket engine includes a combustion chamber with a regenerative cooling path, a turbopump unit TNA with oxidizer and fuel pumps, the output lines of which are connected to the combustion chamber head, the main turbine and the main turbine drive circuit. The main turbine drive circuit includes a fuel pump connected in series and a regenerative cooling circuit of the combustion chamber connected to the main turbine inlet. The output from the TNA turbine is connected to the input of the second stage of the fuel pump.

This engine has a significant drawback. Bypassing the fuel heated in the regenerative cooling path of the combustion chamber to the inlet to the second stage of the fuel pump will lead to its cavitation and the consequences indicated above. Most LREs use propellant components such that the oxidizer consumption is almost always greater than the fuel consumption. Therefore, for powerful liquid-propellant rocket engines with high thrust and high pressure in the combustion chamber, this scheme is unacceptable, because fuel consumption will not be enough to cool the combustion chamber and drive the main turbine.

In addition, the launch system of the rocket engine, the system for igniting the fuel components and the system for shutting down the rocket engine and cleaning it from fuel residues in the regenerative cooling path of the combustion chamber have not been worked out.

Known liquid rocket engine and a method of launching it according to the RF patent for the invention No. 2232915, publ. 09/10/2003 (prototype), which contains a combustion chamber, a turbopump unit, a gas generator, a starting system, means for igniting fuel components and fuel lines. The outlet of the oxidizer pump is connected to the inlet to the gas generator. The output of the first stage of the fuel pump is connected to the regenerative cooling channels of the chamber and to the mixing head. The output of the second stage of the fuel pump (auxiliary fuel pump) is connected to an electrically driven flow controller. The other input of the regulator is connected to the starting tank with standard fuel. The output from the regulator is connected to the gas generator. The outlet from the gas generator is connected to the inlet to the turbine of the turbopump unit, the outlet of which is connected to the mixing head. The flow regulator is equipped with a hydraulic drive of the preliminary stage, which is connected to the starting tank with standard fuel through a cavitating jet and a hydraulic relay. The hydraulic relay is connected to the second stage of the fuel pump. The throttle installed at the outlet of the first stage of the fuel pump is made together with a pre-stage controlled valve.

The disadvantage of this scheme is a fire or explosion of the HP and the rocket at launch or in flight due to the low reliability of the seal between the turbine and the oxidizer pump, between the oxidizer and fuel pump, and also between the fuel pump and the additional fuel pump due to the large pressure drop acting on them: 300 ... 400 kgf / cm 2 for modern rocket engines. For example, when hydrogen and oxygen are used as rocket fuel components, the smallest leaks of these components lead to the formation of an “explosive mixture” and almost always to a rocket explosion.

The objectives of the invention: to prevent the explosion of TNA or rocket at the start or in flight.

The solution to this problem is achieved due to the fact that the turbopump unit of a liquid-propellant rocket engine, containing the parts of the rotor of the turbopump unit mounted on the shaft: the impeller of the oxidizer pump, the impeller of the fuel pump and the impeller of the turbine, the impeller of the additional fuel pump with the shaft and the impeller of the additional fuel pump, differs in that a magnetic coupling is installed between the turbine impeller and the impeller of the oxidizer pump. A magnetic coupling can also be installed between the oxidizer pump and the fuel pump. A magnetic coupling can also be installed between the fuel pump and the auxiliary fuel pump.

The conducted patent studies have shown that the proposed technical solution has novelty, inventive step and industrial applicability. The novelty is confirmed by the conducted patent research, the inventive step is the achievement of a new effect - the absolute tightness of the connections between the turbine and pumps, as well as between the pumps and the prevention of the explosion of the HP and the rocket at the start or in flight.

Industrial applicability is due to the fact that all the elements included in the layout of the HP are known from the prior art and are widely used in engine building.

The essence of the invention is illustrated in Fig.1...3, where:

Figure 1 shows a diagram of the first version of TNA,

Figure 2 shows a diagram of the second version of TNA,

Figure 3 shows a diagram of the third version of TNA.

The turbopump assembly of the liquid-propellant rocket engine TNA 1 (figure 1) contains a fuel pump shaft 2, an oxidizer pump shaft 3. An impeller of an oxidizer pump 4 is installed on the shaft of the oxidizer pump 3, an impeller of the fuel pump 5 is installed on the shaft of the fuel pump 2. The impeller of the turbine 6 is installed at the top of the TNA. All parts of the HP rotor are located inside the HP housing 7. The additional fuel pump 8, which has the impeller of the additional fuel pump 9 and the shaft of the additional fuel pump 10, is made coaxially with the HP 1 and is installed on the side opposite to the turbine impeller 6. The impeller of the additional fuel pump 9 is installed in housing of the additional fuel pump 11, the cavity of which "B" is sealed relative to the cavity of the TPU "A". A magnetic clutch 12 and a multiplier 13 are installed between the impeller of the fuel pump 5 and the additional fuel pump 8 in the HP housing 7. The magnetic clutch 12 and all other magnetic clutches (if they are used in the design) consist of a magnetic clutch drive disk of a magnetic clutch driven disk, and disks of the magnetic coupling is made of a partition of non-magnetic material, such as non-magnetic steel (figure 1...3 not shown). The turbine impeller is mounted on the turbine shaft 14.

The gas generator 15 is mounted coaxially with the TNA 1 above the nozzle apparatus of the turbine 16. The gas generator 15 contains the head of the gas generator 17, inside which there is an outer plate 18 and an inner plate 19 with a cavity "B" above them and a cavity "G" between them. Inside the head of the gas generator 17, oxidizer nozzles 20 and fuel nozzles 21 are installed. The oxidizer nozzles 20 communicate the cavity "B" with the internal cavity of the gas generator "D", and the fuel nozzles 21 communicate the cavity "G" with the internal cavity of the gas generator "D". A fuel manifold 22 is installed on the outer surface of the gas generator 15, to which a high-pressure fuel line 23 from an additional fuel pump 8 is connected. A high-pressure valve 24 and a flow regulator 25 with a flow regulator drive 26 are installed in the high-pressure pipeline line 23. The output from the impeller of the fuel pump 5 connected by a pipeline 27 with the entrance to the additional fuel pump 8 and with the combustion chamber (combustion chamber in figure 1 is not shown).

The outlet from the impeller of the oxidizer pump 4 is connected by the oxidizer pipeline 28 through the oxidizer valve 29 to the cavity "B" of the gas generator 15. One or more ignition devices 30 are installed on the gas generator 15. The control unit 31 is connected by electrical connections with the ignition devices 30, the high pressure valve 24, the valve oxidizer 29 and flow regulator drive 26.

When the LRE is started from the control unit 31, electrical signals are sent to the valves 24 and 29 and the ignition (ignition) device 30. The oxidizer and fuel from the impellers of pumps 4, 5 and 8 by gravity enters the gas generator 15, where it ignites, the combustion products spin the turbine wheel 6 mounted on shaft 14.

In the first variant (figure 1) through the magnetic coupling 12 and the multiplier 13, the shaft of the oxidizer pump 3 is untwisted. The pressure at the outlet of the impellers of the pumps 4 and 5 increases. Part of the fuel (about 10%) enters the additional fuel pump 8, where its pressure increases significantly. The additional fuel pump 8 is driven and has the same speed as the impeller of the oxidizer pump 4 and the impeller of the fuel pump 5 (figure 1).

According to the second option (figure 2), the torque from the shaft of the oxidizer pump 3 is transmitted to the shaft of the fuel pump 2 through the magnetic coupling 12 and the multiplier 13. In this case, the impeller of the fuel pump 5 will have a higher speed than the impeller of the oxidizer pump 4. The shaft of the additional pump fuel 10 is connected to the shaft of the fuel pump 2 directly.

According to the third variant (figure 3), in addition to two magnetic couplings with multipliers, a third magnetic coupling with a multiplier is used in the TNA design. As a result, due to the lack of a seal along the shaft of the additional fuel pump 10, its reliability increases. At the pressure at the inlet to the impeller of the fuel pump 4 order P 1 = 4 ... 5 kgf / cm 2, at the outlet of the impeller of the fuel pumps 4 P 2 = 300 kgf / cm 2 and at the pressure at the outlet of the additional fuel pump 8 approximately P 3 \u003d 900 kgf / cm 2 the pressure difference that has arisen between them of approximately 600 kgf / cm 2 is perceived by a partition made of non-magnetic material 14. The pressure at the inlet to the oxidizer pump P 4 \u003d 4 ... 5 kgf / cm 2, at the outlet of the oxidizer pump P 5 \u003d 400 kgf / cm 2, at the inlet to the combustion chamber P 6 \u003d 300 kgf / cm 2. The presence of magnetic couplings between the pumps and the oxidizer pump and the turbine ensures complete tightness of all modules relative to each other, the presence of multipliers ensures the coordination of the rotational speed of the turbine and pumps and, at the same time, the modularity of the design.

As a result, it became possible to design all the main HP units: the turbine and the pump for optimal parameters, including rotational speeds, and to coordinate the rotation speeds by using one multiplier between the turbine and pumps or several multipliers, and this made it possible to minimize the weight of the HP, which is of decisive importance in rocket technology.

The application of the invention allowed:

1. To prevent the explosion of the HP and the rocket during launch or in flight due to the contact of the oxidizer and fuel in the cavity between the pumps or the penetration of combustion products from the turbine into one of the fuel components, if oxygen and hydrogen or other aggressive components are used as propellant components.

2. Ensure the modularity of the TPU design.

1) Study of the scheme and principle of operation of a liquid-propellant rocket engine (LRE).

2) Determination of the change in the parameters of the working fluid along the path of the LRE chamber.

1. GENERAL INFORMATION ABOUT LRE

2.1. The composition of the rocket engine

A jet engine is a technical device that creates thrust as a result of the outflow of a working fluid from it. Jet engines provide acceleration of moving vehicles of various types.

A rocket engine is a jet engine that uses only the substances and energy sources that are stored on board a moving vehicle.

A liquid-propellant rocket engine (LRE) is a rocket engine that uses fuel (primary energy source and working fluid) that is in a liquid state of aggregation for operation.

LRE generally consists of:

2- turbopump units (TPU);

3- gas generators;

4 pipelines;

5- automation units;

6- auxiliary devices

One or more liquid-propellant rocket engines, together with a pneumatic-hydraulic system (PGS) for supplying fuel to the engine chambers and auxiliary units of the rocket stage, constitute a liquid-propellant rocket propulsion system (LPRE).

As a liquid propellant (LFR), a substance or several substances (oxidizer, fuel) are used, which are capable of forming high-temperature combustion (decomposition) products as a result of exothermic chemical reactions. These products are the working body of the engine.

Each LRE chamber consists of a combustion chamber and a nozzle. In the LRE chamber, the primary chemical energy liquid fuel is converted into the final kinetic energy of the gaseous working fluid, as a result of which the reactive force of the chamber is created.

A separate turbopump unit of the LRE consists of pumps and a turbine that drives them. TNA provides the supply of liquid fuel components to the chambers and gas generators of the LRE.

The LRE gas generator is a unit in which the main or auxiliary fuel is converted into gas generation products used as the working fluid of the turbine and the working fluids of the pressurization system for tanks with LRE components.

The LRE automation system is a set of devices (valves, regulators, sensors, etc.) of various types: electrical, mechanical, hydraulic, pneumatic, pyrotechnic, etc. Automation units provide starting, control, regulation and shutdown of the LRE.

LRE parameters

The main traction parameters of the LRE are:

The reactive force of the LRE - R is the resultant gas and hydrodynamic forces acting on the internal surfaces of the rocket engine during the outflow of matter from it;

LRE thrust - R - the resultant of the reactive force of the LRE (R) and all pressure forces environment, which act on the outer surfaces of the engine with the exception of the forces of external aerodynamic resistance;

LRE thrust impulse - I - integral of the LRE thrust over the time of its operation;

The specific thrust impulse of the LRE - I y - the ratio of thrust (P) to the mass fuel consumption () of the LRE.

The main parameters that characterize the processes occurring in the LRE chamber are pressure (p), temperature (T) and flow rate (W) of the products of combustion (decomposition) of liquid rocket fuel. In this case, the values ​​of the parameters at the nozzle inlet (section index “c”), as well as in the critical (“*”) and outlet (“a”) sections of the nozzle are highlighted.

Calculation of parameter values ​​in various sections of the LRE nozzle tract and determination of the thrust parameters of the engine is carried out according to the corresponding equations of thermogasdynamics. An approximate methodology for such a calculation is discussed in Section 4 of this manual.

1. SCHEME AND PRINCIPLE OF OPERATION LRE "RD-214"

3.1. general characteristics LRE "RD-214"

The RD-214 liquid-propellant rocket engine has been used in domestic practice since 1957. Since 1962, it has been installed on the 1st stage of the Kosmos multi-stage launch vehicles, with the help of which many satellites of the Kosmos and Interkomos series have been launched into near-Earth orbits.

LRE "RD-214" has a pumping fuel supply system. The engine runs on a high-boiling nitric acid oxidizer (a solution of nitrogen oxides in nitric acid) and hydrocarbon fuel (kerosene processing products). A special component is used for the gas generator - liquid hydrogen peroxide.

The main parameters of the engine have the following meanings:

Thrust in the void R p = 726 kN;

The specific impulse of thrust in the void I yn = 2590 N×s/kg;

Gas pressure in the combustion chamber p k = 4.4 MPa;

Degree of gas expansion in the nozzle e = 64

LRE "RD-214", (Fig. 1) consists of:

Four chambers (pos. 6);

One turbopump unit (TPU) (pos. 1, 2, 3, 4);

Gas generator (pos. 5);

pipeline;

Automation units (pos. 7, 8)

The THA of the engine consists of an oxidizer pump (pos. 2), a fuel pump (pos. 3), a hydrogen peroxide pump (pos. 4) and a turbine (pos. 1). The rotors (rotating parts) of the pumps and the turbine are connected by a single shaft.

Units and units that provide the supply of components to the engine chamber, gas generator and turbine are combined into three separate systems - lines:

Oxidizer supply system

fuel supply system

Hydrogen peroxide steam and gas generation system.

Fig.1. Schematic of a liquid propellant rocket engine

1 - turbine; 2 – oxidizer pump; 3 - fuel pump;

4 – hydrogen peroxide pump; 5 – gas generator (reactor);

6 – engine chamber; 7, 8 - elements of automation.

3.2. Characteristics of the LRE units "RD-214"

3.2.1. LRE chamber

Four LRE chambers are connected into a single block along two sections with the help of bolts.

Each LRE chamber (pos. 6) consists of a mixing head and a body. The mixing head includes top, middle and bottom (firing) bottoms. A cavity for the oxidizer is formed between the upper and middle bottoms, and a cavity for fuel is formed between the middle and fire bottoms. Each of the cavities is connected with the internal volume of the engine housing by means of the corresponding injectors.

In the process of LRE operation, liquid fuel components are supplied, sprayed and mixed through the mixing head and its nozzles.

LRE chamber housing includes part of combustion chamber and nozzle. The liquid-propellant rocket engine nozzle is supersonic, has converging and diverging parts.

The housing of the LRE chamber is double-walled. The inner (fire) and outer (power) walls of the body are interconnected by spacers. At the same time, with the help of spacers, channels of the housing liquid cooling path are formed between the walls. Fuel is used as a coolant.

During engine operation, fuel is supplied to the cooling path through special manifold pipes located at the end of the nozzle. Having passed the cooling path, the fuel enters the corresponding cavity of the mixing head and is introduced through the nozzles into the combustion chamber. At the same time, through another cavity of the mixing head and the corresponding nozzles, an oxidizer enters the combustion chamber.

In the volume of the combustion chamber, spraying, mixing and combustion of liquid fuel components takes place. As a result, a high-temperature gaseous working fluid of the engine is formed.

Then, in the supersonic nozzle, the thermal energy of the working fluid is converted into the kinetic energy of its jet, upon expiration of which the LRE thrust is created.

3.2.2. Gas generator and turbopump unit

The gas generator (Fig. 1, item 5) is a unit in which liquid hydrogen peroxide is converted into a high-temperature vaporous working fluid of the turbine as a result of exothermic decomposition.

The turbopump unit provides pressure supply of liquid fuel components to the chamber and engine gas generator.

THA consists of (Fig. 1):

Screw-centrifugal oxidizer pump (pos. 2);

Screw-centrifugal fuel pump (pos. 3);

Hydrogen peroxide centrifugal pump (item 4);

Gas turbine (pos. 1).

Each pump and turbine has a fixed stator and a rotating rotor. The rotors of pumps and turbines have a common shaft, which consists of two parts, which are connected by a spring.

The turbine (pos. 1) serves as a pump drive. The main elements of the turbine stator are the housing and the nozzle apparatus, and the main elements of the rotor are the shaft and the impeller with blades. During operation, peroxide vapor gas is supplied to the turbine from the gas generator. When the steam gas passes through the nozzle apparatus and the blades of the turbine impeller, its thermal energy is converted into mechanical energy of rotation of the wheel and the turbine rotor shaft. The exhaust steam gas is collected in the outlet manifold of the turbine housing and discharged into the atmosphere through special waste nozzles. This creates some additional thrust LRE.

Pumps for oxidizer (pos. 2) and fuel (pos. 3) are screw-centrifugal type. The main elements of each of the pumps are the housing and the rotor. The rotor has a shaft, an auger and a centrifugal wheel with blades. During operation, mechanical energy is supplied from the turbine to the pump through a common shaft, which ensures the rotation of the pump rotor. As a result of the action of the screw blades and the centrifugal wheel on the liquid (fuel component) pumped by the pumps, the mechanical energy of rotation of the pump rotor is converted into potential energy of the liquid pressure, which ensures the supply of the component to the engine chamber. An auger in front of the centrifugal impeller of the pump is installed to preliminarily increase the pressure of the liquid at the inlet to the interblade channels of the impeller in order to prevent cold boiling of the liquid (cavitation) and disruption of its continuity. Disturbances in the continuity of the flow of the component can cause instability of the fuel combustion process in the engine chamber, and, consequently, the instability of the LRE as a whole.

A centrifugal pump (pos. 4) is used to supply hydrogen peroxide to the gas generator. The relatively low consumption of the component creates conditions for non-cavitational operation of a centrifugal pump without installing a screw prepump in front of it.

3.3. The principle of the engine

Start, control and stop of the engine is carried out automatically by electrical commands from the rocket board to the corresponding automation elements.

For the initial ignition of the fuel components, a special starting fuel is used, self-igniting with an oxidizer. Starting fuel initially fills a small section of the pipeline in front of the fuel pump. At the moment of launching the LRE, starting fuel and oxidizer enter the chamber, they spontaneously ignite, and only then do the main components of the fuel begin to enter the chamber.

During engine operation, the oxidizer sequentially passes through the elements and assemblies of the line (system), including:

Dividing valve;

Oxidizer pump;

Oxidizer valve;

Mixing head chamber motor.

The flow of fuel flows through the line, including:

Dividing valves;

fuel pump;

Collector and path for cooling the engine chamber;

mixing head chamber.

Hydrogen peroxide and the resulting vapor gas sequentially pass through the elements and units of the steam and gas generation system, including:

Dividing valve;

Hydrogen peroxide pump;

Hydraulic reducer;

gas generator;

Turbine nozzle apparatus;

Turbine impeller blades;

turbine manifold;

Waste nozzles.

As a result of the continuous supply of fuel components by the turbopump unit to the engine chamber, their combustion with the formation of a high-temperature working fluid and the expiration of the working fluid from the chamber, the LRE thrust is created.

Variation of the thrust value of the engine during its operation is provided by changing the flow rate of hydrogen peroxide supplied to the gas generator. This changes the power of the turbine and pumps, and, consequently, the supply of fuel components to the engine chamber.

Shutdown of the LRE is carried out in two stages with the help of automation elements. From the main mode, the engine is first switched to the final mode of operation with less thrust and only then is completely switched off.

1. WORK METHODOLOGY

4.1. Scope and order of work

In the course of the work, the following actions are sequentially performed.

1) The scheme of the RD-214 rocket engine is being studied. The purpose and composition of the LRE, the design of the units, the principle of operation of the engine are considered.

2) The geometrical parameters of the LRE nozzle are measured. The diameter of the inlet ("c"), critical ("*") and outlet ("a") sections of the nozzle (D c, D * , D a) is found.

3) The value of the parameters of the LRE working fluid in the inlet, critical and outlet sections of the LRE nozzle is calculated.

Based on the results of the calculations, a generalized graph of the change in temperature (T), pressure (p) and velocity (W) of the working fluid along the nozzle path (L) of the LRE is constructed.

4) The thrust parameters of the liquid-propellant rocket engine are determined at the design mode of operation of the nozzle ().

4.2. Initial data for calculating the parameters of the rocket engine "RD-214"

Gas pressure in the chamber (see option)

Temperature of gases in the chamber

Gas constant

Isentropic exponent

Function

It is assumed that the processes in the chamber proceed without energy losses. In this case, the energy loss coefficients in the combustion chamber and nozzle, respectively, are

The nozzle operation mode is calculated (index " r»).

The measurement determines:

Nozzle throat diameter ;

Nozzle outlet diameter .

4.3. Sequence of calculation of LRE parameters

A) The parameters in the outlet section of the nozzle ("a") are determined in the following sequence.

1) Nozzle exit area

2) Nozzle throat area

3) Geometric degree of gas expansion

CONTROL QUESTIONS

1. What is the significance of the RD-214 rocket engine?

2. List the main systems of the studied LRE.

3. What is the purpose of the LRE chamber, what parts does it consist of?

4. What is the purpose of TNA, list its main units?

5. What is the purpose and composition of the steam and gas generation system of the LRE "RD-214"?

6. Describe the sequence of passage of the working fluid of the turbine.

7. List the main traction parameters of the rocket engine; name their values ​​for LRE "RD-214".

TNA are divided into single-shaft and multi-shaft. In single-shaft HPs, the turbine and pumps are located on the same shaft. The advantage of TNA, made according to this scheme, is the simplicity of design and low weight. As a disadvantage, it should be noted that only one of the pumps (usually the oxidizer pump) operates at the optimum speed. In this case, the fuel pump is operated at reduced efficiency values.

There are the following layout diagrams of TNA, Fig.57.

With a three-shaft TPU scheme, the speeds of the pumps and the turbine are independent of each other and can be selected from the conditions for the optimal operation of the pumps. However, the presence of gearboxes operating in difficult conditions (high peripheral speeds, difficulty in providing an effective lubrication and cooling system) in some cases minimizes the gain from increasing pump efficiency values.

Single shaft

Three-shaft

Layout diagrams of TNA

Single-shaft schemes of HP are the most widely used in LRE.

5.3. Centrifugal pump device

In TNA LRE, centrifugal pumps are usually used as the main ones. The main advantages that determine the predominant use of these types of pumps in LRE are:

Ensuring high supply pressures and productivity with small dimensions and weight;

Ability to work on aggressive and low-boiling components;

The ability to work with a large number of revolutions and the convenience of using a turbine to drive them.

Figure 58 shows a diagram of a single-stage centrifugal pump. The liquid is supplied through the inlet pipe 1 to the rotating wheel (impeller) 2. In the pump wheel, the liquid moves through the channel formed by the walls of the wheel and the blades 3. The force acting from the side of the wheel blades on the liquid makes it move in such a way that the energy supply per unit mass of the liquid increases. In this case, both the potential energy (static pressure) and the kinetic energy of the liquid increase.

Fig.58

Scheme of a centrifugal pump:

1 - inlet pipe; 2 - pump wheel (impeller); 3 - blades;

4 - diffuser; 5 - diffuser blades; 6 - collection or snail; 7 - front seal;

8 - shaft bearing; 9 - bearing seal

At the outlet of the wheel, the liquid enters the diffuser 4, where its absolute velocity decreases and the pressure additionally increases. The simplest diffuser consists of smooth discs that make up its walls, and is called bladeless. The blade diffuser has fixed blades 5 (shown in dotted lines in Fig. 58), which contribute to a faster damping of the flow rate. After passing the diffuser, the liquid enters the spiral channel (cochlea) 6, the purpose of which is to collect the fluid coming out of the wheel, as well as to reduce its speed. The liquid is supplied to the network through the discharge pipe.

To reduce the flow of fluid from the high-pressure cavity (diffuser, volute) to the low-pressure area, seals are made in the pump 7.

Fig.59

Schemes of centrifugal pumps:

a-c axial entry; b- with spiral entry;

V- with a double-sided entrance; G- multistage pump

Centrifugal pumps are available with axial, scroll and double inlet, single and multi-stage. Choice of axial or spiral entry (fig.59, a, b) is determined primarily by the layout conditions of the HP and the propulsion system. Double entry (fig.59, V) are performed at high flow rates to reduce the inlet velocity and thereby improve the anti-cavitation properties of the pump. Multistage pumps (fig.59, G) are used when it is necessary to obtain particularly high pressures.

Typically, pump casings are cast from high-strength aluminum alloys, and in the case of high pressures, from steel. The number of profiled impeller blades is no more than 8, and their thickness is in the range of 2 ¸ 5 mm.

5.4. Pump impellers

There are impellers, open and closed types, Fig. 60 (a, b).

The open impeller is used in pumps with low component flow and pressure. An impeller of this type is characterized by significant losses due to the flow of the component from the area of ​​high pressure (at the outlet of the pump) to the area of ​​low pressure (at the inlet to the pump). The impeller consists of disk 1 and blades 2 made on it.

In closed impellers, a cover 3 is installed on the end surfaces of the blades, which can be made integral with the impeller. In impellers of this type, the flow loss of the component is much less than in open impellers. Usually the impellers are made by casting. The number of profiled blades, as a rule, does not exceed 8, and their thickness is less than 5 mm. The impellers shown in Fig. 60 refer to impellers with one-sided component supply.

To reduce the flow rate of the component through the blade channel of the impeller (in order to exclude the occurrence of the cavitation process), impellers with a two-sided supply of the component are used, Fig.61.

Fig.60

One-sided impellers:

a- open type; b - closed type

Fig.61

Double sided impeller

8.5. Impeller seals

In order to reduce fluid leakage, the following types of seals are installed in the pump impellers: slotted, labyrinth and floating, Fig. 62 a, b, c, respectively.

The principle of operation of slotted seals is based on providing a high hydraulic resistance of the annular gap between the graphite liner installed in the pump housing and the groove made in the inlet section of the disc. The design of this seal allows up to 15% leakage of the volume of the pumped liquid, while the labyrinth, Fig.62b, and floating (a set of fluoroplastic and aluminum washers installed in the inlet section of the impeller), Fig.62c, up to 10% and 5%, respectively.

Fig.62

Impeller seals:

a - slit; b - labyrinth; c - floating

5.5. Turbine THA

One of the main elements of TNA is a gas turbine. In the turbine, the potential energy of the combustion products from the gas generator or coolant vapor is converted into the mechanical work of the turbine. The turbine is designed to drive the HP pumps. The turbine consists of a nozzle apparatus 1, an impeller 2 with two rows of rotor blades 3 and 4, a guide vane 5 and a turbine housing 6 with an outlet pipe 7, Fig.75.

The first stage of the turbine is a set of nozzle apparatus 1 and impeller blades 3, the second is formed by fixed blades of the guide vane 5 and the second row of rotor blades 4.

The transformation of the enthalpy of the gas flow into the mechanical energy of the rotation of the shaft is carried out in two stages: the enthalpy of the gas flow - into the kinetic energy of the jet (in the nozzle apparatus); kinetic energy of the jet - into the mechanical energy of rotation of the shaft (on the impeller).

Fig.75

THA turbine design

Shafts of turbopump units (TPU) operate at high loads and high speeds. To lighten the weight, they are made hollow. The greatest alternating stresses in the metal of the shaft occur on its outer surface. In this case, any kind of sharp transitions, traces from the cutting tool and other surface defects are stress concentrators. In these places, cracks may form during operation, which will lead to breakage of the shaft. Therefore, special attention is paid to the cleanliness of the surface finish of the shaft with the introduction in some cases of hardening operations. Not only places for bearings, seals, landings, but also all other parts of the shaft that are not mated with other parts are subjected to finishing.

Large number of revolutions (10000-20000 rpm and more) force the designer to assign very tight tolerances for the alignment of the necks and seats, the accuracy of the location of the axial hole, the difference in wall thickness and other dimensions. The slightest geometric errors lead to an uneven distribution of the rotating masses of metal, which causes vibrations and shaking of the TPU.

5.6. Requirements for gas generators

The thrust value of LRE, as is known, is a linear function of the second fuel consumption. The fuel consumption per second for any particular engine with a pumped component supply system depends on the power developed by the turbine. The power of the turbine is completely determined by the flow rate per second and the parameters of the working fluid at the inlet to the turbine, i.e., at the outlet of the gas generator. Therefore, the gas generator is a device that sets the operating mode of the entire propulsion system. This circumstance determines the special requirements for this link in the fuel supply system (in addition to the general requirements for all LRE units, regardless of the specifics of their work). These requirements are as follows.

1. High stability of work. This means that the gas generator in all engine operating modes must provide the specified second gas flow rate as accurately as possible and, at the same time, the values ​​of the gas parameters (composition, pressure, temperature, etc.) must not go beyond certain (permissible) limits. The more stable the operation of the gas generator, the less stress the engine control systems experience in flight, and this increases the reliability of the engine and the accuracy of fire.

Especially important is the stability of the gas generator for rockets with unregulated rocket engines and rockets whose flight range is controlled only by the flight speed at the end of the active part of the trajectory. In the latter case, the deviation of the coordinates of the end of the active section of the trajectory, caused by the deviation of the engine thrust from the calculated value, due to the unstable operation of the gas generator, will completely turn into a deviation of the point of impact of the rocket from the target.

2. Ease of managing the workflow in a wide range of changes in its parameters. This requirement is also due to the regulating effect of the gas generator on the engine and the need to change the engine operating mode during one start (when regulating thrust during launch and in flight, when switching from the main thrust stage to the final one, etc.).

3. High working capacity of generator gas, which determines either the minimum energy consumption (and, accordingly, the minimum fuel consumption) for the HPP drive, or an increase in HPP power. This requirement is put forward due to the fact that the specific impulse of the engine is determined by the ratio of thrust to the entire second flow rate of the discarded mass. The concept of “discarded mass” includes both the combustion products of fuel in the chamber and the exhaust gas after the turbine. For liquid-propellant rocket engines, in which this gas is released into the atmosphere and develops a specific impulse that is smaller than the fuel combustion products flowing out of the engine chamber, the decisive condition for increasing the engine efficiency is to reduce the fuel consumption for the HPU drive. For an LRE with generator gas afterburning, the main thing is to increase the power of the HP, since this allows you to increase the pressure in the chamber and, at a given pressure at the nozzle exit, increase the degree of expansion of the ejected combustion products, i.e., increase the thermal efficiency of the chamber. The reduction in fuel consumption for the HPP drive and the increase in HPP power depend on the amount of energy given to the turbine by one kilogram of the working fluid. This energy is equal, as is known, to the product of the relative effective efficiency of the turbine and the available adiabatic heat drop.

5.7. Classification of gas generators

The basis for the classification of gas generators is the method of producing generator gas. Currently, three methods of gas generation are common.

1. Decomposition (with the help of catalysts or without them) a substance capable, after an external initiating action, to go on to further stable spontaneous decay, accompanied by the release of a significant amount of thermal energy and gaseous decomposition products. Such a substance can be either a component of the main engine fuel or a special means of gas generation, stored only for this purpose on board the rocket. Gas generators in which this process is implemented are called single-component. In the future, they are distinguished mainly by the type of decomposable substance (hydroperoxide, hydrazine, solid fuel, etc.).

2. Incineration liquid fuel, consisting of two components. It is best to use the main engine fuel for this purpose, since this significantly simplifies its supply to the gas generator and improves the operating conditions of the rocket. Gas generators of this type are called two-component.

3. Liquid evaporation in the cooling path of the engine chamber. With this method of obtaining the working fluid of the turbine, the problem of cooling the walls of the engine chamber is simultaneously solved. Gas generators of this type are called steam generators, and engine circuits are called generatorless. Schemes of steam generators are divided into circulating and with a change in the working fluid. In the first, an arbitrary working fluid (for example, water) circulates in a closed circuit "chamber cooling path - turbine - condenser - pump - chamber cooling path", turning alternately into vapor, then into liquid in its various parts. In schemes with a change in the working fluid, this circulation is absent. The working fluid after the turbine is removed from the cycle. Obviously, the direct release of exhaust gas into the atmosphere would noticeably worsen the efficiency of the engine, since the specific thrust of the exhaust pipes is always less than the specific thrust of the engine chamber. To eliminate these losses, one of the fuel components is usually sent to the chamber cooling path. After evaporation and operation in the turbine, it is sent to the engine chamber, where it is burned together with the second component. Thus, generatorless engines are made according to the scheme with afterburning of the working fluid of the turbine.

By design, gas generation systems differ significantly from each other, but nevertheless, the following common main elements can be distinguished in each of them:

gas generator;

fuel supply devices;

Automation.

In the gas generator (sometimes called a reactor), the working fluid of the turbine is directly formed - gas or steam of given parameters. Fuel supply devices ensure the supply of gas generation means (source materials) to the reactor. Automation regulates the working process, as well as starting and switching off the gas generator. Sometimes (for example, when operating on the main fuel), the gas generation system does not have independent fuel supply devices. In this case, the gas generator is supplied with fuel by the engine supply system.

The following types of gas generators (GG) have found application in LRE:

Solid fuel (TGG);

Hybrid (THG);

One-component liquid (single-component JGG);

Two-component liquid (two-component ZHGG);

Evaporative liquid (evaporative JGG);