Orbital Mechanics

tahanson43206 · 2024-11-08 11:49:23

For GW Johnson re #200

Thanks for taking a look at the site. Someone(s) put a lot of effort into that. As near as I can tell from a preliminary look, it uses JavaScript for the on-browser performance, which explains how the system could perform so well.

I found a page of documentation, and will attempt to make it available here In a form that our members might be able to use. I am a very long way from being able to deliver anything like that, but it is good to have an example to study.

***
Regarding the Starship landing. Thanks for the detailed analysis and prediction that the Starship ** has ** to lay itself on it's side, to survive the landing. The scenarios you've painted look reasonable to my eye, and at first reading. Hopefully SpaceX will have cameras at the landing site as they did last time, so the behavior of a stainless steel cylinder rotating to horizontal while embedded a short distance in water will be interesting to see.

I made a screenshot and saved it on Dropbox.
https://www.dropbox.com/scl/fi/8lzg8dk5 … 2mfvd&dl=0

You can download the image and (hopefully) print it.

I seem to have the program running, with these setups:

1) I put the letter "a" in the box for "enter orbit name" ( I think this can be anything but that's just a guess.
2) I clicked on the add launch button and chose Kennedy Space Center from the drop down list
3) There is a red "fire" button at the bottom of the page. I clicked that and it turned to "firing"
4) I changed the time value at the top of the page to x100, so now the simulation is running 100 times faster than real time (or at least I ** think ** so)

In any case, the default orbit is heading west, and right now the space craft is over the Suez Canal.

I sure hope there is someone else in the NewMars community who is interested in trying this and can make better informed choices.

The copyright notice says: GrayBits LLC

https://graybits.biz/about

I am guessing this Orbital Mechanics Simulation web site may be a subtle advertisement for the capabilities of the team.

The multiple listings via Google seem to indicate this is a relatively small company, with 12 employees.

(th)

tahanson43206 · 2024-12-10 07:21:38

For all ....

In a topic about design of components for a large deep space vessel, kbd512 has introduced early consideration of flight planning, based upon work done by GW Johnson.

https://newmars.com/forums/viewtopic.ph … 29#p228429

This post is intended to provide a focus for the hybrid flight plan suggested by kbd512

The concept to which I would like to draw attention is the fact that the ion drive proposed by kbd512 will provide a steady thrust over an extended period.

The chemical Space Tug flight plan is designed to put the customer vessel at near escape velocity, so that the Tug itself will return to LEO, and thus be able to achieve reusability.

The customer vessel must make up the small dV needed to achieve escape from Earth's gravity.

The strength of the thrust provided by the ion drive may require multiple circuits of the ellipse it is on, but eventually the accumulation of dV will be sufficient, and the vessel will enter into the larger ellipse that takes it to Mars.

This topic is available if we have a member who can put numbers to this concept.

The mass of the customer vessel, and the strength of the ion drive thrust, will determine the time and distance required.

It seems to me likely that the amount of thrust applied at various points in the ellipse must be adjusted to insure the customer vessel ends up pointed in the right direction as the moment of escape occurs.

While the customer vessel is accumulating velocity, the Moon, Mars, the Earth and the Sun will all be pulling on the mass of the vessel.

Software able to perform the calculations needed is likely to exist today, because several deep space probes have employed ion drives to achieve remarkable performance. Whatever that software is needs to be identified and permission secured to use if for this application.

(th)

kbd512 · 2024-12-10 11:16:47

tahanson43206,

My flight plan for the ITV starts in LEO. The space tug first injects the ITV into a halo orbit, the VASIMR main engine completes the push to escape velocity, whereupon the ITV is coasting towards Mars. VASIMR performs any mid-course correction burns required to intercept Mars with great precision. Near the vicinity of Mars, the crew boards the LLM and separates from the ITV. The crew strips the ITV of any consumables they can take with them aboard the LLM.

The LLM continues towards Mars and performs a reentry, descent, and landing on Mars. The ITV continues to follow its nominal two year free return trajectory back to Earth. It never stops in orbit, because then we'd have to wait even longer to reuse the ITV, and time is money. As the much lighter and empty ITV approaches Earth, it uses its VASIMR main engine to slowly spiral back into LEO, where it will be refurbished or repaired, as well as readied for its next outbound flight to Mars.

If aerocapture back at Earth ever proves practical using a ceramic composite hull design and careful flight trajectory optimization, then the ITV is not going to spiral back into Earth, either, since that requires a significant tonnage of additional propellant to accomplish. Nobody will be aboard the ITV when it returns, so risking the vehicle may be acceptable. However, risking people on the ground is not, so the trajectory needs to ensure that the ITV will reenter over an ocean if the aerodynamics involved in the atmospheric braking and capture maneuver is not completely dependable.

If the habitation rings and the rest of the ITV hull was made from RCC or a Sylramic fiber toughened ceramic vs Titanium, then it might be possible to further reduce the HRM and CNM structural mass over Titanium alloys using a material that's also capable of surviving reentry heating. This would entail fabrication of a more sophisticated and costly ITV, even though it would substantially reduce operating cost. Some kind of trade study on aerobraking and aerocapture back at Earth needs to be made. I have a rough general idea of the tonnage of RCC required, and what fabrication would cost, but no final numbers on how much a Starship flight would cost to deliver more Argon or Hydrogen propellant.

tahanson43206 · 2024-12-10 12:10:45

For kbd512 re #203

Thanks for this helpful additional information.

There is more than one halo orbit. There is the one chosen by NASA because there SLS rocket is insufficient to do better.

Then there is the halo orbit that GW Johnson recommends.

I don't know the answer to this question, but am looking forward to seeing answers as they become available:

Can the Space Tug we are talking about deliver the ship you are talking about to an orbit around the Moon?

It's certainly an interesting idea. By using the Moon as a staging area for the ion engine phase of the flight, you avoid the risks associated with falling back toward Earth with results that would be unpredictable.

If it turned out to be necessary, would you be willing to carry a chemical rocket to insure a successful transition into the lunar orbit you have in mind?

(th)

kbd512 · 2024-12-10 12:32:25

tahanson43206,

The issue with starting the flight in orbit around the moon is that you have to push all the mass of the consumables plus passengers plus propellants to the moon, in addition to the mass of the in-space stages required to do that. That only serves to increase cost and complicate the resupply logistics, unless the moon is producing the logistical materials, and then it might be a more practical proposition. Shielding the passengers from the radiation of the Van Allen Belts is a lot easier to do, it requires less mass in orbit, and is therefore cheaper. After someone builds a lunar base producing most of the propellants, food, water, and other consumables, then there's a much better argument to be made for using the moon. That said, I'm designing for the goal of sending lots of people to Mars, because Mars has an atmosphere, however thin, it has more gravity, and more materials to work with than the moon provides.

tahanson43206 · 2024-12-10 13:03:53

For kbd512...

I think everyone understands you need the services of the space tug to put our ship at or near escape. The question I am hoping Gw can answer is if the Space Tug can put your vessel into one of the lunar halo orbits, so you can start the ion engine and begin the slow gradual acceleration it provides.

Your hint of a suggestion of an idea of using the Moon as a safety staging area is interesting. This topic (Orbital Mechanics) needs a NewMars member who is up to speed on modern computer based flight planning software.

(th)

tahanson43206 · 2024-12-10 13:51:28

For all...

GW Johnson provided an image to show the difference between spiral out from Earth with ion vs bust out from Earth which chemicals.

This image does NOT consider kbd512's interesting idea of using the Moon as an anchor.

(th)

tahanson43206 · 2024-12-10 13:56:20

For kbd512 ...

Via email, GW Johnson has endorsed your idea of using the Moon as a staging platform for an ion drive vessel.

He also added that while the Space Tug is ideally used in the toss-and-return mode, the Tug could expend extra energy to put the customer in any of the lunar orbits, and then spend more extra energy to leave the Moon and return to Earth.

The customer will pay the bills, so it appears your idea is practical.

(th)

Oldfart1939 · 2024-12-10 17:28:50

One way of managing this problem of using ion propulsion, would be as a "4th stage," after a chemically powered orbital meet-up with a nuclear thermal 3rd stage, and the ion propulsion would become the deep space only propulsion, by using the nuclear thermal propulsion to get through the Van Allen Belts quickly. I see chemical propulsion as the only effective way to get out of the atmosphere, and where a nuclear thermal powered stage could provide escape velocity, after which ion drive could be utilized to reduce interplanetary transit times.

This is overly complex, but possible in the future. Maybe there will be advances in ion propulsion which increase thrust from mere pounds to tons?

kbd512 · 2024-12-10 19:13:19

Oldfart1939,

Thrust per unit mass ejected is dependent upon the mass of the exhaust species and exhaust velocity. Heat (increased average kinetic energy) adds velocity to ejected particles. In order for an ion engine to generate more force / thrust, you must either ionize / convert into a plasma a greater total mass, or add a lot more heat. VASIMR already achieves temperatures above those predicted to exist in the core of the Sun, quite similar to a fusion reactor. How much hotter do you think we can make the exhaust before we start melting things?

The solution is to eject a lot more hot plasma mass per unit time with the same average kinetic energy level / exhaust velocity. That implies adding a lot more input electrical energy to ionize a lot more mass. At 200kW output level, producing 5N of thrust, mass flow is measured in milligrams per second. To produce 1kN of thrust using VASIMR, your input power requirement starts a hair above 40MWe.

VASIMR is a "little stroke of genius" because it uses microwaves to very rapidly superheat the plasma and electromagnets to confine the plasma to enhance thrust density. It does not need an electromagnet to prevent the superheated plasma from accelerating towards the front of the thrust chamber, another good feature, and unlike most ion engines. Beyond that, it's more significantly more compact and slightly more efficient than a more conventional ion engine.

The Aerojet Rocketdyne RL-10A1 LOX/LH2 expander cycle engine produces 110,100N of thrust, with an Isp of 465.5s.

VASIMR requires 40,000W per 1N of thrust generated, using Argon propellant, with an Isp of 4,900s at that output level.

Therefore, VASIMR would have to receive an input power of 4,404,000,000W (4.404GiaWatts) to generate the same 110.1kN of thrust that the RL-10A1 engine produces. The more practical alternative to generating 1.21 Jigawatts of power (unless Plutonium becomes available at every corner store in the near future), is to slam a mucho faster moving (very high-Isp) exhaust gas product, such as Hydrogen, into Oxygen in the electromagnetic exhaust nozzle, thereby producing the "afterburner effect" of a "plasma jet" engine. Your Isp takes a big hit, but thrust also goes up by an equivalent amount. If your Isp started at 32,000s, as it would for a VASIMR consuming Hydrogen, then you can afford to knock that back down to 5,000s using an O2 afterburner, thereby gaining 6.4X greater thrust (by increasing the exhaust mass flow).

At 6,250W/1N of thrust generated by the combination of the H2 slurping VASIMR equipped with an O2 afterburner, you only need to deliver 688,125,000We. That is still a stupid amount of electrical power, but at 1AU and 612.45We/m^2, you only need 1,123,561m^2 of solar reflector area to generate that much power, so 1,060m by 1,060m solar collector. At 20.4g/m^2 for Aluminized mylar, a 1,123,600m^2 collector only weighs in at 22,921.44kg. At 1MW/kg, the sCO2 turbine rotor weighs in at 688.125kg. The turbine housing is about 4X the turbine weight. Inconel turbine rotors generate 200kW/kg (at a 50MWe to 300MWe output level), so a fiber reinforced ceramic that's 1/4 the weight of Inconel bumps that up to 800kW/kg all by itself, and then the additional 200kW/kg is provided by operating at 1,000C vs 715C, which will also decrease the total reflector surface area. Using supercritical Xenon vs Argon or CO2 would also do the trick, but if we're generating close to 1GWe, then we may as well operate at a higher temperature, and Sylramic fiber was designed to operate at 1,200C, and is presently used in jet engine exhaust con-di nozzles.

23,362kg for the solar thermal collector panels and power turbine, figure on doubling that for the heat exchanger, recompression turbine, and plumbing. NASA's high temp Carbon Fiber (not CFRP) and Inconel radiator is 1kg/m^2. The radiator to perform heat exchange needs to be 2,510m^2 at 1,000C inlet temp, so that adds 2,510kg for the radiator. I've no idea how much supercritical CO2 or Xenon is required, but there's additional mass to add for that and storing it as a highly compressed gas. That said, everything but the working fluid itself and gas storage tanks weighs in at 49,234kg, so that's ~13,977W/kg of power mass for a solar thermal power system operating at 1AU. That does not include the mass of the electric generator or reduction gearing, though (reduction gearing may not be required at this power level, because the turbine will be spinning a lot slower than much smaller models). At 15kW/kg, the electric generator (Aluminum, Copper, Iron, permanent magnets, liquid-cooled, and based upon Wright Electric's 4MW electric aircraft motor) weighs in at 45,875kg. From there, you have to add power cabling and power conditioning mass. Even at 100,000kg for your 688.125MWe solution, ~6,881W/kg of mass is not too shabby for an incredibly powerful non-nuclear / solar thermal electric power plant. I'm sure additional mass for radiators operating at lower temperatures and power conditioning will be required, though.

Frankly, I'd rather reduce the mass of the payload so we can still economically use standard chemical propulsion and singular space tug launches. We employ ion engines only when the mass of propellant makes the overall solution unaffordable. I think that as payload mass approaches or exceeds 1,000t, you have to commit to ion engines, because the total mass of a solar thermal electric power and propulsion solution rapidly becomes a minor fraction of the mass of chemical propellants.

tahanson43206 · 2024-12-10 19:33:27

For kbd512 re commitment to ion propulsion....

Would you accept a push all the way to escape, if GW makes the necessary adjustments?

Oldfart1939 shows us the way forward, and it provides a solution for ** all ** future deep space vessels that are limited to ion propulsion for whatever reason.

It seems to me that there is likely to be a solution that meets all needs.

For Oldfart1939: Thanks again for your helpful input.

(th)

kbd512 · 2024-12-10 19:53:43

tahanson43206,

If we really do commit to producing tens of megawatts of power, then we can also actively shield a spacecraft from the Van Allen Belts, so the entire impediment to using ion engines (the radiation dose associated with extended duration transits through the Van Allen Belts) is summarily removed. At that point, we no longer have much use for a chemical engine space tug, because it won't be a mass-effective propulsion solution.

Deployed Electromagnetic Radiation Deflector Shield (KSC-TOPS-97)

Creating a Zone of Minimum Radiation and Magnetic/Plasma Effects for Spacecraft
Overview
Spaceflight outside of the Earth's protective magnetic field is dangerous from a cosmic radiation perspective. Inside Earth's magnetic field, where the manned International Space Station (ISS) orbits, the radiation encountered is minimal and almost all is deflected by our planet's magnetic fields. However, outside that protective shield, the Sun's solar wind (high energy radiation, solar energetic particles or SEPs) consisting of protons, electrons, alpha particles and plasmas continuously bombards the spacecraft for the months or years of spaceflight. On occasion the Sun produces a CME (Coronal Mass Ejection) that vastly increases the energy and volume of this radiation. These particles damage human DNA as well as living tissue and can destroy sensitive electronics. The Deployed Electromagnetic Radiation Deflector Shield (DERDS) provides a magnetic field that will deflect SEPs and CMEs and other harmful solar and cosmic rays away from a manned spacecraft, robotic spacecraft, or manned extra-planetary base stations using an electromagnet that is deployed between the spacecraft/station and the source of radiation and creates a magnetosphere or zone of minimal radiation in which the spacecraft or base station would reside.
The Technology
The typical remedy for space radiation hazards has been to harden the space vehicle's electronics and software from these high-speed particles and placing heavy shielding in these manned or sensitive areas. This adds considerable weight and cost to the launch vehicle, reduces payload capacity, and is passive in nature. SEPs and CMEs can be deflected by a magnetic field to pass around the spacecraft and not be absorbed by it. This deflection of solar wind and radiation is due to the Lorentz force. However, a magnetic field that is attached to the spacecraft and enclosing it would cause other shielding issues with equipment and would require much more electrical power to operate due to the need to enclose the entire spacecraft within that attached magnetic field. This would also perturb the data collection and transmissions of the spacecraft. Additionally, the generated magnetic field can capture some of the solar radiation as a plasma within the magnetic torus further impeding scientific data collection due to its close position to the spacecraft. DERDS is deployed in space away from the protected spacecraft to remain between the radiation source and the spacecraft. It utilizes on board cosmic ray sensors to note the need for the strength of the magnetic field, and on-board sensors to position itself directly in line between the radiation source and the protected spacecraft or station. Onboard computers and thrusters maintain the required position, so that the magnetic field is positioned for the best deflection angle based on incoming radiation. The DERDS magnetic field can be varied in both direction, intensity, and time by use of both AC and DC electrical power inputs and varied as needed to optimize the deflection angle and power requirements. The magnetic field can also be perturbed in irregular or set patterns by onboard computers and sensors as needed to maintain the proper deflection of SEPs, CMEs, and other cosmic rays. DERDS creates a magnetic field that will not need to be so large (with a much larger power requirement) as to encompass the entire protected spacecraft, and the magnetic field will not interfere with the spacecraft.

tahanson43206 · 2024-12-11 10:16:37

This post is for kbd512, and for any of our readers who may be thinking about doing an ion drive voyage to Mars.

GW Johnson took a look at an idea/suggestion from kbd512, to enlist the Moon to assist with the Space Tug departure maneuver from LEO.

GW found a solution that allows the deep space transport to settle into a lunar orbit (not a halo orbit... a ** real ** orbit) after boost from LEO and before departure for Mars ... The quotation below is about how the deep space transport would proceed from there.

I would note that a stop at the Moon would be worth doing for several reasons:

1) Pick up supplies from the Moon, such as water, oxygen and other useful materials.
2) Pick up fresh produce and other food supplies grown and produced on the Moon
3) Drop off passengers who took advantage of the transport's trip from LEO to travel in comfort instead of a tin can
4) Pick up passengers who've been on the Moon and are prepared for the long trip to Mars
5) Drop off passengers who lost confidence and want to bail out of the Mars trip

In short, a stop at the Moon before continuing on to Mars would be advantageous above and beyond the celestial mechanics factors.

Here is GW's discussion of departure from the Moon:

If you slow down and go into lunar orbit, there is a fairly small dV requirement to slow down, into lunar orbit. It is on the order of lunar escape minus lunar circular, which is around 0.7-ish to 0.8-ish km/s. There might need to be some sort of rendezvous or orbit adjustment budget (perhaps as much as 0.2 km/s, but likely less), but there is no need for going to that ridiculous halo orbit! At low orbit altitudes, your speed around the moon is already higher than anything associated with that halo idiocy! (Neither NASA's nor my more-stable one.) You can just turn on your electric to thrust continuously, and spiral out past lunar escape into a super-extended Earth orbit, and from there, spiral out just a little further into Earth escape onto a sun-centered trajectory to Mars. It's not an ellipse, but an accelerating spiral about the sun, if you continue to thrust. Somewhere about halfway, you switch to deceleration thrust, and switch to a different decelerating spiral, in order to capture gravitationally into a spiral-in decelerating trajectory about Mars.

(th)

Oldfart1939 · 2024-12-11 12:40:05

As an aside here, Robert Zubrin has written of the VASIMIR system as being a "hoax." It seems promising ON THE FIRST GLANCE, but as kbd512 has illustrated by the numbers, it requires an incredible amount of energy to in reality, accomplish very little for a massive interplanetary vehicle.

So--kbd512 is correct to state that we still need to rely on chemical propulsion, and in a short time, move on to nuclear thermal for achieving shorter travel times for manned interplanetary travel to Mars, Ceres, Ganymede, and Callisto. The Isp from chemical propulsion seems maxed out at 450 sec for hydrolox; There's about a 100% Isp increase for a NERVA style nuclear thermal engine up to around 850-900 sec. , so that will reduce travel time enough to actually consider manned flight to Callisto and the Jovian system, but starting from Mars.

The rocket equation as tyrannical and the velocity of a vehicle is limited to ~2x exhaust velocity.

Last edited by Oldfart1939 (2024-12-11 12:42:03)

Oldfart1939 · 2024-12-11 12:58:34

tahanson43206-
Maybe if you wanna go back a number of years, you will find a mission architecture predating Elon's BFR concept, that I suggested for what is now being referred to as a "space tug." I was trying to expand on the Mars Direct concept at the time, by suggesting a "modular approach" to getting a larger than Dragon size spacecraft on it's way to Mars, by launching "modules" roughly the size of Falcon upper stages into orbit that would contain one of the Hydrazine type compounds (MMH, or UDMH), and NTO, both of which wouldn't have evaporation issues, and would require no igniter systems. Just have them parked there for "getting outta Dodge," and having one of them also launched to Mars orbit for a return flight. It might be helpful for you to locate that thread? it would be in roughly the 2017 +/- window.

Last edited by Oldfart1939 (2024-12-11 12:59:39)

kbd512 · 2024-12-11 14:42:12

Oldfart1939,

The only real accomplishment of VASIMR is producing a very large Delta-V with very little propellant mass expenditure. It's merely on-par with other better-established ion engine technologies, with some important key differences that make it more flexible. To my knowledge, only VASIMR and Aerojet Rocketdyne X3 ion engines have operated at or above 200kW for significant periods of time, using multiple propellant types, appear to have sustained very little damage in the process, and have reasonable masses for engines of their type. Whereas X3 requires some significant design tweaks to consume different kinds of propellants, VASIMR does not.

I'm not greatly enamored with any kind of ion engine, I just need a solution which delivers more dead-heat payload mass than propellant. If the electrostatic or electromagnetic deflector shielding proves practical, then taking longer to spiral out from Earth is not a major problem. As of right now, we have a tested electrostatic design that consumes 100We and deflects about 50% of the incoming charged particles for a very nominal mass and power penalty. Electromagnetics are still being iterated through various advanced designs, but are starting to bear fruit. An active radiation shielding system is a required tech item for greatly extended duration missions into deep space, so we need to pursue it regardless of how we protect the passengers aboard the ITV.

As far as NTR tech is concerned, a flight-tested prototype engine is at least 10 years away. I think we should pursue it because high-thrust moderately high-Isp is an enabler, but it won't be viable for civil ITVs if it uses HEU. Thus far, all such engines use HEU. Zero Boil-Off (ZBO) and Hydrogen-impermeable membrane tech is just starting to become a reality.

NTO / Hydrazine propellants are great for government missions where cost is no object, but ZBO and higher-Isp propellants are still going to provide greater performance. Incidentally, NASA did trade studies on various hydrocarbon propellants as an upgrade to the Space Shuttle Orbiter's OMS and RCS, but lacked the funding to pursue the project. Heck, they even spec'd out the use of Lithium-ion batteries and electric motors for feed pumps, which finally became practical 40 years later, as implemented by Rocket Labs Electron.

LOX/Hydrocarbon Auxiliary Propulsion System Study - Final Summary Report

Oldfart1939 · 2024-12-11 21:47:50

kbd512-
I have quite a bit of professional experience with hydrazine, although as hydrazine hydrate. Up until around 1985, it was reasonably cheap. The anhydrous form is more difficult to manufacture, but on the industrial scale--still a cheap chemical. The major expense in using this nasty stuff is the cost and regulations of transport. It was being trucked around the country in tanker truck quantities until the EPA became involved. Lots of other previously "cheap" chemicals became almost unavailable overnight. A case in point was carbon tetrachloride, an extremely useful solvent, which my company used in 55 gallon drum quantities and it was around $400 for a drum. Then the EPA said it could no longer be manufactured because it was almost impossible to dispose of. Not burnable for waste disposal. But it was the solvent of choice for photochemical bromination reactions. Now--it's only available in small laboratory quantities for around $500 a gallon.

kbd512 · 2024-12-12 01:58:35

Oldfart1939,

One of the major selling points about moving to Mars is that there's no EPA on Mars.

If we're talking about "best of breed" engines and propellants, then we have a very limited selection of specific impulse ranges and thrust levels to work with. Any fully electric engine has a TWR well below 1, so only suitable for in-space propulsion. However, Isp falls within the range of the most efficient turbofan and turboprop engines, which is very useful for cruising, but far less useful for impulsive maneuvering.

Any electric heating enhanced chemical engine has a TWR of 1 or less, effectively Isp and thrust intermediaries between chemical and full electric or NTR engines, thus also suitable only for in-space propulsion. Resistojet engines can deliver 450s of Isp using electrically heated Hydrogen. These have only ever been implemented at low thrust levels more suitable for satellites than manned spacecraft. Hydrazine monopropellants have lower Isp on-par with small thruster-type NTO/MMH engines, but at much higher thrust than pure Hydrogen. These engines are worthy alternatives to NTO/MMH and use the same Hydrazine fuel. I spent quite a bit of time looking at these electro-thermal engines.

Chemical engines can have TWRs as high as 200:1 for LOX/RP1, but Isp tops out at 475s for "vacuum nozzle" expander cycle LOX/LH2 engines. All the advanced combustion concept engines such as aerospike nozzles and supersonic / detonation wave combustion, add but a few percentage points to Isp, or no net improvement at all, merely simplified mechanical operation. For example, the rotating detonation wave engines are mechanically exceptionally simple, but horrifically complex from a fluid dynamics perspective. They could be very cheap to manufacture, if 3D printed. This seems to indicate that chemical reaction alone is a very inefficient propulsion method for in-space maneuvering. Somewhere between 75% and 95% of all tonnage delivered to orbit will then be devoted to propellants for human missions to other planets. This is acceptable for infrequent exploration missions, but untenable for colonization.

Solid core nuclear thermal rocket engines are useful intermediaries between the Isp of full electric and thrust of full chemical, with TWRs of 50:1 at best. 1,000s of Isp is a big step up from LOX/LH2 expander cycle engines, but in real terms it only cuts the required propellant tonnage for human missions to other planets in half. The net effect is that the propellant mass required to perform an "optimal" mission is approximately equal to the delivered payload mass. This might be acceptable if the propellant mass came from off-planet water sources.

Microwave-heated and confined plasma Ion engine:
Argon-fed VASIMR: 4,900s
Hydrogen-fed VASIMR: 32,000s

From reading through the literature in NTRS, the Aerojet-Rocketdyne X3 ion engine and all MPD thruster technologies (exclusively university-led engine development experiments) all have similarly high power inputs for a given propulsive force. We can look at the various other kinds of electric heating enhanced chemical engines, but thrust levels and Isp ranges are all remarkably similar for competing products. X3 has a modest input power advantage over VASIMR since X3 is not using microwave heating of the propellant, at the expense of being heavily optimized to run on Xenon or Argon, and significantly worse Isp than VASIMR.

solid core nuclear thermal engine:
LH2 NTR: 850s to 1,000s

pump-fed liquid chemical bi-propellant engines:
LOX/LH2: 475s
LOX/LCH4: 385s
LOX/LC3H8: 375s
LOX/RP1: 365s
NTO/MMH: 350s

Solid hybrid rocket motor:
LOX/Paraffin Wax (w/30% Aluminum powder): 340s

I would like to propose direct solar thermal heating as an alternative to nuclear heating we don't presently have. Achieving solid core NTR temperatures is not a problem, merely a matter of using mirrors to concentrate sunlight over a given surface area.

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
F = mdot * Ve + Ae(Pe - P0) <- Something for GW to calculate and optimize

Online JavaScript compiler / tester for executing the script shown below:
Programiz JavaScript Online Compiler

JavaScript Program for Computing Acceleration Time
// BEGIN PROGRAM
var velocity_initial = 7800;
var velocity_final = 10900;
var velocity_current = velocity_initial;
var mass_payload = 750000;
var mass_propellant = 500000;
var acceleration = 0;
var force = 8825.985;
var elapsed_acceleration_time = 0;
for(; velocity_current < velocity_final; elapsed_acceleration_time++, mass_propellant--){
if (mass_propellant === 0){
throw new Error("All propellant expended.");
break;
};
acceleration = force / (mass_payload + mass_propellant);
velocity_current += acceleration;
};
// Account for the fact that solar thermal only provides input power in the Sun
elapsed_acceleration_time = elapsed_acceleration_time * 2;
console.log("Elapsed acceleration time (seconds): " + elapsed_acceleration_time);
console.log("Propellant remaining (kg): " + mass_propellant);
console.log("Final velocity (m/s): " + velocity_current);
// END PROGRAM

// BEGIN SAMPLE OUTPUT
Elapsed acceleration time (seconds): 740456
Propellant remaining (kg): 129772
Final velocity (m/s): 10900.00259928668
// END SAMPLE OUTPUT

kbd512 · 2024-12-12 03:24:18

I'm introducing the concept of the Solar Thermal Reactor (STR) here, as a substitute for the Nuclear Thermal Reactor (NTR).

STR uses the same Hydrogen fuel NTR. A radial array of lightweight solar thermal mirrors tightly focus light from the Sun onto the surface of a Carbon / graphite thermal mass containing numerous internal Ultra-High Temperature Ceramic (UHTC) lined channels for Hydrogen gas to flow through to absorb thermal energy. The UHTC of choice for this application is Zirconium Carbide (ZrC), because it's proven to withstand hot flowing Hydrogen gas in NTRs. STR will also contain a small amount of supercritical Argon (sAr), acting as both a thermal energy distribution fluid to evenly heat the entire reactor pressure vessel interior to prevent cracking, but also to drive a pump to feed Hydrogen propellant into the reactor. There is no radiation shielding required for this STR, because it uses direct solar thermal input power from the Sun. However, it is still operating at exceptionally high temperatures and large thermal sinks are required to prevent the thermalized solar energy from literally melting the reactor core.

At 1AU distance from the Sun, we get 1,361W/m^2 from the Sun.
Aluminized mylar provides 95% reflectivity, but I'm asserting 90% of the radiation can be reflected and thermalized, so 1,224.9W/m^2.
We need to deliver 39,039,000Wth to superheat 1kg/s of LH2 propellant to 2,750K.
We'll round-up to 40MWth, just to be sure we supply enough input thermal power.
40,000,000W / 1,224.9W/m^2 = 32,656m^2
Aluminized mylar is 20.4g/m^2
32,656m^2 * 20.4g/m^2 = 666,182.4g = 666kg

Carbon / graphite bulk density: 2.26g/cm^3
ZrC bulk density: 6.56g/cm^3
Carbon sublimation point: 3,915K / 3,642C
ZrC melting point: 3,673K / 3,400C (923C above the nominal reactor operating temperature of 2,750K / 2477C)
STR operating temperature to achieve 900s Isp from LH2 propellant: 2,750K / 2,477C

Even if we require 1m^3 of thermal mass to prevent cracking or sublimation / "burn through", 2,260kg of engine mass plus 666kg for solar thermal collector material is quite reasonable.

Calliban · 2024-12-14 23:00:31

The simultaneous achievement of high thrust and high ISP, inevitably requires enormous power. The reason is really quite simple. ISP is just exhaust velocity divided by the force of Earth gravity. So ISP is a direct statement of exhaust velocity in a closed system. The kinetic energy of exhaust scales with the square of exhaust velocity.

KE = 0.5 x m x Ve^2

We can modify the equation slightly to convert KE into engine power:

Pe = (0.5 x m_dot x Ve^2)/n

Where m_dot is mass flow rate of propellant (kg/s) and n is engine efficiency.

Thrust can be expressed as the product of mass flowrate x exhaust velocity: m_dot x Ve. Substituting gives:

Pe = (0.5/n) x thrust x Ve

For any specific power level, we can have high thrust at low Ve, or low thrust at high Ve. But to have both simultaneously, requires high power.

In a chemical rocket, the kinetic energy of exhaust is derived directly from the thermal energy of reacting propellants. As no heat transfer is taking place, power can be extremely high. Efficiency, in terms of chemical energy converted into kinetic energy of exhaust, can approach 100%. Most other propulsion systems that we tend to consider, such as nuclear thermal, solar thermal and nuclear / solar electric, tend to require both or either heat transfer and energy transitions from electrical to thermal and kinetic energy. This intriduces both inefficiencies and rate limiting mechanisms, given the limits of practical energy flux without destroying materials. Systems that combine high thrust and ISP must store energy in the propellant itself. The effective ISP of the system can then be calculated as being proportional to the square root of the mass energy density of the propellant.

ISP = (K/2g) x Q^0.5, where Q is energy density in J/kg

For hydrogen-oxygen bipropellant, Q is 15.7MJ/kg. For 239Pu, Q is 83.6million MJ/kg. That is a 5.3 million times greater energy density in the fission fuel. This tells us that nuclear fission fuel could theoretically provide 2300x the exhaust velocity of H2-O2 bipropellant. Assuming complete fission, Ve would be 10,400km/s, which is just short of 3.5% C. That is the sort of performance needed for cheap interplanetary travel and interstellar travel in human timeframes. Fusion fuel has roughly 4x the mass energy density of fission fuel, implying Ve of 7% C. The maximum velicity achievable for a rocket accelerating from zero, is usually estimated as being twice exhaust velocity. Staging increases that to a limited extent. Under perfect conditions, fission fuels may allow construction of vehicles that can reach 10% of C. Fusion fuels might allow somewhat more, but will top out at 20% C.

In real life, only a fraction of the energy in propellant will be harvestable. And propellant may carry dead weight. So for fission rockets, 10% C is an upper limit that we are unlikely to approach. But even 1% C is ample velocity change to allow complete access to the solar system.

Last edited by Calliban (2024-12-14 23:52:59)

GW Johnson · 2024-12-15 09:52:31

There is one other thing to think about, and that is the pressure at which your process or system does its thing, whatever that might be. Energies and power levels set the speed of the ejected stream. But the pressure at which the energy is transformed or transferred sets the density in the process, in concert with the temperatures involved. If the process density is low, the thrust is going to be low from a device of any given size, period! THAT is why the electric propulsion devices so far have given whispers of thrust (milli-N to at most N) from hardware weighing 1 to many tons, not including the KW-to-MW power supplies they need. The pressure inside the chamber where the ions are accelerated is almost indistinguishable from vacuum. Why would one ever expect a different outcome?

The solar thermal rocket must take that into account. You must do the heating of the hydrogen at significant pressure, since you are getting your exhaust velocity from a compressible flow expansion process. The NERVA XE reportedly did this at 560 psia "chamber pressure", although there would be a strong pressure drop through the reactor core. I'm only guessing the "chamber pressure" to be the aft pressure that feeds into the nozzle. Whatever, the higher that nozzle feed pressure, the better. That notion applies to ALL rockets that use nozzles: chemical, nuclear, or solar!

You only get a lot of force out of a reasonably-sized exit area if the exit plane density is high, so that there is a very appreciable massflow through that exit area (at the very high velocity determined by the energy transformations reflected in chamber temperature). That exit plane density is proportional to the chamber density, in a ratio fixed by the expansion ratio of the nozzle (as are the exit plane pressure and temperature and Mach number). Simple as that.

GW

Last edited by GW Johnson (2024-12-15 09:58:54)

tahanson43206 · 2024-12-15 11:14:41

For all... This topic is about Orbital Mechanics....

It is intended to become a collection of knowledge about Orbital Mechanics.

Rocket engine design is now available as a topic: Rocket Engine Performance.

In that topic, all rocket engine designs are welcome for presentation, discussion and evaluation.

This topic is intended for discussion of the exact procedures spacecraft navigators and mission planners must follow to place their charges in the desired end point, despite the many efforts of the Universe to prevent success.

Thrust of a rocket engine, directed opposite to a particular star, is one of the key methods a spacecraft navigator needs to achieve mission objectives.

The shape of the rocket is not important. What IS important is the amount of thrust available, and the time that thrust is available.

(th)

kbd512 · 2024-12-17 00:21:53

Delta-V Guide to the Solar System

The Solar System - A Subway Map

Edit:
The image linked above comes from the following Space Exploration Stack Exchange Post:
Calculating the delta V budget from Earth to Mercury

You can compute delta-V budgets, at least for Hohmann transfers, by applying what the respondent named "SE stop firing the good guys" told the OP about "where he/she went wrong"

This user has provided a lot of orbital mechanics responses to questions posted in the Space Exploration Stack Exchange:
User - SE - stop firing the good guys

In his "About" section of his profile, it simply says "Patching my conics." I think that says its all. The dude likes orbital mechanics. He's been on the Space Exploration Stack Exchange for almost 10 years and remains an active member. I can't really tell, but he may have been a forum moderator at one point.

Last edited by kbd512 (2024-12-17 00:36:52)

GW Johnson · 2024-12-18 10:33:07

That "subway map" has reasonably good numbers, based on assumptions not explicitly given. Changing those assumptions changes the numbers, but not greatly. You're in the ballpark for getting started. You are much better off using more precise dV estimates, once you really start to size vehicles, though. And don't forget to look at vehicle thrust/weight at every burn, not just rocket equation dV, as that is how you make sure your burns are "impulsive" enough to use orbital dV's without gravity loss factoring. The thrust sets engine masses, which in turn affect your inert mass estimates. Those in turn critically affect your achievable mass ratios. Like everything else, even the rocket equation is subject to "GIGO" (Garbage In, Garbage Out).

GW

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