Author Topic: Basic Rocket Science Q & A  (Read 270870 times)

Offline Jim

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Re: Basic Rocket Science Q & A
« Reply #100 on: 03/10/2009 09:39 PM »
It depends on the propellants. 

Offline cgrunska

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Re: Basic Rocket Science Q & A
« Reply #101 on: 03/10/2009 10:27 PM »
a good point

chemical combustion engine vs electrical engine then?
I assume a nuclear propelled spacecraft is an electrical engine and not an explosive one.

Offline Jim

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Re: Basic Rocket Science Q & A
« Reply #102 on: 03/10/2009 10:36 PM »
I was referring to differences in chemical propellants.  LO2/LH2, RP1/LOX, N2O4/UDMH

Offline mmeijeri

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Re: Basic Rocket Science Q & A
« Reply #103 on: 03/10/2009 11:03 PM »
I have a question about the corrosiveness of N2O4. I have read that adding a few percent of NO reduces the corrosiveness and that the resulting mixture is called MON, which stands for mixed oxides of nitrogen. Various mixture ratios are in use.

So how much of a difference does this make and how much of a problem is this corrosiveness? I'm particularly interested in the effect on the near-term feasibility of orbital hypergolic propellant depots. How long can you realistically store MMH/MON in space before corrosion renders your depot inoperable?
« Last Edit: 03/10/2009 11:04 PM by mmeijeri »
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Offline nomadd22

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Re: Basic Rocket Science Q & A
« Reply #104 on: 03/10/2009 11:06 PM »
 Not sure why you think there wouldn't be fire coming out the backside in space. Near the engine it would look about the same as in the air. Yellow/white flame for most kerosene type and flame so blue you can hardly see it for H2/O2.
 I can't say i know what the water trail for a hydrogen engine looks like in vacuum.
 If you go to Spacex.com you can watch the film of the 4th Falcon 1 launch. They had a camera on the second stage picking up it's flight through the vacuum.

Offline nomadd22

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Re: Basic Rocket Science Q & A
« Reply #105 on: 03/10/2009 11:22 PM »
I have a question about heat transfer and cooling/heating of sattelites.
In many sources, especially in main stream media, it sounds like an unbelievably big problem, that in space, the sun side is so much hotter than the shadow. And they talk about huge figures hundreds of degrees celsius of temperature difference.
On the other hand, my common sense tells me, that because of the low pressure there is probably very little convection and therefor there is only radiation for heat transfer.
Therefor, i'd think that though the air moluecules in LEO are technically really hot, they transfer very little energy to a spacecraft.
So how is the cooling/heating acomplished? I'd guess that making the spacecraft reflective on the outside would block most of the energy coming from the sun, an because of little convection the spacecraft should cool very little when provided with basic insulation against heat transfer to the outer hull (which would remove the energy by way of infrared radiation).
Then I'd guess the electronics or other systems (e.g. a human in a spacecraft) would provide enough heat to keep the sacecraft from freezing.
So what, would be left would be to provide a way of radiating exactly as much energy out of the spacecraft as needed o keep constant temperature, so how would I do that?
Temperature in space doesn't mean much because there's not much there to be hot or cold, so you're right, and convection isn't a factor.
 You just have to worry about direct solar radiation and heat produced by spacecraft systems.
 Satellites in equatorial or low inclination orbits stay evenly baked since they have to rotate once an orbit to keep the same face toward the earth. They can keep the electronics just the right temp by adjusting the design of the satellite to control how much heat it picks up and how much it radiates back out.
 The small lower powered ones radiate enough heat on the dark sides to make up for what the sunlit sides pick up.
 Big ones, like the space station can have big honkin heat pump systems that usually use ammonia for refrigerant and big radiators kept edge on to the sun.
« Last Edit: 03/10/2009 11:22 PM by nomadd22 »

Offline Jim

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Re: Basic Rocket Science Q & A
« Reply #106 on: 03/10/2009 11:28 PM »
The flame becomes very diffuse and thin in the vacuum.  Look at the rocketcams on a Delta first stage.  Youtube has many

Offline Jim

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Re: Basic Rocket Science Q & A
« Reply #107 on: 03/11/2009 12:22 AM »
I have a question about heat transfer and cooling/heating of sattelites.
In many sources, especially in main stream media, it sounds like an unbelievably big problem, that in space, the sun side is so much hotter than the shadow. And they talk about huge figures hundreds of degrees celsius of temperature difference.
On the other hand, my common sense tells me, that because of the low pressure there is probably very little convection and therefor there is only radiation for heat transfer.
Therefor, i'd think that though the air moluecules in LEO are technically really hot, they transfer very little energy to a spacecraft.
So how is the cooling/heating acomplished? I'd guess that making the spacecraft reflective on the outside would block most of the energy coming from the sun, an because of little convection the spacecraft should cool very little when provided with basic insulation against heat transfer to the outer hull (which would remove the energy by way of infrared radiation).
Then I'd guess the electronics or other systems (e.g. a human in a spacecraft) would provide enough heat to keep the sacecraft from freezing.
So what, would be left would be to provide a way of radiating exactly as much energy out of the spacecraft as needed o keep constant temperature, so how would I do that?

There is no air pressure and few to none air molecules where most spacecraft operate, hence no convection.  Sunward side of the spacecraft experiences high radiative heat input.  The sides facing  away would radiate heat into the black body of deep space.   MLI (multi layered insulation) is used to cover the spacecraft and comes in many types, with many different outer layers (white beta cloth, aluminized mylar, kapton, etc).  MLI both retains and reflect heat.  Since there is no convention, electronics would overheat and fluids would freeze. The electronics are usually mounted on radiators or connected to them by heatpipes.  Heaters are employed to keep parts like propellant tanks warm. 
« Last Edit: 03/11/2009 12:27 AM by Jim »

Offline charlieb

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Re: Basic Rocket Science Q & A
« Reply #108 on: 03/11/2009 12:42 AM »
I have a question about heat transfer and cooling/heating of sattelites.
In many sources, especially in main stream media, it sounds like an unbelievably big problem, that in space, the sun side is so much hotter than the shadow. And they talk about huge figures hundreds of degrees celsius of temperature difference.
On the other hand, my common sense tells me, that because of the low pressure there is probably very little convection and therefor there is only radiation for heat transfer.
Therefor, i'd think that though the air moluecules in LEO are technically really hot, they transfer very little energy to a spacecraft.
So how is the cooling/heating acomplished? I'd guess that making the spacecraft reflective on the outside would block most of the energy coming from the sun, an because of little convection the spacecraft should cool very little when provided with basic insulation against heat transfer to the outer hull (which would remove the energy by way of infrared radiation).
Then I'd guess the electronics or other systems (e.g. a human in a spacecraft) would provide enough heat to keep the sacecraft from freezing.
So what, would be left would be to provide a way of radiating exactly as much energy out of the spacecraft as needed o keep constant temperature, so how would I do that?

The following concerns GEO Comm type-satellites, but generically covers any spacecraft to a degree:

Satellite thermal configurations are designed to maintain all spacecraft equipment within allowable temperature limits during transfer orbit and throughout their life on orbit. Thermal control is achieved using both passive and active elements. Passive elements include thermal blankets, coatings, shields, radiators, and heat pipes. Active elements consist of ground and autonomously controlled heaters. Thermistors provide temperature data that are used for thermal control and health monitoring.

Blankets, coatings, and solar reflectors are used to control radiative heat transfer to and from the spacecraft, as well as among units within the spacecraft bus. Multi-Layer Insulation (MLI) blankets are used externally on the east, west, anti-Earth, and Earth spacecraft surfaces to reduce diurnal temperature variations. Optical Solar Reflectors (OSRs), mounted on the north and south communication panels, subsystem module  radiator panels, battery panel radiators, and the Earth sensor radiator panel, minimize solar absorption during periods when there is significant solar flux on these surfaces. The OSRs also increase the spacecraft thermal emissivity, thereby cooling the radiator panels, which contain the majority of high-heat dissipation units.

Internal surface coatings help control radiative heat transfer within the satellite and aid in maintaining elements within limits.

The solar array temperature is controlled entirely by passive means. Absorption of solar energy on the front (cell) side of the array is emitted by the highly emissive, black graphite back surface of the array.

The antenna reflectors are subjected to a diurnal solar flux. MLI blankets covering the reflector backup structures help reduce temperature variations and thermally isolate the antennas from the spacecraft mainbody. MLI blankets are also used to maintain a constant temperature of the antenna tower structure(s).

Heat pipes are embedded in the spacecraft radiator panels (north and south communication panels, subsystem module panels, and battery panels) to minimize temperature variations and to convect heat from high-dissipation units to cooler areas of the panel. Heat pipes consist of a pair of ammonia-filled tubes that convect heat, and thus provide additional paths for heat transfer to take full advantage of the OSR-covered radiator panels.

Thermistors are used to sense the temperatures of all major and critical subsystems, indicating the thermal state of health of the spacecraft and providing performance data and diagnostic capability. Thermistors also allow positive verification of the proper operation of the spacecraft during all phases of operations and test and allow timely detection of anomalies
Former Shuttle Mission Ops Eng  (In them days DF24 - INCO GROUP/COMMS, Now DS231-AVIONICS BRANCH).

Offline ugordan

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Re: Basic Rocket Science Q & A
« Reply #109 on: 03/11/2009 08:25 AM »
From what I gather, LOX/LH2 exhaust is invisible in vacuum. If you looked into the engine nozzle and chamber you'd see a brilliant glow, but past the nozzle water vapor is invisible. Even in darkness (nighttime launches), the exhaust from the engine is invisible.

LOX/RP1 has apparently a more particulate exhaust (sooty particles/unburned fuel) that's visible even in vacuum, although it doesn't "burn", it's a dark color as it expands and cools rapidly. In nighttime operation the exhaust near the nozzle is presumably illuminated by combustion chamber glow diffusion - see Falcon 1 flight 4 2nd stage restart footage.

As for N2O4/UDMH - Delta II 2nd stage footage suggests it's also invisible in vacuum.

Offline Jim

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Re: Basic Rocket Science Q & A
« Reply #110 on: 03/11/2009 11:47 AM »
I have a question about the corrosiveness of N2O4. I have read that adding a few percent of NO reduces the corrosiveness and that the resulting mixture is called MON, which stands for mixed oxides of nitrogen. Various mixture ratios are in use.

So how much of a difference does this make and how much of a problem is this corrosiveness? I'm particularly interested in the effect on the near-term feasibility of orbital hypergolic propellant depots. How long can you realistically store MMH/MON in space before corrosion renders your depot inoperable?

It isn't a problem.  Comsats have serviceable lives of more than 15 years with MON.  Cassini is now over ten.  MGS was over ten

Offline mmeijeri

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Re: Basic Rocket Science Q & A
« Reply #111 on: 03/11/2009 11:51 AM »
It isn't a problem.  Comsats have serviceable lives of more than 15 years with MON.  Cassini is now over ten.  MGS was over ten

Thanks again Jim. Also: long live the internet, I've just found a scanned copy of a Rocketdyne nitrogen tetroxide handling manual from 1961 :-)
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Offline strangequark

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Re: Basic Rocket Science Q & A
« Reply #112 on: 03/19/2009 02:37 AM »
Was wondering if anyone could help me find reference material on mass ratio scaling laws for first stages based on propellant density. Specifically, I'm curious as to how CH4-LOX stacks up against LH2-LOX and Kerolox for a first stage. It looks like the average propellant density nearly evens out with Kerolox versus Methlox, so I just wonder.


I've just found a scanned copy of a Rocketdyne nitrogen tetroxide handling manual from 1961 :-)

Ironic it's called a "handling" manual. Somehow, I think a strictly literal interpretation would be bad :-0 .
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Offline Steven Pietrobon

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Re: Basic Rocket Science Q & A
« Reply #113 on: 03/19/2009 04:10 AM »
strangequark, I wrote two papers on this subject which are attached. In a first stage performance is limited by propellant volume. I show in my paper that the criteria for choosing a fuel for a first stage is its impulse density Id, equal to the product of the propellant density (kg/L) and the exhuast speed (m/s). A list of propellants is given below

Propellants  MR   dp (kg/L)  ve (m/s) Id (Ns/L)
O2/H2        5.0  0.3251     4455     1448
O2/H2        6.0  0.3622     4444     1610
O2/H2        7.5  0.4120     4365     1798
O2/CH4       3.6  0.8376     3656     3062
O2/C2H6      3.2  0.9252     3634     3362
O2/C3H8      3.1  0.9304     3613     3362
O2/C3H4      2.4  0.9666     3696     3573
O2/RP–1      2.8  1.0307     3554     3663
O2/C7H8      2.4  1.0954     3628     3974
HTP/C3H4     6.5  1.2553     3319     4166
HTP/RP–1     7.3  1.3059     3223     4209
HTP/C7H8     6.6  1.3496     3288     4437

HTP is 98% hydrogen peroxide, RP-1 is rocket grade kerosene and C7H9 is quadricyclene or RP-X2 (an exotic hydrocarbon fuel). O2/H2 (liquid oxygen and liguid hydrogen) has the best exhaust speed, but a very poor density, which makes it a bad choice as a first stage propellant (that's why the Delta-IV is so huge). A good combination is O2/RP-1, but there are better combinations, such as HTP/RP-1 which will give 15% more performance, plus it has the advantage of being non-cryogenic at the disadvantage of being unstable in the presence of impurities.

So in answer to your specific question about O2/CH4, that performs worse than O2/RP-1 in a first stage. This means your fuel tanks will need to be about 20% larger in order to have the same performance as O2/RP-1. Against O2/H2 it performs much better. Your fuel tanks will be about 47% smaller (the actual percentage depends on the required delta-v for the first stage, larger values will decrease this amount, but even to orbital speeds, O2/CH4 will still perform better).

For an upper stage, mass is all important, so the high exhaust speed of O2/H2 means that's the best propellant to use.

By the way, for single stage to orbit, my second paper shows that any combination O2 or HTP with any other hydrocarbon fuel will outperform O2/H2. A good choice is O2/RP-1, but O2/C7H8 gives 13% more performance.
« Last Edit: 03/19/2009 04:46 AM by Steven Pietrobon »
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Offline Eerie

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Re: Basic Rocket Science Q & A
« Reply #114 on: 03/20/2009 01:40 PM »
Then why all SSTO projects seems to use O2/H2?

Offline gospacex

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Re: Basic Rocket Science Q & A
« Reply #115 on: 03/20/2009 10:40 PM »
strangequark, I wrote two papers on this subject which are attached. In a first stage performance is limited by propellant volume. I show in my paper that the criteria for choosing a fuel for a first stage is its impulse density Id, equal to the product of the propellant density (kg/L) and the exhuast speed (m/s). A list of propellants is given below

Propellants  MR   dp (kg/L)  ve (m/s) Id (Ns/L)
O2/H2        5.0  0.3251     4455     1448
O2/H2        6.0  0.3622     4444     1610
O2/H2        7.5  0.4120     4365     1798
O2/CH4       3.6  0.8376     3656     3062
O2/C2H6      3.2  0.9252     3634     3362
O2/C3H8      3.1  0.9304     3613     3362
O2/C3H4      2.4  0.9666     3696     3573
O2/RP–1      2.8  1.0307     3554     3663
O2/C7H8      2.4  1.0954     3628     3974
HTP/C3H4     6.5  1.2553     3319     4166
HTP/RP–1     7.3  1.3059     3223     4209
HTP/C7H8     6.6  1.3496     3288     4437


HTP is 98% hydrogen peroxide

It is an explosion hazard, and is more expensive than LOX.

I think O2/RP–1 looks good (no wonder it is such a widely used combination).

O2/C3H8 (that's LOX/propane) has some advantages non-obvious from this table: O2 and C3H8 are cryocompatible (easy to build common bulkhead tank for them), and C3H8 is somewhat easier on engines than RP-1: less sooting, and no nasty kerosene residue. I don't know for sure, but maybe it's possible to build an engine in RD-180 class without oxygen-rich preburner for this combination (with RP-1 it's not possible). Also both are cheap.

Offline Steven Pietrobon

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Re: Basic Rocket Science Q & A
« Reply #116 on: 03/23/2009 04:02 AM »
HTP is 98% hydrogen peroxide

It is an explosion hazard, and is more expensive than LOX.

So are solid rocket motors, which are widely used. Provided that proper precautions are followed, HTP is safe to use, as shown by prior British and US experience. Also, a HTP stage only needs to be filled just prior to launch, unlike solids where an explosion hazard is always present from the time the stage is filled to its use on the launch pad.

Nitrous Oxide (N2O) is also an explosion hazard and is going to be used on a crewed spacecraft. HTP is much less sensitive to shock (unlike N2O), practically requiring a small explosion in order to set it off.

Propellant cost is a small fraction of launch costs. The 15% performance increase more than makes up for any propellant cost increase.
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Offline Steven Pietrobon

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Re: Basic Rocket Science Q & A
« Reply #117 on: 03/23/2009 04:18 AM »
Then why all SSTO projects seems to use O2/H2?

That's a very good question. I believe this is because people only look at exhuast speed ve (divide by g = 9.8065 m/s^2 to get specific impulse) as the performance criteria. O2/H2 has one of the highest ve and is thus thought to be the best propellant. However, it also has the poorest density. O2/RP-1 has a density which is 185% greater than O2/H2, while having an exhaust speed that is only 20% less. This means that in a first stage, you can reduce propellant volume by up to 56%. If you do the sums, a vertical takeoff SSTO vehicle will always have much better performance with O2/RP-1 than with O2/H2.

For horizontal takeoff, then O2/H2 may have better performance due to its much lower total liftoff weight.
« Last Edit: 03/23/2009 04:19 AM by Steven Pietrobon »
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Offline gospacex

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Re: Basic Rocket Science Q & A
« Reply #118 on: 03/23/2009 04:33 AM »
HTP is 98% hydrogen peroxide

It is an explosion hazard, and is more expensive than LOX.

So are solid rocket motors, which are widely used.

I don't like solids either. It doesn't make me like HTP. Imagine a manned rocket at the pad fueled by HTP. Unlike LOX one, which can only burn, HTP *can* explode, LAS wouldn't be able to save the crew.

Quote
Provided that proper precautions are followed, HTP is safe to use, as shown by prior British and US experience.

With enough care, *anything* is safe to use. Like solids or rocketplanes with fragile TPS...

Quote
Propellant cost is a small fraction of launch costs. The 15% performance increase more than makes up for any propellant cost increase.

BTW, do you have data on HTP/C3H8 pair?

Offline sbt

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Re: Basic Rocket Science Q & A
« Reply #119 on: 03/23/2009 07:07 AM »

To add a data point:

The British Gamma decomposed the HTP to Water + O2 by passing it over a catalyst (which had the disadvantage in Gammas case of needing insertion 2 hrs before launch due to a short 'shelf' life). Steam plus O2 at 500 Celsius is Hypergolic with RP1 so no igniter was needed.

By choosing a Silver-plated Nickel Wire catalyst on cost grounds the UK limited its HTP to 85% H2O2. Higher concentrations would have required something like Platinum.

As well as being creating a Hypergolic engine the decomposition of HTP can also be used to power the fuel and oxidiser turbo-pumps, as with Gamma, giving you a Closed Cycle engine for much less fuss than Staged Combustion.

Finally, although you might not think it, HTP can be used in a Regenerative design – as in Gamma.

The reason cost was more of a factor in selecting the catalyst than might be expected in a launcher program was Gamma was a derivative of one of, and related to many of, motors intended for JATO, Rocket and/or Rocket Assisted fighters and the Blue Steel Stand Off Missile.

In addition to performance HTP was chosen for its relatively benign handling on airfields. It didn't need Cryogenic storage or handling and if spilled it had no dangerous off gassing and could be safely washed away with copious water.

The big question in my mind is how the Catalytic Decomposition approach and Regenerative cooling would scale up beyond Gamma and the larger Stentor Large Chamber (24,000 lb Thrust).

There is also the issue of Freezing at relatively high temperatures (It does, after all, contain water).)

On the Explosion Hazard and LAS. If the Ares I LAS can pull a crew away from an exploding SRM then I wouldn't be so sure that one couldn't be devised to protect a crew from a HTP stage. It might be more difficult for a Single Stage system though, as you loose the separation and buffer provided by the 2nd Stage.

As to why its not proposed more often? One big factor is, I suspect, that LOX/LH and LOX/RP1 are regarded as 'developed technologies' whereas HTP based bi-propellant engines haven't been demonstrated much beyond the British program.

Finally the UK Bloodhound SSC record breaking car project will use a HTP/HTPB Hybrid for main propulsion – again using decomposition and no igniter.

Rick
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