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Shadowalker
Unpretentious Rocket Scientist
Reged: 11/23/04
Posts: 3554
Loc: Poplarville, MS, USA
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Occasionally the subject comes up in this forum so I thought I’d give some explanations of some rocket engine terms.
Most rocket engines make thrust by burning their propellants in a combustion chamber. The thrust out of the nozzle is a combination of the reactive mass (all the “stuff” coming out) and the pressure at the nozzle.
The propellants (fuel and oxidizer) must be injected into the combustion chamber. If the chamber pressure is 1000 pounds per square inch (psi), the propellants must be at a higher pressure to flow. There are two ways this can be accomplished in a liquid propellant engine: Pressure-fed or pump-fed.
In a pressure-fed engine the tanks are pressurized to the desired pressure with a reserve of some pressurent gas – helium is often used. The disadvantage is that the tanks must be able to withstand a higher pressure than a non-pressure-fed system. That makes them heavier. The advantage is a simpler engine.
In a pump-fed engine the propellants are pumped into the combustion chamber. A turbo-pump is usually used. The advantage is the requirements of the fuel and oxidizer tanks are less (less pressure in the tanks). The disadvantage is that the engine is more complex and costly.
There are two ways to pump the propellants: Gas generator and staged combustion.
First the gas generator. A gas generator is a small rocket engine that uses a portion of the propellants to drive a turbine. The turbine shaft is coupled to pumps that pump the liquid fuel and oxidizer into the combustion chamber. The turbine exhaust is dumped overboard. Sometimes the turbine exhaust is used for engine vectoring or roll control for the vehicle. RS-68 works this way. The disadvantage is the turbine gas provides no propulsion. As such it’s less efficient than staged combustion. The advantage is that the engine is simpler and the engine pressures are lower.
In a staged combustion engine there are gas generators but they are called “pre-burners.” In the case of the Space Shuttle main engine (SSME) there are two pre-burners. In both all of the fuel propellant (hydrogen) and a small amount of the oxidizer is enjected into both pre-burners. All of the oxidizer is consumed. What is left is a small amount of steam and a large amount of high temperature, high pressure hydrogen. This drives the turbine for the main pumps – One pre-burner drives the hydrogen pump and the other drives the oxygen pump. The turbine outlet (hot hydrogen) flows directly into the combustion chamber, where it is mixed with the output of the oxygen pump for combustion.
So in a staged combustion engine all propellants contribute to thrust. None is dumped overboard. The disadvantage is that the pressures are much higher. The pre-burners must operate at a very high pressure to drive the turbines and then to have enough pressure left over to make its way into the main combustion chamber. The advantage is much higher efficiency than a gas-generator engine.
In the case of SSME there’s another difficulty: The oxygen pump. As I said before, the turbine is driven by hot high pressure hydrogen. This turbine is driving a pump carrying liquid oxygen. Not a good mix! So the seals are extremely important. We don’t want that to mix. The Russians solved that problem by using an oxygen rich pre-burner for the LOX pump side. Instead of running that pre-burner fuel-rich they run it oxygen rich. So the turbine is driven by hot oxygen instead of hot hydrogen. This makes the engine less hazardous to operate. Oh, and the fuel in most Russian engines is RP-1 (Rocket Propellant 1), not hydrogen, but the principle is the same. The RD180 used by the Atlas 5 is such an engine.
From another post, here’s my explanation of Specific Impulse (Isp):
Specific impulse is the measure of efficiency for rocket engines. It's the thrust of the engine divided by the mass flow rate of propellants. Units are in seconds. If an engine produced 1000 pounds of thrust and consumed 10 pounds of propellant per second, the specific impulse would be 1000 divided by 10 or 100 seconds.
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Tom Nicolaides
http://www.first-light.org
My evil self is at that door, and I have no power to stop it
-- Dr. Edward Morbius
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llanitedave
Humble Megalomaniac
Reged: 09/26/05
Posts: 12941
Loc: Amargosa Valley, NV, USA
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That's fascinating information about the pumps, Tom. I had no idea how they were run. That's just one more item in all the tradeoffs and compromises that engineers face when trying to get the best performance out of a system.
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"Since the process of science generates more mysteries than it solves, I predict that we'll never learn everything: and we'll continue to generate new ignorance at the speed of knowledge."
"S.O.E." (Sauron's Other Eye), with 16" Royce conical mirror: A permanent work in progress.
10" Homebuilt dob, old Coulter mirror
Under Construction: The "Eye of Sauron" Observatory!
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Qkslvr
Pooh-Bah
Reged: 06/23/06
Posts: 1316
Loc: NE Ohio, US
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Thanks Tom, I too had wondered how the turbo pumps worked.
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Mike
Onyx 80ED/N8/CG-5/40D
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groz
Pooh-Bah
Reged: 03/14/07
Posts: 1078
Loc: Duncan, BC
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ok, that was rockets 101, let me add the -102 talk 
The thrust of a rocket engine comes specifically from accelerating the hot gasses out the exhaust nozzle. The concept is basically simplicity in itself. Hot gases are generated in the combustion chamber, and fed into the nozzle. The nozzle itself is essentially a venturi. A venturi system has a property that the rockets capitalize on. Slow moving gasses will accelerate thru the exit, but will reach a 'terminal velocity' when the venturi becomes 'choked', and thats when shock waves have developed at the point of smallest diameter, and gasses flowing at this point will reach sonic velocity, but no higher, they are restricted by the shock waves.
The neat thing about a venturi running 'choked' is this, as the gasses pass thru the shock waves at the choke point, and into the expansion chamber, they will accelerate beyond sonic velocity during the expansion. How much they accelerate is a function of a bunch of things, but, the two that are most important, the back pressure (ambient air) behind the nozzle, and the shape of the nozzle itself.
So, the predominant factors in determining how much thrust can come from a rocket run along these lines (very simplified). Combustion in the chamber must produce pressures high enough that the choke point in the venturi remains fully choked. Exactly how much exhaust gas can escape the chamber is a function of the exit diameter, and the sonic velocity for that gas at the prevailing temperature pressure. How much it accelerates is a function of nozzle shape, and back pressure. All of these combine to produce the mass flow of gas, and it's acceleration. Mass flow + acceleration + simple newtonian physics will translate into a number to quantify the thrust produced.
The important things to keeping a rocket running correctly, pressure in the combustion chamber must be kept high enough to maintain 'fully choked' at the chamber exit, ie entrance to the expansion section. From there it's simple newtonian physics combined with not so simple dynamics of high pressure / high temperature gas flow.
The other _slightly important_ detail is that back pressure mentioned above. A nozzle operating in vaccuum has no back pressure, while one operating at sea level has 14.7 psi of back pressure. That makes a HUGE difference in the design of a nozzle to maximize benefit from a given quantity of fuel, and, is the real reason rockets are staged, with one or two stages built for the lift with atmospheric back pressure, and, later stages optimized for zero back pressure in a vaccuum. For reference, just google up some old photos of the saturn V launch stack, compare the shape of the nozzles on the first lifting stage, with that of the nozzles on the third stage use for the trans lunar boost phase. They are shaped dramatically differently, the 5 nozzles in the first stage were optimized for significant atmospheric back pressure, then stage 2 was set up for little/no back pressure, while the third stage and command module nozzles were set up for hard vaccuum.
A lot of the launch stacks today utilize solid boosters for the 'heavy lift' of the initial launch, and any rockets actually burning at that point, have nozzles optimized either for vaccuum, or, close to it. the solid boosters strapped on the sides are doing the heavy lift for getting up out of the 'sticky' part of the atmosphere.
I've always wondered about the ssme nozzles, they have a shape that suggests they are optimized for at least _some_ back pressure, but, I've never found any reference to state just how much ? Are they operating optimally at the initial lift in the lower atmosphere, or, are they operating optimally during the high atmosphere portion of the flight ?
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Shadowalker
Unpretentious Rocket Scientist
Reged: 11/23/04
Posts: 3554
Loc: Poplarville, MS, USA
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Quote:
...Slow moving gasses will accelerate thru the exit, but will reach a 'terminal velocity' when the venturi becomes 'choked'...
...I've always wondered about the ssme nozzles, they have a shape that suggests they are optimized for at least _some_ back pressure, but, I've never found any reference to state just how much ? Are they operating optimally at the initial lift in the lower atmosphere, or, are they operating optimally during the high atmosphere portion of the flight ?
Choked flow through a venturi! Indeed. We also use that property of fluids to regulate fluid flow. Many rocket engine components require a fixed flow rate. To test them we often use a venturi designed to choke at the desired flow rate. Works pretty good. Of course I'm just a dumb old electrical engineer and don't really understand these fluid dynamics things 
You're right about SSME being designed for sea level AND high altitude operation. At 100 percent thrust the nozzle works at sea level, allowing gas expansion to work properly. At thrust levels much below 80% the flame will separate from the surface of the nozzle and cause instability that could and probably would destroy the engine.
At high altitude the back pressure is low enough to prevent this at throttle down. In fact during the altitude of "max Q" (maximum dynamic pressure), the shuttle engines throttle down so the vehicle moves through that altitude more slowly. But by that altitude the atmospheric pressure is low enough.
To test SSMEs at sea level (which is what we do here), we usually run them at just over 100% thrust. To throttle we need to use a diffuser, which is essentially a nozzle extension allowing throttling to about 60% thrust.
I believe SSME is optimized for high altitude flight.
A bit of trivia: In the early 90s we had a project called ASRM - Advanced Solid Rocket Motor. This was designed to correct the flaws in the SRMs (that lead to the Challenger disaster). The plan was to build the correct thrust profile into the ASRM such that the SSMEs would not need to throttle down. So the ASRM would do the throttle down. We built half a rocket factory in northern Mississippi and a complete static test stand in southern Mississippi before that project was cancelled. *sigh*
Back on the altitude stuff. The Saturn upper stages used the J2 engine. The J2 engine has been selected to propel the upper stage of Ares 1 and Ares 5 (also serve as an earth departure stage for the latter). We're building a new test stand to test the new version of the J2, now known as J2X. To test at simulated altitude we're building a giant tube that has a vacuum. The vacuum is generated using chemical steam generators to suck the air out of the long tube (again using the venturi effect). Not really sure how that works. Again, I'm just a dumb old electrical type.
My office sits in the test stand that tests the RS68. It's tested at sea level. They don't attach the nozzle for these tests. This test stand also tested the S1C (Saturn 5 1st stage), the Space Shuttle Main Propulsion test article (a cluster of three SSMEs), single SSMEs, the Delta 4 1st stage and is slated to test the Ares 1 and Ares 5 upper stages.
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Tom Nicolaides
http://www.first-light.org
My evil self is at that door, and I have no power to stop it
-- Dr. Edward Morbius
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imjeffp
Aluminum Falcon
Reged: 09/30/03
Posts: 4902
Loc: Cedar Park, Texas
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Tom--do you think J2-X is really gonna happen?
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Shadowalker
Unpretentious Rocket Scientist
Reged: 11/23/04
Posts: 3554
Loc: Poplarville, MS, USA
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Quote:
Tom--do you think J2-X is really gonna happen?
Yes, I think it is. We've put up the steel for the A3 test stand - a 250 Megabuck project. We've been doing tests on the powerhead (gas generators and turbopumps). Regardless of what manned access to space looks like it will need a robust upper stage. The dual design of 2nd stage for Ares 1 and Ares 5 and earth departure stage helps too.
I'll post a couple of photos of the test stands here.
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Tom Nicolaides
http://www.first-light.org
My evil self is at that door, and I have no power to stop it
-- Dr. Edward Morbius
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Shadowalker
Unpretentious Rocket Scientist
Reged: 11/23/04
Posts: 3554
Loc: Poplarville, MS, USA
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A test of the RS68 on the B1 test stand. B1 is the left side of the stand. The right side is B2, where the S1C , Shuttle MPTA and the Delta 4 was tested. Ares upper stage is scheduled to go into B2. B1 will continue with RS68.
This is about T + 1 second. Photo taken by me with my Palm pilot from about 1500 feet away on top of the Test Control Center.
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Tom Nicolaides
http://www.first-light.org
My evil self is at that door, and I have no power to stop it
-- Dr. Edward Morbius
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llanitedave
Humble Megalomaniac
Reged: 09/26/05
Posts: 12941
Loc: Amargosa Valley, NV, USA
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So what about the idea of a variable-geometry nozzle that can be adjusted for both atmospheric and vacuum operations? I know jet fighters can vary their exhaust geometry -- is the issue for rockets weight or exhaust temperatures, or simply added expense?
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"Since the process of science generates more mysteries than it solves, I predict that we'll never learn everything: and we'll continue to generate new ignorance at the speed of knowledge."
"S.O.E." (Sauron's Other Eye), with 16" Royce conical mirror: A permanent work in progress.
10" Homebuilt dob, old Coulter mirror
Under Construction: The "Eye of Sauron" Observatory!
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Shadowalker
Unpretentious Rocket Scientist
Reged: 11/23/04
Posts: 3554
Loc: Poplarville, MS, USA
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Well, yes. There is such an engine design and we've tested it It's called the Linear Aerospike Engine and the design has been around since the 1960s.
In this Wikipedia Article you can see the basic design and read about the theory. The photo is the engine we tested on the A1 test stand some years ago. With a J2 powerhead (gas generator and turbo pumps), it was supposed to go in the X33 Single Stage to Orbit vehicle. Yet another cancelled program.
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Tom Nicolaides
http://www.first-light.org
My evil self is at that door, and I have no power to stop it
-- Dr. Edward Morbius
Edited by Shadowalker (06/09/09 12:18 PM)
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Joseph Gillman
Carpal Tunnel
Reged: 12/06/05
Posts: 2784
Loc: Aston, PA
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Thanks for the explanations.
As for specific impulse it always bugged me that the units were seconds.
isn't thrust in pounds-FORCE while the mass flow rate of propellants in pound-MASS per second? Those units don't directly cancel, what am I missing?
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Shadowalker
Unpretentious Rocket Scientist
Reged: 11/23/04
Posts: 3554
Loc: Poplarville, MS, USA
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Quote:
As for specific impulse it always bugged me that the units were seconds.
isn't thrust in pounds-FORCE while the mass flow rate of propellants in pound-MASS per second? Those units don't directly cancel, what am I missing?
Absolutely correct. Thrust is in pounds-force. Mass is in pounds mass. They play a little loose with the units. Technically thrust should be in pounds and mass in either pounds-mass or slugs. Which would give Isp in units of pound-seconds per pound-mass or pound-seconds per slug. Or in metric units, newton-seconds per kilogram.
Those units probably reduce further. Lemme see. Let's stick with metric. A newton is a kilogram-meter per second squared. So a newton-second per kilogram would reduce to meters per second. Hmmmm same as speed. That doesn't make sense. But that's what the units come to unless I'm wrong
Still, it works. And everyone knows what we're talking about. Even if we don't sometimes.
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Tom Nicolaides
http://www.first-light.org
My evil self is at that door, and I have no power to stop it
-- Dr. Edward Morbius
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Shadowalker
Unpretentious Rocket Scientist
Reged: 11/23/04
Posts: 3554
Loc: Poplarville, MS, USA
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Here are just a few:
Operational rocket engines, Thrust, Isp, Cycle, Country
F1, 1.5M lbs, 260, LOX-RP-1, GG, US
J2, 230K lbs, 419, LOX-Hydrogen, GG, US
SSME, 450K lbs, 450, LOX-Hydrogen, SC, US
RS-68, 700K lbs, 410, LOX-Hydrogen, GG, US
RS-27, 235K lbs, 302, LOX-RP-1, GG, US
Merlin, 138K lbs, 303, LOX-RP-1, GG, US
NK33/AJ26, 394 K lbs, 334, LOX-RP-1, SC, Russia
RD253, 400K lbs, 316, Hydrazine, SC, Russia
RD-180, 933K lbs, 338, LOX-RP-1, SC, Russia
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Tom Nicolaides
http://www.first-light.org
My evil self is at that door, and I have no power to stop it
-- Dr. Edward Morbius
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Qkslvr
Pooh-Bah
Reged: 06/23/06
Posts: 1316
Loc: NE Ohio, US
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Just for grins, you should add New Horizon's ion engine ISP to you list....
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Mike
Onyx 80ED/N8/CG-5/40D
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JAT Observatory
Got Cookies?
Reged: 02/20/05
Posts: 6794
Loc: Eastern PA
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One of the reasons Isp is used as it allows the OA to keep track of the propellant budget by counting the truster on pulses. Between counting the thruster on pulses or the total thruster on time you can calculate the amount of fuel used during a maneuver.
Provided the REA is operating within known efficiencies and you know fuel tank pressures, fuel tank and fuel line temps along with the spacecraft attitude and CG you can calculate your Delta V based on the total number of thruster pulses or total thruster on times.
It should be noted that a 5 minute continuous burn will produce a different Delta V than let's say 300 one second on pulses at a 100% duty cycle.
If you count individual thruster pulses in an attempt to calculate total thruster on time, you just have to be sure you model the ramp up and ramp down of the thruster so you get a real account of the actual thruster on time during that 1 second thruster pulse.
Because of the ramp up and ramp down time the actual on pulse time at the effective efficiency will be something less than 1 second.
There are a number of other very interesting spacecraft related topics we could discuss here. Things like attitude control systems. They consist of a wide range designs that employ devices such as thrusters, momentum wheels, magnetic torquers, solar trim tabs (solar sails), sun sensors, earth sensors, gyros, star trackers and devices. One of the keys to telling a spacecraft where to go is you have to know where it is are first.
Deployment mechanisms is another one. Things like pyros, memory metals and my favorite..frangibolts.
Spacecraft, manned or unmanned have become so common that people more or less take their launch and operation for granted. But the environment they have to operate in is far from kind. I still think the design and engineering of them is fascinating.
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-Marcus
The problem with free speech is even the stupid have a voice.
http://www.jatobservatory.org
Currently a SV102A, WO 66mm & 10" LX200GPS OTA on a Paramount ME
(these OTAs subject to change without notice)
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Shadowalker
Unpretentious Rocket Scientist
Reged: 11/23/04
Posts: 3554
Loc: Poplarville, MS, USA
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Quote:
Just for grins, you should add New Horizon's ion engine ISP to you list....
Actually, I think New Horizons is on a free-fall trajectory- No in-transit propulsion. It does have some hydrazine thrusters for attitude control, though.
Deep Space 1 - the comet explorer has ion propulsion. It was launched over 10 years ago and is no longer in service. The in-space engine was a Hughes/Boeing built xenon reactive mass ion propulsion engine. It produced about 50 mili-Newtons of thrust with an Isp of about 3100 seconds - 10 times more than traditional chemical propulsion. 50 mili-newtons is about 12 mili-pounds of thrust. Two tenths of an ounce of thrust. Not much But then it could operate for days or weeks. That makes up for a lot! Read about it here.
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Tom Nicolaides
http://www.first-light.org
My evil self is at that door, and I have no power to stop it
-- Dr. Edward Morbius
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Shadowalker
Unpretentious Rocket Scientist
Reged: 11/23/04
Posts: 3554
Loc: Poplarville, MS, USA
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Quote:
There are a number of other very interesting spacecraft related topics we could discuss here. Things like attitude control systems. They consist of a wide range designs that employ devices such as thrusters, momentum wheels, magnetic torquers, solar trim tabs (solar sails), sun sensors, earth sensors, gyros, star trackers and devices. One of the keys to telling a spacecraft where to go is you have to know where it is are first.
Deployment mechanisms is another one. Things like pyros, memory metals and my favorite..frangibolts.
Spacecraft, manned or unmanned have become so common that people more or less take their launch and operation for granted. But the environment they have to operate in is far from kind. I still think the design and engineering of them is fascinating.
True. To that list I'd add propellants and ignition systems too. LOX-Hydrogen, LOX-RP-1 are by no means the only fuel/oxydizers out there. Another class of propellants are storable. The disadvantage of LOX/Hydrogen is that they need to be kept at very low temperatures: -423 Deg F for LH and -300 Deg F for LOX. RP-1 is a petrolium distillate and is liquid at normal temperatures. There are others including Hydrazine - a particularly nasty chemical, it's storable at room temperature and is commonly used for propulsion. Another is Hydrogen Peroxide. Also nasty stuff.
Then there's ignition. Some engines use spark plugs. Others use detonators. Some use a small amount of a hypergolic chemical - chemicals that when mixed catch on fire. Some engines use hypergolic propellants. Open the valves and the engine comes on - a model of simplicity. The engines on the Apollo lunar lander used such engines. When you're in space that engine better come on when you need it to.
Solid rocket engines. Most use powdered aluminum as the fuel and ammonium perchlorate as the oxidizer. Efficiency is about the same as LOX-RP.
Hybrid rocket motors - Uses a solid fuel and liquid oxidizer.
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Tom Nicolaides
http://www.first-light.org
My evil self is at that door, and I have no power to stop it
-- Dr. Edward Morbius
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gazerjim
Postmaster
Reged: 02/12/04
Posts: 8166
Loc: n-dimensional space
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I was a rocket nut as a kid and young man. I still find them fascinating.
I've heard that the number of mach diamonds visible in the rocket exhaust is a function of the mach number of the exhaust velocity. Is this true? What produces mach diamonds to begin with?
I used to fiddle with building small solid fuel AP/HTPB motors. I chickened out (common sense) at around 2 ounce total grain size, so never made anything large.
One of the problems plaguing my small motors was "chuffing", and you probably recognize that if I am using correct terminology. The problem was largely cured by the addition of powdered Al. Even though my motors were too small for the Al to up the Isp before exiting the nozzle.
Thanks for all of the fascinating info.
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Jim Fisher
"How wonderful that we have met with a paradox. Now we have some hope of making progress."
Niels Bohr
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astrotrf
professor emeritus
Reged: 09/30/07
Posts: 709
Loc: Rodeo, NM
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Quote:
Quote:
Just for grins, you should add New Horizon's ion engine ISP to you list....
Actually, I think New Horizons is on a free-fall trajectory- No in-transit propulsion. It does have some hydrazine thrusters for attitude control, though.
Qkslvr was probably thinking of Dawn, which coincidentally re-ignited its ion engine just the day before he posted.
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Terry (astrotrf)
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Qkslvr
Pooh-Bah
Reged: 06/23/06
Posts: 1316
Loc: NE Ohio, US
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Thanks Tom!
10x chemical rockets, now we might be getting somewhere, literally!
One last question, When they start the ssme, it looks like they start it fuel rich, then up lean the mixture out, correct?
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Mike
Onyx 80ED/N8/CG-5/40D
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