Secrets of the D-30F6 Turbofan

[overscan (PaulMM)]

I've always wondered about how the Perm D-30F6 turbofan is able to power the MiG-31 to Mach 2.83, when it appeared to be a relatively conventional adaptation of the D-30 commercial turbofan family.

Source is ENGINES OF RUSSIAN COMBAT AIRCRAFT by Kotelnikov V. R., Khrobystova O. V., Zrelov V. A., Ponomarev V. A. (Mediarost, 2020)

The D-30 was the first experimental domestic turbofan engine with a mixing-type afterburner, which was designated D-30F (product "38"). It was developed and tested in 1965-1969. The aerodynamic appearance of the compressor remained the same as that of the civil prototype, only the materials in the turbocompressor section were replaced. In order to significantly increase the thrust, it was necessary to significantly increase the gas temperature in front of the turbine. The engine developed 11,500 kgf in afterburner. But this was not enough - a thrust of 15,500 kgf was required. The creation of such an engine was associated with high technical risks associated with the general insufficient level of theoretical and practical knowledge on the use of turbofan engines on a supersonic aircraft.

"It was unknown whether the engine compressor, "used" to work with a subsonic unshaded air intake, would retain its functionality when it interacted with a supersonic inlet... - writes V. G. Avgustinovich. - Would the turbine withstand the increased gas temperature, provided that the cooling air temperature reached 700 ° C." According to Valery Grigorievich, the OKB-19 designers had to disprove several theoretical dogmas that scientists at that time considered unshakable. One of them stated that the maximum gas temperature in front of the turbine should be maintained starting from takeoff mode near the ground. But it was not possible to obtain frontal thrust at a speed of 3000 km / h, which was specified in the technical specifications for the aircraft.

The necessary solution had already been tested on the RD36-41 single-shaft turbojet engine with afterburner. To obtain the required thrust at an altitude of 20 km and a speed of 3,200 km/h, the Rybinsk designers had to increase the gas temperature in this mode by 75° compared to takeoff. This was not enough for the turbofan engine. It was necessary to add 150° to the gas temperature in the takeoff mode. For the demonstration version of the future D-30F6, a program was developed to increase the gas temperature in front of the turbine with an increase in the flight speed of the aircraft, which later became known as "temperature spin-up". Thanks to it, the required thrust was provided at an altitude of 20 km at a speed of 2,500 km/h. Now the dogma has been refuted for dual-circuit engines with afterburner as well.

However, the TsIAM scientists still considered the creation of a bypass engine with afterburner with such high parameters for such high flight speeds to be impossible in principle. They wrote about this in the official negative conclusion on the project. But in spite of everything, the general designer of the Mikoyan Design Bureau R. A. Belyakov decided to take a risk and accepted P. A. Solovyov's offer. According to Pavel Aleksandrovich's memoirs, "we were still terribly afraid. For example, meetings with Ustinov [the secretary of the CPSU Central Committee responsible for defense. - Author's note] began with a discussion - is it possible to make such an engine?
They did not believe it. All the time they raised one question, then another... But Batitsky [the commander-in-chief of the Air Defense. - Author's note] auth.] was pushing hard, and Ustinov, apparently, wanted to have such a machine. And he announced there that we would make this engine. And Tumansky's engine was put aside." In the fall of 1971, bench tests of the D-30F6 prototype began. On September 16, 1975, the E-155MP aircraft with two D-30F6s made its first flight (test pilot A. V. Fedotov). During the flight research period, two aircraft were lost - the first prototype and the first production one. There were no casualties, and this, in comparison with the results of other machines being created, was quite a good indicator. However, both incidents were due to the fault of the engines. During eight years of testing, starting with the first bench test, the D-30F6 was deeply fine-tuned, its units were adapted to real operating conditions (35 prototypes were used).

As a result, the dimensions determined back in 1969 were preserved, but many things underwent changes. First of all, the materials: the engine was now completely made only of titanium and nickel. Much effort was put into improving the afterburner, its mixer and the front device, so that improve the ignition system and increase the stability of the combustion process in high-altitude conditions. In 1977, the Perm Engine Plant began serial production of the D-30F6, and two years later it passed state tests.

The D-30F6 engine has a modular design that improves manufacturability and maintainability. The low-pressure compressor consists of five stages and has a total pressure increase ratio of 3, the high-pressure compressor is 10-stage with a total pressure increase ratio of 7.05 (at the design point - 8). The inlet guide vane of the HPC is rotary.

The peripheral speed at the tips of the working blades of the first impeller is 325 m/s.

The rotor of the high-pressure compressor is of a disk type with reinforcement of the last stages. The compressors are relatively low-pressure, which made it possible to obtain the required stability margin. The combustion chamber is a straight-through tubular-annular one, with 12 fire tubes.

The high-pressure turbine is a cooled two-stage one. Air for cooling the nozzle and working blades of both stages is taken after the fifth and last stages of the high-pressure compressor. An air-to-air heat exchanger was installed to reduce the temperature of the cooling air. This technical solution was brought to series production for the first time in the world. To reduce the temperature of the main disks and blades, cover deflectors are attached to the disks.

Two intermediate disks with labyrinth seals are installed between the main disks. The low-pressure turbine is a two-stage uncooled one. For the first time in a domestic serial engine, a mixing-type afterburner and an adjustable all-mode nozzle of a flap design are used, the subsonic part of which is controlled autonomously forcibly , and the supersonic part is aerodynamically. An interesting feature of the nozzle design are the holes made on the second-row spacers, closed from the flow-through side by hinged valves. The holes with valves ensure the suppression of harmful gas oscillations in off-design modes. The critical cross-sectional area of the nozzle changes depending on the engine operating mode by turning the first and second row flaps using 18 hydraulic cylinders. The afterburner with four annular flame stabilizers is ignited by the "fire track" method. This system was previously implemented on the experimental turbojet engine RD-36-41 of the Rybinsk OKB-36.

So, essentially its a conventional low bypass ratio turbofan, with improved cooling of the high pressure turbine using a heat exchanger to cool the bleed air taken from the high pressure compressor in the bypass air flow so it is cooler when it arrives at the HP turbine. This allows the turbine entry temperature to increase from the sea level maximum when flying at high speeds and altitude.
Heat exchanger is shown below:



A somewhat different heat exchanger for turbine cooling air was fitted on the AL-31F




but it appears to be intended to reduce specific fuel consumption rather than increase high speed thrust. This might explain the disparity in SFC between AL-31F and RD-33, which are otherwise fairly similar technically.

As a side note the «огневой дорожки» 'fire track" method of igniting the afterburner was also used on the RD-33, where it was a contributing factor to low TBO and limited engine life as outlined in https://www.sciencedirect.com/science/article/pii/S1350630722000620. The AL-31F used a similar system.

Essentially rather than having separate igniter plugs in the afterburner, the afterburner is lit by a secondary pilot jet nozzle flame which travels through the turbine blades and mixer to the afterburner section.

The advantage of this ignition mix is very reliable afterburner operation, at higher speeds and altitudes in particular.

The disadvantage is the stators in the flame path get extra wear from the flame when afterburners are lit, as do the turbine blades to a lesser degree because they are rotating through it.



Other than the cooling boost from the heat exchanger allowing the engine to run hotter at high speed and temperature, the only other "secret" to high speed performance was careful attention to the air intake and exhaust, which are very important in high speed flight.

 

[paralay]

A large heat exchanger is installed in front of the MiG-31 engine. Smooth surface in the air intake channel

 

[F119Doctor]

The AB ignition system is what P&W called a Hot Streak ignition, used on J57, J75, and all models of the TF30 until the TF30-P-100, which went to high energy spark ignition. On the TF30, there was a single fuel squirt injector in one of the combustor cans, with a longer duration injector aft of the LPT to keep the ignition flame going until the Zone 1 spraying fuel ignited. It was a relatively reliable AB ignition system that provided a high energy ignition source, but did stress the turbine vanes in line with the hot streak, along with complicating the fuel control system to provide the hot streak fuel at the time of AB initiation.

All P&W afterburning engines since have used high voltage capacitive discharge AB ignition, which is lighter, simpler, and easier on the turbine, but is more difficult to get right for reliable ignition across the flight envelope. Electric sparks are much lower energy than the hot streak, and require the fuel /air ratio in the area of the spark to be very near stoichiometric to work.

 

[paralay]

Газотурбинные_двигатели_Иноземцев_А_А_,_Сандрацкий_В_Л_2006.djvu

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disk.yandex.ru

 

[F119Doctor]

A few things that got my attention:

1. How far aft the AB is located behind the turbine. It looks like they were trying to get a lot of passive mixing of the core and bypass air before the AB diffuser / flameholder section. Makes for a very long engine, although this may have been desirable for adapting the engine CG position in the airframe (See duct extension on the F110-400 for the F-14B/D).

2. The use of fan duct heat exchangers for the turbine cooling air. I haven’t seen that on western engines, although I have thought that it might be a useful feature.

3. Hydraulic actuators for each of the 18 convergent segments. I wonder how they synchronized the movement of those 18 cylinders?

4. The nozzle appears to be a conventional ejector configuration in the convergent section closed position, but the ejector path appears to close off partially or completely with the nozzle open. They talk about “valves” in the 2nd stage segment - they don’t show them in the drawings, but I wonder if these are the little triangular segments we see floating on the divergent flaps on the MiG-29 and Su-27 engine nozzles. These could let bay air fill the divergent section when it is over expanded subsonic to prevent flow instability, and then close off when the divergent area ratio becomes correct under higher nozzle pressure ratio conditions at the M2.8 cruise speed.

 

[overscan (PaulMM)]

A few things that got my attention:

1. How far aft the AB is located behind the turbine. It looks like they were trying to get a lot of passive mixing of the core and bypass air before the AB diffuser / flameholder section. Makes for a very long engine, although this may have been desirable for adapting the engine CG position in the airframe (See duct extension on the F110-400 for the F-14B/D).

It would make sense for length to be similar to R15B-300 engine for CG reasons but it's even longer. Perhaps due to the weight of the Zaslon radar etc in the nose.

2. The use of fan duct heat exchangers for the turbine cooling air. I haven’t seen that on western engines, although I have thought that it might be a useful feature.

I hadn't seen that before. According to Viktor Chepkin (Lyulka) it added 20kg in weight for 170 deg K reduction in cooling air temperature.

4. The nozzle appears to be a conventional ejector configuration in the convergent section closed position, but the ejector path appears to close off partially or completely with the nozzle open. They talk about “valves” in the 2nd stage segment - they don’t show them in the drawings, but I wonder if these are the little triangular segments we see floating on the divergent flaps on the MiG-29 and Su-27 engine nozzles. These could let bay air fill the divergent section when it is over expanded subsonic to prevent flow instability, and then close off when the divergent area ratio becomes correct under higher nozzle pressure ratio conditions at the M2.8 cruise speed.

You mean these doors? (D-30f6 on Su-47 Berkut)

The WS-10 has similar doors

 

[F119Doctor]

Yes - those little triangular flaps hanging from the upper part of the nozzle. I couldn’t understand their function on the WS-10 and the Russian engines with the short divergent sections, but might make sense in a nozzle that is over expanded in slower parts of the flight envelope

 

[Trident]

Kuznetsov afterburning turbofans (NK-32, NK-22, NK-144) also have the fan duct heat exchanger, though it takes up considerably less of the duct cross section (probably because BPR is higher). I think some of these Russian engines can further partially shut off cooling air supply to the turbine when not necessary to improve compressor performance/efficiency, which AFAIK the LEAP-1 also does. That may be where the fuel consumption advantage of the AL-31F is coming from.

As a side note the «огневой дорожки» 'fire track" method of igniting the afterburner was also used on the RD-33, where it was a contributing factor to low TBO and limited engine life as outlined in https://www.sciencedirect.com/science/article/pii/S1350630722000620. The AL-31F used a similar system.

Essentially rather than having separate igniter plugs in the afterburner, the afterburner is lit by a secondary pilot jet nozzle flame which travels through the turbine blades and mixer to the afterburner section.

The advantage of this ignition mix is very reliable afterburner operation, at higher speeds and altitudes in particular.

The disadvantage is the stators in the flame path get extra wear from the flame when afterburners are lit, as do the turbine blades to a lesser degree because they are rotating through it.

As F119Doctor mentioned, hot streak reheat ignition is a widespread solution, also found on the RB.199 for instance. Just to be clear, the ignitor flame is transient, maintained only until afterburner light-off is achieved, so the thermal stress on the turbine blading isn't quite as bad as it may seem.

 

[F119Doctor]

Kuznetsov afterburning turbofans (NK-32, NK-22, NK-144) also have the fan duct heat exchanger, though it takes up considerably less of the duct cross section (probably because BPR is higher). I think some of these Russian engines can further partially shut off cooling air supply to the turbine when not necessary to improve compressor performance/efficiency, which AFAIK the LEAP-1 also does. That may be where the fuel consumption advantage of the AL-31F is coming from.

The F119 has one small fan duct heat exchanger, but it is used to cool bleed air used to pressurize / seal the counter rotating rear bearing compartment, not for turbine cooling air. And the F135 has massive fan duct heat exchangers in addition to the bearing compartment unit, but these are part of the ECS system for the airframe.

The F119 uses HPC discharge (9th stage) air to cool the HPT, but uses 7th stage to cool the LPT. During development, they tested a valve to reduce the LPT cooling air at part power conditions to improve SFC, but the weight and complexity wasn’t worth the minimal gain in SFC, so it was deleted from the configuration.

 

[Trident]

It may very well be that the F119, thanks to more advanced materials and blade passage design, doesn't require as much cooling bleed as the AL-31F in the first place, so the savings potential would have been less.

 

[overscan (PaulMM)]

As F119Doctor mentioned, hot streak reheat ignition is a widespread solution, also found on the RB.199 for instance. Just to be clear, the ignitor flame is transient, maintained only until afterburner light-off is achieved, so the thermal stress on the turbine blading isn't quite as bad as it may seem.

Yep, but combined with less than stellar manufacturing quality control on the RD-33 turbine stators and rotors, it contributes to that engine's lower TBO. The report authors found a good correlation between number of afterburner actuations and premature wear. That's why more recent engines don't typically use it I think.

 

[paralay]

D30 engine: https://disk.yandex.ru/d/sJMl33f19KdNwQ

 

[F119Doctor]

D30 engine: https://disk.yandex.ru/d/sJMl33f19KdNwQ

Interesting that these three drawing all are the same length, but the turbomachinery section is different length in each one, with the augmentor section being lengthened or shortened to even out the total length. I’m guessing some intel gamesmanship is going on. Still a nice representation of the basic construction layout of the engine.

 

[overscan (PaulMM)]

I think its three different versions. Second is D30F11 and drawn by paralay himself - that was intended for the S-47 Berkut.

 

[Ronny]

A large heat exchanger is installed in front of the MiG-31 engine. Smooth surface in the air intake channel

why the engine skin look so strange?, it not smooth like other aircraft

 

[TSARb]

(D-30f6 on Su-47 Berkut)

Berkut uses the D-30F11 engine with a shorter afterburner, doesn't she? So, the nozzle may differ either.

 

[paralay]

D-30F6 is a unique engine for a unique aircraft

When designing the D-30F6, to increase thrust, a gas generator in the dimension of the D-30KU engine was adopted (without the first stage of the high–pressure compressor), and a low-pressure compressor from the D-30 engine with the addition of one stage ahead for an air consumption of 150 kg/s.

In the late 70s - 80s, on the basis of the D–30F-6, the D-30F-9 and D-30F-8 turbofan engines with increased thrust and shorter length were being worked out (they were not implemented).

Modifications

D-30F-6M (ed. 64, 1986) is a variant with increased thrust at altitude (takeoff thrust increased to 16,500 kgf) for the upgraded MiG-31M interceptor. At the end of the 80s, an experimental batch was released. MiG-31M aircraft with such engines have been undergoing flight tests since 1986.

D-30F-11 (ed. 70, 1997) – turbofan engine with increased thrust, reduced afterburner and jet nozzle length for the experimental C37-1 (Su-47) aircraft The Golden Eagle. The Su-47 aircraft with two such engines has been undergoing flight tests since 1997.

PS-30V-12 (ed. 75, 1988) is an afterburnerless high–altitude modification of the D-30F-6 with a take-off thrust of 5000 kgf for the M-55 high-altitude aircraft. In the late 80s and early 90s, an experimental batch was released. Since 1988, the engines have been operated on M-55 aircraft.

 

[paralay]

This is probably the most accurate drawing of the engine

 

[zjz]

Газотурбинные_двигатели_Иноземцев_А_А_,_Сандрацкий_В_Л_2006.djvu

Посмотреть и скачать с Яндекс Диска

disk.yandex.ru

Wow, thanks for the link.

I do not read Cyrillic, and I was wondering if there is an equivalent text on this topic written in English??

 

[paralay]

I'm afraid that's not possible. You can try to translate the document through a translator, but a lot of special terms and abbreviations will prevent you from understanding the meaning of the translated text

 

[zjz]

Thanks .... BTW, an observation : pages 8 to 11 (within the table of contents) are missing .....

 

[paralay]

Probably "top secret"