ApolloHoax.net
Apollo Discussions => The Reality of Apollo => Topic started by: Mag40 on December 31, 2012, 09:50:04 AM
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On the Apollo 11 footage post landing docking, we see the LM doing a series of maneuvres. One claim I've seen(can't recall where) states that these moves.....followed by immediate stops are impossible. Now I think I know how it's done, but would love this clarified. I've been reading the Apollo experience report and couldn't see where the astronaut control tied in. Is this a matter of each joystick 'blip' left or right etc..... administering a measured amount, or was this simply down to the astronauts being extremely accurate? What I mean is, say a rotation is initiated of 60 degrees, then the LM stops rotating immediately, how is this performed dead stop.
This is a youtube video for reference, showing some Apollo 11 highlights....at the 5 minute mark
What would also be cool is if someone explained how they managed to move it sideways without rotating, is it a combination of joysticks and other controls?
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I think I've seen the same claims (on Youtube) that these motions were somehow impossible. Obviously, they weren't!
Two important points. First, the 16mm movie camera that took these sequences ran at only 6 fps (I think), so at the nominal 24 fps playback speed the movie plays back 4x faster than real time. (Don't quote me on these numbers before checking them.)
Second, the LM was extremely light at this point. At docking it has lost the entire descent stage and nearly all of its ascent propellants so its moments and products of inertia are all very low. Since the RCS thrusters are fixed at 100 lbf of thrust, the crews said the LM handled like a sports car during the docking. A "pure" pitch, roll or yaw maneuver required 2 or 4 thrusters, i.e., without introducing translation. A pure X axis translation required 4 thrusters, while a pure Y or Z axis translation required either 2 or 4 depending on the center of mass.
A while ago I looked up the LM's mass properties from a table in the mission report, computed the torques produced by the RCS thrusters, and worked out the maximum linear and angular accelerations that they could provide at docking. Taking the movie frame rate into account I found that the maneuvers shown were well within the LM's capability.
The LM was manually controlled with two separate joysticks, one in each hand, each with three degrees of motion. The conventional-looking stick on the right was for attitude control: pitch, roll and yaw. The T-shaped handle on the left was for translation: X (up/down), Y (left/right) and Z (forward/backward). The CSM pilot used the same two types of sticks.
With very few exceptions the LM and CSM operated in a "fly by wire" mode. The sticks sent signals to the computer which in turn decided which thrusters to fire. This was the case even in the so-called "manual mode" used just before landing by every Apollo commander. The "manual control mode" that Neil Armstrong famously used to fly over the boulder field was more accurately a semi-automatic mode. He actually put the computer into an "attitude hold" mode so that when he let go of the stick, the computer automatically fired whatever thrusters were needed to maintain the attitude he had selected.
Even in this so-called manual mode the computer controlled the descent engine, adjusting its gimbals to keep the thrust vector through the center of mass and adjusting the throttle to maintain a fixed descent rate while compensating for the LM's rapidly decreasing weight as it burned descent propellants. The commander adjusted the descent rate with a momentary toggle switch that worked very much like the cruise control in a car: flicking it in one direction slowed the descent rate by a foot per second and the other direction increased it by the same amount.
Except for the final phase of landing, LM powered flight was almost entirely automatic.
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Thanks Ka9Q! The signals to fire the thrusters....do they involve fixed amount of fuel, or is it astronaut skill?
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Don't know about the specific claim you've heard, but on similar issues, it's worth noting that due to the low frame rate of recording that ka9q mentions, many of the RCS bursts didn't get caught so the LM looks like it's just randomly changing motion without cause, which looks strange.
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Thanks Ka9Q! The signals to fire the thrusters....do they involve fixed amount of fuel, or is it astronaut skill?
I know this is an old thread, but I just noticed you asked a question that I never answered.
The RCS thrusters had constant thrust -- 100 lbf each. Impulse was controlled by how long you held the propellant valves open. You could also effectively vary thrust by rapidly cycling them with a particular duty cycle. But because the valves are electromechanical, and because it takes finite time for the propellants to mix, ignite and build thrust, there is a practical "minimum impulse" for each firing. The computer took this into account when computing which thrusters to fire, when and for how long. This is one of the main reasons for "dead bands", an allowable margin of attitude error within which the computer would not try to correct the error further.
The computer could be instructed to maintain a given attitude within the dead band, which it would do as efficiently as possible, or an astronaut could fire them manually in which case fuel consumption depended on his skill. Transposition and docking was performed manually shortly after TLI, and the CMPs seemed to compete with each other for bragging rights as to who could do it with the least amount of gas.
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LM RCS could be operated in "pulse mode," which is a minimum-impulse mode. The servovalves were simply cycled as fast as possible such that a deflection of the hand controller would result (in some settings) in the fastest possible open-close cycle, or (in other settings) a rate-controlled series of pulses over time. This was the answer to an RCS that had to maneuver the fully-fueled LM at rates that were fast enough to facilitate a safe landing, while later maneuvering the ascent stage only at controllable rates.
The basic autopilot mode was ATT HOLD, which will pulse or fire the RCS jets to maintain space-fixed orientation within a deadband. The desired orientation is set in most modes with the rotational hand controller. The size of the deadband was selectable. There is no equivalent "position hold," so translations using the other hand controller had to be manually nulled. However, ATT HOLD will null the rotational residuals.
If you have Orbiter, there is an Apollo add-on that will let you practice CSM transposition and docking, however the CSM dynamics feel markedly different than the LM. For transposition and docking the CSM was the active vehicle. For LOR, the LM was the active vehicle -- hence the jinking around you see on the film.
Regarding film speed, the Maurer on the CSM was indeed set to 6 fps during ascent and rendezvous. Otherwise the CMP would have needed to change magazines right about the time the rendezvous got interesting. As such the maneuvers seem abrupt. Given the oversized RCS at this point, the starts and stops would be abrupt anyway. Ed Mitchell told me that when Shepard let him fly the LM ascent module, he described its performance as "sporty." The LM RCS is unquestionably in pulse mode during the ascent, and the pulses are short enough (something like 6-10 milliseconds) that you rarely catch one on a frame, nor would you even at full 24 fps.
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Speaking of catching a pulse of the thrusters on a frame of film...
In Voyage to the Moon the CGI depicted the thrusters as leaving ice crystals (I think - could be other white dots) behind after firing, with no other exhaust artifact, such as flame. I understand from reading here and elsewhere that the hypergolics used are virtually invisible when burning - IIRC, the visible flames from the Titan series (other than SRB's) were mostly from contaminants in the exhaust stream, from the engine bells, environment, etc.
Not being an organic chemist by any stretch of the imagination, what would have been the residue from the thrusters? Would water have been a component of the exhaust (I'm guessing it would be)? Would enough have been present to make visible crystals? Or was that an artifact added by the visual team in the series to show that the thrusters were firing?
Just as an aside, it does boggle the mind the sheer number of rocket engines used in the entire Saturn V/Apollo stack if you take into account thrusters, ullage motors, etc., etc....
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Not being an organic chemist by any stretch of the imagination, what would have been the residue from the thrusters? Would water have been a component of the exhaust (I'm guessing it would be)?
Water is definitely a large constituent of the exhaust. The main gases produce by the combustion of nitrogen tetroxide and Aerozine 50 are water vapor, nitrogen, hydrogen, and carbon monoxide.
Would enough have been present to make visible crystals? Or was that an artifact added by the visual team in the series to show that the thrusters were firing?
Since the water is expelled as a gas, I doubt ice crystals would form. If the water were expelled as a liquid then, yes, some of it would freeze and form ice crystals. However, I have a hard time seeing how it would go from gas phase to solid phase (though I could be wrong).
Just as an aside, it does boggle the mind the sheer number of rocket engines used in the entire Saturn V/Apollo stack if you take into account thrusters, ullage motors, etc., etc....
Yes it is surprising how many engines/motors there are. I added them all up once and I think it is over 80.
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Those particular hypergols tend to throw an incandescent plume during in the ignition transient, but then burn without incandescence at steady state. In pulse mode it's all about the transient, so if you were watching the thrusters live you could expect to see a very short visible burst.
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Now, we're talking hypergolic engines - is it possible to calculate the temperature of the exhaust of the LM's descent engine? This relates to the hoaxer idea of scorching on the surface. I know the exhaust wasn't capable of burning anything, but a solid calculation would be nice. I've tried myself, but got an improbable result, due to my lack of the correct formulae.
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Now, we're talking hypergolic engines - is it possible to calculate the temperature of the exhaust of the LM's descent engine? This relates to the hoaxer idea of scorching on the surface. I know the exhaust wasn't capable of burning anything, but a solid calculation would be nice. I've tried myself, but got an improbable result, due to my lack of the correct formulae.
Yes, it's possible. The gas expansion is isentropic, thus the effect of pressure on temperature is described by the following equation,
T2 / T1 = (P2 / P1)1 – 1 / k
where k is the specific heat ratio.
The descent engine combustion chamber pressure was 713 kPa (103.4 psia). According to my own calculations, I estimate the combustion chamber temperature at 3,060 K and the specific heat ratio at 1.23. I further estimate the pressure at the nozzle exit to be about 1 kPa. Therefore we can calculate the temperature at the nozzle exit,
T2 = 3060 x (1 / 713)1 – 1 / 1.23 = 896 K
Of course the gas continues to expand after leaving the nozzle, thus the pressure and temperature drops further, and quite rapidly.
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Yes, that was what I thought - I was way off. That equation you use, is that a derivative of the PV=nRT?
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Yes, that was what I thought - I was way off. That equation you use, is that a derivative of the PV=nRT?
I think there's a derivation of it here,
http://www.grc.nasa.gov/WWW/K-12/airplane/compexp.html
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Yup - same equation. It was just me using the wrong chamber pressure, temperature, expansion ratio - basically all the important parts.
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Wouldn't you want to use the stagnation temperature and pressure if you're trying to find the temperature and pressure at the surface?
The plume expands after leaving the nozzle extension so it wouldn't reach the temperature in the combustion chamber, but it should be at least a little hotter than at the nozzle exit.
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Well, in my previous math-attempt I discovered the stagnation pressure. Now I have the exhaust temperature at the same point. And the pressure and temperature drops rapidly after it exits the nozzle.
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In pulse mode it's all about the transient, so if you were watching the thrusters live you could expect to see a very short visible burst.
If ice crystals were observed, perhaps it's unreacted propellant that has frozen by rapid evaporation. Aerozine-50 is a mix of UDMH and N2H4. The N2H4 has the higher freezing point (+2C) so it would be the first to freeze out. N2O4 freezes at -11C and a higher vapor pressure than N2H4 so it's another candidate.
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...it should be at least a little hotter than at the nozzle exit.
Temperature at the surface should be substantially colder than at the nozzle exit plane.
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If ice crystals were observed, perhaps it's unreacted propellant that has frozen by rapid evaporation. Aerozine-50 is a mix of UDMH and N2H4. The N2H4 has the higher freezing point (+2C) so it would be the first to freeze out. N2O4 freezes at -11C and a higher vapor pressure than N2H4 so it's another candidate.
If there are in fact ice crystals, I would believe this explanation before I would believe it to be water ice. But I still have a hard time believing anything would freeze. Given the temperature at which the thrusters operate, I would think all of the expelled matter would be in a gaseous state, even unreacted propellant. I'm not sure if it's even possible to go from gas to solid under these conditions. Both pressure and temperature are dropping so, looking at a phase diagram, we'd be moving down and to the left, which I think will keep us in the gaseous state.
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A chamber pressure of only 7.13 bar? Is that in the range usually used for rockets?
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A chamber pressure of only 7.13 bar? Is that in the range usually used for rockets?
For pressure fed engines, yes. For pump fed, no.
For a pressure fed system it is necessary to pressurized the entire system, including the propellant tanks. This means the tanks walls have to withstand the pressure, which means they must be thick and heavy. To keep the weight down, they operate the system at as low a pressure as practical. This lowers the specific impulse of the engine, but it is made up for in weight savings.
For pumped systems, the high pressure is only downstream of the pumps; therefore the tanks can remain of light construction. You can therefore ramp up the pressure in the engine without suffering a big weight penalty. These systems operate at much higher pressures.
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Given the temperature at which the thrusters operate, I would think all of the expelled matter would be in a gaseous state, even unreacted propellant. I'm not sure if it's even possible to go from gas to solid under these conditions.
The thrusters are hot only when firing. When the valves are first opened, some propellant will escape unreacted before ignition; this is why the plume is momentarily visible.
Liquids cannot exist for long in vacuum. A liquid (water, hydrazine, N2O4) suddenly exposed to vacuum immediately starts to boil, removing enough heat that the rest of it may freeze. This certainly happens when water is dumped to space through a simple nozzle without adding enough heat to vaporize it all.
I recently had a long argument with some of Boy Blunder's sycophants about why Apollo water dumps produced ice streams while the sublimators did not.
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A chamber pressure of only 7.13 bar? Is that in the range usually used for rockets?
Yes, for pressure fed engines. I'm familiar with the 400N thruster used on the AMSAT Phase III spacecraft; it was made by MBB originally for Galileo. The tanks were pressurized with helium to, IIRC, 10 bar. Obviously this must be greater than the combustion chamber pressure.
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When the valves are first opened, some propellant will escape unreacted before ignition
OK, that makes sense.
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Optimal mix occurs only at steady-state. During the transient the mix is not optimal and so some unreacted propellant will naturally escape.
In the larger motors (e.g., the SPS) the oxidizer is intentionally pre-injected to smooth the ignition. This results in a visible, likely incandescent plume for up to 300 milliseconds.
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So when talking about the temperature of the descent engines exhaust, would it be reasonable to say the temperature of the exhaust on the surface is (assuming a 45 degree spread after nozzle exit) 896 K x (area of nozzle/area of impact)? Where impact area is (nozzle diameter/2 + distance between nozzle and ground)^2 x pi?
Around 1.7/4.5 x 896K = 338 K? Or 65 C?
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I think there's more to it than that. When the plume hits the surface, at least some of its kinetic energy is turned back into heat. That's what stagnation temperature is about.
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Compression heating or frictional heating?
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Compression.
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So when talking about the temperature of the descent engines exhaust, would it be reasonable to say the temperature of the exhaust on the surface is (assuming a 45 degree spread after nozzle exit) 896 K x (area of nozzle/area of impact)? Where impact area is (nozzle diameter/2 + distance between nozzle and ground)^2 x pi?
Around 1.7/4.5 x 896K = 338 K? Or 65 C?
Not exactly. The effect of volume on pressure is described by the following equation,
(P2 / P1) = (V1 / V2) k
Combining this with the equation in Reply #10 and we get the effect of volume on temperature,
T2 / T1 = (V1 / V2) k – 1
So in your example, where we have a 1.7 to 4.5 change in volume, the new temperature becomes,
T2 = 896 x (1.7 / 4.5) 1.23 – 1 = 716 K
(ETA) Also note that the change in cross-sectional area doesn't necessarily equate to a proportional change in volume. Area and volume are proportional only if the velocity remains the same. As the gas cools its enthalpy (internal energy) decreases. That decrease in enthalpy is converted to kinetic energy, thus the velocity increases.
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Thank you. My first attempt gave 0.25 K :-[
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Thanks Ka9Q! The signals to fire the thrusters....do they involve fixed amount of fuel, or is it astronaut skill?
I know this is an old thread, but I just noticed you asked a question that I never answered.
The RCS thrusters had constant thrust -- 100 lbf each. Impulse was controlled by how long you held the propellant valves open. You could also effectively vary thrust by rapidly cycling them with a particular duty cycle. But because the valves are electromechanical, and because it takes finite time for the propellants to mix, ignite and build thrust, there is a practical "minimum impulse" for each firing. The computer took this into account when computing which thrusters to fire, when and for how long. This is one of the main reasons for "dead bands", an allowable margin of attitude error within which the computer would not try to correct the error further.
The computer could be instructed to maintain a given attitude within the dead band, which it would do as efficiently as possible, or an astronaut could fire them manually in which case fuel consumption depended on his skill. Transposition and docking was performed manually shortly after TLI, and the CMPs seemed to compete with each other for bragging rights as to who could do it with the least amount of gas.
Thanks again. One of the more interesting upshots of this thread, is finding out that the film speed is not accurate, due to the way it is rendered into digital from 6fps.
I wonder whether there is a site out there that allows interactive docking just like the Lunar Module landing simulator online. :)
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I wonder whether there is a site out there that allows interactive docking just like the Lunar Module landing simulator online. :)
Not exactly online, but some of the Apollo add-ons to Orbiter do this.
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I wonder whether there is a site out there that allows interactive docking just like the Lunar Module landing simulator online. :)
Not exactly online, but some of the Apollo add-ons to Orbiter do this.
Second that. I spent waay too much time with http://nassp.sourceforge.net/wiki/Main_Page (http://nassp.sourceforge.net/wiki/Main_Page) three years ago, but it looks like that project hasn't been updated since 2012. It includes a http://www.ibiblio.org/apollo/ (http://www.ibiblio.org/apollo/) complete AGC emulation (possibly just block 1 though), simulation of a significant portion:
(http://nassp.sourceforge.net/w/images/1/1e/CSM_systems_diagram.gif) of the CSM's internal electrical, thermal, and life support functions, limited failure simulations, and of course realistic fuel and acceleration constraints on all the hardware. It is a bit of a handful, though - check out the http://nassp.sourceforge.net/wiki/Apollo_7_Quickstart_-_Launch (http://nassp.sourceforge.net/wiki/Apollo_7_Quickstart_-_Launch) "quickstart" checklist just for getting A7 to orbit.
For a shallower learning curve, http://www.orbiterwiki.org/wiki/AMSO (http://www.orbiterwiki.org/wiki/AMSO) AMSO is more of an "interactive movie" of an Apollo mission. Much more approachable, but also less depth. It appears to have been updated significantly since 2011, so maybe the "depth" portion has improved.
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complete AGC emulation (possibly just block 1 though), simulation of a significant portion:
of the CSM's internal electrical, thermal, and life support functions, limited failure simulations, and of course realistic fuel and acceleration constraints on all the hardware.
Or at least pre-Apollo 14, when an emergency battery was added to the SM if the fuel cells go offline for some reason. The new battery was identical to one of the LM's descent batteries (400 Ah @ 28V) so it was nowhere near enough to run a whole mission, but it could help keep the CM's entry batteries from being depleted while a fuel cell problem is being worked.
The Skylab models had three such batteries to supply all of the power needed by the CSM from undocking to CM/SM separation, as the fuel cells were dead by the end of the mission.
I've wondered how high the energy density of a battery would have to be to be lighter than the fuel cells, their reactants, tanks and plumbing. On the other hand, an oxygen supply would still be needed for the cabin atmosphere anyway, along with a water supply. The fuel cells produced water as a byproduct, and it could be used both for drinking water and for cooling.
It depends on mission duration, but I suspect that you need solar panels to do better than fuel cells; even lithium primary batteries probably wouldn't cut it alone for more than a very short mission.
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How much information has NASA released about Orion? I'm particularly interested in how its power consumption and thermal load compares with the Apollo CSM. I think it could be considerably less with modern electronics.
Computers, navigation, instrumentation and communications should draw considerably less. The biggest remaining item in the communications power budget would be the RF power amplifier for the downlink and that would depend on the data rate. A lot of HDTV would drive it up.
The environmental control system (life support, cooling) should take somewhat less, as most of its power goes to pumps and blowers that have gotten only moderately more efficient since Apollo.
Lighting should also take somewhat less if LEDs replace Apollo's fluorescent lighting. OTOH, Apollo used glow-in-the-dark radioactive lighting for its switches, which is out of style today.
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The environmental control system (life support, cooling) should take somewhat less, as most of its power goes to pumps and blowers that have gotten only moderately more efficient since Apollo.
True that basically nothing has improved as much as computers in terms of power consumption, but would a contemporary spacecraft have any need for the inverter system and 115 V AC power? As near as I can tell the primary purpose of the system was to operate fans, and modern electronically commutated fan motors would handle the "inversion" for this purpose locally and far more efficiently.
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The environmental control system (life support, cooling) should take somewhat less, as most of its power goes to pumps and blowers that have gotten only moderately more efficient since Apollo.
True that basically nothing has improved as much as computers in terms of power consumption, but would a contemporary spacecraft have any need for the inverter system and 115 V AC power? As near as I can tell the primary purpose of the system was to operate fans, and modern electronically commutated fan motors would handle the "inversion" for this purpose locally and far more efficiently.
What about igniters for the ascent rocket motor. Did they have a banger-box (HEIU) like a jet engine. Most banger boxes I ever worked on was 115v 400Hz. There were 28v ones but they were quite a bit heavier.
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The igniters on the hypergolic ascent engine? ;)
AFAIK, the only AC systems were on the CM. (We need more initialisms!)
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The igniters on the hypergolic ascent engine? ;)
OK, so you are saying that there is no electrical ignition of that motor?
(Please note: What I know about the fine details as regards rocket motors can be scratched on an aspirin with a prybar.....in caps!)
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What about igniters for the ascent rocket motor. Did they have a banger-box (HEIU) like a jet engine. Most banger boxes I ever worked on was 115v 400Hz. There were 28v ones but they were quite a bit heavier.
I don't know what a banger-box is, but the ascent engine on the Apollo LM is hypergolic, like every rocket engine on the CSM and LM, so all you need to start it is to pressurize the propellant tanks and open the valves into the engine. No igniter needed as the propellants ignite spontaneously on contact.
Many hypergolic engine propellant valves are electrically operated with some sort of boosting. The LM ascent engine uses pressurized fuel as a hydraulic fluid to actually move the valves. I'm familiar with a MBB 400N engine in which electrically operated pilot valves gate pressurized helium to operate a second set of valves that actually let the propellants flow.
In any event, the electrical energy requirements of a spacecraft chemical propulsion system are minimal because the engines are operated for such a short time. The largest drains of the Apollo propulsion systems were probably the gimbal motors on the CSM's SPS and the LM's DPS that steered the engines in the desired direction. The SPS gimbals were started a few minutes into ascent from KSC so the SPS could be used to achieve a contingency earth orbit in case the S-IVB shut down early. They drew so much power that the CM's entry batteries had to supplement the fuel cells and be recharged later. I can't find how much power the LM DPS gimbals drew, but it's probably less than the SPS simply because the engine was smaller.
An electric spacecraft propulsion system (e.g., plasma or ion rocket) is a completely different animal; there the electrical power requirements drive the entire electrical power system design.
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OK, so you are saying that there is no electrical ignition of that motor?
No. The word hypergolic, by definition, means that the two propellants ignite spontaneously on contact, with no igniter (heat, spark, etc) needed.
Most hypergolic rockets use some form of hydrazine as fuel and some relative of nitric acid as the oxidizer. Their advantage isn't so much that they don't require an ignition system, but that they can be stored indefinitely as liquids at room temperature. Kerosene can, but oxygen cannot. This is important on a long space mission.
Fuels such as aniline have been used in the past, but since the mid-1960s the three most common hypergolic fuels are straight hydrazine (N2H4), monomethyl hydrazine, MMH (replace one of the hydrogens in hydrazine with a methyl group, CH3), and unsymmetrical dimethyl hydrazine, UDMH (replace both hydrogens on one end with methyl groups). All are highly toxic, carcinogenic and (obviously) flammable. At least they're water soluble, unlike gasoline and other hydrocarbons, so you can flush spills with lots of water.
The big service propulsion system engine on the CSM plus all the engines on the LM used a 50-50 mixture of straight hydrazine and UDMH called Aerozine-50; the CSM RCS used MMH.
Nitric acid itself was used as a hypergolic oxidizer for a time, but since the mid 1960s the oxidizer of choice has been nitrogen tetroxide, N2O4, an evil-looking (and smelling), highly toxic reddish-brown gas or liquid that's one of the major components of photochemical smog. If you look at videos of the recent Proton launch failure in Russia you'll see a big reddish-brown cloud on the edges of the fireball; that's nitrogen tetroxide.
Hypergolic fuels aren't much used in launch vehicles anymore (the Proton is one of the last, and I think the North Koreans are using them) because their extreme toxicity and reactivity make them difficult and expensive to handle. Technicians loading them into spacecraft have to wear special "scape" pressure suits in case of a leak. The Apollo ASTP crew nearly died from nitrogen tetroxide poisoning during descent when some got sucked into the cabin through an open vent. It's so toxic because it forms nitric acid on contact with water, which then eats your lungs out from the inside.
NASA has sponsored development of replacement propellants for some time, and most of the promising combinations use nitrous oxide, N2O as oxidizer. Despite being made of the same elements as nitrogen tetroxide, nitrous oxide is vastly less toxic and corrosive; in fact, it's widely used as an anesthetic. But it's usually not hypergolic, so any engines that use it will need an ignition system. Its big advantage over liquid oxygen is that, like nitrogen tetroxide, it can be stored as a liquid at room temperature but under considerably more pressure.
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What about igniters for the ascent rocket motor. Did they have a banger-box (HEIU) like a jet engine. Most banger boxes I ever worked on was 115v 400Hz. There were 28v ones but they were quite a bit heavier.
I don't know what a banger-box is, but the ascent engine on the Apollo LM is hypergolic, like every rocket engine on the CSM and LM, so all you need to start it is to pressurize the propellant tanks and open the valves into the engine. No igniter needed as the propellants ignite spontaneously on contact.
Many hypergolic engine propellant valves are electrically operated with some sort of boosting. The LM ascent engine uses pressurized fuel as a hydraulic fluid to actually move the valves. I'm familiar with a MBB 400N engine in which electrically operated pilot valves gate pressurized helium to operate a second set of valves that actually let the propellants flow.
In any event, the electrical energy requirements of a spacecraft chemical propulsion system are minimal because the engines are operated for such a short time. The largest drains of the Apollo propulsion systems were probably the gimbal motors on the CSM's SPS and the LM's DPS that steered the engines in the desired direction. The SPS gimbals were started a few minutes into ascent from KSC so the SPS could be used to achieve a contingency earth orbit in case the S-IVB shut down early. They drew so much power that the CM's entry batteries had to supplement the fuel cells and be recharged later. I can't find how much power the LM DPS gimbals drew, but it's probably less than the SPS simply because the engine was smaller.
An electric spacecraft propulsion system (e.g., plasma or ion rocket) is a completely different animal; there the electrical power requirements drive the entire electrical power system design.
OK, thanks, that clear thing up
I had no idea what a hypergolic motor was let along how it works.
NOTE: Banger box is slang for a High Energy Ignition Unit. Think spark plugs for a jet engine, except MUCH more powerful. If you're ever near a jet or turboprop engine when it starts, you can hear the banger box firing away as the turbine become to wind up.
Here a small igniter being tested
The sparks are typically 15KV or greater
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AFAIK, the only AC systems were on the CM. (We need more initialisms!)
The CSM and LM both had 3-phase 115V 400Hz AC power inverters. It was used mostly to drive AC induction motors, but also to produce DC voltages higher than 28V (e.g., in the CM's entry battery charger).
There's nothing really wrong with AC per se. It's just that there's not much need for it anymore on a spacecraft. Even during Apollo the brushless DC motor was already starting to displace the conventional AC induction motor. (A brushless motor is just a 3-phase AC induction motor with its own inverter.) DC-DC converters are now efficient and lightweight, and can minimize I2R losses and wire weight when sending bulk power over lengthy wires.
28V DC was used on Apollo because it's an aviation standard since forever, but it's rather low for a large spacecraft like the ISS so the non-Russian segment standardized on 125V DC for distribution with the solar arrays operating at around 160V. The Russians still use 28V DC.
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Here a small igniter being tested
Ah so. Yeah, I'd say they're pretty powerful.
I don't know large rocket engines as well as spacecraft hypergolics, but I know that some that use non-hypergolic propellants (e.g., RP-1 kerosene + LOX) still used hypergolic igniters. The F-1 on the Saturn V first stage used a cartridge containing a mixture of triethylboron and triethylaluminum. This wicked stuff burns on contact with air to say nothing of LOX. Of course, this means the engine can only be started once but that traditionally hasn't been a problem in a launch vehicle.
The J-2 engine on the S-IVB (though not the S-II) had to be restartable so it did use a spark igniter. (This failed on Apollo 6.) When a spacecraft needs a parking orbit or coast period, most launch vehicles still use hypergolic upper stages to make restarts easy but I think that's starting to change. And SpaceX will restart its first stage kerosene/LOX engines so it can recover them in a soft landing.
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Spark plugs don't work very well in a vacuum. ;D
So yes, large (pump-fed) liquid- and solid-fueled rockets use hypergolic igniters or similar. I've had two separate visitors to the state in the past four weeks, and that has meant two trips up to the ATK rocket park where you can see the length of the grains in some of these large SRM casings. The igniter typically has to produce a flame that travels the length of the grain's internal cavity within a few milliseconds.
I need to take a picture of the shuttle SRM segment without its grain. (With a banana for scale.)
I've spent substantial time around jet engines hearing them start up. I'm quite familiar with the sound of the igniters, but that's the first time I've heard the equipment called a "banger box." I'll add that to my Jay-is-cool vocabulary.
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I normally don't gush, but anyone interested in jet propulsion should check out AgentJayZ's YouTube channel. He's a jet engine mechanic in Canada, and he takes the time to really get into the (literal) nuts and bolts of how jet engines are built.
https://www.youtube.com/channel/UCh57rwk3ySElDpzgCDLh9KA
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It depends on mission duration, but I suspect that you need solar panels to do better than fuel cells; even lithium primary batteries probably wouldn't cut it alone for more than a very short mission.
A metal fuel cell of some variety (something like an aluminum or liquid alkali metal fuel cell) might perform better by avoiding the requirement for high pressure gaseous hydrogen or deeply cryogenic liquid hydrogen tanks. It would definitely be easier to use for long mission durations.
NASA has sponsored development of replacement propellants for some time, and most of the promising combinations use nitrous oxide, N2O as oxidizer. Despite being made of the same elements as nitrogen tetroxide, nitrous oxide is vastly less toxic and corrosive; in fact, it's widely used as an anesthetic. But it's usually not hypergolic, so any engines that use it will need an ignition system. Its big advantage over liquid oxygen is that, like nitrogen tetroxide, it can be stored as a liquid at room temperature but under considerably more pressure.
Though as Virgin Galactic found out, high pressure nitrous oxide can detonate and is shock sensitive. This can be mitigated by mixing in other gases such as nitrogen, but this further dilutes what was already a less effective oxidizer.
One other promising propellant is an ionic liquid based on hydroxylammonium nitrate (http://en.wikipedia.org/wiki/Hydroxylammonium_nitrate), which is toxic, corrosive, and may be carcinogenic (though probably not as bad as hydrazine or N2O4), but has a high enough vapor pressure that there isn't a major fume hazard and is a monopropellant that gives better performance than hydrazine monopropellant.
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Adding to what ka9q said about hypergols:
After WW2 it was believed that ballistic missiles would be a big part of future warfare, so research was conducted to identify the most likely propellant combinations.
For long range missile everybody agreed that LOX was the optimum oxidizer. For the fuel, liquid hydrogen was obviously the best performing, but it was hard to come by, hard to handle, and had an extremely low density. Alcohol, gasoline and kerosene all worked pretty well and were easier to deal with. Kerosene would win out, but new clean-burning formulations had to be developed. And, in time, the US perfected the use of liquid hydrogen.
For uses such as JATO and short-range tactic missiles, storable propellants were desired. Available oxidizers were nitric acid, hydrogen peroxide, and nitrogen tetroxide. Hydrogen peroxide and nitrogen tetroxide both froze at temperatures likely to be experienced at high altitude or in Siberia, thus nitric acid was the leading candidate. Which fuel to use was much less clean-cut. There were many candidates that all had similar performance. Only hydrazine stood out - it was hypergolic with the prospective oxidizers, it had a high density, and its performance was better. But its freezing point (1.5o C) was higher than water.
Prior US experience was with aniline and RFNA (red fuming nitric acid). This stuff was extremely nasty and dangerous. RFNA was so corrosive that is had to be loaded just before firing. It produces dangerous and painful burns, and simply pouring it produces dense clouds of highly poisonous nitrogen dioxide. Aniline was almost as bad. Everyone agreed a better alternative was needed.
Considerable research and experimentation ensued. The first breakthrough came in 1951 when contracts were granted to Metallectro and Aerojet to synthesize certain derivatives of hydrazine and to determine their suitability as rocket propellants. The three derivatives were monomethylhydrazine, symmetrical dimethyl hydrazine, and unsymmetrical dimethyl hydrazine. The hope was that a very slight alteration to the structure might give it a reasonable freezing point without changing the energetics enough to matter.
Symmetrical dimethyl hydrazine turned out to have too high a freezing point, but the freezing points of monomethylhydrazine (MMH) and unsymmetrical dimethyl hydrazine (UDMH) were -52.4o C and -57.2o C. After determining their thermodynamic properties, MMH and UDMH were both found to be very good fuels. MMH was a little denser than UDMH and had slightly higher performance. But UDMH was less liable to catalytic decomposition and had such good thermal stability that it could be used for regenerative cooling. UDMH was also more soluble and would tolerate a large percentage of water in the fuel. Both were hypergolic with nitric acid. The final decision to concentrate on UDMH was made on economic grounds. Military specifications for UDMH were first published in 1955.
RFNA was largely hated by all who used it. Not only for the reasons previously given, but also in cold weather it would react with its aluminum storage drum to produce a slimy white precipitate that would settle to the bottom of the drum. If this sludge got into a rocket engine it would plug up the injectors. If stored in a stainless steel drum the results were even worse - the corrosion was faster and the acid's performance was seriously degraded.
Mixed acids and WFNA (white fuming nitric acid) were tried, but they weren't much better than RFNA, though at least they didn't produce the deadly clouds of nitrogen dioxide.
The fourth possibility was nitrogen tetroxide. Although it is poisonous, if you can avoid handling it in the field, this doesn't matter much. As long as nitrogen tetroxide is kept out of water it is practically noncorrosive to most metals. In fact, ordinary mild steel would do. Thus the tanks of a missile could be filled at the factory and the operators would never have to see, smell, or breathe the nitrogen tetroxide. And it was perfectly stable in storage and didn't build up any pressure. The only problem was that it froze at -9.3o C, which the military would not accept.
A major breakthrough came in 1951 when a chemist named Eric Rau thought that a coating of fluoride on stainless steel might protect it from the acid. It was found that adding 0.5% hydrofluoric acid to WFNA reduced the corrosion rate of steel by a factor of 10 or more. As it turned out, HF was even more effective at inhibiting the corrosion of aluminum, and it was just as good with RFNA as with WFNA.
With the corrosion problem solved, the military now had an oxidizer that could be loaded at the factory and had a low enough freezing point to meet their requirements. Military specifications were written up in 1954. The terms RFNA and WFNA were thrown out and the designation "Nitric Acid, Type I, II, III, and IV" were adopted. These contained 0%, 7%, 14% and 21% nitrogen tetroxide respectively. HF inhibited acid was designated Type I-A, II-A, III-A, and IV-A, and contained 0.6% HF.
It was type III-A that gradually edged out all the others and became the nitric acid oxidizer. Engineers came to call it IRFNA, or Inhibited Red Fuming Nitric Acid.
IRFNA was widely used in tactical missile, where freezing point of the propellant matters. For strategic missiles, which are fired from hardened and heated sites, nitrogen tetroxide, with its somewhat greater performance, is the oxidizer used. Nitrogen tetroxide is typically mixed with nitrogen oxide to depress the freezing point. The resulting oxidizer in called MON (mixed oxides of nitrogen) and is given a number (e.g. MON-10, MON-25) designating the percentage of NO in the mix. When "nitrogen tetroxide" is listed as the oxidizer it is likely actually some form of MON.
For the Titan missile, Aerojet developed Aerozine 50, a 50-50 hydrazine-UDMH mixture. Aerozine 50 has better performance than straight UDMH but is still stable enough to be used for regenerative cooling. Aerozine 50 was also the fuel of choice for Apollo.
For anyone wanting to read more, I recommend the book Ignition, An Informal History of Liquid Rocket Propellants by John D. Clark.
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Hypergolic fuels aren't much used in launch vehicles anymore (the Proton is one of the last, and I think the North Koreans are using them)
The Chinese still use them as well; Chang Zheng 2, 3 & 4 all used nitrogen tetroxide and UDMH in its main stages. It's my understanding that their next generation launch vehicle, Chang Zheng 5, will use LOX and kerosene.
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NASA has sponsored development of replacement propellants for some time, and most of the promising combinations use nitrous oxide, N2O as oxidizer. Despite being made of the same elements as nitrogen tetroxide, nitrous oxide is vastly less toxic and corrosive; in fact, it's widely used as an anesthetic. But it's usually not hypergolic, so any engines that use it will need an ignition system. Its big advantage over liquid oxygen is that, like nitrogen tetroxide, it can be stored as a liquid at room temperature but under considerably more pressure.
N2O is also the poorest performing oxidizer that I've run computations on. All that extra nitrogen drives up the mean molecular weight of the exhaust, thereby hurting performance.
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N2O is also the poorest performing oxidizer that I've run computations on.
Firestar Technologies is claiming an Isp of 300 s for their nitrous oxide fuel blend (NOFBX), a monopropellant consisting of an emulsified mixture of N2O and ethane, ethene or acetylene. It sure sounds explosive to me, but apparently it's pretty stable.
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N2O is also the poorest performing oxidizer that I've run computations on. All that extra nitrogen drives up the mean molecular weight of the exhaust, thereby hurting performance.
But there's plenty of N2 in the exhaust of any hypergolic rocket using a hydrazine and/or N2O4.
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It sure sounds explosive to me, but apparently it's pretty stable.
Sure does sound pretty explosive. All those pi electrons screaming out for attention from the electrophilic nitrogen. Ethene and ethyne are strong Lewis bases and one would expect such a mix to be slightly temperamental. I was quietly surprised that the mix needs a catalyst bed to initiate the reaction.
I was also quite intrigued that N2O mixes were tested by German scientists prior to World War 2, but with ammonia rather than ethene or ethyne.
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But there's plenty of N2 in the exhaust of any hypergolic rocket using a hydrazine and/or N2O4.
There's plenty of N2 but not in the proportion that N2O produces. Consider for instance N2O4 vs. N2O when burned with hydrazine.
N2O4 + 2 N2H4 --> 4 H2O + 3 N2
2 N2O + N2H4 --> 2 H2O + 3 N2
The mean molecular weight of the exhaust in the first example is (4*18+3*28)/7 = 22.3, and in the second case it's (2*18+3*28)/5 = 24.0. That's enough to make about a 4% difference in exhaust velocity, i.e. (22.3/24)1/2 = 0.964. Granted, that's not a big difference but it's there nonetheless. Every sample computation I've made in which N2O was compared against another oxidizer with the same fuel, N2O has had the worst performance.
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Every sample computation I've made in which N2O was compared against another oxidizer with the same fuel, N2O has had the worst performance.
Oh, I can believe that. But 300 s for a monopropellant (or a stable mixture of propellants) is pretty impressive, if true.
It's interesting that the relatively stable N2O has such a high enthalpy of formation: +82.05 kJ/mol vs +50.63 kJ/mol for straight hydrazine (yes, I know their molecular weights aren't the same). Even N2O4 is only +9.16 kJ/mol.
N2O does have the advantage of not producing ammonia when it decomposes. Ammonia has an enthalpy of formation of -46 kJ/mol, so I can see why it's preferentially produced over the elements when hydrazine is decomposed.
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But 300 s for a monopropellant (or a stable mixture of propellants) is pretty impressive, if true.
300 s for a monopropellant sounds awfully high. I'd be interested to know exactly what concoction they say will give that. A few years ago I did some calculations for straight N2O and came up with only about 170 s. Hydrazine as a monopropellant yields about 230-240 s.
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But 300 s for a monopropellant (or a stable mixture of propellants) is pretty impressive, if true.
300 s for a monopropellant sounds awfully high. I'd be interested to know exactly what concoction they say will give that.
It's a claim by Freestar, and it sounded awfully high to me too. I also had trouble believing that their mixture wouldn't be dangerously explosive. Still, with the right fuel N2O (bipropellant or mixed monopropellant) could provide pretty good performance. How about acetylene dissolved in acetone or acetonitrile or some other solvent that would get its volume density acceptably high without pressurizing it too much?
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Bob, as a baseline how do propane (or LPG) and N2O perform as bipropellants? I've always thought they would be a good choice for amateur high-power rocketry since both are readily available, liquids at room temperature under reasonable pressure that can be used instead of pumps, and essentially stable and nontoxic.
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... a monopropellant consisting of an emulsified mixture of N2O and ethane, ethene or acetylene.
I'm not sure that's really a monopropellant. It might be stored in one tank, but both an oxidizer and a fuel are present. N2O is an oxidizer and ethane, ethene and acetylene are fuels. Surely combustion will occur as if we had a bipropellant system. This would explain the high specific impulse.
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I'm not sure that's really a monopropellant.
Well...yes. I had the same thought. But it can be treated as if it were a monopropellant, passed over a catalyst bed to get it going and then fed into a nozzle much like straight hydrazine.
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Bob, as a baseline how do propane (or LPG) and N2O perform as bipropellants? I've always thought they would be a good choice for amateur high-power rocketry since both are readily available, liquids at room temperature under reasonable pressure that can be used instead of pumps, and essentially stable and nontoxic.
I've thought the same thing. I believe I ran some numbers on this in the past but I don't remember what I got. I'll check again and post the results (probably tomorrow).
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One practical complication with N2O that some of the hybrid enthusiasts talk about is that its critical temperature is just above room temperature (like CO2, though at somewhat lower pressure) so its pressure is a sharp, nonlinear function of temperature.
That pressure is about 50 bar, so the tank would be too heavy for a really large rocket. Propane is only about 8 bar. Still, N2O does seem to be the oxidizer of choice for hybrid rockets, so it can't be too terribly difficult to work with.
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It would be monopropellant only in the same sense that ATK's solid fuel is "mono" propellant. Their stuff looks like a pretty homogeneous batch of goop, but it's just a composition of fuel and oxidizer that are chemically distinct until combustion appears.
Monopropellant is traditionally a single compound such as a cold-gas propellant, or something like hydrogen peroxide that undergoes a (possibly catalyzed) energetic decomposition. Still 300 would be impressive if true.
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Like the thrusters on Mercury? They were hydrogen peroxide monopropellent?
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Like the thrusters on Mercury? They were hydrogen peroxide monopropellent?
Correct. When hydrogen peroxide comes in contact with a catalyst it will decompose into hydrogen and steam and will release heat. I believe the catalyst is usually silver. It was used widely in early rocketry. Unfortunately it doesn't have a real high specific impulse, only about 150 s maximum as I recall (depends on H2O2 concentration).
Hydrazine provides much better performance as a monopropellant, about 230-240 s, but early tests with it weren't very successful. After a short period it would eat up the catalyst, making restarts impossible. This wasn't very good for a thruster that has to re-fire over and over. Eventually they found a catalyst that could endure, alumina pellets impregnated with iridium. Once that discovery was made, hydrazine pretty much became the standard. Hydrazine will decompose into some combination of ammonia, nitrogen and hydrogen. The amount of ammonia dissociation effects the temperature and performance of the engine. By varying the design of the catalyst chamber it is possible to control the amount of ammonia dissociation to get the desired performance.
Another type of monopropellant is cold gas, usually nitrogen. This is just pressurized gas released through a nozzle without any decomposition. The performance is low but no heat is generated, which is sometimes more important than high thrust. For example, a manned maneuvering unit.
I've heard of nitrous oxide as a monopropellant but I'm not familiar with any spacecraft that have actually used it. It's much safer than hydrazine and provides better performance than hydrogen peroxide. My own calculations suggest that the specific impulse of pure N2O is about 170 s. What ka9q has presented is a proposal to mix N2O with a fuel that would be stored together. It sounds like the arrangement would resemble a monopropellant system with a single storage tank and a catalyst to initiation the reaction, but it would technically be a bipropellant with separate oxidizer and fuel that would combust inside the engine.
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Like the thrusters on Mercury? They were hydrogen peroxide monopropellent?
Yes. Cold-gas propellant is basically WALL-E with the fire extinguisher. Dirt-simple to make and fly, especially for small satellites. And extremely reliable. But you don't get much from your tank of prepressurized gas.
Most hydrogen peroxide motors are the catalyzed type. As you've noticed from the bubbles, the extra oxygen pops off easily and releases energy as a result. That's the reaction you want, but you want it to happen much, much faster and thereby increase the rate of energy release. You use pure H2O2 (not the highly-diluted mixture you buy at the druggist/chemist), and you spray it across a plate or through a screen made of some transition metal (often platinum or silver) and the decomposition reaction rate ramps up to useful levels. This is dirt-simple too, but it's that really nice potting-soil kind of dirt. Mechanically speaking its just as reliable as cold gas. But you typically want very, very pure catalysts. And the platinum ones are costly.
The hypergols are problematic because the very properties that make them useful as propellants -- chiefly their reactivity -- are the ones that cause problems with handling or system engineering. The Apollo ELS lost a parachute on Apollo 15 because the RCS hot-fire safing procedure burned through a shroud. That in-flight safing procedure was made desirable by the difficulty in handling a fueled CM after splashdown.
Even so, you can never fully empty a propellant feed system simply by running it to fuel "depletion" because the limitations of tanking and plumbing mean some residue always remains. And one of the post-landing CM safing crews suffered injury from just the very little bit of hypergols remaining in the CM RCS. That stuff is just plain evil. I think USAF lost a missile crew or two due to propellant leaks or tanking/detanking mishaps.
Finding propellant formulations that are energetic without being viciously toxic, and which let us have simple, reliable engines is why rocket science is still a going concern.
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Cold-gas propellant is basically WALL-E with the fire extinguisher.
Or Sandra Bullock.
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Don't get me started, Bob. ;D
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Correct. When hydrogen peroxide comes in contact with a catalyst it will decompose into hydrogen and steam and will release heat. I believe the catalyst is usually silver.
A little bit like this:
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You use silver with moderate concentration H2O2 solutions, such as for naval torpedo turbine motors. With high-concentration propellant, you need platinum because it has a higher melting point than silver and isn't contaminated by the stabilizers you have to put into high-concentration propellants. Rocketry generally favors high concentrations due to mass budgets and the ability to delay tanking until just prior to launch.
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At school we used potato or carrot.
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At school we used potato or carrot.
Saturated solution of potassium iodide is a great one, much cleaner than manganese (IV) oxide. Yeast also sets it going too.
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At school we used potato or carrot.
Which can also be used as onboard batteries. :D
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You use silver with moderate concentration H2O2 solutions, such as for naval torpedo turbine motors.
The Royal Navy used such torpedo turbines, but became concerned about the instability of H2O2 and potential oxygen build up. I believe they changed their motors. The Russians did not follow suit, and if I recall it is strongly believed that the likely cause of the Kursk disaster was a result of a fire fueled by oxygen from torpedo motors.
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Indeed, a motor that produces a strong oxidizer as an exhaust is not necessarily good in confined spaces.
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Indeed, a motor that produces a strong oxidizer as an exhaust is not necessarily good in confined spaces.
It happened aboard HMS Sidon (http://en.wikipedia.org/wiki/HMS_Sidon_(P259)).
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And perhaps USS Scorpion.
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And perhaps USS Scorpion.
Interesting reading about the Scorpion. I always found the Indianapolis story interesting, despite the terrible consequences that followed its sinking and the cargo it carried to Tinian.
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My good friend's grandfather was a USS Indianapolis survivor. The only story he ever told us about that vessel was about having his wisdom teeth extracted by the ship's doctor. He has passed away now, but about three years ago I read a book about the tragedy and came away thinking what a stupid way to run a navy.
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Even so, you can never fully empty a propellant feed system simply by running it to fuel "depletion" because the limitations of tanking and plumbing mean some residue always remains. And one of the post-landing CM safing crews suffered injury from just the very little bit of hypergols remaining in the CM RCS. That stuff is just plain evil. I think USAF lost a missile crew or two due to propellant leaks or tanking/detanking mishaps.
Finding propellant formulations that are energetic without being viciously toxic, and which let us have simple, reliable engines is why rocket science is still a going concern.
A good book on the subject is Command and Control by Eric Schlosser. Story of a real bad day at an Arkansas Titan base.
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Bob, as a baseline how do propane (or LPG) and N2O perform as bipropellants? I've always thought they would be a good choice for amateur high-power rocketry since both are readily available, liquids at room temperature under reasonable pressure that can be used instead of pumps, and essentially stable and nontoxic.
I tested liquid nitrous oxide (N2O) with liquid propane (C3H8). It seems to work moderately well if we can run the engine at a fairly high pressure. For example, I selected 34 atm (500 psi) and got a specific impulse of 235.8 s (sea level) with the nozzle expanded to 1 atm pressure. Of course that's a very high pressure to run at for a pressure-fed system. If we run at a pressure of 7 atm (which seems to be typical for many pressure-fed systems), we get an Isp of only about 185 s. That's pretty poor; we'd be better off sticking with solid propellants.
On the other hand, N2O and C3H8 seems viable for a low-pressure, high-expansion ratio engine. I selected a combustion chamber pressure of 7 atm and a expansion ratio of 50:1. Under these conditions I compute a specific impulse of 290.5 s (vacuum). However, I doubt an amateur would ever need an engine with a 50:1 expansion ratio.
Firestar Technologies is claiming an Isp of 300 s for their nitrous oxide fuel blend (NOFBX), a monopropellant consisting of an emulsified mixture of N2O and ethane, ethene or acetylene. It sure sounds explosive to me, but apparently it's pretty stable.
I tested all of these combinations, however I couldn't easily find the thermodynamics properties for liquid phase. Instead I ran the computations assuming gaseous reactants, which results in higher specific impulse. Again I used a pressure of 7 atm and a 50:1 expansion ratio. I first ran the computations for gaseous N2O and C3H8 so we can see how it compares to the liquid phase computations.
N2O + C3H8 ---> 297.6 s (vacuum)
We see that the liquid phase is 2.4% lower. We can probably assume about the same for the other propellants. Under identical conditions, here's what I get for the other combinations:
N2O + C2H6 ---> 299.0 s (vacuum)
N2O + C2H4 ---> 301.8 s (vacuum)
N2O + C2H2 ---> 312.1 s (vacuum)
So we can see that 300 s is definitely feasible.
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If they store acetylene and nitrous oxide already pre-mixed, how do they keep the combustion from running back up the line and into the tank?
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You use silver with moderate concentration H2O2 solutions, such as for naval torpedo turbine motors. With high-concentration propellant, you need platinum because it has a higher melting point than silver and isn't contaminated by the stabilizers you have to put into high-concentration propellants. Rocketry generally favors high concentrations due to mass budgets and the ability to delay tanking until just prior to launch.
IIRC, Armadillo Aerospace's early experiments were with HTP using silver-plated ball bearings for the catalyst. They always seemed to be having trouble getting good flow, decomposition rate, and durability. I think the difficulty of getting reliable performance and the difficulty of obtaining high-concentration hydrogen peroxide drove them to liquid oxygen for their later stuff.
I really wonder about the stability of nitrous oxide fuel blend. Nitrous oxide is widely considered "pretty stable", but is capable of explosive decomposition. The dilution might stabilize it, but diluting it with fuel makes me a bit nervous. There's presumably some degree of insult that'll set it off...so how much does it take? Also, what substances catalyze the reaction? HTP is tricky to handle because of the extreme cleanliness requirements required to keep it from decomposing into water and oxygen...potentially in a dangerous runaway reaction. I also wonder how stable the emulsion is...is this something you have to use just after you mix it?
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Here's some information:
Surrey Research on Nitrous Oxide Catalytic Decomposition for Space Applications (http://digitalcommons.usu.edu/cgi/viewcontent.cgi?article=2109&context=smallsat)
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Correct. When hydrogen peroxide comes in contact with a catalyst it will decompose into hydrogen and steam and will release heat.
Hydrogen peroxide scares me. Almost anything will catalyze its decomposition, whether you want it to or not. And since heat is also a good catalyst, and the decomposition of the pure stuff liberates enough heat to turn the resulting water into superheated steam... You see the problem. How many submarines has this stuff sunk?
In some ways I'd almost prefer to handle hypergols. They're reasonably stable (or at least N2O4 and the organic hydrazines are) so they aren't likely to blow up as long as you keep them apart. So you can protect yourself pretty well with scape suits. There's probably nothing you could wear to protect you against a large tank of high test H2O2 that decides to go up.
At least H2O2 is more environmentally benign. You can hose down a spill with plenty of water, and the stuff is unstable enough that it will probably decompose into harmless water and oxygen by the time it gets out of your fueling facility. Hypergol spills are not usually welcomed by the nearby communities. Or the EPA.
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Hydrogen peroxide scares me. Almost anything will catalyze its decomposition...
Such as shaking the bottle. A marked hidden risk for launches where acoustic loading is the design driver (and often overlooked until the shake table).
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N2O + C2H2 ---> 312.1 s (vacuum)
My eyebrows went up when I saw this as one of Firestar's mixtures. Acetylene would make a great rocket fuel if not for its habit of exploding at pressures above 2 bar absolute. Even if you stored it at low pressure, which would mean a very big tank, you couldn't pump it into an engine. This instability is why tanks of it actually dissolve it in a solvent like acetone.
It would be ironic (and surprising) indeed if dissolving acetylene in an oxidizer like N2O actually stabilized it, given that N2O has a vapor pressure of 51.5 bar at 20C, far above the explosion pressure for pure acetylene.
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Speaking of explosive nitrous oxide mixtures, I vividly remember an incident in Berkeley Heights, NJ when I lived there in the 1980s. A small lab in a commercial part of town specialized in analyzing gas mixtures for the semiconductor industry. Silane, SiH4 is commonly used in fabrication, and it is often mixed with non-reactive carrier gases (e.g., helium) to dilute it. Several gas supply companies will mix up whatever you want to order.
Gould Semiconductor in Idaho had trouble with a cylinder of this stuff so they shipped it cross-country to New Jersey for analysis. The technician soon realized that the cylinder actually contained a mixture of silane and nitrous oxide. He tried to move the cylinder outside, but it blew up before he could do so. He and several others were killed and the building was set on fire. Because additional tanks with unknown contents were nearby, the entire surrounding area was evacuated for a day or so while they let the fire burn out.
A week or so later, I drove by the burned-out building. Tacked on the remains of the door was a nice, new crisp piece of paper: a violation notice from the EPA. Nice to know they're on the job.
Here are some old news stories about the incident. There were evacuations and bomb squad calls at several other places where these cylinders might have been shipped.
http://www.apnewsarchive.com/1988/Gas-Distributor-In-Fatal-Explosion-Issues-Nationwide-Warning/id-59352939c678357731c90cd23b23f5da
http://www.nytimes.com/1988/03/18/nyregion/3-die-in-blast-at-a-jersey-testing-plant.html
http://articles.latimes.com/1988-03-24/news/mn-388_1_gas-canisters
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I tested liquid nitrous oxide (N2O) with liquid propane (C3H8). It seems to work moderately well if we can run the engine at a fairly high pressure. For example, I selected 34 atm (500 psi) and got a specific impulse of 235.8 s (sea level) with the nozzle expanded to 1 atm pressure. Of course that's a very high pressure to run at for a pressure-fed system. If we run at a pressure of 7 atm (which seems to be typical for many pressure-fed systems), we get an Isp of only about 185 s. That's pretty poor; we'd be better off sticking with solid propellants.
Why not run at the higher pressure? N2O has a vapor pressure of 51.5 bar @ 20C, so you might as well use it. Propane is only about 8 bar, but you could conceivably use N2O's much higher vapor pressure to push it out. As long as the propellants leave their tanks as liquids, the propane tank would absorb no heat and relatively little heat would be absorbed by the N2O tank to boil enough liquid to make the gas to fill the ullage space.
Edited to add: Or you could use a small amount of CO2 to pressurize the propane tank. Its vapor pressure is roughly the same as N2O, a little higher actually.
Despite what Firestar says, I'd want a bladder or something to keep them from mixing in the tank...
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Why not run at the higher pressure? N2O has a vapor pressure of 51.5 bar @ 20C, so you might as well use it. Propane is only about 8 bar, but you could conceivably use N2O's much higher vapor pressure to push it out. As long as the propellants leave their tanks as liquids, the propane tank would absorb no heat and relatively little heat would be absorbed by the N2O tank to boil enough liquid to make the gas to fill the ullage space.
Solid or hybrid propellants just sound like a much simpler solution for about the same specific impulse.
However, if we're going to use N2O to pressurize the fuel, then fuel wouldn't necessarily have to be propane. Perhaps something else would produce a higher specific impulse that is still relatively safe to handle.
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You have a very good point there. What's the Isp of N2O oxidizing acrylic or PVC, two of the more popular hybrid fuels? (I think I'd prefer acrylic as it wouldn't generate HCl or possibly chlorinated hydrocarbons).
One of the main problems with these hybrids is igniting them. N2O is so stable that it can be difficult. The units I saw in the desert maybe 10 years ago were lit with steel wool and oxygen gas. That seemed a little crude, but it worked. Sort of.
I like Surrey's heated catalyst, but an arcjet seems like it should be even better. They make a big deal about being able to turn off the power after ignition occurs, and I can see how that's important for a small satellite, but if the power can be found I'd think keeping an arc on during the burn would increase Isp substantially. It would be a hybrid (so to speak) of chemical and electric propulsion.
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Well, thanks everyone.
As a result of this thread I've now learned about both Bayesian Search Theory and SCAPE suits.
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Why not run at the higher pressure? N2O has a vapor pressure of 51.5 bar @ 20C, so you might as well use it. Propane is only about 8 bar, but you could conceivably use N2O's much higher vapor pressure to push it out.
Well, higher pressure tanks are heavier. I don't know what tank pressures are typical for pressure-fed systems, it presumably depends on size.
Here's a useful reference: http://encyclopedia.airliquide.com/Encyclopedia.asp?LanguageID=11&CountryID=19&Formula=&GasID=55
The vapor pressure varies widely with temperature, and there's a critical point at 36.37 °C, 72.45 bar. It might be a bit...optimistic to keep it as a liquid in a 51.5 bar tank.
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Yes, the vapor pressure of nitrous oxide varies sharply with temperature because it's close to its critical point. The same is true for carbon dioxide.
I wonder how large hybrid rockets like those on Space Ship One got around the oxidizer tank weight problem.
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Yes, the vapor pressure of nitrous oxide varies sharply with temperature because it's close to its critical point. The same is true for carbon dioxide.
I wonder how large hybrid rockets like those on Space Ship One got around the oxidizer tank weight problem.
They didn't. They use gaseous nitrous oxide, and don't get much performance from it. SS1 only had to hop up above 100 km altitude, and it appears SS2 can't even do that. There's rumors that they're looking at all-liquid systems now.
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I don't know what tank pressures are typical for pressure-fed systems, it presumably depends on size.
Here are a couple examples.
LM DPS
Helium regulator inlet pressure: 320 to 1750 psia
Helium regulator outlet pressure: 245 +/- 3 psia
Nominal propellant tank ullage pressure: 235 psia (full throttle)
Propellant tank proof pressure: 360 psia
Nominal combustion chamber pressure: 103.4 psia
LM APS
Helium regulator inlet pressure: 400 to 3500 psia
Helium regulator outlet pressure: 176-190 +/- 4 psia (depending on which regulator)
Nominal propellant tank ullage pressure: 184 psia
Propellant tank proof pressure: 333 psia
Injector inlet pressure: 145 psia (steady-state operation)
Nominal combustion chamber pressure: 120 psia
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Here are a couple examples.
LM DPS
Helium regulator inlet pressure: 320 to 1750 psia [22-120.7 bar]
Helium regulator outlet pressure: 245 +/- 3 psia [16.9 +/- 0.2 bar]
Nominal propellant tank ullage pressure: 235 psia (full throttle) [16.2 bar]
Propellant tank proof pressure: 360 psia [24.8 bar]
Nominal combustion chamber pressure: 103.4 psia [7.13 bar]
LM APS
Helium regulator inlet pressure: 400 to 3500 psia [27.6 - 241 bar]
Helium regulator outlet pressure: 176-190 +/- 4 psia (depending on which regulator) [12.1 - 13.1 +/- 0.3 bar]
Nominal propellant tank ullage pressure: 184 psia [ 12.7 bar]
Propellant tank proof pressure: 333 psia [23 bar]
Injector inlet pressure: 145 psia (steady-state operation) [10 bar]
Nominal combustion chamber pressure: 120 psia [8.27 bar]
There. Now I can read it.
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They didn't. They use gaseous nitrous oxide, and don't get much performance from it. SS1 only had to hop up above 100 km altitude, and it appears SS2 can't even do that. There's rumors that they're looking at all-liquid systems now.
Really? How could you get anywhere near enough gaseous N2O on board to burn the fuel? The fuel grain usually requires a liquid N2O tank of roughly the same size to burn it.
Yes, the hype around SpaceShip One really bugged me, especially all the breastbeating about how great the private sector was. People simply don't understand that just getting to altitudes we call 'space' takes only about 4% of the energy required to stay there, i.e., achieve orbit. Nearly all of the energy in low earth orbit is kinetic, not potential.
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It might be a bit...optimistic to keep it as a liquid in a 51.5 bar tank.
Well, I suppose you could always cool it. Not the first time that's been done in rocketry.
The vapor pressure of N2O is 16 bar at 242K. That's the ullage pressure Bob gives for the LM descent engine. 242K is only -31C, not that cold.
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There. Now I can read it.
I just gave it to you as it appears in the NASA News Reference. :)
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However, if we're going to use N2O to pressurize the fuel, then fuel wouldn't necessarily have to be propane. Perhaps something else would produce a higher specific impulse that is still relatively safe to handle.
You have a very good point there. What's the Isp of N2O oxidizing acrylic or PVC, two of the more popular hybrid fuels? (I think I'd prefer acrylic as it wouldn't generate HCl or possibly chlorinated hydrocarbons).
I've run a few sample computations and it looks like just about any hydrocarbon fuel is going to give a similar performance. It looks like those with a slightly higher ratio of hydrogen to carbon are a little better (i.e. methane the best), but they all seem have specific impluses within a few seconds of each other. Alcohols are a little worse, but still in the ballpark. It looks like the only fuels that provide a significant increase in performance are the hydrazines, but that creates the safety issues that we were trying to avoid in the first place.
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Silane, SiH4 is commonly used in fabrication...
To bring the discussion full circle, Silane is hypergolic in air (pyrophoric) at STP (presumably even more readily hypergolic in pure O2), which incidentally makes pipe leaks easy to locate. Bob, what kind of performance could one get from SiH4 and the usual oxidizer suspects?
PS - Wikipedia tells me that with SiH4 as a fuel, CO2 is an oxidizer. The more you know...
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I tested liquid nitrous oxide (N2O) with liquid propane (C3H8). It seems to work moderately well
Hey, while you're at it, why not try cyanogen and hydrogen cyanide? Cyanogen's enthalpy of formation is +309.07 kJ/mol and HCN's is +109.9 kJ/mol. That's still less on a weight basis, but HCN boils at a convenient +26C while cyanogen boils at -21C so it would have to be kept under pressure to stay liquid at room temperature.
There is the minor problem of their affinities for cytochrome c oxidase, but hey, if you're willing to play with hydrazine and nitrogen tetroxide...
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To bring the discussion full circle, Silane is hypergolic in air (pyrophoric) at STP (presumably even more readily hypergolic in pure O2)...
Well, there's experimental evidence to show that it is not pyrophoric with N2O...
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I just gave it to you as it appears in the NASA News Reference. :)
Oh, I know. I marvel that we were ever able to reach the moon using English units. How come the hoaxers don't pick up on this? I mean, it's a bigger handicap than 100 kHz computers with 30 kB of ROM.
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Well, there's experimental evidence to show that it is not pyrophoric with N2O...
Provided the mixture is not shocked, heated, agitated, moved, catalyzed, or acknowledged.
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After I posted that, I realized that cyanogen would probably not be a good rocket fuel despite its extremely high flame temperature in oxygen. It contains no hydrogen. HCN would probably be better, and it's probably safer than hydrazine...
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Bob, what kind of performance could one get from SiH4 and the usual oxidizer suspects?
I can't do that one. The computer program I use doesn't currently contain any silicon compounds. If I had a real need I could build the data tables, but that's too much work for a simple 'what if' experiment.
Hey, while you're at it, why not try cyanogen and hydrogen cyanide?
With what oxidizer?
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I can't do that one. The computer program I use doesn't currently contain any silicon compounds. If I had a real need I could build the data tables, but that's too much work for a simple 'what if' experiment.
If you wanted to PM me a link to a dropbox with the data table format, I might take a crack at it if I get the spare time and inclination (no promises) - the NASA-mentioned possibility of it being a possible in situ Mars fuel is intriguing (are there any other in situ hypergols)?
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After I posted that, I realized that cyanogen would probably not be a good rocket fuel despite its extremely high flame temperature in oxygen. It contains no hydrogen. HCN would probably be better, and it's probably safer than hydrazine...
HCN would be better in terms of molecular weight, but temperature is just as big a factor. Exhaust gas velocity is proportional to (T/M)1/2 so the higher the temperature and the lower the molecular mass, the better the propellant. Of course specific heat ratio is also a big factor.
I determine the best mixture ratio by experimentation, but a good starting place is to include enough oxygen to oxidize the carbon to CO and the hydrogen to H2O. For cyanogen and HCN that's,
C2N2 + O2 --> 2 CO + N2
2 HCN + 1.5 O2 --> 2 CO + H2O + N2
That's an average molecular weight of 28 for cyanogen and 25.5 for HCN. Neither is great but HCN is better.
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If you wanted to PM me a link to a dropbox with the data table format, I might take a crack at it if I get the spare time and inclination (no promises) - the NASA-mentioned possibility of it being a possible in situ Mars fuel is intriguing (are there any other in situ hypergols)?
If you go to http://www.braeunig.us/space/thermo.htm you'll see the required format at the bottom of the page under a section titled "STANJAN". However, I've done some searching and it looks like there are very few compounds to even consider. It looks like SiO2 and maybe SiO are it. With only two tables to create, I can manage that myself.
(ETA) I've already found all the data I need. I just need to get it into the right format.
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Hey, while you're at it, why not try cyanogen and hydrogen cyanide?
With what oxidizer?
N2O, of course. We're trying to find safer alternatives to the hydrazines plus N2O4, remember? ;D
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They didn't. They use gaseous nitrous oxide, and don't get much performance from it. SS1 only had to hop up above 100 km altitude, and it appears SS2 can't even do that. There's rumors that they're looking at all-liquid systems now.
Really? How could you get anywhere near enough gaseous N2O on board to burn the fuel? The fuel grain usually requires a liquid N2O tank of roughly the same size to burn it
I've seen several references to them using gaseous nitrous oxide, and I remember someone specifically stating it was gas for some reason or other, though they may have been misinformed or misquoted. The diagrams I've seen have a tank with considerably greater volume than the burned fuel. The reported temperature and pressure from the Mojave accident were impossible for a tank containing liquid (70 F/21 C and 360 psi/25 bar). Some Scaled materials refer to liquid N2O and things like slosh baffles, though.
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Hey, while you're at it, why not try cyanogen and hydrogen cyanide?
With what oxidizer?
N2O, of course. We're trying to find safer alternatives to the hydrazines plus N2O4, remember? ;D
Have you considered liquid oxygen?
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Hey, while you're at it, why not try cyanogen and hydrogen cyanide?
With what oxidizer?
N2O, of course. We're trying to find safer alternatives to the hydrazines plus N2O4, remember? ;D
To be consistent with previous calculations, I used a chamber pressure of 7 atm and a 50:1 expansion ratio.
N2O + HCN ---> 299.1 s (vacuum)
N2O + C2N2 ---> 311.8 s (vacuum)
C2N2 is far better than I expected but that's because it burns incredibly hot, probably too hot. While everything else burns at about 3000-3400 K, N2O and C2N2 burns at about 4200 K. Of course we could run oxidizer-rich and bring the temperature down to a more manageable level, but that will lower the specific impulse and defeat the purpose. We'd be better off using acetylene.
(ETA) These calculations assume gaseous reactants.
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Hey, while you're at it, why not try cyanogen and hydrogen cyanide?
With what oxidizer?
N2O, of course. We're trying to find safer alternatives to the hydrazines plus N2O4, remember? ;D
Have you considered liquid oxygen?
I think they want a storable oxidizer.
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Have you considered liquid oxygen?
I think they want a storable oxidizer.
I believe that was the initial premise, but to be honest, I've lost track.
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Who cares? This is a great discussion. Makes me wish I'd paid more attention in Chemistry.
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I've arrived home to find that this thread is now examining the use of HCN as a fuel. I don't know whether to laugh or feel alarmed at this deviation.
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I've arrived home to find that this thread is now examining the use of HCN as a fuel. I don't know whether to laugh or feel alarmed at this deviation.
I admit to having my tongue in my cheek when I made that suggestion, but there's a method in my madness. Everybody knows HCN is toxic; it has a long and notorious history. What everybody doesn't know is that many other materials are just as toxic, or worse.
H2S, for example -- it's just about as toxic as HCN, and a major occupational risk in the oil and gas industry. Yet I still encounter teens who think it would be fun to stink up the place with some. The only real difference is that most people can smell H2S more easily at low concentrations than HCN.
Hydrazine and N2O4 are also of roughly equivalent toxicity, although they all kill or injure by different mechanisms.
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Sure, what could possibly go wrong with that?
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C2N2 is far better than I expected but that's because it burns incredibly hot, probably too hot. While everything else burns at about 3000-3400 K, N2O and C2N2 burns at about 4200 K. Of course we could run oxidizer-rich and bring the temperature down to a more manageable level, but that will lower the specific impulse and defeat the purpose. We'd be better off using acetylene.
(ETA) These calculations assume gaseous reactants.
Why not mix in some excess hydrogen? That will cool things down and lower the average molecular weight.
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I've arrived home to find that this thread is now examining the use of HCN as a fuel. I don't know whether to laugh or feel alarmed at this deviation.
HCN is about as toxic as UDMH, isn't it?
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HCN is about as toxic as UDMH, isn't it?
Yeah, but in different ways. The hydrazines are both acutely toxic and carcinogenic. HCN is infamous for being acutely lethal in sufficient concentrations, but we've evolved mechanisms to safely metabolize trace amounts because it occurs in many foods, e.g., apricot pits, apple seeds, and especially cassava. We Westerners know it mainly in tapioca pudding but it is widely consumed in the tropics.
So chronic low level HCN exposure is probably less hazardous than chronic low level hydrazine exposure.
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Who cares? This is a great discussion. Makes me wish I'd paid more attention in Chemistry.
You might say it's been a breath of fresh air.
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Bob, what kind of performance could one get from SiH4 and the usual oxidizer suspects?
To be consistent, I again used a combustion chamber pressure of 7 atm and a 50:1 expansion ratio. Here's what I got:
O2 (liquid) + SiH4 (gas) ---> 334.4 s (vacuum)
N2O4 (liquid) + SiH4 (gas) ---> 322.6 s (vacuum)
N2O (liquid) + SiH4 (gas) ---> 298.3 s (vacuum)
Surprisingly I got almost the exact same Isp using N2O regardless of whether it was liquid or gas. Not quite sure why it worked out that way.
What I found a bit unusual with this fuel is that the mixture ratio needed to attain maximum specific impulse is extremely fuel-rich, more so than usual. It's better to leave most of the hydrogen unoxidized to drive down the molecular weight, even though this results in a significant drop in temperature. At its optimum mixture ratio the combustion temperature is only about 2700 K. With O2 the optimum formula is (ignoring dissociation),
SiH4 + 0.65 O2 --> SiO + 0.3 H2O + 1.7 H2
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What I found a bit unusual with this fuel is that the mixture ratio needed to attain maximum specific impulse is extremely fuel-rich, more so than usual. It's better to leave most of the hydrogen unoxidized to drive down the molecular weight, even though this results in a significant drop in temperature.
It's difficult to write this without sounding sarcastic, but I genuinely found this snippet very interesting. It's counter intuitive, but makes perfect sense. After reading these discussions I'm beginning to wish I took a different career path.
I'm beginning to feel that I lost sight of why I studied physics. I took a left turn instead of a right as this is much more interesting compared to the technical route I took.
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It's difficult to write this without sounding sarcastic, but I genuinely found this snippet very interesting. It's counter intuitive, but makes perfect sense.
Yes, it can be a little counter intuitive. Adding more oxygen drives up the temperature (until reaching a stoichiometric mixture), but that results in heavier products (e.g. CO2 rather than CO). Exhaust gas velocity is proportional to (T/M)1/2, so somewhere there is an optimum point that's hot enough and light enough to produce the maximum velocity.
I generally initially balance the combustion equation with enough oxygen such that all the carbon and hydrogen produce CO and H2O. I'll then adjust the amount of oxygen up or down (usually down) until I find the optimum point. In the case of SiH4, I kept adjusting down and down, and the exhaust velocity keep going up and up. I ended with less than half as much oxygen as I started with. All rockets run fuel-rich but this was extreme.
After reading these discussions I'm beginning to wish I took a different career path.
I don't have any regrets with the career I chose, but I definitely wonder sometimes how things would have worked out had I decided to go into aerospace instead.
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H2S, for example -- it's just about as toxic as HCN, and a major occupational risk in the oil and gas industry. Yet I still encounter teens who think it would be fun to stink up the place with some. The only real difference is that most people can smell H2S more easily at low concentrations than HCN.
Having nearly killed myself with the stuff, I have to agree - it's worth noting that at fatal concentrations you can't smell H2S at all...
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I don't have any regrets with the career I chose, but I definitely wonder sometimes how things would have worked out had I decided to go into aerospace instead.
Anything you do for a living becomes your occupation, and it results inevitably into normalizing its erstwhile glamor down to daily drudgery, and associating with it all the subordinate and attendant problems that accompany anything you do for a living. If you like doing something, and you're good at it, you don't necessarily have to make it your profession in order to enjoy it. In fact, you may enjoy it more if you keep it "pure" and unsullied by occupational concerns.
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Anything you do for a living becomes your occupation, and it results inevitably into normalizing its erstwhile glamor down to daily drudgery, and associating with it all the subordinate and attendant problems that accompany anything you do for a living. If you like doing something, and you're good at it, you don't necessarily have to make it your profession in order to enjoy it. In fact, you may enjoy it more if you keep it "pure" and unsullied by occupational concerns.
Yep, I agree with that. I've heard it said that if you really love doing something, make it a hobby and not a profession. When I first started working full time I really enjoyed it and was a workaholic. After a while though it just became a job. That's when I started to develop hobbies outside of my profession.
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I don't have any regrets with the career I chose, but I definitely wonder sometimes how things would have worked out had I decided to go into aerospace instead.
In fact, you may enjoy it more if you keep it "pure" and unsullied by occupational concerns.
Absolutely. That's why I enjoy reading these boards, I learn so much. More so when hoax theorists turn up. Well, with the exception of awe130. That episode did not really stimulate much more than the blueprint issue, but I understood that already from Clavius and your treatment of Collier's claims.
My original 'regret' was nothing more than 'the stuff here amazes me, I wish I had taken it further earlier in life.'
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Having nearly killed myself with the stuff, I have to agree - it's worth noting that at fatal concentrations you can't smell H2S at all...
Phosgene is an interesting gas from a similar perspective. The troops in WW1 would run through clouds of the stuff unaffected, unlike chlorine which has an immediate action on the lungs. It was not until a few hours after exposure that troops had an onset of symptoms.
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And N2O4, for that matter. It is certainly acutely quite irritating, but the real damage doesn't appear for a day or so when the pulmonary edema sets in.
The ASTP astronauts brushed off their encounter with it after landing until they happened to mention it casually to a flight surgeon. He knew his stuff, so off they went to the ship's clinic and then a two-week stay in Honolulu for observation. Good call.
The stuff is just plain evil. It even looks evil.
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To be consistent, I again used a combustion chamber pressure of 7 atm and a 50:1 expansion ratio. Here's what I got:
O2 (liquid) + SiH4 (gas) ---> 334.4 s (vacuum)
N2O4 (liquid) + SiH4 (gas) ---> 322.6 s (vacuum)
N2O (liquid) + SiH4 (gas) ---> 298.3 s (vacuum)
Surprisingly I got almost the exact same Isp using N2O regardless of whether it was liquid or gas. Not quite sure why it worked out that way.
What I found a bit unusual with this fuel is that the mixture ratio needed to attain maximum specific impulse is extremely fuel-rich, more so than usual. It's better to leave most of the hydrogen unoxidized to drive down the molecular weight, even though this results in a significant drop in temperature. At its optimum mixture ratio the combustion temperature is only about 2700 K. With O2 the optimum formula is (ignoring dissociation),
SiH4 + 0.65 O2 --> SiO + 0.3 H2O + 1.7 H2
Awesome, thanks! Looks like similar performance to ethane, but with a cooler 10 bar boiling point (210 K vs 235 K for ethane). Not sure if SiO is stable at combustion chamber temperatures. SiO and SiO2 (silica) are both ceramics with melting points above 1600C so there could be issues with fouling the exhaust bell - another reason to run the mixture fuel-rich. I guess the advantages to a silane engine would be carbon-free fuel chemistry for in situ fuel generation and less-toxic hypergolic (maybe) operation.
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Not sure if SiO is stable at combustion chamber temperatures. SiO and SiO2 (silica) are both ceramics with melting points above 1600C so there could be issues with fouling the exhaust bell - another reason to run the mixture fuel-rich.
The calculations definitely show that some SiO2 will be present in condensed phase.
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C2N2 is far better than I expected but that's because it burns incredibly hot, probably too hot. While everything else burns at about 3000-3400 K, N2O and C2N2 burns at about 4200 K. Of course we could run oxidizer-rich and bring the temperature down to a more manageable level, but that will lower the specific impulse and defeat the purpose. We'd be better off using acetylene.
Why not mix in some excess hydrogen? That will cool things down and lower the average molecular weight.
I tried this by simply adding hydrogen without making any changes in the initial amount of N2O and C2N2. Interestingly, the relative change in T/M is nearly constant as the proportion of hydrogen is increased. That is, the specific impulse stays about the same. So while it brings the temperature down, the added complexity of a tri-propellant system doesn't seem worth it in terms of performance.
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They didn't. They use gaseous nitrous oxide, and don't get much performance from it. SS1 only had to hop up above 100 km altitude, and it appears SS2 can't even do that. There's rumors that they're looking at all-liquid systems now.
Really? How could you get anywhere near enough gaseous N2O on board to burn the fuel? The fuel grain usually requires a liquid N2O tank of roughly the same size to burn it.
Yes, the hype around SpaceShip One really bugged me, especially all the breastbeating about how great the private sector was. People simply don't understand that just getting to altitudes we call 'space' takes only about 4% of the energy required to stay there, i.e., achieve orbit. Nearly all of the energy in low earth orbit is kinetic, not potential.
Is it fair to be bugged about this, though? I was certainly one of the large number of ordinary members of the public who had no idea that the percentage was as low as 4%. If I'd been pressed for a figure I would have said something closer to 20%.
The thing is, there are all sorts of things about any profession that non-experts simply won't know. When people make uninformed statements about that profession, isn't the best response is to educate people rather than get bugged about their uninformed statements?
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Is it fair to be bugged about this, though? I was certainly one of the large number of ordinary members of the public who had no idea that the percentage was as low as 4%. If I'd been pressed for a figure I would have said something closer to 20%.
It's not just random members of the public who are spouting misinformation like this. Branson really should know better than to say nonsense like:
We have reduced the cost of somebody going into space from something like two weeks of New York’s electricity supply to less than the cost of an economy round-trip from Singapore to London.
They haven't reduced anything, they aren't getting anywhere close to orbit with their approach to "spaceflight", yet Branson is bragging about the supposed superiority of SpaceShipTwo.
For what it's worth, even the energy comparison exaggerates their achievement. The rocket equation means they don't just need 25 times the energy input. SpaceShipTwo has a mass ratio of somewhere around 2, half the takeoff mass being propellant. Reaching orbit with a single stage and similar propulsion system would require a mass ratio closer to 40-50. There's a reason SSTO concepts all use exotic materials and liquid hydrogen fuel.
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Apropos to this discussion of hybrid rockets and especially of N2O as an oxidizer, the recent SpaceShipTwo disaster got me reading up on the hazards of this stuff.
It's not quite as benign as I had thought. While remarkably stable at STP, it can be decomposed with sufficient pressure and temperature -- and its positive enthalpy of formation means it will happily continue the process once started.
The necessary pressure can be supplied by the N2O itself (because of its high vapor pressure at room temperature) and the ignition temperature can be supplied by any number of unintentional mechanisms that are often all too obvious only after an accident has occurred: adiabatic compression of gas in a pipeline, or the implosion of bubbles in a mixture of gas and liquid.
I know it's much too early to know for sure, but I already strongly suspect that autocatalytic decomposition of N2O had a lot to do with it. Scaled Composites had one prior fatal accident with the stuff in 2007 during a ground test that simply filled a tank and then expelled it through an injector, with no intended ignition. Three people were killed and more were injured.
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The latest news is that SpaceShipTwo broke apart when its feathering device accidentally deployed during acceleration. The reason why is not yet known. There apparently was no explosion.
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I've seen that too. If the rocket engine and oxidizer tank were recovered intact, then I guess they couldn't have exploded.
Was the rocket still burning when the breakup occurred? At what point?
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The latest news is that SpaceShipTwo broke apart when its feathering device accidentally deployed during acceleration. The reason why is not yet known. There apparently was no explosion.
I arrived home from work (at 7 pm :() to the news. According to ITN news, the feathering device should deploy at Mach 1.4, but deployed earlier in the flight at a little over Mach 1.
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Was the rocket still burning when the breakup occurred? At what point?
I believe that it was, but don't quote me on that. The breakup occurred just 11 seconds after detaching from its carrier aircraft.
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I've seen that too. If the rocket engine and oxidizer tank were recovered intact, then I guess they couldn't have exploded.
Was the rocket still burning when the breakup occurred? At what point?
That doesn't mean much: the engine's a big nearly indestructible mass of plastic, and the tank's a heavy duty pressure vessel. They both withstood impact with the ground and stayed individually intact. The weak point was the connection between them, an explosion outside the oxidizer tank could easily have separated them without causing much damage to either.
As far as I've heard, they still don't have a reason for the aircraft entering the feathered configuration: the mechanism was unlocked, but the feathering action was not commanded. It might have been forced by the aerodynamic environment just above Mach 1, but it's also possible that an issue with the rocket motor caused it.
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An explosion outside the oxidizer tank wouldn't be very big. The real disaster happens when a detonation wave somehow ripples through the tank itself. N2O is capable of detonating under the right (wrong?) conditions.
Scaled Composites put out a lengthy white paper on N2O safety after their fatal 2007 accident. It does seem they got very serious about it.