Merlin-1

[Salo]

http://nssphoenix.wordpress.com/2010/06/08/pintle-injector-rocket-engines/

Pintle Injector Rocket Engines[/size]

Posted by drdave on June 8, 2010

We have had several queries concerning "pintle injectors" (make sure you read the last paragraph of this post), as these are mentioned in the Space-X page on the Falcon 9, where it refers to the Merlin rocket engine and the "pintle style injector":

    The main engine, called Merlin 1C, was developed internally at Space-X, drawing upon a long heritage of space proven engines. The pintle style injector at the heart of Merlin 1C was first used in the Apollo Moon program for the Lunar Excursion Module (LEM) landing engine, one of the most critical phases of the mission.

Based on the queries and the Space-X information, we went sleuthing. First, we came across the fact that TRW built the LEM descent engine, which used the pintle injector. We ran across David Meerman Scott's blog apolloartifacts for a discussion and look at a model of the famous Lunar Module Descent Engine (LMDE). The engine was made famous by the Apollo 13 mission, where:  the Service Propulsion System (SPS) was never used subsequent to the cryotank stir/explosion. Because the extent of damage to the SPS was unknown, there was great concern at the time that collateral damage could have caused a catastrophic malfunction (if the engine was fired). Instead the LMDE was used for the return burn and subsequent course correction. Quite a famous engine.

In 2000, TRW demonstrated the TR-106 engine (pintle injector) using LOX / LH2 at NASA's John C. Stennis Space Center . The engine generated 650,000 pounds of thrust, more than the 400,000 pounds of thrust generated by the Space Shuttle Main Engine SSME. Al Frew, vice president and general manager, TRW Space & Technology Division stated: "Most engines are designed for maximum performance and minimum weight, but we deliberately set out to develop an engine that minimizes cost while retaining excellent performance. We believe this engine will cost 50 to 75 percent less than comparable liquid hydrogen boosters. By reducing engine costs, which make up almost half of the cost of a launch vehicle, we will reduce the cost of launch vehicles and access to space for government and commercial customers."

Despite the promise the motor demonstrated, NASA canceled further work.

The pintle injector engines have a long history in the former Soviet Union. The NK-33 was the successor to the NK-15 engines used in the failed Soviet N1 Moon launcher. NK-33 have been used with the Russian Proton launch system. An interesting discussion of the Soviet Moon rocket, its engines and the NK-33 successor can be found here, along with spectacular video of the launch and explosion. Orbital Sciences has now contracted with Aerojet (owner of the NK-33 engines) to finish developing and testing the NK-33 engines, now designated as AJ26-58 for the Taurus II.

Jonathon Goff, at Masten Space Systems, had a commentary at Selenianboondocks on the 2006 Space-X change from an ablative Merlin engine to a regenerative engine. Jon states that the "engine related problems are interrelated, and that they have to do with the combination of using a high chamber pressure engine design with a pintle-injector and an ablatively cooled chamber wall." That is, the flame produced by the cone of fuel and oxidizer hits the wall of the chamber and overheats the wall.

Included in the commentary is a simplified image of a pintle injector rocket engine, which illustrates the flow of liquid oxygen and fuel (RP-1 or liquid hydrogen) through the pintle injector into a cone shaped spray in the combustion chamber.

The replacement of the ablative chamber with a regenerative chamber eliminates the overheating.


Pintle Injector
Image Credit: Forschungsgruppe Alternative Raumfahrtkonzepte


Lunar Module Descent Engine. The business end of the LEM Descent engine, showing the Pintle Injector:
Image Credit: jurvetson on Flickr    


Merlin Engine with Pintle Injector. An image by Warren W. Thompson at the unveiling of Space-X's Falcon 1 at the Air & Space Museum on 4 December 2003.
Image Credit: Warren W. Thompson

Finally, while explaining the Pintle Injector to a friend, I realized that almost everybody who has a garden or tends a lawn has personal experience with pintles. You all use a nozzle on the end of your watering hose. Crank it down and you get a steady, narrow stream of water shooting out in a long arc. Crank it back the other way when you want to shut it off, and you get a wide, cone shaped fan spray. Now, turn off the water and look at the business end of the garden hose nozzle (please shut the water off first). There in the middle is a round pintle that moves back and forth as you crank the outer casing one way or the other. And the fan shaped spray of water with which you are familiar is what the fuel and oxidizer spray looks like inside the rocket engine. So take another look at the two images above and imagine the fan shaped spray. The only difference is that your spray of water doesn't explosively combust and throw a rocket into space.

[Salo]

Немного об одном из прародителей Мерлина  LMDE:

http://www.flickr.com/photos/jurvetson/4464220730/

The Eagle has Landed[/size]

I gasped when I looked out the window and saw this enormous crate at the front door.

This Apollo era rocket engine was originally designed for the Lunar Module Descent Engine (LMDE or DPS), and then employed as the 2nd-stage engine on the Delta Space Launch Vehicle. This may be the only complete unit in existence.

During the Apollo 13 emergency, the LMDE brought the spacecraft back to earth from the moon in an untested manner. Since an earlier explosion took out the main oxygen tanks, they improvised and used the LMDE— the engine from the lunar lander, designed to slow its decent to the moon — to instead push the crippled Command Module and reentry capsule for the return burn and subsequent course corrections. (Here are some Interesting details from Lovell and Haise)

This engine uses hypergolic fuels – noxious chemicals that combust on contact and make for simple, reliable engines that you can use repeatedly in the vacuum of space. The pump for these fuels on the LM ascent engine was the focal point of Neil Armstrong's nightmares before the Apollo 11 launch.

I'll post some unpacking photos and details as I gather them below.

Big thanks to Spaceaholic for the following details, engraved in a plaque that now hangs above it in our lobby:

"TRW TR-201 Bipropellant Rocket Engine. The thrust chamber was initially developed for the Apollo Lunar Module and was subsequently adopted for the Delta Expendable Launch Vehicle 2nd stage. The engine made 10 flights during the Apollo program and 77 during its Delta career between 1974-1988. This TR-201 has been configured as a fixed thrust version of the Lunar Module Descent Engine (LMDE) for Delta's stage 2. Multi start operation is adjustable up to 55.6 kN and propellant throughput up to 7,711 kg; and the engine can be adapted to optional expansion ratio nozzles. Development of the innovative thrust chamber and pintle design is credited to TRW Aerospace Engineer Dr. Peter Staudhammer.

The combustion chamber consists of an ablative-lined titanium alloy case to the 16:1 area ratio. Fabrication of the 6A1-4V alloy titanium case was accomplished by machining the chamber portion and the exit cone portion from forgings and welding them into one unit at the throat centerline. The ablative liner is fabricated in two segments and installed from either end. The shape of the nozzle extension (not installed on the example in this collection) is such that the ablative liner is retained in the exit cone during transportation, launch and boost. During engine firing, thrust loads force the exit cone liner against the case. The titanium head end assembly which contains the Pintle Injector and propellant valve subcomponents is attached with thirty-six A-286 steel

[Salo]

И наконец непосредственный предшественник Мерлина:

http://www.nasa.gov/centers/marshall/news/background/facts/fastrac.html

Fact sheet number: FS-1999-02-002-MSFC
Release date: 02/99

Fastrac Engine -- A Boost for Low-cost Space Launch[/size]

Photo description: Fastrac engine hot-fire test at Marshall Center.
Fastrac engine hot-fire test at Marshall Center (NASA/MSFC)
Engineers at NASA's Marshall Space Flight Center in Huntsville, Ala., are designing what may be one of the world's simplest turbopump rocket engines. It will be quite different from the Space Shuttle Main Engine, which was designed at Marshall in the 1970s and is considered by many to be the world's most sophisticated reusable rocket engine.

But rocket design for the 21st century breaks with aerospace Fastrac engine hot-fire test at Marshall Center tradition and embraces a new challenge to build engines that are cheaper and better.

The new Marshall-developed engine, called Fastrac, is - true to its name - on a fast track to propelling the next generation of launch vehicles. While some of the concepts for the Fastrac engine have been around for decades, actual technology development and design began in early 1996 and the engine's first flight is planned for late 1999 - a much faster-than-usual design cycle. For example, it took nine years to develop and design the Space Shuttle Main Engine.

Fastrac is only the second American-made engine of the 29 new rocket engines developed in the last 25 years. The simple, robust, easy-to-build engine is part of the Low Cost Technologies effort, one element of NASA's Advanced Space Transportation Program managed at Marshall. The program is paving the highway to space by developing technologies that will dramatically reduce the cost of getting to space.

Low Cost

Using the Fastrac engine as the propulsion system for a future launch vehicle is one means to achieve NASA's goal of making space launch affordable. Because much of the research and small payload market has been "locked out" of space by high launch costs, NASA is trying to open the door to cheaper space travel. The Fastrac provides 60,000 pounds of thrust to boost payloads weighing up to 500 pounds.

Each Fastrac engine will initially cost approximately $1.2 million - about one-fifth of the cost of similar engines. That price is expected to drop even more within a few years as design enhancements are discovered and incorporated through industry participation and flight experience. As improvements are made, the cost is expected to drop to $350,000 per engine. Fastrac incorporates drastic reductions in the cost of turbomachinery, which uses pumps to increase propellant pressure and a turbine to provide energy to turn the pump. Initially, each turbopump will cost about $300,000 - one-tenth of the average cost of a current rocket engine turbopump. That cost also is expected to drop to about $90,000.

Technology Development

Photo description: Marshall employees assemble Fastrac engine by mating the injector with the nozzle.
Marshall employees assemble Fastrac engine by mating the injector with the nozzle. (NASA/MSFC)

In a salient departure from traditional engine design, NASA and its business partners have adapted commercial, off-the-shelf technologies and common manufacturing methods to develop the Fastrac engine. Significant involvement by small business has aided in broadening the competition and producing lower cost hardware.

For example, Barber-Nichols, Inc. of Arvada, Colo., worked alongside Marshall engineers to design and manufacture the turbopump. The Colorado-based company is experienced in building turbomachinery for the automotive industry and chemical plants, and not traditionally associated with the aerospace industry. The company helped design a turbopump for the Fastrac engine that can be built easily using commercial manufacturing techniques.

The Fastrac engine is 7 feet long and 4 feet wide, and weighs almost 2,000 pounds.

How It Works

The engine is fueled by a mixture of liquid oxygen and kerosene, the same propellants used for the largest rocket engine ever built - the Saturn F1. Kerosene doesn't provide the same kick as hydrogen - which combines with liquid oxygen to fuel the Space Shuttle - but is cheaper and easier to handle and store.

The engine is started with a hypergolic igniter - a starter fluid that spontaneously ignites when oxygen is fed to the chamber. Once the kerosene is injected, the engine is running. The propellants are then supplied to the gas generator and thrust chamber assembly for mixing and burning.

The engine uses a gas generator cycle, which burns a small amount of kerosene and oxygen to provide gas to drive the turbine and then exhausts the spent fuel. That's the same cycle that was used on the Saturn rockets, but much more simplistic than the Shuttle engine system.

The Fastrac engine is a much simpler piece of machinery than previous American-made rocket engines. It has significantly fewer parts than engines that have driven other American spacecraft.

The reduced number of parts is a result of selecting technologies and design concepts that use simple manufacturing and assembly processes. For example, casting might be preferred over machining, because the latter method could require fabrication in several pieces due to shapes.

Chamber pressure is supplied by a single turbopump, unlike a Shuttle main engine which has four turbopumps. The Fastrac turbopump features only two pumps - one for fuel and one for liquid oxygen.

Another design feature that keeps the engine simple and inexpensive is its avionic - or electronic - control system. Typically a sophisticated, expensive part of a rocket engine, the avionics of the Fastrac engine are supplied from the vehicle's computer and are only used to open and close the valves. The thrust and mixture ratio is set during ground calibration. That's much simpler and cheaper than most rocket engine avionics, which continually modify the amount of propellants flowing into the chamber as changes in thrust are observed by on-board computers.

A typical rocket launch produces a temperature in the range of 5,500 degrees Fahrenheit, hot enough to melt almost any material. A common solution to keep the engine from overheating is regenerative cooling, which circulates liquid fuel around the engine chamber and nozzle through hundreds of feet of tediously welded tubing.

The Fastrac engine design avoids complex plumbing - opting instead to cool the chamber by charring or scorching its inside surface as the engine heats. The process is called ablative cooling. Layers of silica-phenolic composite material form a liner inside the chamber. The liner decomposes to prevent excessive heat build-up.

Nearly all of the engine's parts are reusable. The ablative chamber nozzle and the hypergolic ignition cartridge will be replaced after each flight. The chamber nozzle's protective, interior liner is damaged by intense heat inside the chamber. The hypergolic ignition cartridge must be refilled with propellant and replaced after each flight.

Schedule

The first vehicle scheduled to be powered by the Fastrac engine is the X-34, a technology testbed vehicle to demonstrate key vehicle and operational technologies applicable to future low-cost Reusable Launch Vehicles (RLVs). The engine is scheduled for delivery to the X-34 program in mid-1999 and its first flight is scheduled for late-1999.

The engine is now undergoing development and reliability testing. Individual components, such as the gas generator, turbopump assembly and thrust chamber assembly, are being tested at Marshall. In August 1998 the Marshall Center shipped the first complete engine system to NASA's Stennis Space Center in Mississippi with plans to conduct about 85 hot firings. These tests simulate launch of the X-34 vehicle and potentially a first stage booster, with the tanks and engine assembled.

The Fastrac engine is being designed at Marshall and built by NASA's industry partners, including several small businesses. Major subcontractors include Summa Technology Inc. of Huntsville, which builds components such as the gas generator, propellant lines,ducts and brackets; Allied Signal Inc. of Tempe, Ariz., and Marotta Scientific Controls Inc. of Montville, N.J., which supply valves; Barber-Nichols Inc., which builds the turbopump; and Thiokol Propulsion, a divison of Cordant Technologies Inc. of Salt Lake City, Utah, which builds the ablative chamber nozzle.


Fastrac engine hot-fire test at Marshall Center (NASA/MSFC)



Marshall employees assemble Fastrac engine by mating the injector with the nozzle. (NASA/MSFC)

[Salo]

http://techtran.msfc.nasa.gov/tech_ops/Fastrac_Engine.pdf

Fastrac Engine[/size]

The National Aeronautics and Space Administration (NASA) at Marshall Space Flight Center (MSFC) seeks qualified companies to further develop and commercialize the Fastrac turbopump rocket engine. The Fastrac engine can be built for less than $1 million using commercially available off-the-shelf (COTS) components and simplified manufacturing techniques. Fastrac provides 60,000 pounds of thrust and has many potential launch system applications. Having been tested successfully, the engine will propel NASA's X-34 flight demonstrator vehicle.

Market Possibilities

According to the Federal Aviation Administration report 1998 LEO Commercial Market Projections,approximately 400 to 500 launches are anticipated through the year 2010. MSFC's Fastrac engine could provide an alternative launch vehicle. Because Fastrac may provide a less expensive launch platform, market sizemay increase. The Fastrac engine can be used as a reusable launch vehicle thrust systemfor the 150 to 200 lower weight launches. Further developments will enable  Fastrac to beused as a nupper stage engine or scaled to accommodate larger size payloads.

Benefits - Low Cost of Manufacture

• Reduced complexity of engine design:
– Simple cycle: liquid oxygen rocket propellant, gas generator
– Simple control system: open loop sequencer
– Simplified geometry: easy to machine
– Fewer parts than previous American-made rockets.

• Use of commercially available off-the-shelf com-
ponents technology.

• Use of low-cost, high-performance materials.

The Technology

NASA's goal is to develop a launch infrastructure that reduces the cost-to-orbit of a pound of payload fromthe current $10,000 to $1,000. This goal has helped define the major attributes of a new generation of low-cost rocket engine technologies that are key components of the new MSFC Fastrac engine. The Fastrac engine is being designed to cost approximately $1 million, about one-fifth the cost of other engines of similar size and performance. The Fastrac engine provides 60,000 pounds of thrust and hasmany potential launch system applications.

The Fastrac engine uses a gas generator cycle to drive the turbine. A mixture of liquid oxygen and kerosene fuels the engine, which has significantly fewer parts than previous American-made rocket engines. The Fastrac engine is 7 feet long and 4 feet wide, and it weighs less
than 2,000 pounds. Among the innovative elements of the Fastrac engine are a new low-cost combustion chamber and a low-cost injector.

Combustion Chamber

The combustion chamber features an ablative cooling layer that decomposes as it absorbs the heat of combustion.The chamber is integratedwith themain nozzle assembly into a unitized structuremade of state-of-the-art ablative and refractory materials. High-performance
silicaphenolic tapemakesupthe ablative liner,whichis overwrappedwith graphite epoxy to formthe complete chamber/nozzle assembly. The ablative behavior of the liner is used to both cool and insulate the metal nozzle shell by resin boil-off and char layer formation. The ablative layer can be replaced after each flight of the engine's expected 7-launch life.

Injector

The new innovative injector design comprises only three parts: an injector core, a liquid oxygen dome close-out, and a faceplate. The geometry of each part is designed to be relatively easy to machine. The novel geometry of the injector core eliminates the need for fuel delivery manifolds largely due to the combination of a large monolithic faceplate with unique fuel channel patterns inthe core.This three-part assembly configuration greatly reduces the complexity of the overall injector and virtually eliminates the need for fuel downcomer holes.

Commercial Opportunities with NASA

Patent application shave been filed for three elements of the engine technology: (1) the rocket nozzle and combustion chamber structure, (2) the fuel injector, and (3) the combination of the combustion chamber/nozzle with the injector to form the thrust chamber assembly.
Commercial opportunities exist through licensing and cooperative development opportunities with NASA.

[Salo]

http://www.nasa.gov/centers/marshall/pdf/100364main_Fastrac_Tech_Brief1.pdf

[Salo]

http://www.barber-nichols.com/products/rocket_engine_turbopumps/default.asp

Fastrac LOX/RP-1 Turbopump

BNI teamed with NASA's Marshall Space Flight Center (MSFC) to design and build the turbopump for the Fastrac LOX/RP-1 Engine. The Fastrac Engine produces 60,000 pounds of thrust. Barber-Nichols consulted on the engine design and manufactured six turbopumps. The Fastrac is part of NASA's Low CostFastrac Hotfire Booster Technologies (LCBT) Program and the development of this engine initially cost approximately $1.2 million - about 1/5 the cost of a similar engine. A hot fire test of the Fastrac Engine was performed at NASA's Stennis Space Center in March 1999 and the first engine was installed on the X-34 A1 vehicle that was unveiled at NASA's Dryden Flight Research Center on April 30, 1999 .


 

Merlin LOX/RP-1 Turbopump

BNI designed and manufactured the Merlin Turbopump for the SpaceX Falcon Launch Vehicle. The Merlin Engine produces more than 100,000 pounds of thrust at sea level and the turbopump is the lightest in its thrust class. Barber-Nichols used its experience gained on the Fastrac and Bantam projects to rapidly develop the Merlin Turbopump. The first unit was delivered to the customer less than one year after design work began. BNI also used Design for Manufacturability principals to increase reliability and significantly cut costs. The Falcon Launch Vehicle had its first successful launch in March 2007.

[Salo]

Ещё фото LMDE:
http://www.apolloartifacts.com/page/2/

[Потусторонний]

Информация из новой презентации Спейсекса про Драгон
1st Stage Engines
Sea Level Thrust : 556 kN (125,000 lbf)
Vacuum Thrust: 617 kN (138,800 lbf)
Sea Level Isp: 275 s
2nd Stage Vacuum Engines
Vacuum Isp: 304 s
То есть повторили инфу про перспективный вариант М-1С :lol: из документа про Ф9(Блок2).

[Salo]

Повторюсь: не нашёл ни одного пресс-релиза в котором говорилось бы о ОСИ Merlin 1С с такими параметрами.

[Salo]

Статья по штифтовым форсункам:
http://www.lpre.de/resources/articles/TRW_PINTLE_ENGINE.pdf

[Salo]

Реферат статьи о самом мощном двигателе с штифтовой форсункой:
http://pdf.aiaa.org/preview/2001/PV2001_3987.pdf

ABSTRACT
Successful testing of the TRW Low Cost Pintle Engine (LCPE) Thrust Chamber Assembly (TCA) at the NASA Stennis Space Center E1 test facility has demonstrated the feasibility of this design as a solution to dramatically reduce the cost of access to space. Using the 650,000 Ibf thrust LOX/LH2 hardware developed for the EELV program, TRW performed this test program as a cooperative agreement with NASA-MSFC. The overall objective of this program was to demonstrate that a low cost pintle engine could substantially lower the costs of U.S. launch vehicles and thus enhance America's competitiveness in the international space transportation market.
The test article, consisting of a pressure fed, film-cooled ablative thrust chamber assembly, was tested at the NASA Stennis E1 test facility in Mississippi during the spring and summer of 2000. Low enthalpy hydrogen (64 - 78 °R, 675-1035 psia) and oxygen (180 - 186 °R, 848 - 1035 psia propellants were directly injected into the thrust chamber, pressure fed from high-pressure facility run tanks. The 650 Klbf TCA, scaled from a previously tested 40 Klbf TCA, demonstrated good performance, with a combustion efficiency of 96%, and total absence of combustion instabilities. TCA design chamber pressure was 700 psia, yielding 645,000 pounds thrust at sea level with a 3:1 expansion ratio test nozzle. The fixed-element injector based on TRW's unique coaxial pintle design was operated at 60% and 100% thrust levels by adjusting facility feed pressures.

BACKGROUND
Since the 1960s, TRW has been advocating a shift to engines with lower chamber pressures, simple cycles, low parts count and either pressure fed or pump fed propellant supplies as a way to lower engine production costs by over an order of magnitude (~$5/lbf)2. The largest of these earlier earth storable propellants burning at 300 psia in an ablative (DC-93-104) lined combustion chamber. In the 1970's, TRW tested similar engines with LOX/RP-1 and LOX/propane at 2,000 Ibf and 50,000 Ibf with combustion pressures of 300 psia. During the 1990's, with support from, McDonnell-Douglas, TRW tested 16,500 Ibf and 40,000 Ibf thrust LOX/LH2 engines at chamber pressures ranging from 280 psia to 380 psia. These tests demonstrated high performance (>95%) and stable operation. Several stability tests were conducted using 40 grain bombs. No combustion instabilities were detected in these tests or in any other test of a pintle injector engine.
In 1995-1996, a 650,000 Ibf thrust, LOX/LH2 scale-up of these early engines was designed and built during the first phase of the EELV program. This test engine, shown in Figure 1, represents a critical step in the development of the LCPE. The testing objectives were to verify engine performance, combustion stability and ablative life at a full-scale booster size. Under a Cooperative Agreement with NASA-MSFC, TRW tested the Thrust Chamber Assembly of the LCPE at NASA-SSC. The TCA, consisting of the injector and combustion chamber, demonstrated the capabilities of the pintle injector, ablative thrust chamber and nozzle in a pressure-fed configuration.

[SpaceR]

Информация из новой презентации Спейсекса про Драгон
1st Stage Engines
Sea Level Thrust : 556 kN (125,000 lbf)
Vacuum Thrust: 617 kN (138,800 lbf)
Sea Level Isp: 275 s
2nd Stage Vacuum Engines
Vacuum Isp: 304 s
То есть повторили инфу про перспективный вариант М-1С :lol: из документа про Ф9(Блок2).

Хм, небольшая неточность - параметров по 2й ступени здесь нет, оба значения УИ касаются только Мерлина-1С.

И небольшое расхождение в цифрах - если верить тяге в фунтах и вакуумному УИ, то земной должен быть 273,775 с.

[Salo]

Давление в КС:
Merlin 1A - 5.39 МПа;
Merlin 1С для F1 с тягой 78400 lbf на уровне моря  - 5.69 МПа;
Merlin 1C для F9 с тягой 95000 lbf на уровне моря  - 6.77 МПа.

Еще надо отметить что в документах SpaceX прописаны разрабатываемые версии на 115000, 125000 lbf (так называемый Merlin 1C+) и 140000 lbf (Merlin 1D)
Merlin 1C с тягой 115000 lbf упоминается в документах про Falcon 1e
Merlin 1C с тягой 125000 lbf упоминается в документах про Falcon 9 (Block II)
Merlin 1D с тягой 140000 lbf упоминается в документах про Falcon Heavy

Тяги на уровне моря. Из чего следует, что при неизменной геометрии КС и сопла мы получим примерно следующие давления в камере:

Merlin 1C с тягой 115000 lbf - 8,20 МПа;
Merlin 1C с тягой 125000 lbf - 8,91 МПа;
Merlin 1D с тягой 140000 lbf - 9,98 МПа.

Для сравнения у кислородно-керосинового РД-111 давление в КС было 7,84 Мпа,  а у вонючего РД-261 - 8,33 МПа.

Интересная картина: TRW довела давление в КС на водородном TR-106 до 50МПа.
SpaceX на Merlin 1C для F1 была вынуждена отказаться от абляционной камеры и перейти на регенеративное охлаждение.  Дальнейшее повышение давления в КС может привести к проблемам с высокой частотой. Да и насосы в ТНА одноступенчатыми уже не будут.

[Salo]

Глушко пытался бороться с высокой частотой на РД-111 улучшением смесеобразования при гладкой ФГ.  Результата он добился, но времени потратил очень много.
На РД-107А и РД-108А использовали перегородки из форсунок, чтобы разделить зону горения на части.

Что можно сделать в случае с штифтовой форсункой?

[Потусторонний]

Не подвержена ВЧ не в смысле абсолютно, а в существенно меньшей степени?

[Bell]

Тип форсунки не имеет значения.

[SpaceR]

Глушко пытался бороться с высокой частотой на РД-111 улучшением смесеобразования при гладкой ФГ.  Результата он добился, но времени потратил очень много.
На РД-107А и РД-108А использовали перегородки из форсунок, чтобы разделить зону горения на части.

Что можно сделать в случае с штифтовой форсункой?

Другой впрыск - другие особенности.
Возможно, могло бы помочь добавление завес на стенках КС (за счет асимметрии их струй).
Асимметрия струй на самой пинтл-форсунке, как я понимаю, уже давно используется.

[Uriy]

Глушко пытался бороться с высокой частотой на РД-111 улучшением смесеобразования при гладкой ФГ.  Результата он добился, но времени потратил очень много.
На РД-107А и РД-108А использовали перегородки из форсунок, чтобы разделить зону горения на части.

Что можно сделать в случае с штифтовой форсункой?

На штифтовой форсунке по-моему создаётся хорошая зона обратных токов,которая омывает наружную часть конуса распыла.За счёт этого и хорошая устойчивость.

[Salo]

Тип форсунки не имеет значения.

Как это, как это?

[Salo]

Глушко пытался бороться с высокой частотой на РД-111 улучшением смесеобразования при гладкой ФГ.  Результата он добился, но времени потратил очень много.
На РД-107А и РД-108А использовали перегородки из форсунок, чтобы разделить зону горения на части.

Что можно сделать в случае с штифтовой форсункой?

Другой впрыск - другие особенности.
Возможно, могло бы помочь добавление завес на стенках КС (за счет асимметрии их струй).
Асимметрия струй на самой пинтл-форсунке, как я понимаю, уже давно используется.

Если бы можно было добавить завесу на стенках, от абляционной камеры наверное не отказались бы.