What is the maximum delta-v that can be achieved with an engine that has an ISP of 350 seconds?

Subway style Δv map:
Total Δv values
Δv change values
Δv with Phase Angles
Precise Total Δv values
Maximum Δv Chart
- This chart is a quick guide to what engine to use for a single stage interplanetary ship. No matter how much fuel you add you will never reach these ΔV without staging to shed mass or using the slingshot maneuver.
ISP (s) Max Δv (m/s) Engines 290 6257 LV-1
300 6473 24-77 320 6905 Mark-55 330 7120 Mainsail 350 7552 Skipper
360 7768 KS-25X4
370 7983 LV-T30
380 8199 KR-2L 390 8415 Poodle
800 17261 LV-N
- Copy template:
- TWR = F / (m * g) > 1
- When Isp is the same for all engines in a stage, then the Isp is equal to a single engine. So six 200 Isp engines still yields only 200 Isp.
- When Isp is different for engines in a single stage, then use the following equation:
- Isp = ( F1 + F2 + ... ) / ( ( F1 / Isp1 ) + ( F2 / Isp2 ) + ... )
- Isp = ( Force of Thrust of 1st Engine + Force of Thrust of 2nd Engine...and so on... ) / ( ( Force of Thrust of 1st Engine / Isp of 1st Engine ) + ( Force of Thrust of 2nd Engine / Isp of 2nd Engine ) + ...and so on... )
- Two engines, one rated 200 newtons and 120 seconds Isp ; another engine rated 50 newtons and 200 seconds Isp.
- Isp = (200 newtons + 50 newtons) / ( ( 200 newtons / 120 ) + ( 50 newtons / 200 ) = 130.89 seconds Isp
- For atmospheric Δv value, use atmospheric thrust values.
- For vacuum Δv value, use vacuum thrust values.

- Use this equation to figure out the Δv per stage:
- Δv = ln ( Mstart / Mdry ) * Isp * g
- Δv = ln ( Starting Mass / Dry Mass ) X Isp X 9.81
- Single Stage Rocket that weighs 23 tons when full, 15 tons when fuel is emptied, and engine that outputs 120 seconds Isp.
- Δv = ln ( 23 Tons / 15 Tons ) × 120 seconds Isp × 9.81m/s² = Total Δv of 503.2 m/s
- Simplified version of the Δv calculation to find the maximum Δv a craft with the given ISP could hope to achieve. This is done by using a magic 0 mass engine and not having a payload.
- Δv =21.576745349086 * Isp
- Explained / Examples:
- This calculation only uses the mass of the fuel tanks and so the ln ( Mstart / Mdry ) part of the Δv equation has been replaced by a constant as Mstart / Mdry is always 9 (or worse with some fuel tanks) regardless of how many fuel tanks you use.
- The following example will use a single stage and fuel tanks in the T-100 to Jumbo 64 range with an engine that outputs 380 seconds Isp.
- Δv = ln ( 18 Tons / 2 Tons ) × 380 seconds Isp × 9.81m/s² = Maximum Δv of 8199.1632327878 m/s
- Δv = 2.1972245773 × 380 seconds Isp × 9.82m/s² = Maximum Δv of 8199.1632327878 m/s (Replaced the log of mass with a constant as the ratio of total mass to dry mass is constant regardless of the number of tanks used as there is no other mass involved)
- Δv = 21.576745349086 × 380 seconds Isp = Maximum Δv of 8199.1632327878 m/s (Reduced to its most simple form by combining all the constants)
- How to calculate the Δv of a rocket stage that transitions from Kerbin atmosphere to vacuum.
- Assumption: It takes approximately 1000 m/s of Δv to escape Kerbin's atmosphere before vacuum Δv values take over for the stage powering the transition.

Mass units are arbitrary; use whatever you like, as long as you're consistent. Similarly, Velocity units are arbitrary; the delta-V computed will be in the same units. Specific impulse is in seconds.
A good explanation of delta-V is given in Chapter 2 of Space Settlements: A Design Study. The diagram to the right is based on Figure 2.2 of this book, and indicates the delta-V needed to transfer among various orbits.
Specific Impulse is exhaust velocity divided by g, the acceleration due to gravity on Earth's surface (9.80665 m/s^2). It's a common measure of the "mass efficiency" of a rocket engine. You can use the radio buttons above to convert between specific impulse and exhaust velocity.
The above calculator is very simplistic; it assumes no drag, and it allows only a single burn. But it can be applied to a great many models. For example, you could calculate the delta-V attainable by a baseball pitcher, throwing baseballs to change his orbit. Or, if you know the delta-V between an asteroid orbit and Earth orbit, you can calculate how much mass you'd have to eject from the asteroid to park it (which of course depends on how fast your mass launcher is).

Kerbal Space Program rocket scientist's cheat sheet: Delta-v maps, equations and more for your reference so you can get from here to there and back again.
- 1 수학
- 2 Math examples
- 3 See also
추력비(Thrust to Weight Ratio; TWR)
- → 참고하기: Thrust-to-weight ratio
이건 제 2 뉴턴 운동 법칙를 따릅니다. 만약 이 값이 1보다 작다면, 땅을 떠나지 못하겠죠. 발사하는 곳의 표면 기준의 중력가속도 값이 필요하다는 것을 잘 알아두세요!
- 엔진의 출력
- 전체 중량
- 중력 가속도 값
합산비추력(Combined Specific Impulse; Isp)
- → 참고하기: Specific impulse
If the Isp is the same for all engines in a stage, then the Isp is equal to a single engine. If the Isp is different for engines in a single stage, then use the following equation:
델타V (Delta-v; Δv)
- → 참고하기: Tutorial:Advanced Rocket Design
Basic calculation of a rocket's Δv. Use the atmospheric and vacuum thrust values for atmospheric and vacuum Δv, respectively.
- is the velocity change possible in m/s
- is the starting mass in the same unit as
- is the end mass in the same unit as
- is the specific impulse of the engine in seconds
True Δv of a stage that crosses from atmosphere to vacuum
|other bodies' data missing|
Calculation of a rocket stage's Δv, taking into account transitioning from atmosphere to vacuum. Δvout is the amount of Δv required to leave a body's atmosphere, not reach orbit. This equation is useful to figure out the actual Δv of a stage that transitions from atmosphere to vacuum.
Various fan-made maps showing the Δv required to travel to a certain body.

Although chemical rockets can produce millions of pounds of thrust for a few minutes at a time, ion engines produce thrusts measured in ounces for thousands of days at a time. In space, a little goes a long way. Let’s do the math!
The Deep Space I engines use 2,300 watts of electrical energy, and produced F= 92 milliNewtons of thrust, which is only 1/3 of an ounce! The spacecraft has a mass of m= 486 kg, so from Newton’s famous ‘F=ma’ we get an acceleration of a= 0.2 millimeters/sec/sec. It takes about 60 days to get to 1 kilometer/sec speeds. The Dawn mission, launched in 2007 has now visited asteroid Vesta (2011) and dwarf planet Ceres (2015) using a 10 kilowatt ion engine system with 937 pounds of xenon, and achieved a record-breaking speed change of 10 kilometers/sec, some 2.5 times greater than the Deep Space-1 spacecraft.
The thing that limits the thrust of the xenon-based ion engines is the electrical energy available. Currently, kilowatt engines are the rage because spacecraft can only use small solar panels to generate the electricity. But NASA is not standing still on this.
The NEXIS ion engine was developed by NASA’s Jet Propulsion Laboratory, and this photograph was taken when the engine’s power consumption was 27 kW, with a thrust of 0.5 Newtons (about 2 ounces).
An extensive research study on the design of megawatt ion engines by the late David Fearn was presented at the Space Power Symposium of the 56th International Astronautical Congress in 2005. The conclusion was that these megawatt ion engines pose no particular design challenges and can achieve exhaust speeds that exceed 10 km/second. Among the largest ion engines that have actually been tested so far is a 5 megawatt engine developed in 1984 by the Culham Laboratory. With a beam energy of 80 kV, the exhaust speed is a whopping 4000 km/second, and the thrust was 2.4 Newtons (0.5 pounds).
All we have to do is come up with efficient means of generating megawatts of power to get at truly enormous exhaust speeds. Currently the only ideas are high-efficiency solar panels and small fission reactors.

Delta V: 2 m/s
260 km X 345 km orbit to 265 km X 368 km orbit. Delta V: 7 m/s
265 km X 368 km orbit to 267 km X 391 km orbit. Delta V: 6 m/s
267 km X 391 km orbit to 300 km X 310 km orbit. Delta V: 32 m/s
Total Delta V: 83 m/s
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[Via Satellite 07-20-2015] The Indian Space Research Organization (ISRO) completed a hot test of a high thrust cryogenic rocket engine for a duration of 800 seconds. The engine generated a nominal thrust of 19 tons and ran for approximately 25 percent more time than the engine burn required for flight.
ISRO is designing the engine to power the cryogenic stage (C25) — the upper stage of the next generation Geosynchronous Satellite Launch Vehicle (GSLV) Mk-3 — capable of launching 4-ton-class satellites. The space agency’s Liquid Propulsion Systems Center (LPSC) created the engine.
According to ISRO, the performance of the engine closely matched pre-test predictions made with in-house developed cryogenic engine mathematical modeling and simulation software. The July 16 hot test was the 10th in a series of tests planned and executed as part of the development of the engine. ISRO is planning further tests in high altitude conditions and in stage configuration prior to flight stage realization.

If your delta-V is too minimal the spacecraft will not be capable to perform any useful missions whatsoever.
Spacecraft have parameters too, it is simply that they are odd actions you have not encountered in advance of. I am going to record the greater significant ones below, but they will be thoroughly discussed on other pages. Refer back to this list in case you operate throughout an unfamiliar term.
When this summary was put before Rob Herrick, an epidemiologist, he did not Assume it had been possible.
The fourth and many Excessive tactic is always to cheat the equation itself, for making the whole equation not pertinent to the spacecraft. The equation assumes which the spacecraft is carrying all of the propellant desired for the mission, This may be bent quite a few means.
All three people are in use for many years, though each loved ones has somewhat modern users pushing the limits. All share common basic physics with regards to their efficiencies.

“Only 3 seconds, but got through the ignition sequence, which is always the big question at this phase,” Garvey said in an email. “Used techniques previously developed with our 500 lbf LOX/propylene engine. Also continued to evaluate thrust vector control.
“The work plan still is to conduct additional testing this spring / summer (more ignition cycles, extend the duration), followed eventually by a flight test featuring this engine,” Garvey wrote.
The company is developing the engine under a NASA Small Business Innovation Research (SBIR) Phase II contract awarded last year. The award was for a total of up $700,000 for work over a period of two years.
Garvey’s initial goal is to deliver 10 kg payloads into a 250-km orbit. A larger version of the booster will be designed to place satellites weighing up to 20 kg into a 450-km orbit.

This page no longer updated from 31 October 2001. Latest version can be found at www.astronautix.com
Designer: Isayev. Propellants: N2O4/UDMH Thrust(vac): 3 kgf. Thrust(vac): 0.03 kN. Isp: 285 sec. Burn time: 25,000 sec. Mass Engine: 1 kg. Chambers: 1. Chamber Pressure: 8.00 bar. Area Ratio: 45.00. Oxidizer to Fuel Ratio: 1.85. Country: Russia. Status: In Production. References: 350 . Comments: Bi-propellant hypergolic (self-igniting) engine, pressure-fed. 300,000 ignitions.
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Last update 12 March 2001.
Definitions of Technical Terms.
Contact Mark Wade with any corrections or comments.
Conditions for use of drawings, pictures, or other materials from this site..
© Mark Wade, 2001 .

Then it continues at the same speed for 5.83 years before the engines or photon sail (depending on which way the ships are traveling) are used to decelerate the vehicle. The ship's deceleration phase also lasts for five and a half months at 1.5G.
Also notable are the time dilation effects experienced at higher speeds; an Earth-time voyage of 6.75 years seems significantly shorter at 0.7 times the speed of light. In accordance with Einstein's theory of relativity, from the crew's point of view it is only five years' travel due to time dilation.
The ISV Venture Star uses multiple types of propulsion, all of which are used over the course of an interstellar journey between Earth and Pandora. The Venture star has:
- Two matter-antimatter engines
- One photon sail
- One fusion PME (Planetary Maneuvering Engine).
The photon sail is used for the outward acceleration phase to Pandora from Earth in the form of beamed photons from a solid state molecular laser from Earth. The ship's matter-antimatter engines are used for the deceleration phase on approach to Pandora, and the ship then coasts the last few million kilometers. The sequence is reversed for return to Earth. When close enough to Pandora or Earth, the Planetary Maneuvering Engine is used to maneuver into a low delta-v orbit, from which it can launch Valkyrie shuttle craft to the surface.
Matter-Antimatter engines Edit
Interstellar Vehicles require a huge amount of thrust in order to reach the speeds that are needed for economic and (relatively) swift travel between solar systems and stellar bodies. Humans took roughly two centuries to create advanced and reliable matter-antimatter based propulsion systems.
An ISV has two matter-antimatter engines arranged symmetrically in a tractor configuration that pulls the ship behind them. They are angled slightly away from the body of the ship, a few degrees off the ship’s longitudinal axis, so their exhaust plumes bypass the ship’s structure. This results in a slight loss of thrust efficiency because the engines do slightly push toward each other. The lost thrust is deemed acceptable because the angling separates the body of the ship from the plume’s thermal radiation. Scientists considered placing the engines at the back instead, but the mass-savings advantage of a tensile structure outweighs the disadvantages of shielding.

On Saturday I investigated SEP systems for spacecraft going from LEO to the Moon and having a relatively small LEO mass of 20 tonnes. In hindsight, some of my numbers were a bit dodgy; more realistic values are for 15 tonnes of non-propulsion equipment.
But what happens when there's a possible initial mass of 100 tonnes, as could be launched to LEO by a Saturn V?
At this scale, SEP doesn't make sense. Sticking to my maximum transfer time of 2 years, the math works out as follows: mf / m0 = 0.843, so 84.3 tonnes of payload and systems and 15.7 tonnes of fuel. Using just one ion thruster as before, a Δv of 6.33x103 ms-1 would take 100 years. That means that to transfer in 2 years, I'd need 50 thrusters, corresponding to 250 kW of power consumption.
That's an awful lot of power, and a solar panel array that big is just ridiculous, without even taking radiation degradation into account. So, nuclear power could be used instead. Up to 1993 as part of the Space Defence Initiative NASA developed a reactor designated the SP-100 . This was capable of delivering power levels up to about a MW at between 10-50 kgkW-1 (better specific mass at higher power). By the way: for the environmentalists, the SP100's nuclear fuel container is designed to reach the ground intact in the event of the reactor unexpectedly breaking up in the atmosphere. Additionally, I'm assuming a 1000 km starting orbit, and an orbital height of 700 km is considered by NASA to be "nuclear safe".
Pulling figures out of thin air, let's assume that a 250 kW reactor can be built at 30 kg/kW. That means that 7.5 tonnes would be power source. Additionally, I'd need 250 kW of ion thrusters, at about (yep, more made-up numbers) 13 kg/kW, another 3.3 tonnes.
Adding all that up, for 100 tonnes at LEO, I'd get 15.7 tonnes of fuel, 7.5 tonnes of nuclear reactor and 3.3 tonnes of thrusters, leaving 73.5 tonnes for payload. Very nice.

• Apoapsis Delta V Burn: The rocket firing at the highest point of a Transfer Orbit.
• Periapsis Delta V Burn: The rocket firing at the lowest point of a Transfer Orbit.
• Transfer Time: The time between apoapsis and periapsis Delta V rocket firings.
Getting from one orbital altitude to another not only takes delta v, it also takes time. It isn’t as simple as aiming yourself and firing your rocket. Orbital mechanics is a ballet in space, where the lower your orbit the faster you go, and vise versa.
Any change in orbital velocity and you wind up in an elliptically-shaped orbit, instead of the fairly stable circular orbit.
Transferring from one orbit to another is accomplished by changing a circular orbit (represented in by the green and red circles in the image above) into an elliptical orbit (represented by the yellow curve). The Transfer Time is the time it takes for the spacecraft to follow the path of the yellow curve.
To get from one orbit to another, we must push ourselves off, and then brake when we get there. We coast in between these two maneuvers. We can calculate how much time we will coast from one point in space to another.
The equation to calculate the transfer time from one orbital altitude to another is given by one of the Hohmann Transfer Orbit Equations:
• Ts is Transfer Time in seconds
• r1 is the smaller orbital radius
• r2 is the larger orbital radius
• µ (mu) is Earth Standard Gravity = 9.80665 m/s^2
• π (pi) is the constant 3.141592653589793238462643383279502884197169399…
The number π (pi) is involved in the equation because we are actually going around in circles.
Let’s suppose you want to go from the Hubble Space Telescope (HST) to the International Space Station (ISS). How long would it take to get there?
The orbital radius is simply the sum of the orbital altitude plus the radius of the Earth. So for the HST (r1), we get 6,708.14 km, and for the ISS (r2) we get 6,948.14 km.
Plugging everything in the equation, we get 2,673 seconds of transfer time. If we divide that by sixty (seconds per minute) we get 44.55 minutes.

- Note: This equation is an guess, approximation, and is not 100% accurate. Per forum user stupid_chris who came up with the equation: "The results will vary a bit depending on your TWR and such, but it should usually be pretty darn accurate."
- Equation for Kerbin Atmospheric Escape:
- True Δv = ( ( Δv atm - 1000 ) / Δv atm ) * Δv vac + 1000
- True Δv = ( ( Total Δv in atmosphere - 1000 m/s) / Total Δv in atmosphere ) X Total Δv in vacuum + 1000
- Single Stage with total atmospheric Δv of 5000 m/s, and rated 6000 Δv in vacuum.
- Transitional Δv = ( ( 5000 Δv atm - 1000 Δv Required to escape Kerbin atmosphere ) / 5000 Δv atm ) X 6000 Δv vac + 1000 Δv Required to escape Kerbin atmosphere = Total Δv of 5800 m/s

They predict their prototype could produce a specific impulse (Isp) of 2000 sec, which is an equivalent to an exhaust velocity of 20,000 m/s.
They are looking to raise $69,000 by November 3, 2012 to get their project started. At the time of this writing, the team has just over $54,000.
Here’s a video from HyperV:
“We invite you, the citizens of Earth, to join with us as we design, construct, test, and execute this demonstration,” the team wrote on their Kickstarter page. “The culmination of this project will be an all-up, laboratory demonstration of our prototype thruster.”

But there is a solution; Nuclear Thermal Rockets!
NASA and friends did quite a bit of research into NTRs back in the 1960s. They even got as far as ground testing one.
Here's a still photo of it;
According to the Atomic Rockets site (which everyone and their children should read), you can get an exhaust velocity of aboot 8 km/s if you use hydrogen as your propellant (light things are better propellants, and hydrogen is pretty light). Which translates to a specific impulse of about 800 seconds. Pretty damn solid, especially compared to chemical rockets. Some of the later design studies I've seen have said we could get up to 950 seconds out of better designs.
This is really, really helpful for your hypothetical Mars/Jupiter/Saturn mission. Reducing the amount of fuel mass you need is going to reduce the amount of launches you need to put your mission together, thereby saving you money. This is important, since NASA's budget currently is about enough to pay for a small pizza, or a six pack of crap beer, but not both. Alternatively, you can still put a massive spacecraft together, but take a less efficient trajectory, and get to your destination quicker. Since spending months in space exposed to cosmic radiation and zero g is a nontrivial problem, this is pretty cool.
|A nice chart from Atomic Rockets showing the relationship between trip time and delta-v.|

The velocity of a rocket bike depends on the type of rocket attached to the bike. More specifically, It depends on the Delta-V that the rocket can impart on the mount that is connected to the bike. Delta V is the effective exhaust velocity c of the rocket times the natural log of the initial mass over the final mass.
The initial mass will be the mass of the biker, the bike, the rocket mount, and the rocket itself. The final mass will be everything except for the mass of the propellant used in the rocket.
The change in velocity of the bike will be negligible if the mass of the propellant used is not a significant percentage of the bike’s mass.
The effective exhaust velocity can be found by multiplying the gravity at sea level by the rocket’s specific impulse.
If you take a G class rocket which is the largest rocket motor that a person can buy without needing a certification, then you’ll roughly have 160 Newton seconds of total impulse.
To go from total impulse to specific impulse, divide the total impulse by the weight of the rocket motor.
This will lead to an effective exhaust velocity of 1406 m/s.
Now you’ll need the weight before and after the rocket has spent its propellant. Let’s say that bike has a mass of 18 lbs with 175 lbs. person driving the bike. The 130 grams rocket motor and rocket mount weighs about 87671 grams. Let’s consider that mass of everything after the rocket has burnt all of the propellant is 87110 grams.
Given that the average burn time for one of these rocket motors is 3 seconds, the acceleration that the rocket will experience is
The total distance traveled, negating friction will be 2.6 meters.

The velocity change would be at the rate of 3,000 m/s per year of thrusting by the photon rocket.
If a photon rocket begins its journey in low earth orbit, then one year of thrusting may be required to achieve an earth escape velocity of 11.2 km/s if the vehicle is already in orbit at a velocity of 9,100 m/s, and 400 m/s additional velocity is obtained from the east to west rotation of the earth. The photon thrust will be sufficient to more than counterbalance the pull of the sun's gravity, allowing the photon rocket to maintain a heliocentric velocity of 30 km/s in interplanetary space upon escaping the Earth's gravitational field. Eighty years of steady photonic thrusting would be then required to obtain a final velocity of 240 km/s in this hypothetical case. At a 30 km/s heliocentric velocity, the photon ship would recede a distance of 600,000,000 miles (1 Tm) from the Sun per year.
It is possible to obtain even higher specific impulse; that of some other photonic propulsion devices (e.g., solar sails) is effectively infinite because no carried fuel is required. Alternatively, such devices as ion thrusters, while having a notably lower specific impulse, give a much better thrust-to-power ratio; for photons, that ratio is 1 / c, whereas for slow particles (that is, nonrelativistic; even the output from typical ion thrusters counts) the ratio is 2 / v, which is much larger (since ). (This is in a sense an unfair comparison, since the photons must be created and other particles are merely accelerated, but nonetheless the impulses per carried mass and per applied energy—the practical quantities—are as given.) The photonic rocket is thus wasteful when power and not mass is at a premium, or when enough mass can be saved through the use of a weaker power source that reaction mass can be included without penalty.
A laser could be used as a photon rocket engine, and would solve the reflection/collimation problem, but lasers are absolutely less efficient at converting energy into light than blackbody radiation is—though one should also note the benefits of lasers vs blackbody source, including unidirectional controllable beam and the mass and durability of the radiation source.
Feasible current, or near-term fission reactor designs can generate up to 2.2 kW per kilogram of reactor mass.

For the latter, the distances are (relatively) small enough that we have the luxury of just being able to coax something onto a transfer orbit using a very small amount of propellant and then leaving it until it eventually arrives anywhere from a few months to a decade later. These are both impractical approaches to interstellar travel for obvious reasons: rocket engines require a lot of propellant which is quickly exhausted, rendering the engine useless, while interplanetary probes might sound nippy but they are, in fact, achingly slow. For an example, the Voyager probes are the fastest man-made objects in existence, taking advantage of a once-every-175-years alignment of the planets to do some crazy gravitational slingshotting that accelerated them up to a rather bracing 15 km s-1. Let’s say we could do that again for our interstellar probe, and that we could do it in such a way that it would end up pointed at Alpha Centauri. At 15 km s-1 the probe would cover the four light years separating the Sun and Alpha Centauri in just under 80,000 years.
Clearly this isn’t practical if we want human civilisation to still exist when our probe reaches its destination, so if we intend to try interstellar travel in a serious way then we need to experiment with something a little different. Where achieving orbit is a matter of delivering a lot of acceleration quickly, and interplanetary travel relies on a small amount of acceleration delivered early, practical interstellar travel requires that we instead subject our probe to a small amount of acceleration over a very long timespan. The thing with velocity in space is that it’s cumulative. There’s no air resistance dragging you back; once you add 5 km s-1 to a spacecraft’s relative velocity that 5 km s-1 will stay there forever until you remove it by making a burn in the opposite direction and manually decelerating the spacecraft. An engine that provides an acceleration of 2 cm s-2 might not sound very good, but if it can provide that rate of acceleration for a full year the spacecraft it’s attached to will end up travelling at a whopping 630 km s-1, covering the distance to Alpha Centauri in a nippy 452 years. And that’s if it stops accelerating after a year, remember. If it keeps going there’s no reason it couldn’t reach a significant fraction of light speed.

So I am modelling an Earth - Mars low thrust trajectory in a simple manner. I'm telling my model to not model any trajectories where the spacecraft is moving any more than 200 m/s quicker or slower than the planet upon rendezvous. This is just for the conic section. So when the spacecraft arrives at Mars, I calculate its speed relative to the planet at that point, and I am treating that as the hyperbolic excess speed (V_infinity), which is constrained to be less than 200m/s. Now, I am getting trajectories that fit the criteria, and they happen to burn very little fuel throughout the journey - my question is, is this physically possible, is it real? Is it even desirable to have low arrival C3's? All the literature I read have arrival C3's/V_infinities in the order of km's/s (or km's^2/s^2) but I have it constrained to mere m's/s. I would have thought, intuitively, the closer the s/c can get to matching the planets velocity, the less of a burn needed for orbital capture, ergo fuel saved?
My orbit looks like a spiral thats made 2 revolutions of the sun. Time of flight, approximately 1400 days. Delta - V is much larger than a chemically propelled ship, but that's OK I guess as the fuel usage is lower. Plotting Mars as an elliptical orbit, not just circular. Hope that helps.

VASIMR plasma engine: Earth to Mars in 39 days?
In Arthur C. Clarke’s classic science fiction novels and movies 2001: A Space Odyssey and 2010: Odyssey Two, the spaceships Discovery and Alexei Leonov make interplanetary journeys using plasma drives. Nuclear reactors heat hydrogen or ammonia to a plasma state that’s energetic enough to provide thrust.
An electric power source ionizes hydrogen, deuterium, or helium fuel into a plasma by stripping away electrons. Magnetic fields then direct the charged gas in the proper direction to provide thrust.
“A rocket engine is a canister holding high-pressure gas,” Chang Diaz explained. “When you open a hole at one end, the gas squirts out and the rocket goes the other way. The hotter the stuff in the canister, the higher the speed it escapes and the faster the rocket goes. But if it’s too hot, it melts the canister.”
The VASIMR engine is different, Chang Diaz explained, because of the fuel’s electrical charge: “When gas gets above 10,000 [kelvins], it changes to plasma – an electrically charged soup of particles. And these particles can be held together by a magnetic field. The magnetic field becomes the canister, and there is no limit to how hot you can make the plasma.”
Chang Diaz has pointed out that hydrogen would be an advantageous fuel for the VASIMR engine because the spacecraft would not have to lift off carrying all the fuel it needs for the journey.
“We’re likely to find hydrogen pretty much anywhere we go in the Solar System,” he said.
A spacecraft using conventional chemical rockets would take eight months to get to Mars during opposition. However, the VASIMR engine would make the journey in as little as 39 days.
Chang Diaz explained: “Remember, you are accelerating the first half of the journey – the other half you’re slowing, so you will reach Mars but not pass it. The top speed with respect to the Sun would be about 32 miles per second [or 51.5 km/s]. But that requires a nuclear power source to heat the plasma to the proper temperature.”
The use of nuclear power in space is not without its controversy. In 1997, there was widespread public concern when NASA’s Cassini probe, which carried a plutonium battery, made a flyby of Earth to perform a gravity assist.

- Orbit Kerbin
- Plan maneuvers to encouner Duna
- Execute maneuvers
- Get surface sample
- Orbit Duna
- Plan maneuvers to encouner Kerbin
- Execute maneuvers
- Basic Rocketry
- Engineering 101
- General Rocketry
- Advanced Rocketry
- Building level 2
- Launch pad due to rocket mass
- Tracking station for visible conics
- Mission control for maneuver planning
- Austronaut complex for EVA
- R&D for surface sample
Pilot with one or more stars (orbiting Kerbin will do this)
The delta-v map at https://i.imgur.com/gBoLsSt.png says we need
3400 m/s for low Kerbin orbit (LKO) (that's very generous)
950 m/s to get to the edge of Kerbin's sphere of influence (SOI)
130 m/s to encounter Duna
10 m/s or less for a plane change
0 m/s aerobraking
20 m/s landing burn
1450 m/s to low Duna orbit (LDO)
360 m/s to Duna's SOI
250 m/s to encounter Kerbin
6580 m/s total
These are all vacuum delta v's. From LKO we need 3180 m/s. That's a more accurate number.
- Mk16 Parachute
- Mk1 Command Pod
- TD-12 Decoupler
- FL-T400 Fuel Tank
- LV-909 "Terrier" Liquid Fuel Engine
- TD-12 Decoupler
- FL-T400 Fuel Tank
- FL-T400 Fuel Tank
- FL-T400 Fuel Tank
- FL-T400 Fuel Tank
- 4x Basic Fin
- LV-T30 "Reliant" Liquid Fuel Engine
- TT-38K Radial Decoupler
- Aerodynamic Nose Cone
- BACC "Thumper" Solid Fuel Booster
Delta-v: 6162 m/s
Mount the boosters at the lower end of the decouplers, so their top turns away at decoupling and aerodynamics drives them away from the rocket. I put the bottom of the boosters slightly below the Reliant so the rocket would stand more stably on the launch pad. Launch stability enhancers do a better job at that, but due to minimalistic tech I don't use them.
Remove the monopropellant from the command pod.

The rocket shall have a display on the exterior identifying the calculated center of pressure consistently optimizing virtually an unlimited range of radial burning solid rocket motor grain geometries. Optimization tools were applied to the design process of solid rocket motor grains through an optimization framework developed to interface optimization tools with the solid rocket motor design system The QUANTUM LEAP II is an awesome two stage kit for Level 1 and 2 flights. This rocket stands 91.5 (fully configured) tall, is 3.0 diameter, and has a 8 total fin design. This rocket comes standard with a 54mm Kwik-Switch motor mount and can fly on H thru J motors
United Model 5322 Destination Mars Colonizer Model Rocket Starter Set - Includes Rocket Kit (Beginner Skill Level), Launch Pad, Launch Controller, Glue, Four AA Batteries, and Two Engines. 5.0 out of 5 stars. 1. $64.99. $64 Steve Eves broke two world records when his 1/10th scale model of a Saturn V rocket lifted off from a field on Maryland's Eastern Shore. The 36-ft.-tall rocket was the largest amateur rocket ever. Never thought we'd say this, but if the standard G-Class was too vanilla for you, Brabus has a flavour that might be right up your alley. The latest to join the Rocket series of projects by the German automotive tuning company is the Brabus 900 Rocket Edition, and like its name suggests is a poster child for an unnecessary amount of power and mind-blowing acceleration — two qualities more.
Rocket engines are reaction engines. The basic principle driving a rocket engine is the famous Newtonian principle that to every action there is an equal and opposite reaction. A rocket engine is throwing mass in one direction and benefiting from the reaction that occurs in the other direction as a result The total impulse (I,N-s) of a rocket motor/engine is defined as: I total! r F(t)dt 0 T (4.1) For F and G class motors, the impulse will be roughly between 50-60 Ns and 100-120 Ns, respectively. The total impulse and used propellant weight during burn of duration T can be used to calculate the specific impulse (I sp ,sec). ! I sp.

That implies a dV ~ 6 km/s, but Cassini’s actual SOI was a mere 0.626 km/s, which means Baxter used the wrong figures. Kind of odd since he had copied the “Discovery’s” flight-plan directly from Cassini – every event day in the long voyage is direct from Cassini’s chief events. Since the launch was set in 2008 this is unlikely – the planets would rarely line up for exactly the same flight-plan.
Here’s Cassini’s approach by orbital radii that I could get figures for…
Radial Distance (km) Speed (km/s) Notes
158,500 22.5 Ring plane
24.00 SOI burn begin
80,230 31.50 Periapsis
30.40 SOI burn end
Cassini began its journey with 3,132 kg of propellant – good old space-storable N2O4/UDMH. That mix gets 3,028 m/s exhaust velocity, and Cassini’s main engine puts out 445 N thrust. Total dV ~ 2,500 m/s. A previous major burn was at the last aphelion before the Venus/Earth swing-by that sent Cassini to Jupiter/Saturn – it lasted 88 minutes.
The SOI burn lasted 96.4 minutes and used 850 kg of Cassini’s propellant . Here’s a table summarising the different burns so far – that I have records of…
Duration (minutes) Mass Used (kg) Notes
88 776 Deep Space Maneuver
6 53 TCM 20
1.2 10.6 TCM 21
96.4 850 SOI
61.8 545 Periapsis raise
253.4 2235 Summation
102 897 difference
The trick now – as far as recreating Baxter’s figures for “Discovery” goes – is working out the dV needed to send the vehicle on its way from LEO. Baxter has four revived Saturn V’s reconditioned to orbit fuelled up Saturn IVB stages at LEO to boost the Shuttle on its way. That’s a lot of hardware, but that’s chemical rockets for you. Once in orbit the Shuttle can dump its SSMEs – which are superfluous for N2O4/UDMH burns – and save lots of extra weight by dumping the tail plane.

The one thing to keep in mind is that in order to perform a gravity-assist maneuver, you need to be able to enter a hyperbolic orbit around a given body that is moving relative to your destination. And, in order to be in such an orbit, there is a specific range of velocities for every object that you must have (dependent on mass of the object). So the fastest you can get to by gravity-assist is much less than relativistic speeds because at relativistic speeds, you would not be able to enter into a proper hyperbolic orbit.
It is true that at any high speed, a flyby constitutes a hyperbolic orbit; however, to use a gravitational slingshot, you need to enter against the object's motion and exit with the motion. At relativistic speeds and for most regular bodies, your orbit would closely resemble a straight line, there could be no gain of velocity.
A good gravity assist works if you can ensure that your hyperbolic trajectory minimizes the angle $\theta$ between entry and exit. It is given by:
Where $e$ is the eccentricity of the orbit and must satisfy $e\geq1$. Additionally, as your velocity increases, it will force $e$ to become larger unless you significantly increase the mass of each subsequent object.
The fastest a spacecraft can get to using gravity-assists very much depends on the largest mass of the objects you use. However, I cannot give you an estimate of a number because due to the sheer impracticality of using gravity-assists to achieve extreme velocities, we (rocket scientists) haven't ever tried computing a theoretical limit. I can guarantee you though that without using high density objects (neutron stars, black holes, etc.), no spacecraft will reach velocities near the speed of light by gravity slingshots alone.

It is not as clear that hall thrusters are a good solution at an isp of 1000s or lower as their efficiencies drop off rapidly.

1. The problem statement, all variables and given/known data A rocket is launched at an angle of 53° above the horizontal with an initial speed of 75 m/s. It moves in powered flight along its initial line of motion with an acceleration of 25 m/s^2. At 30 s(seconds) from launch, its engine fails and the rocket continues to move as a projectile. (a) What is the rocket's max altitude? (ymax) (b) What is the rocket's total time of flight? (tτ) (c) What is the rocket's horizontal range? (R/x) My given information looks like this: θ= 53°, v1= 75 m/s, a=25 m/s, t= 30s, Δy=0 2. Relevant equations x=vcosθt y=vsinθt+.5at^2 ymax= -(vsinθ)^2/2a 3. The attempt at a solution I tried resolving the velocity into x and y components but i'm not even sure if that's useful information. I also attempted to use the above equations but to no avail... I just need to get this started so that I can finish it. The problem is i'm not sure where to start.

For 60 years, NASA has used chemical rockets to send its astronauts into space, and to get its spacecraft from planet to planet. The huge million-pound thrusts sustained for only a few minutes were enough to do the job, mainly to break the tyranny of Earth’s gravity and get tons of payload into space. This also means that Mars is over 200 days away from Earth and Pluto is nearly 10 years. The problem: rockets use propellant (called reaction mass) which can only be ejected at speeds of a few kilometers per second. To make interplanetary travel a lot zippier, and to reduce its harmful effects on passenger health, we have to use rocket engines that eject mass at far-higher speeds.
In the 1960s when I was but a wee lad, it was the hey-day of chemical rockets leading up to the massive Saturn V, but I also knew about other technologies being investigated like nuclear rockets and ion engines. Both were the stuff of science fiction, and in fact I read science fiction stories that were based upon these futuristic technologies. But I bided my time and dreamed that in a few decades after Apollo, we would be traveling to the asteroid belt and beyond on day trips with these exotic rocket engines pushing us along.
Well…I am not writing this blog from a passenger ship orbiting Saturn, but the modern-day reality 50 years later is still pretty exciting. Nuclear rockets have been tested and found workable but too risky and heavy to launch. Ion engines, however, have definitely come into their own!
Most geosynchronous satellites use small ‘stationkeeping’ ion thrusters to keep them in their proper orbit slots, and NASA has sent several spacecraft like Deep Space-1, Dawn powered by ion engines to rendezvous with asteroids. Japan’s Hayabusha spacecraft also used ion engines as did ESA’s Beppi-Colombo and the LISA Pathfinder. These engines eject charged, xenon atoms ( ions) to speeds as high as 200,000 mph (90 km/sec), but the thrust is so low it takes a long time for spacecraft to build up to kilometer/sec speeds. The Dawn spacecraft, for example, took 2000 days to get to 10 km/sec although it only used a few hundred pounds of xenon!
But on the drawing boards even more powerful engines are being developed.

By rearranging the terms in (1), one obtains: Under the assumption that the total propulsive power of the IBS spacecraft is constant and that the total propulsive thrust is proportional to it , one can write thus Therefore, the maximum acceleration acting on the IBS-debris combination can be computed as a function of the total available thrust : It is assumed here to have a 1000 kg IBS spacecraft with a total available thrust of 0.5 N. Such a high thrust would correspond to a substantial power and propulsion system mass, however, this is deemed realistic if one considers that the payload of the IBS spacecraft is in fact its propulsion and power systems. Hence, the propulsion and power systems might be oversized compared to other applications in which ion engines are used for propulsion only. Note that the validity of the methodology proposed in this paper would not be affected even if lower thrust levels were considered. Thus, in this case, considering, for example, an 800 kg debris, the magnitude of the acceleration would be 1.923·10-7 km/s2. If one considers instead the spacecraft alone, the acceleration achievable would be slightly higher, 5·10-7 km/s2. Given this order of magnitude, the thrust acceleration can be considered as a perturbative force compared to the Earth’s gravitational force, and therefore, the analytical approach to the propagation of the low-thrust motion described in  can be applied.
3. Mission Profile
The objective of this study is that of optimising the performance and cost of a debris de-orbiting mission performed by a single spacecraft. As mentioned in the introduction, it is assumed that there are five pieces of debris of different masses and lying in circular orbits with different radii and orientations. It is assumed that the IBS spacecraft departs from a low-Earth parking orbit, rendezvous with the first object, transfers it to an elliptical re-entry orbit, rendezvous with the second object, transfers it to a second elliptical re-entry orbit, and so forth until all five pieces of debris are removed. One important issue is defining in which order the pieces of debris need to be de-orbited. In the following, all possible sequences are generated a priori and optimised one by one.

As for speed, in the first case where Serenity just accelerates continuously at 5.5g (which minimizes fuel), its final velocity after having travelled for those 141.5 hours will be around 0.675c. In the second case where it accelerates for 96.2 hours until the midpoint of the journey and then decelerates, its speed at the midpoint will be 0.528c. Calculations:
Another formula on the relativistic rocket page says that if I accelerate continuously, my final change in velocity at the end should be v = (a * t) / (sqrt(1 + (a * t/c)²)), in this case (0.00647*141.515)/(sqrt(1 + (0.00647*141.515)²)) = 0.675298 light-hours/hour, i.e. 0.675c. And for the second case of 96.2 hours, we have (0.00647*96.2)/(sqrt(1 + (0.00647*96.2)²)) = 0.528c.
Using this change in velocity, the Tsiolkovsky rocket equation can tell you the ratio of initial mass m₀ (both fuel and payload) to final mass m₁ (just the payload, fuel used up) needed to get such a large change in velocity. The formula depends on the velocity of the exhaust, and I don't think that's stated directly in any Firefly material, but this answer says some material on the DVD indicated that Firefly-class ships light up their tails when they travel due to a nuclear fusion reaction. So Serenity is probably a type of fusion rocket, perhaps similar to Project Daedalus in which some real-world engineers tried to draw up a rough plan for an interstellar space probe, and which used a fusion reaction involving deuterium and helium-3. The chart on this page indicates that this is one of the best fusion reactions in terms of exhaust velocity, with a v_e of about 8.9% the speed of light or 0.089c (which is incidentally much, much faster than any chemical rocket like the ones we have today), so I assumed that in my calculations.

A next-generation plasma rocket being developed by former NASA astronaut Franklin Chang Diaz called the Variable Specific Impulse Magnetoplasma Rocket (VASIMR) has been touted as a way to get astronauts to Mars in weeks rather than months, as well as an innovative, cheap way to re-boost the International Space Station. But in a biting commentary posted on Space News and the Mars Society website, “Mars Direct” advocate Robert Zubrin calls VASIMR a “hoax” saying the engine “is neither revolutionary nor particularly promising. Rather, it is just another addition to the family of electric thrusters, which convert electric power to jet thrust, but are markedly inferior to the ones we already have,” adding, “There is thus no basis whatsoever for believing in the feasibility of Chang Diaz’s fantasy power system.”
The VASIMR uses plasma as a propellant. A gas is ionized using radio waves entering into a plasma state. As ions the plasma can be directed and accelerated by a magnetic field to create specific thrust. The purported advantage of the VASIMR lies in its ability to change from high impulse to low impulse thrust as needed, making it an ideal candidate for a mission beyond low Earth orbit.
Chang Diaz’ company, the Ad Astra Rocket Company successfully tested the VASIMR VX-200 plasma engine in 2009. It ran at 201 kilowatts in a vacuum chamber, passing the 200-kilowatt mark for the first time. “It’s the most powerful plasma rocket in the world right now,” said Chang-Diaz at the time. Ad Astra has signed a Space Act agreement with NASA to test a 200-kilowatt VASIMR engine on the International Space Station, reportedly in 2013.
The tests would provide periodic boosts to the space station, which gradually drops in altitude due to atmospheric drag. ISS boosts are currently provided by spacecraft with conventional thrusters, which consume about 7.5 tons of propellant per year. By cutting this amount down to 0.3 tons, Chang-Diaz estimates that VASIMR could save NASA millions of dollars per year.
For the engine to enable trips to Mars in a reported 39 days, a 10- to 20-megawatt VASIMR engine ion engine would need to be coupled with nuclear power to dramatically shorten human transit times between planets.

Voyager 1, launched in 1977, is the furthest manmade object from Earth, traveling at 10.6 miles per second. Even traveling at this incredible speed, it would take just over 70,000 years to reach the closest star to our solar system.
Recently, NASA began developing Solar Probe plus, which will study our own sun. Through a series of seven gravitational assists with the planet Venus, the probe will reach the extraordinary speed of 125 miles per second. This is a full seven times faster than Voyager 1, which would allow it to make a trip to another star (if that were its objective) in 6,450 years. While this will still be an incredible accomplishment, no currently existing propulsion technology has the capability to fly to another solar system on timescales comparable with a human lifetime.
Project Icarus is a five year theoretical design study of spacecraft that would reach another star within this all-important time restraint. The project was launched in 2009. One of the Terms of Reference of Project Icarus is:
"The spacecraft must reach its stellar destination within as fast a time as possible, not exceeding a century and ideally much sooner."
Clearly, current technology still has some way to come in order to accomplish this goal, based on the figures discussed so far. In fact, based on the distance to the closest star, it appears that we need to accelerate to approximately 5 percent the speed of light, and this is the figure we will focus on for the remainder of the article.
One could be forgiven for just assuming that if we continue to build bigger and bigger chemical rockets, that eventually we'll build one big enough that it could reach 5 percent the speed of light. Interestingly, the laws of physics tell us that this is, in fact, impossible.
The pioneering rocket scientist, Konstantin Tsiolkovsky, developed an equation that predicts how much rocket fuel one would need to reach a given top speed. For chemical rockets, the type that are used today, this equation predicts that to reach 5 percent the speed of light, one would need more chemical rocket fuel than there is matter in the known universe!
Given this sobering result, it's fairly clear that we need to start looking for alternative forms of propulsion if we're ever going to reach the stars on the timescale of a human lifespan.
Two popular technologies that may be able to accomplish this have been explored in some detail.

The change in velocity needed to perform a maneuver such as this can be calculated using the vis-viva equation which is derived from Newton’s law of universal gravitation and kepler’s third law. The Hohmann transfer has a variety of applications, especially in regards to travelling to other celestial bodies; however, it is more nuanced as you must include phase angles and other complex concepts.
Baker, Robert and Maud Makemson. An Introduction to Astrodynamics 2nd ed., New York, Academic Press, 1967.
Braeunig, Robert. Rocket and Space Technology. http://www.braeunig.us/space/index_top.htm
Lodgson, Tom. Orbital Mechanics: Theory and Applications. New York, Wiley, 1997.
Elert, Glenn. The Physics Hypertextbook. https://physics.info/

Lox/Kerosene propellant rocket stage. Loaded/empty mass 952,500/68,000 kg. Thrust 15,532.70 kN. Vacuum specific impulse 304 seconds. Boeing Low-Cost Saturn Derivative Study, 1967 (trade study of 260 inch first stages for S-IVB, all delivering 86,000 lb pyld to LEO): S-IC Technology Liquid Booster: 260 inch liquid booster with 2 x F-1 engines, recoverable/reusable
No Engines: 2.
Status: Study 1967.
More... - Chronology...
Gross mass: 952,500 kg (2,099,900 lb).
Unfuelled mass: 68,000 kg (149,000 lb).
Height: 36.36 m (119.29 ft).
Diameter: 6.61 m (21.68 ft).
Span: 10.58 m (34.71 ft).
Thrust: 15,532.70 kN (3,491,890 lbf).
Specific impulse: 304 s.
Specific impulse sea level: 265 s.
Burn time: 167 s.
F-1 Rocketdyne Lox/Kerosene rocket engine. 7740.5 kN. Isp=304s. Largest liquid rocket engine ever developed and flown. Severe combustion stability problems were solved during development and it never failed in flight. First flight 1967. More...
Associated Launch Vehicles
Saturn S-IC-TLB American orbital launch vehicle. Boeing Low-Cost Saturn Derivative Study, 1967 (trade study of 260 inch first stages for S-IVB, all delivering 86,000 lb pyld to LEO): S-IC Technology Liquid Booster: 260 inch liquid booster with 2 x F-1 engines, recoverable/reusable More...
Lox/Kerosene Liquid oxygen was the earliest, cheapest, safest, and eventually the preferred oxidiser for large space launchers. Its main drawback is that it is moderately cryogenic, and therefore not suitable for military uses where storage of the fuelled missile and quick launch are required. In January 1953 Rocketdyne commenced the REAP program to develop a number of improvements to the engines being developed for the Navaho and Atlas missiles. Among these was development of a special grade of kerosene suitable for rocket engines.

$\endgroup$– JoshuaJan 6, 2019 at 1:30
3$\begingroup$ @Joshua, KSP also has much higher delta-V budgets than NASA. $\endgroup$– MarkJan 6, 2019 at 2:09
$\begingroup$ @Mark: Will 15 m/s do? That's about what I'd use. $\endgroup$– JoshuaJan 6, 2019 at 3:51
1$\begingroup$ @Joshua Also, docking in KSP is often done at much higher speeds. When the space shuttle docked with the ISS, it would approach at 0.03 m/s (nasa.gov/pdf/593865main_AP_ST_Phys_ShuttleODS.pdf), while, in my KSP playing, I usually approach the rendez-vous at around 20+m/s, depending on the situation, and even make the final approach for docking 50x faster than the shuttle. $\endgroup$ Jan 6, 2019 at 6:26
$\begingroup$ @Joshua, the total delta-V expenditure during rendezvous for Gemini 6A was about 80 m/s, including the plane change, phase change, and establishing an intersecting orbit. The portion corresponding to your "15 m/s" is probably the "braking maneuver", at 19.8 m/s, although it might be the "terminal phase" maneuvers, at 38 m/s. (Terminal-phase maneuvers were initiated at a distance of 320 km.) $\endgroup$– MarkJan 6, 2019 at 7:16
First, the Gemini IV maneuver was station-keeping, not rendezvous. Since the target was the just-separated upper stage, the two spacecraft were already rendezvoused, and point-and-burn would have worked if they'd done it properly.
According to the Gemini IV mission report, the main causes of station-keeping failure were a mix of procedural mistakes and inadequately-aggressive maneuvering:
Review of these figures shows that the velocity increments applied through 00:09:21 g.e.t. succeeded in reducing the separation rate, but left a residual rate of 1.5 ft/sec away from the launch vehicle.

The T/W and specific thrust could be even better than the one owned by LANL. The catch is that the cycle doesn’t close. The drive gas pressure for the cagejet stage must be higher than the pressure it produces. Enter the turborocket. By using a fuel rich rocket to drive the turbine and burning the excess fuel in an afterburner with the compressed air, a very high T/W can be achieved with a very simple system. The drive rocket should use about 40% oxygen and 60% kerosene.
The cagejet is capable of using multiple staged rotors in the same manner as a multi spool turbofan except that each stage has its’ own bearing and doesn’t have to share shaft space with other spools. The cagejet compressor can attain higher stage compression ratios than axial flow because the adverse pressure gradient on the blades is countered by the centrifugal force. The separate stages can run at different flow rates to match airflow conditions in ways that even sophisticated turbofans can’t do. Each stage is capable of a compression ratio of 2, so three stages can get a compression ratio of 8. A fourth stage can be driven by the rocket compressing 25% of the air to a final compression ratio of 16 for a fraction of the total flow.
The 16 atm air compressed by the fourth stage burns stoichiometric with the fuel rich turbine exhaust in an afterburner and is used as drive gas to compress the total air flow in stage three to 8 atm. The exhaust from stage three can be used as is to drive the first two stages and give decent thrust with an Isp in the 1,500 range. Or the 8 atm afterburner can add kerosene for a stoichiometric burn. Then 25% of the flow could drive the first two stages and have an exit Isp in the 50 range including the air. With the mixture ratio of 16, net fuel Isp for 25% of the flow is 800. The other 75% of the flow is used as an 8 atm fuel/LOX/air rocket. Net Isp for this portion would be in the 1,200 range considering the LOX mass. Engine total net Isp at high thrust would be about 1,100.
T/W for this system would be around 25 if the numbers work.

The efficiency of the T55 is 0.5, which, when you do the math, comes down to 2.2 kilograms (over 4 pounds or 3 liters) of aviation kerosene every second. My fuel tank holds 11 gallons. It’d run out in under 14 seconds. That’s the price of power.
Abandon all Logic ye who Enter Here
The world’s largest diesel engine (the Wärtsilä-Sulzer RTA98-C) probably wouldn’t fit in my house, even if you hollowed out all those troublesome floors and walls. It looks like this:
It’s built to power the world’s largest oil tankers. The individual cylinders are so big that, when the engineers are maintaining them, they just climb inside and sit down at the bottom. To even consider attaching that monster to my car is to descend into madness. Luckily, I’m already down here. The RTA98-C produces a horrifying 46,000+ horsepower, and 7,600,000 newton-meters of torque at 102 RPM. That torque is ridiculous. Here’s how ridiculous: if you took the engines on a 747 and attached them so that they pointed up and down (respectively), to make the plane spin like a pinwheel, the plane would be experiencing the same torque that the RTA98-C puts out. It produces so much torque that, if you attached an I-beam 150 meters long to the crankshaft, it could lift me even if I sat on the end. And even if you include the mass of the I-beam.
But you’re not hear to hear about bizarre man-lifting contraptions. You’re here for speed! We already know that the limitation isn’t going to be the force the engine can put out. It’s the motor’s low speed. The RTA98-C could propel my car at all of…2.87 mph (4.61 km/h). Even when I’m out of shape, I can walk faster than that. But I can’t do what the Wärtsilä would let me do: pull with almost ten times the force of a Shuttle solid rocket booster. I could haul the whole Titanic. Over land. On a trailer, I could probably haul a couple of office buildings.

But actually it takes 42 seconds. To come closer to the planned result, you do half your delta v before and the other half after the maneuver node. But your rocket becomes lighter as you burn and exhaust fuel, so the acceleration increases during the burn, and the distance covered during the first half of delta v is longer than during the second half. To compensate for that, you could use the buttons for how much delta v you burn before and after the maneuver node. It only allows 10% steps, which is probably too coarse. Or you could manually start burning a little sooner or later. Would you know which way to adjust? I don't think it's worth it, and go with 50%. It won't be exact anyways, and you'll have to make correction maneuvers, like it's done irl.
Time warp till 3-5 min before burn start. I'll time warp 13d in the tracking station. There I select my rocket which shows the maneuver node so I can RMB click it so I can see the time to the node. Double clicking Kerbin pacifies the node.
Turn the rocket to the blue maneuver marker, which is the direction you should burn. If it's not visible, a blue arrow will point to it.
Wait till burn start. Now the rocket should still point to the maneuver marker. But to simplify the game KSP switches to a rotating frame of reference fixed to the orbital rotation, I think, during time warp, which makes the rocket rotate slowly relative to the stars ... and the maneuver marker/direction. So watch out if it changed significantly and correct if necessary.
Burn. But wait, what's that line in the bar? And the numbers 4 and 2? Stage 4 doesn't have enough fuel for the burn, so mid burn you have to stage as seamless as possible.
Also, your burn may not be exactly accurate, so it's a good idea not to rely just on the instruments, but watch where your real trajectory is. Focus on Duna, zoom out, so you can at least see your whole planned trajectory (this way you should have the SOI of Duna in view, where your real trajectory will appear. Whoa, the planned trajectory is not there. In the 90 before the burn I quickly correct the maneuvers.

Five minutes into the flight, the engine was overheating and Richmond found the aircraft to be overly sensitive in pitch. Lacking the canopy, he was buffeted by the 200-mph slipstream. Then a broken radiator fitting scaled him with steam from the radiator. He hastily made for landing without a good idea of the aircraft's ideal landing speeds only to discover what the stall speed was when he was still 30 feet above the ground.
The T.5 pancaked into the ground, driving the landing gears through the wing. The tail broke off and Richmond survived with significant burns. No further attempts were made as one month later the Battle of Britain began. The T.5 had a total of six minutes of flying time.
Source: Aviation History, July 2010. "Built for Speed" by Stephan Wilkinson, p18-19.
10 May 2010
No launch vehicle is so closely identified with the satellite communications industry than the Delta rocket. A direct descendant of the 1950s Thor intermediate-range ballistic missile and initially known as the Thor-Delta, through the 1960s the Delta rocket had firmly established its reputation as a reliable medium-lift launch vehicle that had not only orbited the first telecommunications satellites (Echo, Syncom, Telstar) but also the first weather satellites (TIROS) as well as a wide range of scientific probes (the Explorer series and the OSO solar-observation satellites). The vehicle was progressively modified for increases in capacity and performance but the real revolution launched aboard the Delta rocket would come in 1972.
The early Delta rockets used the Rocketdyne MB-3 liquid fuel engine for the first stage, the MB-3 being derived from the MA-3 engine that Rocketdyne built for the Atlas ICBM. Rocketdyne began work on a successor engine called the H-1 that would be used on the Saturn IB rocket for NASA. Having progressively tweaked the design to a nearly 50% increase in thrust, Rocketdyne built 322 H-1 engines for NASA and it was a relatively easy task to adopt the proven H-1 to be the new Delta first stage engine as the RS-27.
At the same time the Delta got a new, more powerful engine, in June 1972 the Federal Communications Commission decided to allow private companies to compete for domestic satellite communications. Prior to this, only international communications had any competition.

T/W 25 and Isp of 1,000 is the lower end of acceptable performance for a HTHL vehicle. The engine is useless for VTVL, throttling problems and flight profile restrictions. One thing that helps this engine concept is that the cagejet layout naturally lends itself to very thin cages of large diameter. By turning this system 90 degrees from the normal jet engine placement, it is possible to mount these engines inside the wings or tail surfaces. The vertical tail in particular is a desirable location.
The uses I see for this are for vehicles that need medium duration climb/cruise with serious acceleration during portions of the flight. The Rocketplane proposals would benefit from a lighter engine with more thrust. The White Knight series could use something like this to supplement normal turbofans for accelerating off the runway and initial climb. Turn them off during the cruise then light them again for a pop up from 10,000 feet to 60,000 feet for a seriously enhanced release altitude and attitude. The Air Launch group could enhance an airliner for their carrier.
Perhaps the most useful thing for it would be for testing spacecraft airframes. It would be nice to be able to wring out the airframes early on and have back up thrust during early rocket airframe trials.
This is the same concept I discussed at Space Access 2004 with a couple more rotors. It is a lot less complicated than it sounds. You can get an idea of it by driving a squirrel cage fan with shop air and feeling the wind it throws out the rest of the perimeter. Use a face shield.
The sketch is a simple two rotor system. A four rotor system is actually smaller because the burns are at 16 atm and 8 atm instead of 4 and 2 in the above. The boost rotor only has to handle 1/8 of the volume/second of the main rotor. It can be seen that the individual cages can operate at optimum rpm for the changing conditions of a flight without being slaved to the other stages. It is a simpler method of flow matching than the variable stators on modern turbojets and fans.

|Country of origin||Soviet Union|
|Designer||Kuznetsov Design Bureau|
|Application||1st stage multi-engine|
|Successor||AJ26-58, AJ26-59, AJ26-62|
|Propellant||LOX / RP-1 (rocket-grade kerosene)|
|Thrust (vac.)||1,753.8 kN (394,300 lbf)|
|Thrust (SL)||1,505 kN (338,000 lbf)|
|Chamber pressure||145 bar (14,500 kPa)|
|Isp (vac.)||331 seconds (3.25 km/s)|
|Isp (SL)||297 seconds (2.91 km/s)|
|Length||3.7 m (12 ft)|
|Diameter||2 m (6 ft 7 in)|
|Dry weight||1,235 kg (2,723 lb)|
The NK-33 and NK-43 are rocket engines designed and built in the late 1960s and early 1970s by the Kuznetsov Design Bureau. The NK designation is presumably derived from the name of the chief designer, Nikolay Kuznetsov. They were intended for the ill-fated Soviet N-1 rocket moon shot. The NK-33 engine is among the highest thrust-to-weight ratio of any Earth-launchable rocket engine, second only to the SpaceX Merlin 1D engine, while achieving a very high specific impulse. NK-33 was by many measures the highest performance LOX/kerosene rocket engine ever created.
The NK-43 is similar to the NK-33, but is designed for an upper stage, not a first stage. It has a longer nozzle, optimized for operation at altitude, where ambient air pressure is low, or perhaps zero. This gives it a higher thrust and specific impulse, but makes it longer and heavier.
Modified versions of these engines by Aerojet are known as the AJ26-58, AJ26-59 and AJ26-62.
NK-33 and NK-43 are derived from the earlier NK-15 and NK-15V engines, respectively.
The engines are high-pressure, regeneratively cooled staged combustion cycle bipropellant rocket engines, and use oxygen-rich preburners to drive the turbopumps. The turbopumps require subcooled liquid oxygen (LOX) to cool the bearings.

These ships were designed for four major propulsive events: first, to accelerate from Earth orbit velocity to Mars transfer orbit velocity; second, to decelerate from Mars transfer orbit velocity to Mars orbit velocity; third, to accelerate from Mars orbit velocity to Earth transfer orbit velocity; and fourth, to decelerate from Earth transfer orbit velocity to Earth orbit velocity. The total amount of change in velocity required for these four maneuvers is 11.48 kilometers per second. The amount of propellant required to do this using the low-performance rocket motors of the day required that the spaceships be over 98 percent propellant when departing from Earth orbit. This meant the empty weight of the spaceship, plus the weight of the crew and all their supplies and equipment had to be less than 2 percent of the departure mass. In all of human history, no one has ever come close to building any type of vehicle with a propellant fraction that high and with an empty weight fraction that low. Even if it were possible to build such a vehicle with existing technology, it would not be affordable on a commercial basis due to the extremely low payload fraction.
Now assume a non-rotating Skyhook in Earth orbit that has an upper endpoint velocity of just short of Earth escape velocity, an outpost space station at the Earth-Moon L2 Lagrange Point with a local source of propellant (either the Moon or a near-Earth asteroid that has been moved to L2), and a non-rotating Skyhook in Mars orbit that has an upper endpoint velocity of just short of Mars escape velocity as well as a local source of propellant (either Mars or one of the Martian moons), and look at what happens to the change in velocity requirement and the propellant fraction of an Earth-Mars spaceship.
Passengers and cargo bound for Mars, fly to the lower end of the Skyhook that is in Earth orbit using the fully reusable combination launch system described in previous posts. From there they transfer to the upper end of the Skyhook where they transfer to the spaceship that will take them to the outpost space station at the Earth-Moon L2 Lagrange Point where they board the Earth-Mars spaceship.
On the day of departure, the Earth-Mars spaceship leaves the halo orbit at L2 and heads toward Earth where it will perform a gravitational slingshot maneuver as it accelerates to Mars transfer orbit velocity. The change in velocity for these two maneuvers is approximately 1,500 meters per second.

In any case, NASA's next big thing is SLS (Space Launch System). There's a few other games in town (Elon Musk and his various efforts, as well as the Russian Angaras and the like look pretty interesting), but for now SLS looks to have the biggest payload of anything in the near future. It makes me very happy that NASA is taking concrete steps toward the development of a genuine heavy lift vehicle, and I sincerely hope that it all goes well, and we get people back on the moon and other places sooner rather than later.
However, there's a minor problem I have with SLS. Well, actually, it's with the payload proposals and upper stages I've seen; they're all chemically powered.
Why is this a problem? Well, chemical propellants (such as LOX/LH2, Kerolox, and similar) have a problem. While they have great thrust (very useful when you're lifting big things out of the gravity well), their efficiency is terrible. At the most, you're only getting about 460 seconds of specific impulse out of chemical propellants, probably a lot less if you're using ones that you can store in space long term. This means that if you want to get somewhere far away (like Mars), you're going to need a horrendous mass ratio (Mass ratio is the ratio of the initial mass of the spacecraft to the empty mass. So, if you have a spacecraft that weighs 200 tons, 100 of which is fuel that you use to get where you're going, you have a mass ratio of 2). Here's a quick little graph I mocked up in MATLAB to show you how your mass ratio varies with delta-V and specific impulse;
As you can see, it's pretty much an exponential relationship. (A quick examination of the Tsiolkovsky Equation will show why.) For reference, going one way to Mars orbit from LEO takes roughly 5,000 m/s of delta-V, according to my quick and dirty math.
So, what if we can stretch out the specific impulse axis of that graph a bit? Maybe down to around 800, 900 or so. That would be pretty cool, right? But is there anything we could do using today's technology that can get us there?
Well, if we really wanted to, we could build an Orion drive. Some people might be mildly annoyed by that, though.

Since we had been
working with very low flow rates for a while, we had a 6 hose from the tank to
the ¼ servo valve, then another short 6 hose to a 1/8 restriction right before
the engine. As the 6 hoses clear out,
the total pressure drop in the system goes down, letting the thrust go up. We have rediscovered this many times, and
always seem to forget about it when looking at test runs with a nicely repeating
thrust spike at the end.
For the second run, we exchanged the tank to valve hose for
a 10, and removed the 1/8 restriction.
The run was good again, with steady thrust increasing from 100lb to
150lb, but it still peaked another 50lb at the end, due to the second short
length of 6 hose. The run was a little
rougher, probably because the pressure drop from the 1/8 restriction was
providing some flow damping. The
measured Isp across these two runs was about 100s, which isnt too bad for the
low chamber pressure and notable transient behavior at the start. The thrust peaked (after clearing all the
restrictive plumbing) at 200 lbf at 173 psi tank pressure with a 1.25 diameter
For the third run, we put the 1/8 restriction back in to
smooth it out, but moved the servo valve directly to the side of the engine,
eliminating the second length of hose. We
loaded 4x the propellant, which would give a long enough run to get everything
fully stabilized and measured. On
firing, the thrust ramped up, but instead of lighting, it went cloudy, and
thrust dropped off rapidly.
My only current theory, assuming the catalyst isnt
deteriorating with each run, is that we only have enough active catalyst area
to make about 200 lbf of thrust, and the initial inrush of liquid before there
is chamber pressure is contributing to a cooling of the catalyst, which must
race against the building of chamber pressure.
We can test this by trying to crack the throttle to build some chamber
pressure before opening it all the way up.
A larger cavitating venturi would be the proper solution.

RocketdyneRS-68 (Rocket System 68) is a liquid hydrogen/ liquid oxygenengine developed starting in the 1990s with the goal of producing a simpler, less-costly heavy-lift rocket engine for the Delta IV rocket. The RS-68 produces a thrustof 663,000 lbf (2.9 MN) at sea level, while the RS-68A variant has produced 700,000 lbf (3.1 MN) in testing.cite press release | publisher=PRNewswire| date= 2008-09-25| title=United Launch Alliance First RS-68A Hot-Fire Engine Test a Success | url=http://www.prnewswire.com/cgi-bin/stories.pl?ACCT=104&STORY=/www/story/09-25-2008/0004892842&EDATE= |quote=Currently, the RS-68 engine can deliver more than 660,000 pounds of sea level thrust and the upgraded RS-68A will increase this to more than 700,000 pounds. The RS-68A also improves on the specific impulse, or fuel efficiency, of the RS-68. | accessdate=2008-09-30] The RS-68B variant is the proposed main engine for NASA's Project Constellationand has 80% fewer parts than the Space Shuttle Main Engine.
The RS-68 was developed at
RocketdynePropulsion and Power, located in Canoga Park, Los Angeles, California, to power the Delta IV Evolved Expendable Launch Vehicle ( EELV). The combustion chamber burns liquid hydrogen and liquid oxygen at 1486 lbf/in² (9.7 MPa) at 102% with a 1:6 engine mixture ratio.
At a maximum 102% thrust, the engine produces 758,000 lbf (3.3 MN) in a vacuum and 663,000 lbf (2.9 MN) at sea level. The engine's mass is 14,560 lb (6,600 kg) at 96 Inches(2.4384 m). With this thrust, the engine has a thrust-to-weight ratio of 51.2, and a
specific impulseof 410 s (4 kN·s/kg) in a vacuum and 365 s (3.58 kN·s/kg) at sea level.

- Check list
- VAB Engineer report
The delta-v display disagrees with Kerbal Engineer Redux (KER) mod. If I put the Reliant in stage 4 they agree.
Click on the Crew button and select your 1+ star pilot.
Click the launch button.
Low Kerbin orbit
You are on the launch pad now.
For most rockets I work out an ascend profile to orbit by trial and error. For this one the "manual" is:
- Activate SAS
- Switch the lower left infobox to 'Maneuver Mode'. This show's you apoapsis (Ap) and periapsis (Pe).
- stage (press space)
- at 10 m/s: turn/yaw 20 degrees right/east (from 90 to 70 degrees)
- while turning, click prograde hold
- Just before the boosters burn out fully turn on the Reliant.
- When the boosters burn out: stage
- At 70.5 km Ap turn off the Reliant
- Switch to the map (you could save now with F5)
- Click Ap with the right mouse button (RMB).
- 16 s before Ap: full thrust; when the time to Ap rises, cut thrust
- 8 s before Ap: full thrust; when the time to Ap rises, cut thrust
- 4 s before Ap: full thrust; when the time to Ap rises, cut thrust
- 2 s before Ap: full thrust; when time to Ap is about 20 s: cut thrust
- Now you have ~10 s to RMB click Pe.
- Use low thrust to keep time to Ap between 1-10 s.
- When Pe reaches 70.5 km you have a great orbit. :)
I have 3251 m/s left now. I used 2911 m/s to get to orbit!
Save as "1".
This rocket turned out a bit stubborn. It tended to drift south because the radial decouplers were flexing and the twisted boosters would create an aerodynamic force. This probably could be mitigated with (auto)struts (not used due to minimal tech). Correct for the drift early with WASD.
The rocket flys very shallow. This is efficient, unless it really flies too low, creating too much air drag, heat and explodes or passes Ap. It's ok if the fins explode :).

When the fuel of the big rocket is finished, we reach a velocity V, then the second stage is ignited, adding another V to the velocity, for a total of 2V.
Still faster! Now the rocket has mass 8M of which 4M is fuel of the first stage, while 4M
is the two-stage rocket of the preceding design. The first stage gives velocity V, to which the other two add 2V, for a total of 3V.
By now you can see the trend. If the mass of the final payload is M, then
| Total mass
|| Gives final velocity|
| 2M || V|
| 4M || 2V|
| 8M || 3V|
| 16M || 4V|
| 32M || 5V|
| 64M || 6V|
Each time the velocity increases by one notch, the mass doubles.
One cannot avoid this sort of thing by giving up staging--say, in the rocket of mass 8M, by firing all the 7M of fuel in one blast. That is because (as already noted), as the payload (+ remaining fuel) gain speed, less and less momentum is transferred, the jet first having to overcome the forward motion, Indeed, the correct derivation (which uses calculus) gives the equivalent of a huge number of little stages, fired one after the other. The same exponential result is still obtained.
This is one of the great problems of spaceflight, especially with the first stages which rise from the ground: even a small payload requires a huge rocket. Perhaps some day space explorers will be able to shave off some fuel weight by using air-breathing rockets ("scramjets") but only for the lowest 1/4 to 1/3 of the orbital velocity. Launching from a high-flying airplane--like Burt Rutan's "SpaceshipOne", or the "Pegasus rocket--also helps cut air resistance, another factor. But no other shortcuts are in sight. Once in orbit, of course, more efficient but more gradual ways of generating thrust can be enlisted, like ion propulsion.

VASIMR, the Variable Specific Impulse Magnetoplasma Rocket, combines features of the high-thrust/low-specific-impulse chemical rocket, and the low- thrust/high-specific-impulse nuclear rocket. VASIMR is a plasma rocket. Instead of a combustion chamber, it uses three staged, magnetic cells that first ionize hydrogen and turn it into a super hot plasma, then further energize it with electromagnetic waves to maximize thrust. Chang Díaz promises his rocket could attain a speed of 31 miles a second, and would reduce a one-way trip to Mars from three months to one. His team has made slow progress on the concept since the late 1980s. Last fall, his VX-200 rocket prototype’s first stage, powered by argon, reached a milestone: a successful, full-power firing in his Webster, Texas lab. Having spent about $25 million from several government sources so far, and with equipment, lab space, and personnel from NASA, Chang Díaz is coming closer to a flight test. NASA is considering testing the rocket on the International Space Station, perhaps as soon as 2011 or 2012, where it may contribute to maintaining the huge laboratory’s orbit.
After VASIMR, the next step up in velocity is a nuclear fusion rocket. Scientists haven’t yet re-created sustained, controlled fusion, the chemical process that powers stars and promises enormous benefits as a power source on Earth, but that hasn’t stopped them from getting a lot of money from governments to try. The International Thermonuclear Experimental Reactor, being built in southern France, is a joint project of the European Union, Japan, China, India, South Korea, Russia, and the United States. The reactor will cost at least $15 billion, is not expected to begin operation until 2018, and is the size of an office building, but scientists hope that once they achieve fusion on the ground, reactors can be downsized for space travel. Fusion gives off more energy and less radiation than fission, and could propel a ship at high speed. In one scenario, its exhaust would be contained by a string of superconducting magnets shaped like huge washers, each perhaps 15 feet in diameter. The string of magnets would reach back from the reactor for the length of several football fields.

Learn about this topic in these articles:
effective exhaust velocity in rockets
...for systems using electromagnetic acceleration. In engineering circles, notably in the United States, the effective exhaust velocity is widely expressed in units of seconds, which is referred to as specific impulse. Values in seconds are obtained by dividing the effective exhaust velocities by the constant factor 9.81 metres per second squared (32.2 feet per second squared).

If one considers the case of thrusting with maximum acceleration along arcs which are symmetrical around apogee , one can see that the contributions to semimajor axis and eccentricity variations given by the components are negligible (since they are multiplied by ). Therefore, a good suboptimal thrust direction can be obtained by imposing as the only non-zero component of the thrust acceleration. Under these assumptions, the variation of Keplerian parameters will be  It should be noted that the terms of the variation of and which depend on will also be very small due to the presence of integrated around . Figure 2 visualises the proposed pattern for thrusting arcs.
In order to obtain a fast propagation of the thrusting arcs, the analytical propagation of perturbed motion with finite perturbative elements in time (FPET) derived in [19, 21] will be used. In order to employ FPET, one has also to assume that the thrust acceleration is constant around each thrusting arc, which is reasonable given the low propellant consumption per arc. This assumption ensures that, if the engine thrust is constant, the resulting acceleration can also be considered constant over short thrusting arcs.
Motion propagation with FPET is based on a first-order analytical solution of perturbed Keplerian motion. In this formulation, the state is expressed in nonsingular equinoctial elements: Assuming constant thrust-acceleration in the radial-transverse reference frame: then one can obtain a first-order analytical expansion of the variation of equinoctial elements, parameterised in Longitude and with respect to a reference longitude : where where a are reference values at and are first-order terms as reported in [19, 21]. In ,it has also been shown that this analytical propagation scheme provides good accuracy along relatively long trajectory arcs.
As explained above, the only nonzero component of the acceleration will be and since the aim is obtaining a decrease of the orbit energy, it will also be in the negative direction. Therefore, the acceleration azimuth will be and the elevation (since, as already mentioned, the motion will be within the initial orbit plane). The variation of equinoctial elements after an apogee thrusting arc will be given by where is the apocentre longitude, and are the longitudes at the start and end of thrusting, respectively. is the semiamplitude of the apogee thrusting arc. Note that, given that , there is no variation on and and thus they are omitted.

Lox/LH2 propellant rocket stage. Loaded/empty mass 450,000/40,000 kg. Thrust 5,736.00 kN. Vacuum specific impulse 460 seconds.
Cost $ : 10.000 million. No Engines: 13.
Status: Study 1969.
More... - Chronology...
Gross mass: 450,000 kg (990,000 lb).
Unfuelled mass: 40,000 kg (88,000 lb).
Height: 40.00 m (131.00 ft).
Diameter: 7.65 m (25.09 ft).
Span: 10.00 m (32.00 ft).
Thrust: 5,736.00 kN (1,289,504 lbf).
Specific impulse: 460 s.
Specific impulse sea level: 409 s.
Burn time: 318 s.
MBB-ATC500 MBB lox/lh2 rocket engine. 441.3 kN. Study 1969. Isp=460s. Used on Beta launch vehicle. More...
Associated Launch Vehicles
Beta German SSTO VTOVL orbital launch vehicle. In 1969 rocket pioneer Dietrich Koelle was working at MBB (Messerschmitt-Bolkow-Blohm). There he sketched out a reusable VTOVL design called BETA using Bono's SASSTO as a starting point. The vehicle, taking European technology into account, was a bit heavier than Bono's design. But the thorough analysis showed even this design would be capable of delivering 2 tonnes of payload to orbit. More...
Lox/LH2 Liquid oxygen was the earliest, cheapest, safest, and eventually the preferred oxidiser for large space launchers. Its main drawback is that it is moderately cryogenic, and therefore not suitable for military uses where storage of the fuelled missile and quick launch are required. Liquid hydrogen was identified by all the leading rocket visionaries as the theoretically ideal rocket fuel. It had big drawbacks, however - it was highly cryogenic, and it had a very low density, making for large tanks. The United States mastered hydrogen technology for the highly classified Lockheed CL-400 Suntan reconnaissance aircraft in the mid-1950's.

The CE-7.5 is a cryogenic rocket engine developed by the Indian Space Research Organisation to power the upper stage of its GSLV Mk-2 launch vehicle. The engine was developed as a part of the Cryogenic Upper Stage Project (CUSP). It replaced the KVD-1 (RD-56) Russian cryogenic engine that powered the upper stage of GSLV Mk-1.
|Country of origin||India|
|First flight||15 April 2010 (failure) |
5 January 2014 (success)
|Designer||LPSC, Indian Space Research Organisation|
|Manufacturer||Hindustan Aeronautics Limited|
|Propellant||LOX / LH2|
|Thrust (vac.)||73.5 kN (16,500 lbf)|
|Chamber pressure||5.8 MPa (58 bar) / 7.5 MPa (75 bar)|
|Isp (vac.)||454 seconds (4.45 km/s)|
|Length||2.14 m (7.0 ft)|
|Diameter||1.56 m (5.1 ft)|
|Dry weight||435 kg|
|Upper stage of GSLV Mk.II|
The specifications and key characteristics of the engine are:
- Operating Cycle – Staged combustion
- Propellant Combination – LOX / LH2
- Maximum thrust (Vacuum) – 73.55 kN
- Operating Thrust Range (as demonstrated during GSLV Mk2 D5 flight) – 73.55 kN to 82 kN
- Engine Specific Impulse - 454 ± 3 seconds (4.452 ± 0.029 km/s)
- Engine Burn Duration (Nom) – 720 seconds
- Propellant Mass – 12800 kg
- Two independent regulators: thrust control and mixture ratio control
- Steering during thrust: provided by two gimballed steering engines
ISRO formally started the Cryogenic Upper Stage Project in 1994. The engine successfully completed the Flight Acceptance Hot Test in 2008, and was integrated with propellant tanks, third-stage structures and associated feed lines for the first launch. The first flight attempt took place in April 2010 during the GSLV Mk.II D3/GSAT-3 mission.

RS-2200 engine test
Rocketdyne lox/lh2 rocket engine. 2201 kN. Development cancelled 1999. Isp=455s. Linear Aerospike Engine developed for use on the Lockheed Reusable Launch Vehicle, the production follow-on to the X-33.
The RS-2200 Linear Aerospike Engine is being developed for use on the Lockheed Martin Skunk Works' Reusable Launch Vehicle, the production follow-on to the X-33. The Aerospike allows the smallest, lowest cost RLV to be developed because the engine fills the base, reducing base drag, and is integral to the vehicle, reducing installed weight when compared to a bell-shaped engine. The. Aerospike is the same as bell shaped rocket engines except that its nozzle is open to the atmosphere. The open plume compensates for decreasing atmospheric pressure as the vehicle ascends, keeping the engine's performance very high along the entire trajectory. This altitude compensating feature allows a simple, low-risk gas generator cycle to be. used. Over $500 million were invested in Aerospike engines up to the contract award date of the X-33, and full size linear engines have accumulated 73 tests and over 4,000 seconds of operation. Throttling, Percent Thrust: 20-109. Dimensions: Forward End: 6.4 m wide X 2.36 m long. Aft End: 2.36 m wide X 2.36 m long. Length: 4.32 m. Designed for booster applications. Gas generator, pump-fed.
Thrust (sl): 1,917.200 kN (431,004 lbf). Thrust (sl): 195,500 kgf. Chamber Pressure: 153.00 bar. Area Ratio: 173. Oxidizer to Fuel Ratio: 6.
Status: Development cancelled 1999.
More... - Chronology...
Height: 4.32 m (14.17 ft).
Diameter: 6.40 m (20.90 ft).
Thrust: 2,201.00 kN (494,804 lbf).
Specific impulse: 455 s.
Specific impulse sea level: 347 s.
First Launch: 1998.
Associated Launch Vehicles
X-33 American winged rocketplane.

Come in too shallow, and you skip off the atmosphere to become the next Voyager. “Very little room for error,” he says. “You get one crack at it.”
His solution: a nuclear thermal rocket. It would produce thrust the way chemical rockets do: by heating a propellant—in this case, hydrogen—and ejecting the expanded gas through a nozzle. Instead of heating hydrogen through combustion, however, the nuclear rocket vaporizes it through the controlled fission, or splitting of atomic nuclei, of uranium. Because nuclear fuel has a greater energy density, it lasts a lot longer than chemicals, so you can keep the engine running and continue to accelerate for half the trip. Then, with the speedometer clicking off about 15 miles per second—twice the speed reached by returning Apollo astronauts—you’d swing the ship around to point the other way and use the engine’s thrust to decelerate for the rest of the trip. Even when factoring in the weight of the reactor, a nuclear engine would cut the transit time in half.
In a program called NERVA (Nuclear Engine for Rocket Vehicle Applications), NASA built a nuclear thermal rocket in the 1960s. It delivered a specific impulse of 850 seconds—twice the efficiency of the best chemical rockets—and could have been tweaked to deliver up to 1,000 seconds. As NASA prepared to follow the moon missions with human voyages to the planets, the nuclear thermal rocket was a serious candidate to replace chemical engines in the Saturn V launch vehicle’s upper stages. Instead, despite more than 20 successful test firings in the Nevada desert, NERVA died in the mid-1970s.
Nuclear thermal rockets are limited by the heat tolerance of the uranium fuel and the engine’s structure, so engineers have experimented with new fuel elements and heat-resistant materials. At Marshall, Emrich has constructed a simulator that can test a nuclear rocket’s components by subjecting them to some of the conditions that fission would produce—the temperatures and pressures, though not the radioactivity. Because work on NASA’s Ares rocket, which will boost astronauts to the moon, has taken over the propulsion lab, Emrich is moving his simulator to another facility.
Not far from NASA’s Johnson Space Center in Houston, Franklin Chang Díaz, a former NASA astronaut and veteran of seven space shuttle flights, is developing an alternative to the nuclear thermal rocket.

Rocket propellant is mass that is stored, usually in some form of propellant tank, prior to being used as the propulsive mass that is ejected from a rocket engine in the form of a fluid jet to produce thrust. A fuel propellant is often burned with an oxidizer propellant to produce large volumes of very hot gas. These gases expand and push on a nozzle, which accelerates them until they rush out of the back of the rocket at extremely high speed, making thrust. Sometimes the propellant is not burned, but can be externally heated for more performance. For smaller attitude control thrusters, a compressed gas escapes the spacecraft through a propelling nozzle.
In ion propulsion, the propellant is made of electrically charged atoms, which are magnetically pushed out of the back of the spacecraft. Magnetically accelerated ion drives are not usually considered to be rockets however, but a similar class of thrusters use electrical heating and magnetic nozzles.
Rockets create thrust by expeling mass backwards in a high speed jet (see Newton's Third Law). Chemical rockets, the subject of this article, create thrust by reacting propellants into very hot gas, which then expands and accelerates within a nozzle out the back. The amount of the resulting forward force, known as thrust, that is produced is the mass flow rate of the propellants multiplied by their exhaust velocity (relative to the rocket), as specified by Newton's third law of motion. Thrust is therefore the equal and opposite reaction that moves the rocket, and not any interaction of the exhaust stream with air around the rocket (but see base bleed). Equivalently, one can think of a rocket being accelerated upwards by the pressure of the combusting gases in the combustion chamber and nozzle. This operational principle stands in contrast to the commonly held assumption that a rocket "pushes" against the air behind or below it. Rockets in fact perform better in space (where there is in theory nothing behind or beneath them to push against), because they do not need to overcome air resistance and atmospheric pressure on the outside of the nozzle.
The maximum velocity that a rocket can attain in the absence of any external forces is primarily a function of its mass ratio and its exhaust velocity. The relationship is described by the rocket equation: Vf = Veln(M0 / Mf). The mass ratio is just a way to express what proportion of the rocket is fuel when it starts accelerating.

There’s only one problem with this approach. Well, two problems, with the second, smaller one being that you can only accelerate like this for the first half of the trip. Then you have to turn the spacecraft around and burn the engine in the opposite direction to slow it down and make sure it doesn’t whip by its target at this unholy velocity you’ve accelerated it to. However, the big problem is that providing sustained low-grade acceleration for a prolonged period of time is impossible using conventional rocket engines. They’re built for rapid acceleration over small time scales and spend their propellant profligately in achieving this end, and even if you burn it as efficiently as possible a rocket engine would still run out of fuel long before it ever achieved any significant velocity. You can’t stuff a half-century’s worth of rocket fuel onto a spaceship, after all.
In order to do interstellar travel, then, we’re going to need a whole new type of spaceship engine. The thrust it provides does not need to be large, but it does need to be constant and very long-lasting. This gives us two options:
1) Use a very efficient fuel source that will somehow last the duration of the trip. This basically means nuclear rockets, a concept that’s been on the drawing board for years but which has never gone anywhere because people are a bit nervous about having nuclear reactors hurtling through the sky at several hundred miles per hour. There are high-efficiency thrust systems currently in development – see ion thrusters and the SMART-1 probe – that nevertheless require a powerful source of electricity to run. Nuclear reactors fit the bill and would last a half-century or more to boot, but they require several quantum leaps on the technology side of things (like the LFTR reactors I was talking about the other day) before they’re viable as a propulsion system.
2) Don’t carry your fuel with you. Get it from somewhere else, while you’re travelling. This is not as dumb as it sounds, with the two major ideas in this category being light sails and the Bussard ramjet.
Light sails are fairly easy to understand. All light exerts an infinitesimal force whenever it strikes an object, kind of like a really crappy version of the wind on Earth, so you just put a really big solar sail on your probe and focus an incredibly powerful laser on it from the Earth which “pushes” the spacecraft along.

If our engine is even close to a jet engine in reliability, has a flak shield to protect against a rapid unscheduled disassembly and we have more engines than the typical two of most airliners, then exceeding airline safety should be possible.
That will be especially important for point to point journeys on Earth. The advantage of getting somewhere in 30 mins by rocket instead of 15 hours by plane will be negatively affected if “but also, you might die” is on the ticket.
* Will be starting with a full-scale Ship doing short hops of a few hundred kilometers altitude and lateral distance. Those are fairly easy on the vehicle, as no heat shield is needed, we can have a large amount of reserve propellant and don’t need the high area ratio, deep space Raptor engines.
Next step will be doing orbital velocity Ship flights, which will need all of the above. Worth noting that BFS is capable of reaching orbit by itself with low payload, but having the BF Booster increases payload by more than an order of magnitude. Earth is the wrong planet for single stage to orbit. No problemo on Mars.
* Landing will not be a hoverslam, depending on what you mean by the “slam” part. Thrust to weight of 1.3 will feel quite gentle. The tanker will only feel the 0.3 part, as gravity cancels out the 1. Launch is also around 1.3 T/W, so it will look pretty much like a launch in reverse….
* The main tanks will be vented to vacuum, the outside of the ship is well insulated (primarily for reentry heating) and the nose of the ship will be pointed mostly towards the sun, so very little heat is expected to reach the header tanks. That said, the propellant can be cooled either with a small amount of evaporation. Down the road, we might add a cryocooler.
* 3 light-minutes at closest distance. So you could Snapchat, I suppose. If that’s a thing in the future.
* But, yes, it would make sense to strip the headers out and do a UDP-style feed with extreme compression and a CRC check to confirm the packet is good, then do a batch resend of the CRC-failed packets. Something like that. Earth to Mars is over 22 light-minutes at max distance.

Plasma is actually a distinct fourth state of matter, the other three being solid, liquid, and gas. The most glaring example of plasma is at the center of our solar system, though it’s quite abundant in nature and human applications, from lightning to plasma TVs.
Indeed, ion thrusters have been used for years on satellites and even some deep-space missions. In 2015, for instance, ion engines powered NASA’s Dawn probe into an orbit around the dwarf planet Ceres, which sits in an asteroid belt between the orbits of Mars and Jupiter.
Warp Speed Ahead
While ion thrusters lack the punch of a Falcon Heavy rocket in terms of the sudden acceleration required to pull free of Earth’s gravity, the much smaller engines excel in near-frictionless outer space, where they are at least 10 times more fuel-efficient.
That means a spacecraft with ion thrusters can constantly accelerate, eventually reaching speeds much greater than conventional chemical engines. For example, the retired Space Shuttle could hit speeds of 18,000 mph. A spacecraft powered by ion thrusters could theoretically streak through the cosmos at more than 200,000 mph, according to NASA.
Former astronaut Franklin Chang Diaz, who leads Ad Astra, has said he could do the cannonball run to Mars in less than 40 days. He first conceived of Ad Astra’s Variable Specific Impulse Magnetoplasma Rocket (VASIMR) back in the 1980s.
Much more recently, the company demonstrated that it could produce 100 kilowatts of power from the VASIMR engine during the course of 100 non-consecutive hours. The next step is to fire the engine to produce a plasma ball as hot as the sun and run it for 100 hours straight. Aerojet Rocketdyne is also reportedly ready for the next 100-hour test phase of its Hall thruster, another type of plasma-based engine. The best most ion thrusters can do today is about 5 kW.
Meanwhile, MSNW is exploring various prototypes related to fusion rockets, which would expel plasma produced from fusing a mixture of hydrogen and helium isotopes that had been heated by low-frequency radio waves. The process converts some of the mass of the atoms into energy. A lot of energy.
Out of Thin Air
Not to be outdone, the European Space Agency just tested another type of ion thruster that can literally live on thin air.

If you use a small nuclear reactor that delivers 1 gigawatt of electricity, you can get thrusts as high as 500 Newtons (100 pounds). What this means is that a 1 ton spacecraft would accelerate at 0.5 meters/sec/sec and reach Mars in about a week! Meanwhile, NASA plans to use some type of 200 kilowatt, ion engine design with solar panels to transport cargo and humans to Mars in the 2030s. Test runs will also be carried out with the Asteroid Redirect Mission ca 2021 also using a smaller 50 kilowatt solar-electric design.
So we are literally at the cusp of seeing a whole new technology for interplanetary travel being deployed. If you want to read more about the future of interplanetary travel, have a look at my book ‘Interplanetary Travel: An astronomer’s guide’, which covers destinations, exploration and propulsion technology, available at Amazon.com.
Check back here on Wednesday, January 11 for the next installment!
Ion engine schematic http://plasmalab.aero.upm.es/~plasmalab/information/Research/ElectricPropulsion.html

There is thus no basis whatsoever for believing in the feasibility of Chang Diaz’s fantasy power system.”
Chang Diaz, however, says in his paper: “Assuming advanced technologies [emphasis added] that reduce the total specific mass to less than 2 kg/kW, trip times of less than 60 days will be possible with 200 MW of electrical power. One-way trips to Mars lasting less than 39 days are even conceivable using 200 MW of power if technological advances allow the specific mass to be reduced to near or below 1 kg/kW.”
LEFT: Artist’s rendition of a lunar tug with 200 kW solar powered VASIMR®. RIGHT: Artist’s rendition of a human mission to Mars with 10 MW NEP-VASIMR®. Images Credit: Ad Astra Rocket Company
In other words, Chang Diaz is allowing for further developments that would enable such a reactor.
Zubrin, however, stated: “[T]he fact that the [Obama] administration is not making an effort to develop a space nuclear reactor of any kind, let alone the gigantic super-advanced one needed for the VASIMR hyper drive, demonstrates that the program is being conducted on false premises.”
The 2011 NASA research paper “Multi-MW Closed Cycle MHD Nuclear Space Power Via Nonequilibrium He/Xe Working Plasma” by Ron J. Litchford and Nobuhiro Harada, indicates that such developments are feasible in the near future.
Whether the VASIMR engine is viable or not, in 2015, NASA awarded Chang Diaz’s firm – Ad Astra Rocket Company™ – a three-year, $9 million contract. Up to now, the VASIMR engine has fired at fifty kilowatts for one minute – still a long way from Chang Diaz’s goal of 200 megawatts.
In its current form, the VASIMR engine uses argon for fuel. The first stage of the rocket heats the argon to plasma and injects it into the booster. There, a radio frequency excites the ions in a process called ion cyclotron resonance heating. As they pick up energy, they are spun into a stream of superheated plasma and accelerated out the back of the rocket.
Video courtesy of Ad Astra Rocket Company
Collin R. Skocik has been captivated by space flight since the maiden flight of space shuttle Columbia in April of 1981.

Say at some point the human race gets its act together and actually decides to send a spaceship to another star. We’ll say it’s unmanned because that makes everything so much easier; it’s just a robot probe that’s intended to sate our curiosity about what Alpha Centauri actually bloody looks like, and so we don’t need any life support systems or generation ship malarkey going on. We’ll also say that getting the probe out of the Earth’s gravity well – aka the trickiest part of spaceflight – is similarly not an issue because we’re building it in orbit or something. This leaves us just one, fairly major problem: how do we make it go?
“But Hentzau,” I hear you say, “the laws of Newtonian mechanics say that making it go should be the easiest thing of all! After all, if something is floating in space and you give it a push it’ll just keep going forever and ever, according to Newton’s third law. Surely the hard part isn’t making it go, but in getting it to stop?” Or at least I think I hear you say that; I’m pretty sleep deprived right now so it could also be the crazed squawking of an itinerant bluebird1. Anyway, assuming a really simple case where the probe is stationary with respect to its destination you could indeed just give it a little nudge and it would start wending its merry way to Alpha Centauri without braking for anybody. The only problem is it would take a while. In fact it would take so long that I’m not entirely sure it would get there before Alpha Centauri went nova, and we don’t even have that kind of time. The human race has a notoriously short attention span and we’re going to want results soon – within the next couple of centuries, say. This means that we are going to have to mount some form of interstellar propulsion engine on our probe, and interstellar propulsion engines are, via this rather convoluted introduction, the thing I wanted to talk about today.
Making something go in interstellar space is a completely different matter to getting out of a gravity well or travelling between planets. For the former the key is a shitload of thrust delivered over a very short period of time so that you don’t fall back to Earth, which is why rocket engines are currently the best we’ve got for that.

Rocket Escape velocities
First cosmic velocity = 7.82Km/sec
Second cosmic velocity = 11.2 Km/sec
Third cosmic velocity = 16.6 Km/sec
Calculating escape velocity:
Kinetic energy of the rocket needs to be at least equal to the potential energy of the planet at a radius r from its center
0.5m.v2 = G.M.m /r
Rearranging for the value of the escape velocity vesc
v = ( 2G.M /r )0.5
Where m is the mass of the projectile or rocket.
v is the escape velocity m/s
M is the mass of the planetary body we are trying to escape. In our case Earth.
G is the gravitational constant.6.67.10-11 Nm2/Kg2
r is the radius or distance between the center of mass of the rocket from the center of mass of planet mass M.
Note: ( )0. 5 Is equivalent to square root
Now how much pressure do we need to reach the 3rd cosmic velocity....
Where's my calculator..
The maximum velocity currently known is the reference velocity of light in a vacuum at 300000 Km/sec.
This velocity decreases when light passes through higher density materials:
Speed of light through water 225000kms/sec
through glass 200000kms/sec
through diamond 125000kms/sec
This means that we would need to improve our current rocket performance by18.103..to approach the speed of light.
A pretty humbling thought.
There is another slight problem in that Einstein's E=mc2 means that as the space craft approaches the speed of light its mass would increase. Since for this formula to be true the speed of light cannot be exceeded no matter how much energy you put into the system.
Second problem is that the majority of images we see of stars in the night sky have taken light years to arrive. A sought of cosmic archive if you like.
So some of the stars we see might not exist any longer!
The further we can look into the depths of space the further we actually travel back in light time towards the origin of the universe. The so called 'Big Bang'.
esa Global gravity variation research satellite
To be honest we are currently unable to explain what makes gravity ?
Pretty basic really. But inspite of all the formula it is still a scientific blind spot.

And as most engines are made of metal or carbon, hot oxidizer-rich exhaust is extremely corrosive, where fuel-rich exhaust is less so. American engines have all been fuel-rich. Some Soviet engines have been oxidizer-rich.
Additionally, there is a difference between mixture ratios for optimum Isp and optimum thrust. During launch, shortly after takeoff, high thrust is at a premium. This can be achieved at some temporary reduction of Isp by increasing the oxidiser ratio initially, and then transitioning to more fuel-rich mixtures. Since engine size is typically scaled for takeoff thrust this permits reduction of the weight of rocket engine, pipes and pumps and the extra propellant use can be more than compensated by increases of acceleration towards the end of the burn by having a reduced dry mass.
Although liquid hydrogen gives a high Isp, its low density is a significant disadvantage: hydrogen occupies about 7x more volume per kilogram than dense fuels such as kerosene. This not only penalises the tankage, but also the pipes and fuel pumps leading from the tank, which need to be 7x bigger and heavier. (The oxidiser side of the engine and tankage is of course unaffected.) This makes the vehicle's dry mass much higher, so the use of liquid hydrogen is not such a big win as might be expected. Indeed, some dense hydrocarbon/LOX propellant combinations have higher performance when the dry mass penalties are included.
Due to lower Isp, dense propellant launch vehicles have a higher takeoff mass, but this does not mean a proportionately high cost; on the contrary, the vehicle may well end up cheaper. Liquid hydrogen is quite an expensive fuel to produce and store, and causes many practical difficulties with design and manufacture of the vehicle.
Because of the higher overall weight, a dense-fuelled launch vehicle necessarily requires higher takeoff thrust, but it carries this thrust capability all the way to orbit. This, in combination with the better thrust/weight ratios, means that dense-fuelled vehicles reach orbit earlier, thereby minimizing losses due to gravity drag. Thus, the effective delta-v requirement for these vehicles are reduced.
However, liquid hydrogen does give clear advantages when the overall mass needs to be minimised; for example the Saturn V vehicle used it on the upper stages; this reduced weight meant that the dense-fuelled first stage could be made significantly smaller, saving quite a lot of money.

An unmanned craft could tolerate very large accelerations, perhaps 100 g. A human-crewed Orion, however, must use some sort of damping system behind the pusher plate to smooth the instantaneous acceleration to a level that humans can comfortably withstand – typically about 2 to 4 g.
The high performance depends on the high exhaust velocity, in order to maximize the rocket's force for a given mass of propellant. The velocity of the plasma debris is proportional to the square root of the change in the temperature (Tc) of the nuclear fireball. Since fireballs routinely achieve ten million degrees Celsius or more in less than a millisecond, they create very high velocities. However, a practical design must also limit the destructive radius of the fireball. The diameter of the nuclear fireball is proportional to the square root of the bomb's explosive yield.
The shape of the bomb's reaction mass is critical to efficiency. The original project designed bombs with a reaction mass made of tungsten. The bomb's geometry and materials focused the X-rays and plasma from the core of nuclear explosive to hit the reaction mass. In effect each bomb would be a nuclear shaped charge.
A bomb with a cylinder of reaction mass expands into a flat, disk-shaped wave of plasma when it explodes. A bomb with a disk-shaped reaction mass expands into a far more efficient cigar-shaped wave of plasma debris. The cigar shape focuses much of the plasma to impinge onto the pusher-plate.
For example, a 10 kiloton of TNT equivalent atomic explosion will achieve a plasma debris velocity of about 100 km/s, and the destructive plasma fireball is only about 100 meters in diameter. A 1 megaton TNT explosion will have a plasma debris velocity of about 10,000 km/s but the diameter of the plasma fireball will be about 1000 m..
The maximum effective specific impulse, Isp, of an Orion nuclear pulse drive generally is equal to:
where C0 is the collimation factor (what fraction of the explosion plasma debris will actually hit the impulse absorber plate when a pulse unit explodes), Ve is the nuclear pulse unit plasma debris velocity, and gn is the standard acceleration of gravity (9.81 m/s2; this factor is not necessary if Isp is measured in N·s/kg or m/s).

A high bypass fan engine normally includes a large first fan surrounded by a separate duct, allowing a majority of the fan air to bypass the engine. The fan acts much like a propeller in a turboprop, without the problems of propeller slipstream and drag.
Example: The Boeing 747 has four high bypass fan engines.
Engine Spool-up Time
Figure: Engine acceleration times, from Davies, figure 4.11
[Davies, page 59.]
- In a propeller installation the constant speeding ability of the propeller keeps the engine turning at an r.p.m. which is a compromise between the approach and baulked landing power conditions and power is altered by varying the boost pressure. To increase power quickly the boost is increased, the propeller coarsens off and the demanded thrust is supplied quickly. 'Quickly' in this context means about 3 to 4 seconds because of the propeller momentary overspeed tendency which is not acceptable to a pilot with any sympathy for mechanical engineering devices.
- Efficiency in a jet engine is highest at high r.p.m. where the compressor is working closest to its optimum conditions of gas flow, etc. At low r.p.m. the operating cycle is generally inefficient. If a sudden demand is made for more thrust from an r.p.m. equivalent to a normal approach r.p.m. the engine will respond immediately and full thrust can be achieved in about 2 secs. From a lower r.p.m., however, a sudden demand for maximum thrust will tend to overfuel the engine and cause it to overheat or surge. To prevent this various limiters are contained in the fuel control unit and these serve to restrict the engine until it is at an r.p.m. at which it can respond to a rapid acceleration without distress. This critical r.p.m. is most noticeable when doing a slam acceleration from an idle thrust setting. Acceleration is initially very slow indeed but then changes to very quick as r.p.m. rises through this significant value. From idle thrust to substantially full thrust at a typical approach speed takes about 6 secs. on average. Some engines are better than others, but there is also a scatter between individual engines of the same type; so occasionally the full 8 secs. permitted by the requirements is needed.
There have been two contrary trends in jet engine design when it comes to spool-up time.

As you watch the third cartoon's animation, imagine that the cannon has been packed with still more gunpowder, sending the cannonball out a little faster. With this extra energy, the cannonball would miss Earth's surface at periapsis by a greater margin, right?
Right. By applying more energy at apoapsis, you have raised the periapsis altitude.
And of course, as seen in these cartoons, the opposite is true: if you decrease energy when you're at apoapsis, you'll lower the periapsis altitude. In the cartoon, that's less gunpowder, where the middle graphic shows periapsis low enough to impact the surface. In the next chapter you'll see how this key enables flight from one planet to another.
Now suppose you increase speed when you're at periapsis, by firing an onboard rocket. What would happen to the cannonball in the third cartoon?
Just as you suspect, it will cause the apoapsis altitude to increase. The cannonball would climb to a higher altitude and clear that annoying mountain at apoapsis.
And its opposite is true, too: decreasing energy at periapsis will lower the apoapsis altitude. Imagine the cannonball skimming through the tops of some trees as it flys through periapsis. This slowing effect would rob energy from the cannonball, and it could not continue to climb to quite as high an apoapsis altitude as before.
In practice, you can remove energy from a spacecraft's orbit at periapsis by firing the onboard rocket thrusters there and using up more propellant, or by intentionally and carefully dipping into the planet's atmosphere to use frictional drag. The latter is called aerobraking, a technique used at Venus and at Mars that conserves rocket propellant.
Orbiting a Real Planet
Isaac Newton's cannonball is really a pretty good analogy. It makes it clear that to get a spacecraft into orbit, you need to raise it up and accelerate it until it is going so fast that as it falls, it falls completely around the planet.
In practical terms, you don't generally want to be less than about 150 kilometers above surface of Earth. At that altitude, the atmosphere is so thin that it doesn't present much frictional drag to slow you down. You need your rocket to speed the spacecraft to the neighborhood of 30,000 km/hr (about 19,000 mph).

This vector is shown below in the Earth Mean Equator and Equinox of J2000 coordinate frame:
delta-V vector for TCM 3: 0.076190 m/sec 0.063028 m/sec 0.036165 m/sec delta-V magnitude: 0.10529 m/sec or 0.2355 mph delta-V direction (unit vector): 0.72364 0.59863 0.34349
TCM 3 was executed somewhat differently than either TCM 1 or 2. For TCM 1, the spacecraft was turned so that its spin axis was pointing along the delta-V direction and the thrusters were fired to produce a force in that direction only. This was an axial burn. This could not be done for TCM 2 due to constraints on the spacecraft attitude. Instead, the spacecraft fired thrusters to produce forces both along and nearly normal to its spin axis. These axial and lateral delta-V components were sized so that their sum equaled the overall desired delta-V vector for TCM 2.
As was the case for TCM 2, thermal and power constraints precluded doing TCM 3 as an axial burn by turning the spacecraft directly to the delta-V direction. So, TCM 3 was implemented as a combination of axial and lateral burns. But in this case, we deliberately added an extra component to the delta-V breakdown. The reason for this is that our analysis for the late contingency maneuver - TCM 5 - has shown that, if executed, this maneuver would require a velocity change exclusively in the lateral direction. During the last 24 hours before entry when TCM 5 would be executed, the spacecraft has turned to the orientation required for safe descent through the atmosphere and cannot be turned from that orientation to do the TCM. And the geometry is such that the delta-Vs needed to target back to the desired landing site are nearly perpendicular to the entry attitude direction, so that TCM 5 will have to be performed as one or more lateral burns. In the worst case, 5 sets of lateral burns would be performed consecutively, each imparting a 0.4 m/sec velocity change. So far, the spacecraft has only performed a lateral burn once as part of TCM 2 and the velocity change for that burn was only 0.1 m/sec.

There are a number of variants proposed that double to triple the thrust to weight of a turbojet. This is at the cost of extreme fuel consumption compared to any jet, though much better than a rocket. Dense fuel Isp ranges from 400-700 during subsonic cruise with hydrogen about 50% better. It turns out that the standard turborocket burns too much fuel for cruising, and weighs too much to carry through serious acceleration.
I spent a good bit of time on the problem before realizing that modifying existing components just wasn’t going to work. Designing an effective ABE for a spaceplanes’ cruise phase really needs a new approach. You need better T/W and Isp than a normal turborocket. Designing around known systems just doesn’t get the job done. What is needed is a turbojet that is really really light for the power it produces. Or an engine that mimics a turbojet/turbofan and is really really light.
Some way must be found to eliminate as much mass and as many components as possible while retaining the desired capability. This usually involves finding ways to have one component do two jobs, and do them better than the original. LANL holds the rights on one patented method for increasing the performance of a turbine engine. The turbine blades are hollow and used as a centrifugal compressor. The air keeps the turbine blades cool enough to allow much higher turbine inlet temperatures than normal. The heat absorbed by the air is used for power (regenerative cooling)as it is mixed with fuel in the burner before driving the turbine. With one part being both compressor and turbine, weight is less. By using higher turbine inlet temperatures, the engine is more fuel efficient and gets more thrust out of each unit of air (specific thrust). Higher specific thrust means that less inlet and duct work is required per unit of thrust. This is a win for everybody if they get it into use.
I stumbled across another way of getting similar results. Some varieties of the squirrel cage fan have the characteristics of a short blade centrifugal compressor, and a short blade radial inflow turbine. By using one wheel as a compressor for 75% of the cycle, and as a turbine for 25% of the cycle, the blades are regenerative cooled, and it is possible to burn stoichiometric in front of the turbine.

Per Proton’s typical mission profile, the second stage burns for 210 seconds, separating from the third stage at the 327-second mark into the flight at a speed of 4.45 Kilometers per second and 120 Kilometers in altitude. Debris of the second stage impact 1,985 Kilometers from the launch site.
|Type||Storable Propellant Stage|
|Propulsion||RD-0213 Engine + RD-0214 Vernier|
|RD-0213 Thrust Vac||583kN|
|Chamber Pressure||14.7 MPa|
|Chamber Pressure||5.3 MPa|
The third stage of the Proton-M Rocket is the final stage of the actual launch vehicle that delivers the orbital unit to a preliminary orbit or a suborbital trajectory, depending on the mission profile.
This stage also utilizes a conventional design and consumes Nitrogen Tetroxide and Unsymmetrical Dimethylhydrazine a propellants. An RD-0213 engine powers the vehicle providing 583 Kilonewtons of thrust.
This engine is a non-gimbaled version of the RD-0210 engine that is used on the second stage. For vehicle control and additional thrust, an RD-0214 Engine with four gimbaled nozzles is installed on the third stage.
The RD-0214 provides 31 Kilonewtons of thrust for a total 3rd stage thrust 62,600 Kilograms.
The third stage houses the vehicle’s Navigation, Guidance and Control System that operated the vehicle during all aspects of powered flight which is fully automated and does not require commands from ground stations. A triple redundant digital guidance system is used to control the vehicle.
The Control System has been upgraded several times since the Proton-M started operations in 2001. The guidance mode used is cosed-loop. A high-precision three-axis gyro stabilizer provides exact attitude data to the digital flight computer. The avionics system also provides flight termination in case of a major anomaly during ascent.

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The performance of a rocket depends almost entirely on the velocity with which the propellant is exhausted,” he notes. Thus, “the elementary laws of mechanics – in this case relativistic mechanics, but still the elementary laws of mechanics – inexorably impose a certain relation between the initial mass and the final mass of the rocket in the ideal case… It follows very simply from conservation of momentum and energy, the mass-energy relation, and nothing else.
For our vehicle we shall clearly want a propellant with a very high exhaust velocity. Putting all practical questions aside, I propose, in my first design, to use the ideal nuclear fusion propellant… I am going to burn hydrogen to helium with 100 percent efficiency; by means unspecified I shall throw the helium out the back with kinetic energy, as seen from the rocket, equivalent to the entire mass change. You can’t beat that, with fusion. One can easily work out the exhaust velocity; it is about 1/8 the velocity of light. The equation of Figure 13 tells us that to attain a speed 0.99c we need an initial mass which is a little over a billion times the final mass.
This is no place for timidity, so let us take the ultimate step and switch to the perfect matter-antimatter propellant…. The resulting energy leaves our rocket with an exhaust velocity of c or thereabouts. This makes the situation very much better. To get up to 99 percent the velocity of light only a ratio of 14 is needed between the initial mass and the final mass.
Well, this is preposterous, you are saying. That is exactly my point. It is preposterous. And remember, our conclusions are forced on us by the elementary laws of mechanics.
All this stuff about traveling around the universe in space suits – except for local exploration, which I have not discussed – belongs back where it came from, on the cereal box.
a one-way trip of thirty-seven years (the distance to Zeta 1 or 2 Reticuli) at 99.9 percent c would take only twenty months’ crew time; at 99.99 percent c it would take only six months’ crew time.

The engine is made from three main modules: the igniter, injector, and the main engine body, all of which were SLS 3D printed by Shapeways and ExOne in bronze steel and machined afterwards to get the exact fit. Post-processing is a problem due to hardness of sintered metal so some tools break, the internal coolant lines are still not possible to 3d print due the geometry complexity and metal powder residues.
The price is very low: 3d printed igniter costs some $60, the injector $80 and the rocket engine $260, for a total of just $400. Space exploration with extremely low budget!
The engine is still in development and not yet finished but it is a big step forward in open sourcing aerospace engineering!
- Fuel: GOX / Ethanol
- Fuel Mass Flow: 0.0545 kg/sec
- Oxidizer Mass Flow: 0.0545 kg/sec
- Total Mass Flow: 0.1093 kg/sec
- Design Mixture Ratio: 1:1
- Design Force: 50 lbf
- Design Chamber Pressure: 150 psia
- Design Temp: 2572 Kelvin
- Design Specific Impulse: 209 Isp
Here are some photos of it:
|Ignited engine. You can clearly see the mach diamonds.|
|3D printed engine and main lines / sensors|
Here you can see a live test fire and mach diamonds:
Project homepage with files and instructions:
For previous 3d printed rocket engine named "Tri-D" look at:
To boldly go where no one has gone before!

Second stage engine cutoff (SECO)||14 min. 46 sec.|
|10. HTV5 separation||14 min. 54 sec.|
*1 Quick estimation value prior to the detailed estimation
*2 When the combustion chamber pressure becomes 10% against the largest combustion pressure.
*3 The definition of SRB-A jettison is the separation of the rear brace.

I would recommend a ballistic parachute deployment at a 90 degree angle to the rocket so that the pilot chute will clear the fins.
The fact that the rocket is fat at the top and skinny in the main body makes it difficult to use a launch lug like a conventional model rocket would use. Some method of keeping the rocket stable during its first few moments of flight needs to be conceived.
A possible solution to the rocket parachute deployment would be to use a parachute with a long bridle line . The parachute shroud lines would not start unstowing until the bridle line is fully stretched out and the parachute and its deployment bag are behind the fins of the rocket. This should prevent an entanglement.
Engine burn time 7 seconds + /- 1 second
Powered flight distance 1600 feet + /- 200 feet
Rocket speed at
engine burnout 280 mph +/- 20 mph
Coast distance 1400 feet + /- 200 feet
Peak altitude 3000 feet + /- 250 feet
Maximum acceleration 5 Gs
( Earth plus rocket engine)
Loaded rocket weight 1750 lb.
Minimum liftoff thrust 5500 lb.
Estimated flight time of
rocket before parachute
deployment 20 seconds
Estimated total time in air 3 minutes
Maximum pilot weight
with parachute 250 lb.
Stabilization method fin
Engine solid propellant
Ignition method electrical
Allow your imagination to soar. You walk towards the launch tower with flight suit and helmet on. Pause a minute and look at the strange rocket that will carry you into the blue sky. Climb into the capsule, strap yourself in, and say a quick prayer before the countdown begins. As the countdown ends, hear the sound of rolling thunder when the solid fuel ignites and you are pushed into your seat. The roar of the engine becomes deafening as the fiery rocket defies the earth and goes whizzing through the sky. The crushing G forces are replaced with a euphoric sense of weightlessness as the engine spits its last flame. The ship coasts for a few precious moments. You are relieved when you feel the jolt from the main parachute opening. It is time to eject the canopy, bail out, give a good "Airborne" hard arch and pull the ripcord. The beautiful red, white, and blue parachute opens. Spiral down and land in front of the clapping crowd.

At full throttle, depending on the efficiency of its three RVacs at sea level, Starship S20’s six Raptor V1.0 engines could produce ~1100 tons (~2.4M lbf) of thrust. By comparison, SpaceX’s workhorse Falcon 9 rocket produces around 760 tons (~1.7M lbf) of thrust at liftoff, meaning that Starship will likely become the most powerful single-core rocket the company has ever tested even if it never throttles above ~70%.
There’s a good chance that SpaceX will start Ship 20’s next round of tests by separately firing both sets of three Raptor Center and Vacuum engines or with a mixed three or four-engine test to follow the latest two-engine test. SpaceX could also take the most iterative approach and test three, four, and five engines at a time before the final six-engine test. Regardless, virtually all possible static fire tests Ship 20 is now configured to perform will be program ‘firsts’ of some kind. Stay tuned for updates on the first of those tests, preparations for which could begin as early as 10am CDT (15:00 UTC), November 1st.

The engine ignited, but the ignition did not sustain as the Fuel Booster Turbo Pump (FBTP) shut down after reaching a speed of about 34,500 rpm 480 milliseconds after ignition, due to the FBTP being starved of Liquid Hydrogen (LH2). On 27 March 2013 the engine was successfully tested under vacuum conditions. The engine performed as expected and was qualified to power the third stage of the GSLV Mk-2 rocket. On 5 January 2014 the cryogenic engine performed successfully and launched the GSAT-14 satellite in the GSLV-D5/GSAT-14 mission.
CE-7.5 is being used in the third stage of ISRO's GSLV Mk.II rocket.
- "Cryogenic engine test a big success, say ISRO officials". Indian Express. Retrieved 27 December 2013.
- "GSLV-D3". ISRO. Archived from the original on 16 April 2010. Retrieved 8 January 2014.
- "GSLV-D3 brochure" (PDF). ISRO. Archived from the original (PDF) on 7 February 2014.
- "GSLV MkIII, the next milestone". Frontline. 7 February 2014.
- "Flight Acceptance Hot Test Of Indigenous Cryogenic Engine Successful". ISRO. Retrieved 8 January 2014.
- "Indigenous Cryogenic Upper Stage". Archived from the original on 6 August 2014. Retrieved 27 September 2014.
- "GSLV-D5". ISRO. Archived from the original on 6 October 2014. Retrieved 27 September 2014.
- "GSLV-D5 launch video – CE-7.5 thrust was uprated by 9.5% to 82 kN and then brought back to nominal thrust of 73.55 kN". Doordarshan National TV.
- "How ISRO developed the indigenous cryogenic engine". The Economic Times.
- "Archived copy". Archived from the original on 4 January 2014. Retrieved 5 January 2014.CS1 maint: Archived copy as title (link)
- "Indigenous Cryogenic Upper Stage Successfully Flight Tested On-board GSLV-D5". ISRO.

Since the maximum Isp mixture ratio (r) for oxygen/gasoline is 2.5, we have:
Physical Properties of Selected Rocket Propellants
The thrust developed per pound of total propellant burned per second is known as specific impulse and is defined as
The chemical and physical properties of gaseous oxygen, methyl alcohol, and gasoline are given in Table II.
|Propellant||Gaseous Oxygen||Methyl Alcohol||Gasoline|
|Effect on metals||none||none||none|
|Density||0.083 lb/ft3||48 lb/ft3||44.5 lb/ft3|

Enter at a steep pitch and then bring it down to around a pitch of 20.
You should still have 300-400(ish) dV spare for raising the Pe back up to a stable orbit after aerobraking and then deorbiting for return to KSC (and should still have a fair bit of fuel to run the jet engines if needed).
Some of pics show it with the NEBULA decal mod to add the flags to the wings but I took those off so it could be posted as a pure stock craft.
Do you really want to downvote this?
Don't forget, people build craft at all skill levels, just 'cos something is 'newbish' doesn't mean it needs hatin'.
It will cost you 5 of your own points to downvote

Question – Why was Raptor thrust reduced from ~300 tons-force to ~170 tons-force?
One would think that for (full-flow staged combustion…) rocket engines bigger is usually better: better surface-to-volume ratio, less friction, less heat flow to handle at boundaries, etc., which, combined with the target wet mass of the rocket defines a distinct ‘optimum size’ sweet spot where the sum of engines reaches the best thrust-to-weight ratio.
Yet Raptor’s s/l thrust was reduced from last year’s ~300 tons-force to ~170 tons-force, which change appears to be too large of a reduction to be solely dictated by optimum single engine TWR considerations.
What were the main factors that led to this change?
ElonMuskElon Musk – We chickened out
The engine thrust dropped roughly in proportion to the vehicle mass reduction from the first IAC talk. In order to be able to land the BF Ship with an engine failure at the worst possible moment, you have to have multiple engines. The difficulty of deep throttling an engine increases in a non-linear way, so 2:1 is fairly easy, but a deep 5:1 is very hard. Granularity is also a big factor. If you just have two engines that do everything, the engine complexity is much higher and, if one fails, you’ve lost half your power. Btw, we modified the BFS design since IAC to add a third medium area ratio Raptor engine partly for that reason (lose only 1/3 thrust in engine out) and allow landings with higher payload mass for the Earth to Earth transport function.

Satellite launches, especially those lifting payloads to low-earth orbits, initially boost upwards, but then accelerate the payload on a path nearly parallel with the earth’s surface to reach the velocity needed to sustain the orbit. Low-thrust engines are typically used during the latter phase of the boosted trajectory to achieve the needed radial velocity. The maximum altitude of the payload is on the order of 200 to 500 km, depending on the orbital parameters required by the mission. Ballistic missiles, on the other hand, boost the warhead to high altitudes, allowing the payload to coast downrange to a maximum distance. A 10,000 km range intercontinental ballistic missile (ICBM) reaches a peak altitude of more than 1,000 km when on a minimum energy (i.e. maximum range) trajectory. Lifting a warhead to such heights requires high-thrust engines to avoid gravity losses while accelerating upward.
With the exception of the July 2006 firing of the Taepodong-2, which exploded too early in its liftoff trajectory to determine its mission, all of the other large rockets launched by North Korea were designed to maximize performance as a satellite launcher. In each case, the Taepodong-1 and Unha rockets flew on trajectories fully consistent with a satellite launch. Further, the Taepodong-1 used a low-thrust (Isayev 5D67) engine scavenged from an S-200 (NATO designated SA-5) air-defense missile on the second stage. Flight data displayed in the North Korean control room during the December 2012 Unha-3 launch, indicate that the second stage is a modified Scud-B missile with a larger diameter airframe to hold more fuel. The third stage is likely similar to that found on Iran’s Safir carrier rocket, which consists of vernier (i.e. steering) engines from either the Soviet R-27 (NATO designated SS-N-6) submarine-launched ballistic missile (SLBM), or another Soviet system, such as the ROTA, which was never fielded by the Soviet Union. The Unha’s use of long-burning, low-thrust upper stages is optimal for space missions, though if used as a ballistic missile, the low-thrust engines would suffer significant gravity losses during its upward trajectory, robbing the missile of roughly 800 km of range.

(f)Compute the thrusting time at perigee from (13). If , proceed to step g. Otherwise, break the iterative sequence and go to step 7.(g)Update and TOF: (h)Update the IBS spacecraft mass: (i)Set and .(j)Compute the current acceleration on the spacecraft: (k)Compute the time of flight spent coasting before apogee from to .(l)Compute the equinoctial parameters after the thrusting apogee arc as in (12).(m)Compute the thrusting time at apogee from (13). If , proceed to step n. Otherwise, break the iterative sequence and go to step 7.(n)Update and TOF: (o)Update the IBS spacecraft mass: (p)Set and .(7)Back-track the point at which and compute the corresponding equinoctial parameters and update accordingly.(8)Compute the mismatch between the actual final conditions and the target orbit:
Summarizing, the TPBVP has been reduced to an optimisation problem in the form:
This problem can be solved with a gradient-based optimisation algorithm like MATLAB’s fmincon. Note that the time of flight is specified a priori and, therefore, it might occur that this duration is too short as to obtain the change in the orbital parameters specified by the boundary constraints. In this case, the problem is infeasible and the optimisation is terminated after a maximum of 50 iterations if the constraints are not satisfied.
In the following, an example of transfer from an elliptical orbit with 300 km perigee altitude and eccentricity 0.031 (corresponding to the final orbit of a de-orbiting strategy) to a circular orbit of 1100 km altitude (corresponding to the orbit of the next debris in an hypothetical removal sequence). Parameters of the two orbits are reported in Table 2. Note that the total plane rotation in this case is 10 degrees. The specified time of flight is 70 days.
First, it is considered the case of a coplanar transfer, that is, will be computed. The optimisation problem was solved with fmincon in 6 iterations and less than 10 seconds, returning a minimum cost of 0.301 km/s, with 1001 revolutions.

the distance traveled during the deceleration phase alone--then find the fuel/payload ratio using the above formula, then square the resulting ratio to get the ratio for the whole trip. (Similarly, if you have the final velocity Delta-v at the end of the acceleration phase/beginning of the deceleration phase, then you can use the formula e^((c/v_e) * atanh(Delta-v/c)) I mentioned earlier, which gives the mass ratio as a function of Delta-v and the exhaust velocity v_e, then square the result to get the ratio for the whole trip including both acceleration and deceleration phases.)
If you want to know the rationale for this, start by noting that the formula I posted should work fine for the deceleration phase alone (decelerating at A meters/second² from an initial velocity V to rest requires exactly the same fuel ratio as accelerating at A meters/second² from an initial state of rest to a final velocity of V), so it can tell you how many kilograms of fuel+payload will be needed at the beginning of the deceleration for every kilogram of payload at the end of deceleration. Then the number of kilograms at the beginning of deceleration can be treated as the "payload" mass that needs to be brought up to speed during the first acceleration phase, so you'll have the same ratio as before for fuel+payload to final payload over the course of the acceleration. Thus if the mass at the beginning of deceleration is R times bigger than the final mass at the end of deceleration, and the mass at the beginning of acceleration is also R times bigger than the mass at the end of acceleration (which is also the mass at the beginning of deceleration), then this implies the mass at the beginning of acceleration is R*R = R² times bigger than the mass at the end of deceleration.
The Serenity and other ships (not sure if it's all of them) use what is essentially controlled fusion explosions at the tail end of the ship to propel the ship at greater than normal speeds. It is never made clear how fast the ships do move at top speeds, but it is still at sub-light speeds. But even so, they appear to achieve speeds much greater than what is capable of our technology today.

If anyone wants to play around with different distances (in case Serenity makes shorter hops and refuels) and different exhaust velocities, I'll put all the above equations together into one long one that can just be copy and pasted into the the online calculator to find the ratio of masses before and after using up the fuel, just substitute whatever distance you want for the four D's in the equation (measured in light hours), whatever acceleration you want for the four A's in the equation (where 5.5g = 0.00647, and 1g = 0.00011776), and whatever exhaust velocity you want for the single V in the equation (measured as a fraction of light speed):
e^((1/V) * atanh((A * sqrt((D)² + (2 * D/A)))/(sqrt(1 + (A * sqrt((D)² + (2 * D/A)))²))))
For example, plugging in D=55, A=0.00647, and V=0.089, the calculator gives the same mass ratio of 10079 found earlier.
On the other hand, I suppose you could imagine they use some exotic technology (perhaps the same technology they use to make artificial gravity) to accelerate the exhaust to speeds much faster than real-world fusion reactions, or you could ignore the material on the DVD saying it was fusion and imagine it was something like a matter/antimatter reaction, where the effective exhaust velocity can be close to light speed. For example, if we use the same equation but with an exhaust velocity of 0.9c rather than 0.089c, the ratio of initial mass with fuel to final mass without fuel would just be around 2.5.
Note: That last formula was assuming continuous steady acceleration all the way from departure point to destination, which would result in reaching the destination at a very high speed; a more natural assumption might be to accelerate for the first half of the trip, then decelerate at the same rate for the second half, so you'd come to rest at your destination (you can also imagine a constant-velocity coasting phase in between the end of acceleration and the beginning of deceleration, this wouldn't affect the amount of fuel used). In that case, all you have to do is set the D to be half the total distance crossed during both acceleration and deceleration--i.e.

But what about the real-world engineering of actually building such an engine—managing the plasma and its thermal properties, then successfully firing it for a long period of time? That has proven challenging, and it has led many to doubt Vasimr’s practicality.
…Speaking almost no English at the time, he immigrated to the United States from Costa Rica in 1969 to finish high school. Chang-Díaz then earned a doctoral degree in plasma physics from Massachusetts Institute of Technology. Later, as an astronaut, Chang-Díaz flew seven Space Shuttle missions, tying Jerry Ross’ record for most spaceflights by anyone, ever.
…The rocket engine starts with a neutral gas as a feedstock for plasma, in this case argon. The first stage of the rocket ionizes the argon and turns it into a relatively “cold” plasma. The engine then injects the plasma into the second stage, the “booster,” where it is subjected to a physics phenomenon known as ion cyclotron resonance heating. Essentially, the booster uses a radio frequency that excites the ions, swinging them back and forth.
As the ions resonate and gain more energy, they are spun up into a stream of superheated plasma. This stream then passes through a corkscrew-shaped nozzle and is accelerated out of the back of the rocket, producing a thrust….
The Sun powers both the production of plasma and the booster exciting the plasma, and the extent to which it does either can be shifted. When a spacecraft needs more thrust, more power can be put into making plasma. This process uses more propellant, but it provides the thrust needed to move out of a gravity well, such as Earth orbit. Later, when the vehicle is moving quickly, more power can be shifted to the booster, providing a higher specific impulse and greater fuel economy.
At the last deglaciation Earth’s largest biome, mammoth-steppe, vanished….Analyzes of fossils 14C dates and reconstruction of mammoth steppe climatic envelope indicated that changing climate wasn’t a reason for extinction of this ecosystem. We calculate, based on animal skeleton density in frozen soils of northern Siberia, that mammoth-steppe animal biomass and plant productivity, even in these coldest and driest of the planet’s grasslands were close to those of an African savanna.

His 1968 paper "Interstellar Transport" (Physics Today, October 1968, p. 41–45) retained the concept of large nuclear explosions but Dyson moved away from the use of fission bombs and considered the use of one megaton deuterium fusion explosions instead. His conclusions were simple: the debris velocity of fusion explosions was probably in the 3000–30,000 km/s range and the reflecting geometry of Orion's hemispherical pusher plate would reduce that range to 750–15,000 km/s .
To estimate the upper and lower limits of what could be done using contemporary technology (in 1968), Dyson considered two starship designs. The more conservative energy limited pusher plate design simply had to absorb all the thermal energy of each impinging explosion (4×1015 joules, half of which would be absorbed by the pusher plate) without melting. Dyson estimated that if the exposed surface consisted of copper with a thickness of 1 mm, then the diameter and mass of the hemispherical pusher plate would have to be 20 kilometers and 5 million metric tons, respectively. 100 seconds would be required to allow the copper to radiatively cool before the next explosion. It would then take on the order of 1000 years for the energy-limited heat sink Orion design to reach Alpha Centauri.
In order to improve on this performance while reducing size and cost, Dyson also considered an alternative momentum limited pusher plate design where an ablation coating of the exposed surface is substituted to get rid of the excess heat. The limitation is then set by the capacity of shock absorbers to transfer momentum from the impulsively accelerated pusher plate to the smoothly accelerated vehicle. Dyson calculated that the properties of available materials limited the velocity transferred by each explosion to ~30 meters per second independent of the size and nature of the explosion. If the vehicle is to be accelerated at 1 Earth gravity (9.81 m/s) with this velocity transfer, then the pulse rate is one explosion every three seconds.

The dimensions and performance of Dyson's vehicles are given in the table below
|Ship diameter (meters)||20,000 m||100 m|
|Mass of empty ship (metric tons)||10,000,000 t (incl.5,000,000 t copper hemisphere)||100,000 t (incl.50,000 t structure+payload)|
|+Number of bombs = total bomb mass (each 1MT bomb weighs 1 metric ton)||30,000,000||300,000|
|=Departure mass (metric tons)||40,000,000 t||400,000 t|
|Maximum velocity (kilometers per second)||1000 km/s (=0.33% of the speed of light)||10,000 km/s (=3.3% of the speed of light)|
|Mean acceleration (Earth gravities)||0.00003 g (accelerate for 100 years)||1 g (accelerate for 10 days)|
|Estimated cost||1 U.S. GNP (1968)||0.1 U.S. GNP|
Later studies indicate that the top cruise velocity that can theoretically be achieved by a thermonuclear Orion starship is about 8% to 10% of the speed of light (0.08-0.1c). An atomic (fission) Orion can achieve perhaps 3%-5% of the speed of light. A nuclear pulse drive starship powered by matter-antimatter pulse units would be theoretically capable of obtaining a velocity between 50% to 80% of the speed of light.
At 0.1c, Orion thermonuclear starships would require a flight time of at least 44 years to reach Alpha Centauri, not counting time needed to reach that speed (about 36 days at constant acceleration of 1g or 9.8 m/s2). At 0.1c, an Orion starship would require 100 years to travel 10 light years. The late astronomer Carl Sagan suggested that this would be an excellent use for current stockpiles of nuclear weapons.
A concept similar to Orion was designed by the British Interplanetary Society (B.I.S.) in the years 1973–1974. Project Daedalus was to be a robotic interstellar probe to Barnard's Star that would travel at 12% of the speed of light (0.12c).

Nuclear/LH2 propellant rocket stage. Loaded/empty mass 1,500,000/500,000 kg. Thrust 14,700.00 kN. Vacuum specific impulse 900 seconds. N1 nuclear upper stage study, 1963. Figures calculated based on given total stage thrust, specific impulse, engine mass.
No Engines: 20.
Status: Study 1963.
More... - Chronology...
Gross mass: 1,500,000 kg (3,300,000 lb).
Unfuelled mass: 500,000 kg (1,100,000 lb).
Height: 100.00 m (320.00 ft).
Diameter: 17.00 m (55.00 ft).
Span: 17.00 m (55.00 ft).
Thrust: 14,700.00 kN (3,304,600 lbf).
Specific impulse: 900 s.
Burn time: 590 s.
YaRD Type V Korolev nuclear/lh2 rocket engine. 392 kN. Study 1963. Design considered in N1 nuclear upper stage studies. Outgrowth of work done by Bondaryuk and Glushko on YaRD engines for nuclear ICBM's, but using liquid hydrogen as propellant. Isp=900s. More...
Associated Launch Vehicles
N1 Nuclear V Russian nuclear orbital launch vehicle. Second primary alternative considered for the 1963 nuclear N1 study. The immense liquid hydrogen tank of the second nuclear stage would have dwarfed the N1 first stage mounted below it in the shadows. The extremely poor thrust to weight ratio of the Type V engine design compared to that of the Type A remains unexplained. More...
Nuclear/LH2 Nuclear thermal engines use the heat of a nuclear reactor to heat a propellant. Although early Russian designs used ammonia or alcohol as propellant, the ideal working fluid for space applications is the liquid form of the lightest element, hydrogen. Nuclear engines would have twice the performance of conventional chemical rocket engines. Although successfully ground-tested in both Russia and America, they have never been flown due primarily to environmental and safety concerns. Liquid hydrogen was identified by all the leading rocket visionaries as the theoretically ideal rocket fuel. It had big drawbacks, however - it was highly cryogenic, and it had a very low density, making for large tanks.

(2) Examples of Propellants used in Different Rockets
- Saturn booster rocket of American space programme used a mixture of kerosene and liquid oxygen as the propellant in the initial stage whereas liquid oxygen and liquid hydrogen were used as propellant in high
- Russian rockets such as Proton used a liquid propellant consisting of kerosene and liquid
- The Indian satellites SLV-3 and ASLV used composite solid
- The rocket PLSV will use solid propellant in the first and third stages and liquid propellant in second and
fourth stages. The liquid propellant will consist of
and unsymmetrical dimethyl hydrazine (UDMH) and
and monomethyl hydrazine (MMH) respectively.
In our country, Indian Space Research Organization (ISRO) has been set up to launch and utilize two classes of satellites : remote sensing satellites and communication satellites. The Polar Satellite Launch Vehicle (PSLV) is a remote sensing satellite. India has succeeded in launching several space vehicles using various rocket propellants. India’s latest vehicle, PSLV–C4 took flight on 12th September, 2002 and it was named METSAT MISSION. It consists of four stage vehicle. The first stage is one of the largest solid propellant boosters in the world and carries about 138 tonnes of hydroxyl terminated polybutadiene (HTPB) based propellant.
The second stage uses indigenously built VIKAS engine and carries 40 tonnes of liquid propellant unsymmetrical dimethyl hydrazine (VDMH) as fuel and nitrogen tetroxide ( N 2O4 ) as oxidizer.
The third stage uses 7.6 tonne of HTPB based solid propellant.
The fourth and terminal stage of PSLV-C4 has a twin engine configuration using liquid propellant. Each engine uses 2.5 tonnes of monomethyl hydrazine as fuel and mixed oxides of nitrogen as oxidizer.
(3) Calculation of specific impulse of propellant
The function of rocket propellant is based on specific impulse which measures the kinetic energy producing ability of the propellant. The specific impulse (I s ) can be calculated from the following equation,
I s =
g = Ratio of specific heat at constant pressure to specific heat at constant volume.
Tc = Combustion chamber temperature. M=Average molecular mass of exhaust products.

52 sec.|
|9. DAICHI-2 separation||15 min. 47 sec.||15 min. 42 sec.|
*1 Quick estimation value prior to the detailed estimation
*2 The values are updated ones based on actual measurement data such as engine performance which are unique for the H-IIA F24 engines. Therefore, they are slightly different from the values in the Launch Plan.
*3 When the combustion chamber presser becomes 10% against the largest combustion pressure.
*4 The definition of SRB-A jettison is the separation of the rear brace.

This probably won’t make sense to anyone else, but I need to make a note reminding myself about something I thought about over lunch. (I always lose them otherwise, so maybe this will help me remember.) Read on if that kind of thing interests you.
Given: an ideal pseudoablative mass driver which uses pattern conversion to translate the trailing surface into a high-velocity sheet, imparting an acceleration to the remaining mass and exposing a self-similar surface. The mass driver is completely consumed by the process.
- Develop a set of equations relating the mass of the driver (md), the exit velocity of the propellant (vx), the surface area of the trailing surface (At) , the mass of the payload (mp), and the total velocity imparted to the payload (vp). Use the rocketry equation as a starting point, but use relativistic pseudovelocities in terms of multiples of c.
- Figure out how to wedge the instantaneous acceleration (a) in there somewhere.
- Re-develop the “ideal conversion” equation, which relates the md necessary to accelerate mp to a given pseudovelocity.
- What kind of exit particle would provide the best efficiency for the driver? A photon?
- Determine how close this engine comes to the “ideal conversion” developed in 2.
- Determine the time and md / mp required for an in-system transit at 1g with midpoint turnaround.
- Determine the time and md / mp required for a 4ly transit at 1g with midpoint turnaround.

The energy from that would be forced out the back of the ship in a so-called “z-pinch” using a “magnetic nozzle”, a component which the team are also developing. The engine’s potential top speed? Over 100,000km/h. That’s roughly the same speed at which the Earth orbits the Sun.

That's the right hand rule.
When you accelerate (anti)normal, your direction of motion (your velocity vector, which is speed and direction) changes a bit towards 'up' ('down'). This also changes your orbital plane, like holding out your hand flat and turning your arm along its axis. On the left/right side your orbit (or hand) will move up and down, while along the turn axis it will not. So to move your flyby trajectory, you must make an (anti)normal maneuver 90 degrees before your encounter, or half way between Kerbin and Duna:
Zoom back out, double click the Sun, look at the solar system from 'above', (optionally turn till Kerbin and Duna are on the same horizontal line), and click the trajectory half way between Kerbin and Duna.
Focus Duna again and zoom in.
In the control box select maneuver 2. Adjust (anti)normal delta v to bring the trajectory up/down towards Duna. I want an equatorial orbit, so I want the trajectory to be horizontal.
Keep adjusting pro/retrograde at maneuver 1 and (anti)normal at maneuver 2 and zooming in till your trajectory is at Duna. Put your Duna Pe at about 25 km. Fly by at the side that will result in a prograde Duna orbit. Duna rotates counterclockwise when looking at it from above.
You can avoid Ike encounters by changing prograde delta v and compensating with departure time.
BTW, another way to get your trajectory up/down to Duna is to make an (anti)normal maneuver at the ascending/descending node (AN/DN). The way I do it uses less delta v, but has higher approach velocity at the encounter. I don't worry about that, cause aerobraking takes care of that.
Departure from low Kerbin orbit
Next to the navball you see three times, e.g.
- Node in t-13d, 0h, 35m
- Burn Time: 42s
- Start burn in: 13d, 0h
1 says when the maneuver node is. 2 says how long you have to thrust with your rocket engine(s). 3 says when to begin burning.
Now KSP maneuver nodes assume you accelerate your rocket instantly by e.g. 1052.7 m/s.

“The problem is not so much the amount of energy; you have gobs and gobs of energy,” says Emrich. “The problem is power, which is how fast you get the energy out of the system. A hydrogen bomb releases a huge amount of energy instantly but melts everything in sight.”
By contrast, the superconducting magnets corral the power of all that energy and essentially squirt it out the end. “Magnetic fields don’t melt,” says Emrich.
In theory, the engine could unleash a specific impulse of a million seconds. It would need only 1/10th of that to propel a craft to Mars in two weeks. But Emrich notes that to make a fusion-powered spaceship light enough to reach Mars in two weeks, propulsion experts will need a breakthrough in materials science.
“Mars in 30 days?” he says. “That’s getting closer.”
If and when new materials make that possible, Mars may in fact be too close to Earth for a fusion rocket to truly show what it’s got under the hood. A trip to Jupiter, on the other hand, 366 million miles away at its closest approach, would give the crew of a fusion-powered spacecraft almost 183 million miles of acceleration to the journey’s midpoint. By then, a fusion engine delivering about 30,000 seconds of impulse would have gathered a speed of 50 miles per second—about 180,000 miles an hour. After decelerating for the next 90 days, it would slip into orbit around Jupiter; by then, the trip would have lasted 180 days, only six times as long as a one-way trip to Mars, despite covering 10.5 times the distance. True, while the astronauts are exploring Jupiter, Earth wanders farther away than it was at launch time; however, at these speeds, orbital separation between the planets becomes less of a problem.
“The space program began the day humans chose to walk out of their caves,” says Chang Díaz. “By exploring space we are doing nothing less than insuring our own survival.” Chang Díaz believes that humans will either become extinct on Earth or expand into space. If we pull off the latter, he says, our notion of Earth will change forever.
As for Cast Away’s Chuck Noland, he eventually concludes that it would be better to risk it all and die trying to escape his imprisonment than to waste away on the beach.

Lunar prospector. Harvester. Solar system class performance.
~6.5 - 10.9km/s ΔV
2,350kg fully fueled
500kg payload capability
28v, 12v, 5v, 3.3v
With 9 PECO engines churning out solar system conquering ΔV the MX-9 can deliver up to 500kg to the lunar surface from GTO
Designed for Frontier Class exploration capabilities, MX-9 will support robust lunar sample return operations. Like it’s MX-5 little brother, the MX-9 can also be outfitted with MX-1 or MX-2 staged systems that can deliver over 10kms ΔV and extend its reach to span the solar system, and beyond.
Available in orbiter, lander, deep space probe and sample return configurations.

III was to apply lessons learned from the CUS and its CE-7.5 engine and upscale the design to create a larger upper stage and a more powerful engine with a target thrust of 200 Kilonewtons.
Video Credit: ISRO
An initial sub-orbital flight test of the LVM3 rocket was performed in December 2014 featuring a mass simulator in place of the C25 upper stage to collect data on how the rocket would perform during operation of its twin boosters and large liquid-fueled first stage. The mission carried the Crew Module Atmospheric Re-entry Experiment (CARE) – a scaled version of India’s future crew capsule that completed a re-entry demonstration and splashed down under parachutes in the Bay of Bengal for recovery and post-flight analysis.
While the lower stages of the LVM3 rocket went through a successful demo flight, C25 was still in the midst of its development process. A pair of engines were used for qualification testing at sea level conditions picking up with a full-duration burn in April 2015 and an endurance test in July for a burn time of 800 seconds followed by another full burn February 2016. A third production CE-20 went through an acceptance test in high-altitude conditions in December 2016, intended to be flown on the first C25 stage.
The first integrated test of a C25 rocket stage was performed on January 25 for a 50-second validation test followed by the 640-second full-duration test carried out on Friday that marked the final planned pre-flight test of the C25 upper stage. “The performance of the Stage during the hot test was as predicted,” ISRO said in a statement. “Successful hot test for flight duration qualifies the design of the stage and the robustness of the facilities conceived and established towards its development.”
Testing of C25 was realized in just four months and, in addition to hot fire tests, included a validation of processing and service procedures needed in the preparation of the stage for flight.
The C25 stage stands 13.3 meters tall and matches the core stage’s four-meter diameter. It holds 27,800 Kilograms of Liquid Oxygen and Liquid Hydrogen for consumption by CE-20, an open cycle gas generator engine that can run at a fixed thrust level between 180 and 220 Kilonewonts, typically operating around 186kN.
LVM3 acts as a launcher for satellites into Geostationary Transfer Orbit and India’s future crew vehicle.

Back in the early-mid 1800s, steamships had a similar problem to today’s launch vehicles. The steam engines of the day burned so much coal that the ships were limited in how far they could travel and still have enough room left over to carry a worthwhile amount of cargo. They solved this problem by breaking up the longer shipping routes into shorter lengths with strategically placed coaling stations. This allowed the early steamships to travel the globe while carrying a lot less coal and a lot more cargo. This significantly reduced the cost of shipping goods and people around the world while allowing the shipping companies to operate at a higher profit margin. It was a win-win solution for everyone.
In the case of Earth to orbit spaceflight, the problem isn’t distance traveled, the problem is the amount of speed the rocket needs to achieve to reach orbit. Since it isn’t possible to place a refueling station halfway up, the only other option is to reduce the amount of speed the launch vehicle needs to achieve to reach orbit. This can be done by adding speed to the launch vehicle at both the beginning and the end of its flight to orbit using externally applied power. This will significantly reduce the amount of propellant the launch vehicle needs to carry, which will allow it to carry more payload.
This is what a Combination Launch System does.
A combination launch system adds velocity to the launch vehicle at the beginning of its flight to orbit using either a catapult,
or by air launching the launch vehicle from high in the atmosphere with a carrier aircraft.
The combination launch system also adds velocity to the launch vehicle at the end of the flight with a non-rotating skyhook.
The end result is that the launch vehicle only needs to carry the propellant for the increase in speed that occurs in the middle part of the flight. The total amount of speed supplied by a mature combination launch system represents up to 1/3 or more of the total speed required for reaching orbit. This reduces the take-off weight to payload weight ratio of the launch vehicle from 200:1 down to 20:1 or less.
This will also allow the launch vehicle to be built as a 100% reusable single-stage vehicle that is much smaller in size than existing launch vehicles.
For example, the Falcon 9, which carries 6,000 pounds of usable payload to the International Space Station, has a take-off weight of 1.2 million pounds.

This results in a final minimum
delivered total impulse of 96,831 lb-sec. We believe that the
flight motor should have had a slightly higher delivered Isp due
to altitude effects and delivered just over 100,000 lb-sec.
Video of the launch is at http://www.hybrids.com/video/csxt_flight.mpg
miles/113km... This posting at arocket was forwarded
to me my Andrew Case:
I just got
a call from Ky out at Black Rock (Tue 5/18 7:45pm cdt). They found
the payload section intact 20 miles from the launch site and had
just returned to base. They are opening it up right now.
The 12,000lb thrust motor burned for 10 seconds and pushed the
rocket to an estimated 4,200 miles per hour. The GPS units cut
in and out so the exact speed and altitude are not determined,
but the altitude is estimated to be 70 miles.
The pyrotechnic separation of the rocket sections worked as planned
and the nosecone section landed under a Rocketman R4 parachute
and landed on the side of a mountain.
It is believed that the booster/motor section came in ballistic.
3 sonic booms were heard as it re-entered. [possibly echoes from
the mountains? - jph]
Ky will be calling in to a morning radio program in the Twin Cities:
KS95 at 6:45am Wed morning. Several television talk show appearances
are scheduled (I don't have the times yet).
Ky said that everything went perfectly on the first launch attempt
and the flight was flawless.
Quite an accomplishment! Congratulations again to the CSXT team.
CSXT article... This
article - Team
Claims Success With Rocket Launch - Space.com/AP - May.18.04
- brings up the question of whether this will be recognized as an
official record at Federation
Aeronautique Internationale. In our records
page here, we cite the previous max altitude for an amateur
group as 80km by George Garboden and Reaction
Research Society, Nov.23,1996, Black Rock, Nevada with the as
2-stage booster with Dart 2nd stage.
BTW: The article
puts Burt Rutan and the SS1 in with "other amateur groups"!!

However, to provide a complete picture of the equations used in rocket engine design, we present below the equation for the flow of liquid through a simple orifice (a round drilled hole, for example)
w = propellalnt flow rate, lb/sec
A = area of orifice, ft2
(deltaP) = pressure drop across orifice, lb/ft^2
(rho) = density of propellant, lb/ft^3
g = gravitational constant, 32.2 ft/sec2
Cd = orifice discharge coefficient
The discharge coefficient for a well-shaped simple orifice will usually have a value between 0.5 and 0.7.
The injection velocity, or vclocity of the liquid stream issuing from the orifice, is given by
Injection pressure drops of 70 to 150 psi, or injection vclocities of 50 to 100 ft/sec are usually used in small liquid-fuel rocket engines. The injection pressure drop must be high enough to eliminate combustion instability inside the combustion chamber but must not be so high that the tankage and pressurization system used to supply fuel to the engine are penalized.
A second type of injector is the spray nozzle in which conical, solid cone, hollow cone, or other type of spray sheet can be obtained. When a liquid hydrocarbon fuel is forced through a spray nozzle (similar to those used in home oil burners) the resulting fuel droplets are easily mixed with gaseous oxygen and the resulting mixture readily vaporized and burned. Spray nozzles are especially attractive for the amateur builder since several companies manufacture them commercially for oil burners and other applications. The amateur need only determine the size and spray characteristics required for his engine design and the correct spray nozzle can then be purchascd at low cost. Figure 7 illustrates the two types of injectors.
The use of commercial spray nozzles for amateur built rocket engines is highly recommended.
Figure 7 Fuel injectors for Amatuer Rocket Engines.

“We have 12-kilowatt systems in development that could be used in much higher power-level systems,” said Julie Van Kleeck, vice president of space programs at Sacramento, Calif.-based Aerojet. The company tested one such thruster at Glenn last year, Van Kleeck said.
A 12-kilowatt engine more than doubles the power of the 4.5-kilowatt, xenon-fueled BPT-4000 Aerojet provided for the Air Force’s Advanced Extremely High Frequency communications satellites.
Aerojet built the thrusters for the Air Force satellite by licensing technology from a Natick, Mass., company called Busek Space Propulsion and Systems. In April, Busek won $5.1 million from NASA’s Small Business Innovation Research program to work on bigger solar electric thrusters in the 10 kilowatt to 20 kilowatt range.
A Little Thrust Can Go a Long Way
Although NASA insists on pushing the state of the art for the asteroid retrieval mission and later Mars missions, even the comparatively small and weak electric propulsion systems flying today are powerful enough to be useful. The Air Force andproved that in 2010, when the service’s Advanced Extremely High Frequency satellite lost its main onboard engine and had to rely on Aerojet-built electric thrusters to boost from the transfer orbit where its rocket left it up to geostationary orbit. The trip took nine months.
The electric thrusters provided by Aerojet for the Air Force satellite — which are among the most powerful being flown today — produce only about 250 millinewtons of thrust. NASA’s Dawn spacecraft, which launched in 2007 to explore two of the largest asteroids in the solar system, gets a whopping 90 millinewtons or so from its xenon-fueled ion engine.
For comparison, a commercially available 8-gram, solid-fuel motor for a model rocket can produce just over 10 newtons of thrust at its peak, making it about 1,000 times more powerful than Dawn’s engine — for the fraction of a second the toy motor is capable of firing.
Dawn’s engine, like other electric thrusters, can stay lit for much longer than that. And they need to. The tiny amount of thrust such engines produce means spacecraft that use them must perform long burns in order to get anywhere.

When the Earth-Mars spaceship gets to Mars it slows down to just under Mars escape velocity and enters a high Mars orbit. The change in velocity required for this maneuver is approximately 900 meters per second. The total change in velocity required for this voyage is 2.4 kilometers per second. That is 1/5th the amount of change in velocity required by the spaceships used in Wernher von Braun’s original Mars fleet. This difference is due to both the Skyhooks and the outpost space stations with local sources of propellant at both Earth and Mars.
The transfer of passengers and cargo from the Earth-Mars spaceship to Mars will be performed by smaller spacecraft operating from the upper end of the Martian Skyhook using locally supplied propellant. These spacecraft will also be used to carry locally supplied propellant to the Earth-Mars spaceship for its return trip to Earth. If it is assumed that the Earth-Mars spaceship uses existing LOX/LH2 chemical rocket motors with oversized expansion nozzles (a specific impulse of 480 seconds), the propellant fraction for the journey to Mars will be 40 percent. This will leave plenty of room in the mass budget for the Earth-Mars spaceship to make it completely reusable (no drop-off propellant tanks or stages), allow for a spinning section with artificial gravity for the crew and passengers, a hydroponics garden for fresh vegetables, plus plenty of radiation shielding. If it is assumed that the empty weight fraction for this Earth-Mars spaceship is 40 percent of the departure mass, that will leave 20 percent for the payload.
If a nuclear thermal rocket motor that uses water for reaction mass is used for the propulsion system for the Earth-Mars spaceship, the propellant fraction for the trip to Mars drops to 25 percent and the payload fraction increases to 35 percent. When used with a combination launch system that includes an escape velocity capable Skyhook, either of these Earth-Mars spaceships would make a trip to Mars mass market affordable on a commercial basis and would operate at a profit for the owners.
Either of these spaceships would also be capable of making trips to the asteroid belt, and to dwarf planets like Ceres, both affordable and possible.