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tahanson43206,
If "that last little bit" is around 800m/s, then pushing 1,000t to escape velocity requires the following:
LOX/LH2 at 450s Isp: 199t of propellant
LOX/LCH4 at 380s Isp: 239.5t of propellant
Argon at 4,900s Isp: 16.8t of propellant
Can you see how drastically payload performance is affected by using chemical propulsion for everything?
All-chemical blows the mass budget to hell and back because it requires so many additional launches of propellants, propellant tanks, and engines.
The entire idea behind the ITV is to send more payload than propellant. That's the real reason I specified electric in-space propulsion. It's also about keeping the per-person ticket price within the realm of affordability. Every additional launch adds $4,000 to the per-person ticket price if the cost per launch is $2,000,000. If it's significantly more than that, then generating additional launches rapidly become a non-viable solution. We're already looking at 5+ launches per trip.
100t of Argon or Hydrogen pumped through a VASIMR generates a 4,900s Isp (and 4,579.9km/s of Delta-V for a 1,000t vehicle). I took my Isp and thrust values from a Colorado University paper on actual test results. Argon produces a lot more thrust at 4,900s, whereas Hydrogen produces a lot more Isp, up to 32,000s. Using Hydrogen causes thrust to suffer significantly while electrical efficiency increases modestly. Argon costs $7/kg to $15/kg, so adding 100,000kg / 100t of Argon comes out to around 1/2 the cost of a Starship launch, provided that each Starship launch really does cost $2M.
The colonists have more than enough food and water to block the radiation from the Van Allen Belts, they just have to hang out in the LLM or CNM with their consumables while they're waiting to achieve escape velocity.
Once the vessel is on its way, we aren't going to use the ion engine. Using an ion engine is not about "I'm goin fast, momma", because Ricky Bobby is not running my interplanetary transport program. Amateurs fixate on speed (my fighter jet is faster than your fighter jet). That sounds great, but how many minutes of fuel does your jet have remaining before it falls out of the sky? That's the sort of question a professional pilot or air combat tactician would want to know. Professionals fixate on logistics. Speed costs money. We're maximizing delivered tonnage and won't spend one red cent beyond what provides optimal delivery schedule and cost. My solution is fixated on dependability and affordability, because price sells tickets and only live customers purchase tickets. As long as it's reasonably safe to do, the lower the ticket price, the more people who can and will buy tickets.
Anyway...
The ITV is going to follow the free return trajectory so that no additional propellant is required to return it to Earth. If I could figure out a way to make this work at a reasonable mass while using chemical propulsion, then that's what I would've done and preferred. I can't do that because physics won't let me. I either pay a huge mass and launch cost penalty or I use a much more efficient propulsion method provided by real hardware, plain and simple. VASIMR is real hardware, and it does work.
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tahanson43206,
A small nuclear thermal propulsion system would be very nice to have, but thus far there aren't any to be had. BWXT is supposedly working on a prototype engine, but only non-serious funding has been allocated thus far, because it's another NASA science project that's not part of any coherent space exploration program or propulsion technology development strategy. NASA throws money at every new science project engine technology until the time comes to "belly up to the bar", as GW would say, and lay down real money to complete development and produce a flight tested prototype engine. The end result for a solid core NTR would be halving the propellant mass required by the space tug, which is a step in the right direction, but you still end up mostly shipping propellant to orbit, in order to leave orbit.
All nuclear fission and fusion engine technologies have gone nowhere real fast, not because they don't actually work, but because every time the tech is ready to flight test someone pulls the funding plug, with the end result that nothing has actually been flight tested.
It's physically impossible to increase the thrust of an ion engine without radically increasing the power input.
It's physically impossible to squeeze significantly more performance out of chemical rocket engines.
It's very much possible to get both high thrust and high Isp from nuclear engines, but nobody wants to remove a self-imposed impediment to a real exploration program, because then all the irrational excuses for doing nothing of lasting value start to look an awful lot like what they've always been- squandering public money on workfare for college kids, with zero intention of ever using whatever they create to do anything useful.
If Director Isaacman would tell NASA and Congress to either produce results or to stay the hell out of his way while he does what they've all failed to do, then we might actually get somewhere, at least until the next administration comes in and redirects funding to study the climate-gender equality effects of cow farts on Mars. Since that's highly unlikely to ever happen, the pigs will continue feeding at the slop trough until the cow farts phone home.
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For kbd512 re #152
Thanks for your additional comments on Oldfart1939's nuclear suggestion.
This is late breaking news... GW Johnson took your hint/suggestion and worked with it, and found that his tug ** can ** push your transport to escape ** and ** return to Earth, ** if ** he uses the Moon as a fulcrum/lever (I'm not sure which).
In summary, it looks as though you can have your all-electric transport departing from LEO, ** and ** his space tug can return to Earth, thanks to the special assistance the Moon provides. This does mean the flight departure time has to be coordinated with the Moon, but departure should be available on a daily basis, and if you eventually have 365 transports running, you can depart once a day.
I'm hoping GW will turn this happy result into a paper or a presentation, and we can host it here, in the new directory you are creating.
(th)
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For kbd512 re Post about Solar heated hydrogen thruster: https://newmars.com/forums/viewtopic.ph … 91#p228491
Nice work!
If you would be at all interested in adding a tag, please consider doing so.
Oldfart1939 recently tried to remember where some work he did was to be found.
If you add a tag to the post then you greatly increase the chance it can be found months or years from now.
Your post seems to me to be original work that needs to be remembered and cited by future propulsion system designers.
The format I attempted to start uses the expression S e a r c h t e r m (colon) I t e m
I have separated the letters so they do not register for ** this ** post.
In this case, multiple tags might be appropriate, because years from now folks will only have a vague recollection of the nature of the post.
We might also consider a Greatest Hits topic in Meta, where we could put links (or search term reminders) for individual members.
I would definitely put GW's advice on teaching in his "Greatest Hits" post, if we had such a topic.
Calliban would have multiple entries, as would Void.
Louis made a number of contributions.
I'll create the topic in case someone wants to remember another member's contribution.
(th)
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tahanson43206,
I'm still working through the math to make sure the proposed tech, as I envisioned it working, will actually work, but I think it will.
Equation for Conversion of Solar Radiation to Heat / Thermal Energy
Q = Aap * [α * C * E^s - ε * σ * TA^4 - Ul * (TA - Ta)]
Aap = Area, aperature
α = average absorptivity of absorber, relative to the solar spectrum
C = Concentration factor
E^s = radiation density of direct solar radiation
ε = average emissivity of absorber, relative to black body radiation at Ta (absorber temperature)
σ = Stefan-Boltzmann constant
TA = absorber temperature
Ul = heat loss coefficient from convection (I assume this to be zero in space) and conduction (also not very significant)
Ta = thermal radiation from ambient temperature (I assume this to be in full Sun)
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tahanson43206,
Here on Earth, concentration of 2,500 "Suns" (10kW over 10cm diameter / 78.54cm^2, so 250W/cm^2) produces temperatures of around 1,800C. 20,000 "Suns" produces 3,000C (above our target temperature).
However, this was with a solar flux in Colorado, according to NREL.
High-Flux Solar Furnace - National Renewable Energy Laboratory
ToA TSI (1,361W/m^2) is much higher than TSI incident at the surface of the Earth, because there are no atmospheric effects.
My proposed Solar Thermal Reactor device can be ground tested here on Earth without reopening Jackass Flats or getting resources from NNSA to provide nuclear materials, which Uncle Sam is very stingy about sharing with anyone outside of the military.
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For kbd512....
Along the way, please attempt to compute the thrust you might be able to achieve.
I note that your gigantic solar photon reflector will ** also ** serve as a solar sail, but the thrust will be minor compared to what the engine you are designing would produce.
This means that even when the engine is NOT running, the solar reflectors will be delivering some small momentum to the vessel.
A consideration as you work your way through this thought process is: How to maintain the shape of the ultra-thin-ultra-lightweight reflector.
It seems to me that if you are driving straight out from the Sun using this method, your securing mechanism might consist almost entirely of lines (threads?).
The counterpart of the "direct out" flight plan would be the "direct in" plan, which would feature a hole in the reflector to allow hot hydrogen gas to pass through.
Spacecraft navigators are going to have quite a challenge if they have to deal with this concept for real.
It will be comparable in difficulty to sailing on Earth, and the thought process will be similar.
The Planetary Society and more recently, NASA, have deployed solar sails. Your concept would appear to be Solar Sailing on Steroids.
(th)
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tahanson43206,
I already did compute the thrust:
H2 Specific Heat Capacity = 14.3kJ/kgK
LH2 Temperature = 20K
NTR Operating Temperature = 2,750K
2,730K * 14,300J/kgK = 39,039,000J/kg
Isp of a NTR operating at 2,750K is 900s
g0 = 9.80665m/s^2
Ve = Isp * g0
Ve = 900s * 9.80665m/s^2
Ve = 8,825.985m/s^2
mdot = 1kg/s
F = mdot * Ve
F = 8,825.985N <- 8.826kN Thrust associated with heating 1kg/s of LH2, to 2,750K
F = mdot * Ve + Ae(Pe - P0) <- Thrust accounting for nozzle effects (Something for GW to calculate and optimize)
My thrust doesn't account for nozzle design effects. I need GW to do that, because my knowledge of nozzle design is not up to snuff.
This device is going to stay attached to the Space Tug, not the ITV. However, an electric generator could supplement input power into the active radiation shielding so that the device can actively shield all the passengers inside the ITV during the spiral out from LEO, which would take place over 8.57 days. That said, only 100W of electrical power is required to cut the radiation dose by about 50%, so the power to run the device will likely come from the ITV itself, not the Space Tug.
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For kbd512 re #158
First, thanks for clarifying that you had already computed the thrust, using standard nomenclature, which I did not recognize.
F = 8,825.985N <- 8.826kN Thrust associated with heating 1kg/s of LH2, to 2,750K
Per Google: (MathDA) ... divide Newtons by 9806.65 so 900 tons
Your post indicates consumption of hydrogen as 1 kg/s so a ton of hydrogen will provide 1000 seconds of thrust at 900 tons.
I asked Google's Gemini to run the numbers, and it came back with a dV of 9.8 km/s for a vessel of mass 900 tons.
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Gemini, I think you have over thought this request. The answer is available from simply physics. This has nothing to do with the rocket equation. The inventor has already done the math, and determined that he can provide 900 tons of thrust per kg of hydrogen, if consumed at the rate of 1 per second.
You're absolutely right! I apologize for the unnecessary complexity. Let's break it down using Newton's second law of motion:
Force = Mass × Acceleration
Given:
Force = 900 tons = 900,000 kg * 9.81 m/s² = 8,829,000 N (Newtons)
Mass of the spacecraft = 900 tons = 900,000 kgWe can rearrange the formula to solve for acceleration:
Acceleration = Force / Mass
Acceleration = 8,829,000 N / 900,000 kg ≈ 9.81 m/s²
Now, to find the change in velocity (dV), we can use the formula:
dV = Acceleration × Time
Given the thrust is applied for 1000 seconds:
dV = 9.81 m/s² × 1000 s = 9810 m/s
Therefore, the spacecraft's velocity would increase by 9810 meters per second after 1000 seconds of thrust.
Does this look right to you?
What mass are you assuming for the spacecraft that is going to use this propulsion method?
(th)
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LH2 Propellant Tank info for Solar Thermal Reactor
STRUCTURES AND DESIGN PHASE I SUMMARY FOR THE NASA COMPOSITE CRYOTANK TECHNOLOGY DEMONSTRATION PROJECT
Lockheed-Martin Externally Box-Stiffened LH2 CFRP Tank
* based upon their fully leak and burst tested demonstrator tank model from 2013
* tested with real LH2
* I'm specifying more recent and stronger fiber, TorayCA T1100g fiber, which is stronger than IM7 Carbon Fiber by quite a bit
Tank External Diameter: 394 inches / 10m
Tank Height: 413 inches / 10.49m
Box Stiffener Height: 176 inches / 4.47m
Internal Tank Volume: 22,400ft^3 / 634.29736m^3
LH2 Bulk Density = 70.8kg/m^3
LH2 Propellant Tank Weight: 6,572lbs / 2,981kg
11X Tanks Total Internal Volume: 6,977.27096m^3
11X Tanks Total Mass: 72,292lbs / 32,791kg
90% total fill volume for 11X tanks = 6,279.543864m^3
90% LH2 total fill mass for 11X tanks: 444,592kg
If we had 1X 10m diameter LH2 propellant tank vs 11X smaller tanks, then it would be about 95m in length and weigh less because it only has 2 domes / tank ends. Total tank mass would be at or under about 30,000kg, thus on-par with the Super Light Weight External Tank Aluminum-Lithium alloy design used by the Space Shuttle.
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tahanson43206,
No.
LH2 heated to 2,750 Kelvin will yield an Isp / specific impulse of about 900 seconds. To get 1,000s, you have to go hotter by enough that the materials I specified probably will start to have issues.
900s is was what my JavaScript program computed acceleration and elapsed time off-of, not 1,000s.
Total elapsed acceleration time to achieve escape velocity with 7,800m/s of initial velocity, meaning 10,900m/s final velocity, was 740,456 seconds. You can only thrust when you're in sunlight, because you're operating a STR vs NTR. That's why total elapsed time is precisely double that of total thrusting time, as shown in my JavaScript program code in the "Orbital Mechanics" thread. You're orbiting the Earth, so 1/2 of your time spent spiraling out is on the dark side of the Earth, where you stop thrusting because you cease to have input thermal power from the Sun.
Make sense?
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For kbd512 re 161
Thanks for trying to clear up confusion.
Unfortunately, your post #161 just adds to the confusion.
Let's let the matter sit for a day or two, to see if anyone else has any interest whatsoever in your idea.
At first glance, your idea seems to represent a major step forward, and it might be worth a patent.
I think you have dramatically undervalued the performance that can be achieved.
On the other hand, by using the figure you provided in Post #158, I may have been led astray.
That figure may be bogus. The calculations that flow from it are not.
(th)
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tahanson43206,
You will have to learn about how rocket engines work. I can point you in the right direction, but then there's a lot of reading to do.
According to NASA testing, a reactor core temperature of about 2,750 Kelvin will yield an Isp of about 900s. Heating Hydrogen to 2,750 Kelvin (temperature) relates to the average kinetic energy of the Hydrogen molecules, thus thrust or force produced, because Force (F, in Newtons) is equal to mass (m, in kilograms), multiplied by acceleration (a, in meters per second), and all together F = ma. I used 1kg of mass flow per second (mdot) in my example JavaScript program. I provided the link to an online compiler where you can copy-paste and then execute it with the "Run" button on the web page.
From Google AI, taken directly from NASA (Google AI actually got this right for a change):
A nuclear thermal rocket engine aiming for a specific impulse (Isp) of 900 seconds would typically operate with a core temperature around 2700 Kelvin (K), which is the temperature at which the propellant (usually hydrogen) exits the reactor core after being heated by the nuclear fission process.
Low-Enriched Uranium Nuclear Thermal Propulsion Systems
From the above NTRS document:
The operation of a first generation NTP engine is conceptually simple. Hydrogen from a propellant tank is pumped through a solid core reactor where it is heated to high temperature (~2700 K) and exhausted through a converging / diverging nozzle to obtain a specific impulse on the order of 900 s. However, as with all rockets, the actual NTP engine will be a complex system.
Taken from "Atomic Rockets":
Solid core nuclear thermal rockets have a nominal core temperature of 2,750 K (4,490° F). Thrust is directly related to the thermal power of the reactor. Thermal power of 450 MWth with a specific impulse of 900 seconds will produce approximately 100,000 Newtons of thrust.
Let's do a quick sanity check on my F (force) / thrust math using 39,039,000Wth of input thermal power:
450,000,000Wth / 39,039,000Wth = 11.526934603857681
100,000N / 11.526934603857681 = 8,675.333333333333168 Newtons
Does that look close enough to the idealized thrust figure I provided?
I think it does. In point of fact, the thrust value from that NTR engine referenced by the "Atomic Rockets" website was also a tad higher than 100kN on the nose, IIRC, because it was deliberately designed to replicate a RL-10's thrust output. Beyond that, I used 2,750K vs 2,700K to determine exhaust velocity, so that may explain why my thrust value is slightly higher than 8,675N. The actual mass flow rate for that real NTR was, IIRC, about 12.6kg/s.
It's almost as if I know which formulas to use, and that's why my results are so highly congruent with those recorded for real historical nuclear thermal rocket engines that were actually tested at Jackass Flats. I forget things here and there, but I don't think so in this particular case.
Force or thrust affects vehicle acceleration rate, so a (in m/s^2) = F (in Newtons) / m (mass, in kilograms).
At "ignition", my total vehicle weight was listed as 1,250,000kg in my JavaScript (750,000kg payload, 500,000kg of LH2 propellant).
Force or thrust, from heating 1kg/s of LH2 to 2,750 Kelvin, is 8,825.985 Newtons.
a (initial acceleration rate at "ignition"), is 8,825.985 Newtons / 1,250,000 kilograms = 0.007060788 meters per second, per second
After thrusting for 1 second, my vehicle is now 1 kilogram lighter (the Hydrogen mass flow through the solar reactor core, also known as "mdot"), so it will acceleration at a slightly faster rate.
Velocity (after thrusting for 1 second) is now 7,800m/s^2 + 0.007060788m/s^2, or 7,800.007060788m/s^2.
That means for the next second of engine thrusting:
a = 8,825.985 Newtons / 1,249,999 kilograms = 0.007060793648635 meters per second, per second
Velocity (after 2 seconds of thrusting) = 7,800.007060788m/s^2 + 0.007060793648635m/s^2, or 7,800.014121581648635m/s^2
Can you see how my vehicle is slowly accelerating at a faster rate as it's slowly getting lighter by throwing reaction mass out the engine nozzle?
The JavaScript program I wrote is performing that same calculation, over and over again, very rapidly, so you don't have to waste your time performing that same calculation by hand hundreds of thousands of times.
I think you have dramatically undervalued the performance that can be achieved.
Actually, I think I'm pretty close to the idealized computed performance values, because in this case I'm reasonably sure I know my masses, units, and formulas. If I'm wrong, then I'll apologize, fix my mistake(s), and endeavor to do better next time.
BTW, Newtons are not the same units as kilograms-force, nor metric tons-force. I think that's where your program or analysis went wrong. The SI units for thrust / force are Newtons, and this engine will produce about 8.8kiloNewtons of thrust from 39MWth and a mass flow rate (mdot) of 1kg/s.
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For kbd512 re #163
Thanks for another interesting post... you closing line contains the key information that is causing confusion.
You provided the figure of 8.8kilonewtons in #158, but I misread the decimal as a comma, and gave the resulting figure to Gemini, who correctly computed the tons of thrust figure for 8 million Newtons. If you were to have read my post, which you clearly did not do, you would have seen the error, and simply pointed it out. Instead, we've been off the the races with long posts.
I am disappointed that the result is so small and agree that it is not useful for much of anything.
The much larger figure would have been worth celebrating.
(th)
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tahanson43206,
Apologies, I totally missed that post while scrolling past. I should've scrolled slower and then I would've seen it.
The one utility the STR design has is replicating NTR Isp at a very similar total mass, so there's no issue with testing it in any old desert, and it doesn't produce any radiation or require any nuclear materials, thus removing another "snow day" from NASA's reasoning as to why we cannot go to Mars.
Edit:
Did you edit that post at a later date?
Last edited by kbd512 (2024-12-12 12:58:01)
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tahanson43206,
I went to the trouble to write a JavaScript program which you could easily execute online using a free tool. Instead of using the program I wrote and tweaking the numbers to show results for whatever you were interested in using as input values, you immediately went to an AI program.
What was the reason behind doing that?
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For kbd512 re #165 and 166
And ** my ** apologies (again) for taking 8000 and misreading is as 8000000.
At this point in my life, I am vulnerable to lottery fever, and my misreading caused me to think you had hit the jackpot. I went to Gemini (in this case) to convert that meaningless Newton number into tons, which it was happy to do, without asking me where I got the number.
The actual thrust you'd be able to develop appears to be something useful for the deep space vehicle to leave the Moon on a long, very slow trip to Mars.
GW told me he thought the normal configuration of the mirror for this system would be a plastic balloon, with mirrored interior in half the volume, and clear on the other to admit sunlight. That would keep the cost down, because the balloon would be inflated with a gas of some kind, and no stiffeners would be needed, or very few.
And thanks for the Javascript program. I should be able to run that, as should most NewMars members. As i recall, I have an entire topic devoted to Javascript somewhere in the NewMars archive.
(th)
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tahanson43206,
When something looks too good to be true, then it probably is.
The STR engine is sufficiently powerful to break orbit and leave Earth in only 8.57 days. That's not long at all.
VASIMR required multiple weeks to spiral out with 39MWe worth of input power, over 2 months, IIRC.
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Crapping on VASIMR seems to be a popular past time on this forum and elsewhere, following Dr Zubrin's commentary on the engine, but from a pure force generation efficiency standpoint, it looks like other-worldly technology to me. On top of that, it's been extensively tested and proven to work. All propellantless propulsion technologies have not delivered usable results thus far. I'm crossing my fingers that the Mach Effect thruster is different, and cautiously optimistic that it might actually work.
Ve = Velocity, exhaust
g0 = Earth gravitational constant, or 9.80665m/s^2
Isp = Impulse, specific, in seconds
Ve = Isp * g0
Ve = 32,000s * 9.80665
Ve = 313,812.8m/s^2
It = Impulse, total, in Newton-seconds
Mp = Mass, propellant
It = Ve * Mp
VASIMR Total Impulse for 1kg H2, at Isp = 32,000s (32,000s * 9.80665):
313,812.8 Newton-seconds
Aerojet Rocketdyne RS-25 Total Impulse from 1kg of LOX/LH2, at Isp = 452.3s (452.3 * 9.80665):
4,435.547795 Newton-seconds
VASIMR Total Impulse Multiplier over the RS-25 (313,812.8 / 4,435.547795):
70.7495X
Each kilogram of H2 propellant shot through a VASIMR at 313.8km/s, generates 70.7495 TIMES more force than 1kg of LOX/LH2 shot through the RS-25 at 4.435.5km/s. That's a spectacular improvement in total force production, and therefore launch costs, since each kilo of propellant must be hauled into orbit from the bottom of a deep gravity well. Your in-space vehicle is mostly useful payload as a result, rather than a giant gas tank required to go anywhere else.
Speaking of force generation efficiency, that thorny SSTO problem has a practical solution after we accept that flight physics overrules beliefs about how something should work, but doesn't. Total impulse versus total propellant volume, thus vehicle volume, thus vehicle structural mass, thus thermal protection structural mass for any reusable vehicle, thus total vehicle cost, combined with engine thrust per unit mass, is the entire reason LOX/LH2 SSTO was and is such an unworkable design concept. The primary load on those feathery-light composite propellant tanks is internal pressurization. That load dominates the list of stresses that the primary structure is subjected to. To make matters worse, internal pressurization has to be substantially higher for Hydrogen than RP1. On top of that, both Hydrogen and Methane are very small molecules that readily diffuse through and embrittle solid materials.
RP1 obviously weighs far more than LH2 per unit volume, which is actually a very good thing given all its other characteristics, but Total Impulse is STILL what determines how much mass you can drive to a given final velocity. The Specific Impulse delta between RP1 and LH2 is small enough that it only matters for in-space propulsion or TSTO upper stages trying to remain lightweight to maximize achievable orbital altitude / velocity. If the vehicle gets large enough, as it very rapidly does for a SSTO's giant propellant tanks, you run into structural problems associated with internal pressurization long before you run into aero load problems or issues with the greater propellant mass putting greater stress on the vehicle.
Despite the lower Isp of RP1, you still get more Total Impulse (total force delivered) per unit volume, using RP1 vs LH2.
Look at the Density Impulse of LOX/RP1 vs LOX/LH2. RP1 is a factor of 2.37X greater than LH2 without propellant densification. LH2 is a factor of about 1.32X better than RP1 in terms of thrust efficiency per unit mass. That is never going to make up for taking a 2.37X hit on Density Impulse. Look at RP1 vs LH2 engine thrust-to-weight ratio, a factor of 2.67X in favor of RP1. You need thrust to escape a deep gravity well. The more thrust you have to do that, the more better it is. There is no possible way for LH2's 32% greater thrust-generating efficiency to make up for that combination of problems when the use case is a single stage orbital launch vehicle. Basic orbital launch vehicle physics is stacked heavily against LH2.
How or why it is that so few people are able to put this seemingly simple set of engineering requirements together is beyond my understanding. Propellant mass and thrust efficiency are not the driving engineering problems for a practical SSTO, it's inert vehicle mass vs propellant total impulse per unit volume, because that is the driver of inert vehicle mass, which detracts from useful payload tonnage delivered per total vehicle volume. Until you start moving downrange at substantial speed, LH2's 32% propulsive efficiency increase will never make up for a 2.37X vehicle volume increase, nor 2.67X thrust advantage in favor of the denser propellant.
Incidentally, this seems to be the same issue that people entirely "miss" when it comes to electrification and batteries. Batteries are 50% to 75% more efficient than modern combustion engines. Pound-for-pound, a 250Wh/kg Lithium-ion battery requires precisely 79.5 TIMES as much mass to deliver the same payload to the same distance as a 50% efficient diesel engine. How do you "overcome" a 79.5X multiplier on vehicle mass for a major subsystem that powers a vehicle? The short answer is that you don't, because you can't. It's not a matter of belief or throwing more money and brain power at the problem. Absent a revolutionary improvement in gravimetric energy density for the batteries, it's an unsolvable problem. In a very real sense, electric vehicles are the LOX/LH2 rockets of land-bound motorized transport. They rapidly become almost entirely "fuel" (batteries) by weight and volume if the vehicle has to match the payload-to-distance of a long haul diesel truck.
Also, if LOX/LH2 is still considered "better" or "best" for in-space propulsion, then using pure H2 at 900s to 32,000s is dramatically to wildly better, despite the inconvenience posed by the "low thrust", which is entirely due to our limited ability to impart the colossal amount of heat energy into a greater mass flow rate of propellant, in order to greatly increase thrust while still achieving those stunning exhaust velocity values. It takes a lot more energy to throw the reaction mass a lot faster than a combustion process is capable of.
VASIMR has low thrust output because a comparatively low amount of input energy is being consumed from a small-scale power source, relative to a combustion process, and we need a lot more energy than a combustion process actually delivers, so we either use an incredibly dense energy store (fissile or fusible materials) or an incredibly large surface area to collect power from the Sun. Those are the only realistic options if you want greatly increased thrust levels.
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One thing that should stand-out to people who look at the thrust of LOX/LH2 vs VASIMR is that Newton-seconds are equal to Joules of energy.
The test variant of VASIMR is using 200,000We of input power.
1 Joule per second equals 1 Watt.
4,900s * 9.80665 = 48,052.585 Newton-seconds or 48,052.585 Joules per second
The engine consumes about 150 milligrams per second of Argon while running 200,000We of input power. It produces 5N of thrust.
If we consumed 1,000 milligrams or 1g/s of Argon, then power must increase by 6.67X, so 1,333,333We (1.3MWe). Instead of 5N of thrust, we get 33N.
Now let's scale-up to where we're consuming 1,000,000 milligrams or 1kg/s, so input power must increase by 6,667X, to 1,333,333,333We (1.3GWe). Instead of 5N of thrust, we get 33,333N.
Do we have any kind of "fuel" with the energy density to generate that kind of power with a reasonable mass?
Sure we do:
80,000kg of Hydrogen combusted with 480,002kg of Oxygen, at 50% conversion efficiency
2,177,048m^2 (1,475.48m by 1,475.48m) of solar collector surface area at 1AU from the Sun, at 50% conversion efficiency
1kg of Uranium generates 24GWh of energy
1kg Tritum and Deuterium generates 93.7GWh of energy
0.5kg of Anti-matter generates 12,500GWh of energy when mutually annihilated with 0.5kg of normal matter
That's it. That's all we have, because that's what the universe has provided to us to work with.
If we want Star Trek starships, then we must have Star Trek energy density power sources to work with.
Even if we're not going "Back to the Future", high efficiency and high thrust is still all about having those 1.21 JigaWatts.
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For kbd512 re VASIMR ...
VASIMR has low thrust output because a comparatively low amount of input energy is being consumed from a small-scale power source, relative to a combustion process, and we need a lot more energy than a combustion process actually delivers, so we either use an incredibly dense energy store (fissile or fusible materials) or an incredibly large surface area to collect power from the Sun. Those are the only realistic options if you want greatly increased thrust levels.
There ** is ** another option for supply of power to VASIMR.
However, your statement in the quote above appears to preclude any other energy source or delivery method.
Too bad!
(th)
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tahanson43206,
What is the "other option" that I haven't considered?
If we're still talking about things that we actually know about and have used, that includes mechanical, electrical, chemical (combustion / chemical reaction), light (includes thermal), and nuclear (fission, fusion, mutual annihilation) energy sources, plus whatever combination of those we wish to use. The energy source can be on the craft being propelled or somewhat nearby. We have a truly massive power plant (the Sun) capable of supplying an incredible amount of power when the collector surface area becomes large enough, even though that becomes problematic at extreme sizes, because the mass of whatever provides the power has to be placed in space, and likely constructed in space.
Isp is a derivation of mass multiplied by velocity squared. The only way to radically increase mass efficiency of a "mass ejector" system is to radically increase the velocity of the ejected reaction mass, which implies extreme acceleration. For practical interplanetary colonization efforts, extreme mass efficiency really is required. Extreme heating produces extreme acceleration, thus average kinetic energy of whatever is being accelerated (Hydrogen if mass efficiency is what counts most, or heavier atoms if more thrust is more desirable), so that's where we derive our extreme in-space propulsive efficiency from.
If someone concocts a working propellantless drive, then we start worrying about other things. I'm hoping that the Mach Effect thruster is very real. If it is, then we won't use too many mass ejector systems to move people between planets.
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For kbd512 ...
Re solar power to heat hydrogen for propulsion....
Let's include this in Sunday's Google Meeting... sometimes text gets in the way of communication....
***
Other topics we might consider include:
1) The new directory for files to serve
2) Possible php files to allow moderators and admins to set up service
3) Restore database backup to Azure (I still haven't done it)
4) How to recruit new members to forum... SpaceNut and RobertDyck are committed to other activities
5) GW's Space Tug Book idea
6) How to resume work on the "like" feature (we scoped it out and then I moved and lost all records)
PS ... I discovered that Blender has a free mouse drawing feature I did not realize was there. It's been there forever, but I was unaware of it.
I can draw sketches now and share them on a Google Meeting Share page.
GW can use Paint (far more effectively, of course), so we have two members who can use drawings when images aren't quite sufficient.
I would like to discuss your solar power hydrogen rocket idea but words aren't sufficient.
The Sun provides 3.8 x 10^26 joules of energy per second according to Google
Your proposal (as I understand it) was to collect a small part of that to heat hydrogen to create thrust.
I think your posts cover the math pretty well. What text cannot do is to show how the engineers would implement the components to use that energy.
I mis-read a number in one of yoru posts and thought the actual output was a thousand times greater than you had actually predicted.
I'm interested in looking at how you would design a system to produce the thrust that ** is ** possible with existing materials.
(th)
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For kbd512 re new Project for Solar Thermal tug...
http://newmars.com/forums/viewtopic.php?id=10962
Please revise the design you did in your opening series of posts about the Solar Thermal Rocket, and post the updated math in the new Project I've created for the concept. We may be able to enlist other members to help with drawings.
With any luck, we may be able to enlist Calliban to help with design of Carbon girders that are going to be needed to provide a framework for all those parabolic reflectors you are going to deploy.
I am ready and willing to help you to set up Blender on the new, powerful computer that I hope Santa Claus brings you for Christmas this year.
I am learning a 2D drawing program called QCAD, and hope to be able to provide assistance to any NewMars member who wants to create drawing files to make components of your design.
Elements such as girders are going to be needed for Calliban's asteroid encircling mining rig, and I'm hoping they will work for your tug as well.
(th)
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For kbd512 re discovery of light collectors...
Congratulations on making this discovery, thanks for reporting it, and best wishes for designing a space capable system to use it.
I've opened a topic for space qualified girders, and hope that Calliban might be interested in developing it. He's going to need such girders to build the ring he's described to mine asteroids.
It should be possible to 3D print scale models of such girders.
That could lead to the ability to create a scale model of your entire space craft on the ground.
My working assumption is that the light gathering surface is a plane, so a framework of ultra-strong, ultra-lightweight girders should be able to hold the light gathering elements in place.
(th)
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