Author Topic: Rocket Engine Q&A  (Read 382958 times)

Offline msat

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Re: Rocket Engine Q&A
« Reply #680 on: 07/14/2015 11:08 pm »
For the small rockets Armadillo and Masten have flown, I don't think turbomachinery would be an option. Though there may be a particular size of rocket where COTS turbocharger turbines could possibly be used, but designing the impellers might be tricky. I suppose that could be contracted out to an experienced company. The heat exchanger could "simply" be a coil running through the bottom of the fuel or oxidizer tank (have to make sure the coolant can't possibly freeze - another point in favor of propane). 

One thing I've been thinking about is the upper limit of thrust an expander cycle engine could generate without making the combustion chamber really long. I wonder if an aerospike nozzle would be a better candidate as the nozzle surface is supposedly exposed to a higher heat flux than a bell (though I'd imagine the total surface area is actually smaller). I don't actually understand it, but that's what all the pros say, so who am I to argue. Anyway, the aerospike may provide the much needed heat to drive a more powerful expander cycle pump. Add in the benefit of reasonably good altitude compensation, and it seems like a win. 

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Re: Rocket Engine Q&A
« Reply #681 on: 07/14/2015 11:26 pm »
For the small rockets Armadillo and Masten have flown, I don't think turbomachinery would be an option.

So far they've stuck with pressure-fed systems, but I suspect they'd love to move beyond that. Steam turbines are an old technology and probably easier than gas generator systems, although CS is moving straight to a gas generator system. Electric pumps could also be an option.
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Offline msat

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Re: Rocket Engine Q&A
« Reply #682 on: 07/15/2015 12:06 am »
A steam turbine shouldn't inherently be different than any other kind of turbine for a given size and pressure ratio. The difference would be in the materials used to accommodate for the chemical makeup and temperature of the gas.   

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Re: Rocket Engine Q&A
« Reply #683 on: 07/15/2015 07:23 am »
The difference would be in the materials used to accommodate for the chemical makeup and temperature of the gas.

Yeah, that's what I meant. You also don't need to design a gas generator. Not show stoppers, I just imagine it make things easier while you're still learning how to do it.
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Offline john smith 19

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Re: Rocket Engine Q&A
« Reply #684 on: 07/15/2015 08:39 am »
A steam turbine shouldn't inherently be different than any other kind of turbine for a given size and pressure ratio. The difference would be in the materials used to accommodate for the chemical makeup and temperature of the gas.
You're right. In fact the first steam turbines were running in the 1900's, the first gas turbines in the 1930's.

IIRC  the team under Eugene Sanger in Germany in WWII planned a steam turbine drive for the turbo pumps for the "Silver Bird" concept. I'm not sure if they actually built the whole engine (they seem to have managed a 100 tonne combustion chamber) at least.

There is a translation of their report done by IIRC the US Navy, complete with (very bad) photographs.

Note that the water used on steam plants is not straight from the tap water. For this application you'd probably want to go with deionized water with no scale forming impurities in it running in a sealed system.

Note that the work of the Whitehead group at LLNL indicates turbines scale down badly below about 5000 lb thrust. The turbines are small, which means the clearances are very tight and the surface finish has to be very good, as the boundary layer will be proportionately thinner.

OTOH a positive displacement reciprocating  pump using say corrugated pistons would be easier to fabricate, eliminate rotating seals and bearings and deliver complete separation of propellant from driving fluid. Also the pump drive fluid is relatively inert (super heated steam should be treated with considerable respect.  :( ).

The problem with all separate pump driver fluid concepts is now you need two heat exchangers, the CC wall and whatever you're dumping the waste head the was not used in the pump drives. This might not be the problem people think it is if the engine stays in the atmosphere. Different goals, different trades.

In vacuum you're looking at either radiators (which get big at low temperature differences, even with a carbon fibre "carpet" around them to enhance emission) or dumping it to the propellant flow IE a 2nd HX.

Historically people have said "No, weight penalty not worth the flexibility of being able to select optimum drive fluid properties"

OTOH amateur developers, or people wanting lower maintenance reusability might prioritize things differently.
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Offline msat

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Re: Rocket Engine Q&A
« Reply #685 on: 07/15/2015 06:54 pm »

Yeah, that's what I meant. You also don't need to design a gas generator. Not show stoppers, I just imagine it make things easier while you're still learning how to do it.

Possibly. I couldn't say with any kind of certainty. I will however counter Jim's assertion that TCF is too complicated. Perhaps more complex than a "standard" expander like the RL10, but not more so than the gas generator cycle and the variations thereof.




Note that the water used on steam plants is not straight from the tap water. For this application you'd probably want to go with deionized water with no scale forming impurities in it running in a sealed system.

Note that the work of the Whitehead group at LLNL indicates turbines scale down badly below about 5000 lb thrust. The turbines are small, which means the clearances are very tight and the surface finish has to be very good, as the boundary layer will be proportionately thinner.

OTOH a positive displacement reciprocating  pump using say corrugated pistons would be easier to fabricate, eliminate rotating seals and bearings and deliver complete separation of propellant from driving fluid. Also the pump drive fluid is relatively inert (super heated steam should be treated with considerable respect.  :( ).

The problem with all separate pump driver fluid concepts is now you need two heat exchangers, the CC wall and whatever you're dumping the waste head the was not used in the pump drives. This might not be the problem people think it is if the engine stays in the atmosphere. Different goals, different trades.

In vacuum you're looking at either radiators (which get big at low temperature differences, even with a carbon fibre "carpet" around them to enhance emission) or dumping it to the propellant flow IE a 2nd HX.

Historically people have said "No, weight penalty not worth the flexibility of being able to select optimum drive fluid properties"

OTOH amateur developers, or people wanting lower maintenance reusability might prioritize things differently.


Of course you wouldn't fill up the rocket with water straight from the river! Distilled, and maybe deionized for sure. You don't want crud buildup or all the various other things that could happen to occur. Eliminating impurities in a third coolant fluid is no less important than doing it for your propellant.

I know that jet engines become less and less efficient as they're scaled down for numerous reasons. The thing that surprises me is the amount of power rocket engine turbines generate relative to their size. Any inefficiencies may be acceptable in light of the other options. But when it comes to rockets, there seems to be a point where active pumps just aren't worthwhile, and pressure fed is more suitable. Out of curiosity, I looked up some turbine efficiency figures which are annoyingly hard to come across: for turbochargers, I found that one of Garrett's largest units has a turbine efficiency of 82%, while GE's GE90 turbofan is 93%. I couldn't find anything about the F1, but if I'm not mistaking, according to http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19850010710.pdf the RL10 turbine has an efficiency of 95%. I wouldn't rule out turbocharger turbines for the next step up from pressure-fed designs in amateur rocketry.

Positive displacement pumps are an option, as Xcor has proven, but they pose some of their own challenges. Pumping isn't steady state, even with multiple pistons. If you're pumping gasses, this isn't much of an issue as you can feed it to a pressure vessel and maintain output with a pressure regulator (at the expense of reduced output pressure). You can do something similar with liquids I suppose, but it's a bit trickier. I'm not entirely convinced that the challenges and potential mass penalties outweigh the use of turbopumps. Even if it's easier to fabricate, it's still more complex.

In the case of the Third Fluid Coolant concept, the idea is to dump the heat into the propellants rather than the atmosphere or space. It's certainly the most compact option. Plus it's desirable to pre-heat the propellants in certain cases.

I agree that maximal efficiency isn't always the be-all-end-all of rocketry, this is especially true of teams with limited experience and budgets. There's so many possible solutions to a given problem, with some of those solutions being better for a team than others. I do appreciate how some very clever solutions are born out of necessity due to limited budgets. 
« Last Edit: 07/15/2015 06:57 pm by msat »

Offline msat

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Re: Rocket Engine Q&A
« Reply #686 on: 07/16/2015 07:27 pm »
I had a thought about a variation of the TFC concept. Well, it's not actually TFC, but it is an expander cycle and uses heat exchangers like one. The fuel should be suitable for an expander cycle and probably subcooled (such as propane at LOX temps). The idea is that the fuel exiting the compressor is split off with one path going to the combustion chamber and the other going to the cooling channels. After the cooling channels, it is expanded and drives the turbine and then routed through a heat exchanger (possibly integrated as part of the injector head?) where it's condensed and then dumped back somewhere upstream of the compressor inlet.

Pros:
1) No need for an additional compressor as a standard TFC requires
2) Higher pressure drop across the turbine like a TFC than possible with a standard expander cycle

Cons:
1) More limited in the fuels that can be used in comparison to a TFC. No H2 either, probably
2) There will be some pressure losses for the propellants entering the combustion chamber as they have to pass through a heat exchanger like a TFC.


It would seem that the most important factor in making the TFC or variants practical is the efficiency and weight of the necessary heat exchangers. I tried finding figures that might give an indication of what's possible (even looking for data on power plants) but came up empty handed. 

Offline Jim

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Re: Rocket Engine Q&A
« Reply #687 on: 07/16/2015 07:32 pm »
I still have to ask why?   Your list of cons is missing many items.

Offline msat

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Re: Rocket Engine Q&A
« Reply #688 on: 07/16/2015 08:36 pm »
I still have to ask why?   Your list of cons is missing many items.

Then help populate it.

Why? Why do anything? Why make rocket engines or anything at all? If there was a single type of engine that is better than every other one in every way, then why are there so many different ones with various different cycles currently in use? It would seem that by your logic, there's only one metric of importance.

I didn't think I needed to state some of the more obvious pros, but I will:

0) No secondary combustion chambers
1) Closed cycle
2) Higher chamber pressure than a standard expander
3) Lower temps and higher pressure drops across turbines, and possibly higher chamber pressures compared to the various staged combustion cycles (not to mention less corrosive than the hot oxidizer rich ones) while being no more complex, and likely simpler.

Those sound like some decent pros to me.

Offline Jim

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Re: Rocket Engine Q&A
« Reply #689 on: 07/16/2015 09:14 pm »
1.  If there was a single type of engine that is better than every other one in every way, then why are there so many different ones with various different cycles currently in use? It would seem that by your logic, there's only one metric of importance.

I didn't think I needed to state some of the more obvious pros, but I will:

0) No secondary combustion chambers
1) Closed cycle
2) Higher chamber pressure than a standard expander
3) Lower temps and higher pressure drops across turbines, and possibly higher chamber pressures compared to the various staged combustion cycles (not to mention less corrosive than the hot oxidizer rich ones) while being no more complex, and likely simpler.

Those sound like some decent pros to me.

1. All the goods ones are in use or have been used.  This one is on sideline because it isn't good.
0) why is that good?

The complexity, extra mass, and losses in the heat exchangers are greater drawbacks.  Not to mention the inability to start such an engine. 
« Last Edit: 07/16/2015 09:15 pm by Jim »

Offline msat

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Re: Rocket Engine Q&A
« Reply #690 on: 07/16/2015 10:44 pm »

1. All the goods ones are in use or have been used.  This one is on sideline because it isn't good.
0) why is that good?

The complexity, extra mass, and losses in the heat exchangers are greater drawbacks.  Not to mention the inability to start such an engine.

It may not be good, and perhaps I'm overlooking something, but saying "all the good ones have been used" is a poor argument and akin to saying everything that could possibly be invented already has. What happened the last time someone made that very claim?

Complexity compared to what? Are you saying staged combustion is less complex? If we want minimal complexity, then every rocket would be pressure fed. After all, pumps are complex. Extra mass compared to what? Pressure fed rockets, GG combustion chambers? I'll agree that there's no free lunch. When it comes to rocket motors, if you remove mass by eliminating one component, chances are you'll have to add mass by adding another. The biggest drawback is definitely the heat exchanger, and I've admitted as much. I also don't know what is to be expected in terms of mass and efficiency. If you have any insight, then please share it. Otherwise how can you claim it's a deal-breaker outright? Every method will have losses somewhere, that's why there's no single engine that rules them all in every way. You have pressure losses prior to [some of] the propellants entering the combustion chamber even in the mighty staged combustion engines, because they're pushed through the cooling system prior to entering the gas generator and turbine (so even more losses!).

It wouldn't be any good if it couldn't be started, right? This isn't significantly different than starting many other types of rockets. The RL10 starts just fine. Vacuum starting would be the most trivial. One option is to have an external pressurant and associated valving ahead of the turbine, as well as a valve behind the turbine that vents to the atmosphere. This pressure difference gets the turbopump started until it's self-sustaining. It's not all that different from many other engines.

From the sound of it, you are involved in the field, but I have no idea what your specialty is. In that light, whose thoughts on the matter of using heat exchangers on a rocket should I value more: yours, or Vladimir Balepin's - who has actual and verifiable professional experience on the subject?

In fact, there has at least been some testing done (and please don't say that just because we haven't heard anything since means it was a failure. Many workable ideas don't get pursued for various reasons as you know)

http://aerospace.xcor.com/rocket-engines/research-development/

Quote
In early 2006, XCOR and ATK GASL received a DARPA contract to "investigate, develop, and demonstrate a novel configuration for a liquid rocket engine, namely a Third Fluid Cooled (TFC) liquid rocket engine," for which we used the Tea Cart to generate superheated steam suitable for driving a Rankine cycle. This steam cycle will allow a turbopump system to develop high chamber pressure in a more efficient way than staged combustion cycles, thus improving the performance and durability for boost propulsion, as well as orbital transfer applications. 

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Re: Rocket Engine Q&A
« Reply #691 on: 07/16/2015 11:45 pm »
DARPA contracts are not relevant indicators for viability of technology.  More like throwing stuff at a wall and see what sticks.

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Re: Rocket Engine Q&A
« Reply #692 on: 07/16/2015 11:50 pm »
The RL10 starts just fine.

It just needs valves opened to start. 

Existing engines don't have a closed system driving the turbo pump so back pressure is not a concern

Offline msat

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Re: Rocket Engine Q&A
« Reply #693 on: 07/17/2015 02:42 am »
DARPA contracts are not relevant indicators for viability of technology.  More like throwing stuff at a wall and see what sticks.

Fairly true, but generally you need some preliminary research before you get funding. I'd imagine many of these questions have been addressed to some degree in Balepin's AIAA paper written on ATK's dime.


The RL10 starts just fine.

It just needs valves opened to start. 

Essentially the same for this

Quote
Existing engines don't have a closed system driving the turbo pump so back pressure is not a concern

Keep in mind, I'm not talking about TFC. The coolant system isn't a closed loop. Instead, once the coolant (which is the same as the fuel) passes through the heat exchanger, it is dumped back into the fuel supply upstream of the compressor.

Pretty much all rocket engines besides pressure-fed/open cycle GGs with ablative thrust chambers & nozzles have some pressure drops relative to the pump output at the very least due to losses from the cooling channels. In the case of staged combustion, you have similar issues as an expander cycle where you have to provide a suitable pressure drop across the turbine to generate sufficient power for the pumps, further lowering the pressure to the thrust chamber relative to pump output. To avoid these pressure drops so to generate more thrust (at the expense of ISP), Mitsubishi built the LE-5A/B as an expander bleed cycle ensuring the engine saw maximum pump pressure. Can't argue that there's no logic behind it.
« Last Edit: 07/17/2015 06:16 am by msat »

Offline ClaytonBirchenough

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Re: Rocket Engine Q&A
« Reply #694 on: 08/06/2015 12:13 am »
So I have questions regarding erosive burning that solid rocket engines experience. This next explanation is from braeunig.us.

Quote
For most propellants, certain levels of local combustion gas velocity (or mass flux) flowing parallel to the burning surface leads to an increased burning rate. This "augmentation" of burn rate is referred to as erosive burning, with the extent varying with propellant type and chamber pressure. For many propellants, a threshold flow velocity exists. Below this flow level, either no augmentation occurs, or a decrease in burn rate is experienced (negative erosive burning).
The effects of erosive burning can be minimized by designing the motor with a sufficiently large port-to-throat area ratio (Aport/At). The port area is the cross-section area of the flow channel in a motor. For a hollow-cylindrical grain, this is the cross-section area of the core. As a rule of thumb, the ratio should be a minimum of 2 for a grain L/D ratio of 6. A greater Aport/At ratio should be used for grains with larger L/D ratios.

So if I understand correctly, if we're using a cylindrical core burning solid rocket engine with a core diameter of 3 inches and an outer grain diameter of 6 inches with a throat diameter of 2 inches, the port-to-throat area ration would be:

3 inches / 2 = 1.5 inch radius

1.5 inches * 1.5 inches * pi = 7.07 inches squared for the port area. And then the throat area would be:

2 inches / 2 = 1 inch radius

1 inch * 1 inch * pi = 3.14 inches squared for the throat area. So then the port-to-throat ratio would be:

7.07 inches squared / 3.14 inches squared = 2.25.

Now here comes my question (assuming I've just got everything right). Going by what I emphasized in the above quote, the grain L/D ratio should be no more than 6. I assume that by "L/D" ratio, the author means grain length to diameter ratio? Does this mean the outside grain diameter? So would the limit on the length of my solid rocket engine be around:

 6 inch diameter * 6 = 36 inches? Or does L/D ratio refer to the rocket engine length to the core diameter? So the max length of my solid rocket engine could only be around:

3 inch diameter * 6 = 18 inches? Can someone help by clarifying the approximations for what the dimensions of a solid rocket engine should be to minimize erosive burning?
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Offline R7

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Re: Rocket Engine Q&A
« Reply #695 on: 08/06/2015 08:46 am »

7.07 inches squared / 3.14 inches squared = 2.25.

Now here comes my question (assuming I've just got everything right).

All good so far but note that you can simplify the math by just squaring the ratio of diameters (or radii, does not matter which). Area is proportional to square of length, that's enough knowledge so no need to calculate the actual areas.

(3/2)2 = 2.25

Quote
Going by what I emphasized in the above quote, the grain L/D ratio should be no more than 6. I assume that by "L/D" ratio, the author means grain length to diameter ratio? Does this mean the outside grain diameter? So would the limit on the length of my solid rocket engine be around:

 6 inch diameter * 6 = 36 inches?

It's the grain L/D, L is grain length and D is outside diameter. I don't know how the thumb rule 2 Ap/At ratio scales with bigger L/Ds.

The combustion area in your simple cylinderical core desing would double at the end of the burn, thrust also. Eventually you want to seek out more complex port geometries for more even thrust.

Have you read Richard Nakka's pages? A lot of information for solid rocket hobbyists there.

http://www.nakka-rocketry.net/th_grain.html
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Offline ClaytonBirchenough

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Re: Rocket Engine Q&A
« Reply #696 on: 08/06/2015 12:34 pm »

7.07 inches squared / 3.14 inches squared = 2.25.

Now here comes my question (assuming I've just got everything right).

All good so far but note that you can simplify the math by just squaring the ratio of diameters (or radii, does not matter which). Area is proportional to square of length, that's enough knowledge so no need to calculate the actual areas.

(3/2)2 = 2.25

Quote
Going by what I emphasized in the above quote, the grain L/D ratio should be no more than 6. I assume that by "L/D" ratio, the author means grain length to diameter ratio? Does this mean the outside grain diameter? So would the limit on the length of my solid rocket engine be around:

 6 inch diameter * 6 = 36 inches?

It's the grain L/D, L is grain length and D is outside diameter. I don't know how the thumb rule 2 Ap/At ratio scales with bigger L/Ds.

The combustion area in your simple cylinderical core desing would double at the end of the burn, thrust also. Eventually you want to seek out more complex port geometries for more even thrust.

Have you read Richard Nakka's pages? A lot of information for solid rocket hobbyists there.

http://www.nakka-rocketry.net/th_grain.html

Thanks for the response!

I understand everything you said, and what you explained is actually what I tried to communicate I was doing haha. In the math I did, I showed that the max length of a 6 inch diameter grain was 36 inches (again, using the 6 L/D ratio where I don't exactly know comes from). This *would be right?

I'm also aware that the grain surface area would double in my design. Not optimal, but makes for a easier fabrication. Also, it may actually be a good thing to minimize erosive burning because the mass flow rate and chamber pressure at ignition happens with a smaller port. In addition, I have read a lot of Nakka's website. He's the man.

But I do get confused on how L/D can become an approximation of the grain length? Is it maybe because they take into account chamber pressure and propellant surface area and assume your Kn (propellant surface area to throat area) is within reason? Thanks R7 for your reply!
Clayton Birchenough

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Re: Rocket Engine Q&A
« Reply #697 on: 11/21/2015 11:36 pm »
Please correct me if I'm wrong:

Was MMS the last usage of the RL-10A-4-2 Centaur engine, or do they have specific manifested flights for the old engine?
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Offline Newton_V

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Re: Rocket Engine Q&A
« Reply #698 on: 11/21/2015 11:41 pm »
Please correct me if I'm wrong:

Was MMS the last usage of the RL-10A-4-2 Centaur engine, or do they have specific manifested flights for the old engine?

OSIRIS-REX and possibly all Dual Engine Centaurs.

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Re: Rocket Engine Q&A
« Reply #699 on: 11/21/2015 11:57 pm »
Please correct me if I'm wrong:

Was MMS the last usage of the RL-10A-4-2 Centaur engine, or do they have specific manifested flights for the old engine?

OSIRIS-REX and possibly all Dual Engine Centaurs.

Thank you.
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