Today I'm taking a break from the alternate ITS missions series, as I'm not yet finished with it. So I've going to post a couple of videos about spaceflight I've found.
First, a video from Scott Manley, about reused spacecraft.
This video is a timelapse of the night sky, but with the motion of the Earth corrected, showing how the Earth moves relative to the stars. Bonus video.
This shows six daylight Falcon 9 landings, timed so they all land at the same time.
And this, from Tom Scott, is a video on how spaceflight hardware and zero-g experiments that do not need enough time in zero-g to require a parabolic flight are tested.
Showing posts with label spacex. Show all posts
Showing posts with label spacex. Show all posts
Sunday, July 30, 2017
Sunday, July 2, 2017
Alternate ITS missions part 3
This post got too long again, so I'm splitting it into yet another section. This will cover how to calculate interplanetary trajectories.
All orbits are conic sections. circles, ellipses, parabolas, or hyperbolas. Circles and ellipses are the orbits that a spacecraft is in when it is in a stable orbit around a planet. Parabolas are when the spacecraft is going at exactly escape velocity, and hyperbolas are when the spacecraft is going faster than escape velocity.
For ellipses (assuming circles are a type of ellipse) the planet the spacecraft is orbiting (or the primary) is always at one of the two foci. The major and minor axes are the longest and the shortest lines that can be drawn through the ellipse, respectively. The semi-major or minor axes are half of the major or minor axes.
The periapsis is the lowest point on an orbit, and the apoapsis is the highest. They are always 180 degrees apart. The closer to the periapsis, the faster the satellite travels, and the closer to apoapsis, the slower.
A quick aside on nodes: nodes are the points at which the orbit passes the plane of the equator or ecliptic. The ascending node is the one the spacecraft passes travelling from south to north, and the descending node is the one the spacecraft passes travelling from north to south.
Eccentricity is the distance between the foci divided by the length of the major axis (always between 0 and 1 except for hyperbolas), and is essentially a measure of how "squished" the ellipse is. 0 is a circle, 1 is parabola, > 1 is a hyperbola.
Inclination is how tilted the orbit is from the equatorial plane (or the ecliptic) of the primary. An inclination of less than 90 degrees indicates a orbit in the same direction as the spin of the primary, or prograde, an inclination of greater than 90 degrees means that the orbit is in the opposite direction of the spin, or retrograde.
I'm ignoring how to calculate these, as that would make this post much longer.
So, let's imagine you want to go from one circular orbit to another. What is the most efficient way of doing this? The answer is a Hohmann transfer.
A Hohmann transfer is the most efficient way to go from a coplanar orbit to another. The image shows a transfer with two circular orbits, however it is the most efficient for any two coplanar orbits where one cannot be reached from the other with only a single burn. For example, in the image, 1 and 3 require a hohmann transfer, while 1 and 2 or 2 and 3 only need a single burn.
The Delta-V required for a hohmann transfer is the sum of the Delta-V required for the burn to go from 1 to 2, and the Delta-V required to go from 2 to 3. It is the same in the opposite direction (3 to 1 = 3 to 2 + 2 to 1), and the burns are the same, except with the spacecraft burning its engines in the opposite direction.
The transfer orbit between two planets is the same thing, with the transfer orbit going from one planet to another, except for the fact that you start in orbit around one of those planets.
Transfer windows are times when the orbits of the planets align such that a spacecraft on a transfer orbit reaches apoapsis (or periapsis) at the same time as the destination planet moves through the part of the orbit that intersects the transfer orbit.
It isn't very important to calculate transfer windows for this post, as they do occur regularly for most planets.
Delta-V for an interplanetary transfer is less than you might expect, as the Oberth effect comes into play on the escape burn. The Oberth effect is, in short, an effect that makes Delta-V count for more when near a large mass. It is why the Delta-V required for a hohmann transfer to GEO is more than the Delta-V required for escape velocity.
The Oberth effect is the effect of the extra velocity of a circular orbit around earth has compared to a circular orbit around the sun at earth's altitude. As the satellite orbits around the earth, it traces out a spiral pattern relative to the sun. When it is orbiting in the same direction around the earth that the earth is orbiting around the sun, the velocity relative to the sun is increased. It isn't as simple as adding the velocities, but we'll get into that later.
On arrival at the destination planet, the Delta-V can be further reduced with aerobraking or aerocapture. This is where the you use the atmosphere of the planet you are arriving at to slow down or speed up without using any fuel. This only works on planets with atmospheres, though, and you cannot capture into a circular orbit without a burn, although it is much less with most of the velocity bled off already.
A very useful equation is the Vis-visa equation:
v^2 = GM(2/r - 1/a)
Where v is the speed of the spacecraft at any point in a orbit, G is the gravitational constant, M is the mass of the primary, r is the distance between the spacecraft and the primary at the point at which you wish to know velocity, and a is the semi-major axis.
The calculations for hohmann transfer orbits assume instantaneous impulses, which lessens delta-v requirements. To compensate, I will multiply the estimate by 1.1.
To calculate Delta-V requirements, you simply need to get the difference in velocities between the orbits at the correct points. This is why the Vis-visa equation is so useful.
We apply this to a hohmann transfer from one planet to another, getting the velocities of the transfer at apoapsis and periapsis. Then, we calculate the velocity of the starting orbit, and calculate the required Delta-V to go to the transfer orbit, accounting for the Oberth effect.
Reversing the rocket equation will then let us calculate payload.
Earth to Venus example
Earth has a mass of 5.9723 * 10^24 kg. It has a semi-major axis of 149.6 * 10^9 meters. Its orbital radius goes from 147.09 * 10^9 m to 152.1 * 10^9 m.
Venus has a mass of 4.8675 * 10^24 kg. It has a semi-major axis of 108.21 * 10^9 m. Its orbital radius goes from 107.48 * 10^9 m to 108.94 * 10^9 m.
The sun has a mass of 1988500 * 10^24 kg.
A transfer orbit between them depends on the position of the planets at departure and arrival. For this example, the position of Venus does not matter, as aerobraking will work at either of those speeds.
Earth should be at apoapsis, because in this case we want to lose, rather than gain, velocity as Venus is in a lower orbit than Earth.
Let's begin. First, the velocity of Earth around the Sun at apoapsis. We don't really need to calculate this, as it is well known: 29290 m/s.
The transfer orbit is an elliptical orbit with its apoapsis at Earth, and its periapsis at Venus. This gives a semi-major axis of:
(Radius of Earth's orbit at departure + Radius of Venus' orbit at arrival)/2.
Or (152.1 * 10^9 m + 108.94 * 10^9 m)/2 = 130.52 * 10^9 m.
Velocity at apoapsis:
v^2 = 1.327 * 10^20 (2/152.1 * 10^9 - 1/130.52 * 10^9)
v = 26985 m/s
At periapsis:
v^2 = 1.327 * 10^20 (2/108.94 * 10^9 - 1/130.52 * 10^9)
v = 37676 m/s
The velocity of the spaceship in Earth orbit can be calculated from the altitude, and the altitude is as low as possible to take maximum advantage of the Oberth effect. I'm going to guess around 200 km.
v^2 = 3.986 * 10^14 (2/6578000 - 1/6578000)
v = 7784 m/s.
To figure out how much of an advantage the Oberth effect gives us, we use something called hyperbolic excess velocity. This is how much extra velocity the spaceship will have when it reaches infinite distance on a hyperbolic escape orbit. If you reach exactly escape velocity, HEV will be zero. The desired HEV is equal to the orbital velocity of the primary - the velocity of the transfer orbit at the point where it begins so when the spacecraft escapes the primary's gravity, the remaining velocity puts it on the transfer orbit.
The equation for HEV is HEV^2 = v^2 - ve^2
Our desired HEV is 29290 - 26985 = 2305.
Rearranging the equation:
v^2 =2305^2 + 11008^2.
(as escape velocity is equal to the velocity of a circular orbit at that altitude times √2)
So v = 11246 m/s.
11246 - 7784 = 3462.
Here you can see the clear advantage of this over a escape burn (3224) + a burn outside of Earth's sphere of influence to put the spacecraft on a transfer orbit (2305), which would be 2067 m/s more.
Venus is inclined 3.39 degrees off of Earth, however this amounts to at most a few m/s if it is corrected for at the very start of the transfer. For this kind of loose approximation, we can ignore this.
So, multiplying by 1.1, we have a Delta-V of 3808 m/s.
Next post, we will calculate transfers for other missions, including approximate landing Delta-V, and use the reverse rocket equation to figure out how much payload can be carried to them.
All orbits are conic sections. circles, ellipses, parabolas, or hyperbolas. Circles and ellipses are the orbits that a spacecraft is in when it is in a stable orbit around a planet. Parabolas are when the spacecraft is going at exactly escape velocity, and hyperbolas are when the spacecraft is going faster than escape velocity.
For ellipses (assuming circles are a type of ellipse) the planet the spacecraft is orbiting (or the primary) is always at one of the two foci. The major and minor axes are the longest and the shortest lines that can be drawn through the ellipse, respectively. The semi-major or minor axes are half of the major or minor axes.
The periapsis is the lowest point on an orbit, and the apoapsis is the highest. They are always 180 degrees apart. The closer to the periapsis, the faster the satellite travels, and the closer to apoapsis, the slower.
A quick aside on nodes: nodes are the points at which the orbit passes the plane of the equator or ecliptic. The ascending node is the one the spacecraft passes travelling from south to north, and the descending node is the one the spacecraft passes travelling from north to south.
Eccentricity is the distance between the foci divided by the length of the major axis (always between 0 and 1 except for hyperbolas), and is essentially a measure of how "squished" the ellipse is. 0 is a circle, 1 is parabola, > 1 is a hyperbola.
Inclination is how tilted the orbit is from the equatorial plane (or the ecliptic) of the primary. An inclination of less than 90 degrees indicates a orbit in the same direction as the spin of the primary, or prograde, an inclination of greater than 90 degrees means that the orbit is in the opposite direction of the spin, or retrograde.
I'm ignoring how to calculate these, as that would make this post much longer.
So, let's imagine you want to go from one circular orbit to another. What is the most efficient way of doing this? The answer is a Hohmann transfer.
Image from Leafnode
The Delta-V required for a hohmann transfer is the sum of the Delta-V required for the burn to go from 1 to 2, and the Delta-V required to go from 2 to 3. It is the same in the opposite direction (3 to 1 = 3 to 2 + 2 to 1), and the burns are the same, except with the spacecraft burning its engines in the opposite direction.
The transfer orbit between two planets is the same thing, with the transfer orbit going from one planet to another, except for the fact that you start in orbit around one of those planets.
Transfer windows are times when the orbits of the planets align such that a spacecraft on a transfer orbit reaches apoapsis (or periapsis) at the same time as the destination planet moves through the part of the orbit that intersects the transfer orbit.
It isn't very important to calculate transfer windows for this post, as they do occur regularly for most planets.
Delta-V for an interplanetary transfer is less than you might expect, as the Oberth effect comes into play on the escape burn. The Oberth effect is, in short, an effect that makes Delta-V count for more when near a large mass. It is why the Delta-V required for a hohmann transfer to GEO is more than the Delta-V required for escape velocity.
The Oberth effect is the effect of the extra velocity of a circular orbit around earth has compared to a circular orbit around the sun at earth's altitude. As the satellite orbits around the earth, it traces out a spiral pattern relative to the sun. When it is orbiting in the same direction around the earth that the earth is orbiting around the sun, the velocity relative to the sun is increased. It isn't as simple as adding the velocities, but we'll get into that later.
On arrival at the destination planet, the Delta-V can be further reduced with aerobraking or aerocapture. This is where the you use the atmosphere of the planet you are arriving at to slow down or speed up without using any fuel. This only works on planets with atmospheres, though, and you cannot capture into a circular orbit without a burn, although it is much less with most of the velocity bled off already.
A very useful equation is the Vis-visa equation:
v^2 = GM(2/r - 1/a)
Where v is the speed of the spacecraft at any point in a orbit, G is the gravitational constant, M is the mass of the primary, r is the distance between the spacecraft and the primary at the point at which you wish to know velocity, and a is the semi-major axis.
The calculations for hohmann transfer orbits assume instantaneous impulses, which lessens delta-v requirements. To compensate, I will multiply the estimate by 1.1.
To calculate Delta-V requirements, you simply need to get the difference in velocities between the orbits at the correct points. This is why the Vis-visa equation is so useful.
We apply this to a hohmann transfer from one planet to another, getting the velocities of the transfer at apoapsis and periapsis. Then, we calculate the velocity of the starting orbit, and calculate the required Delta-V to go to the transfer orbit, accounting for the Oberth effect.
Reversing the rocket equation will then let us calculate payload.
Earth to Venus example
Earth has a mass of 5.9723 * 10^24 kg. It has a semi-major axis of 149.6 * 10^9 meters. Its orbital radius goes from 147.09 * 10^9 m to 152.1 * 10^9 m.
Venus has a mass of 4.8675 * 10^24 kg. It has a semi-major axis of 108.21 * 10^9 m. Its orbital radius goes from 107.48 * 10^9 m to 108.94 * 10^9 m.
The sun has a mass of 1988500 * 10^24 kg.
A transfer orbit between them depends on the position of the planets at departure and arrival. For this example, the position of Venus does not matter, as aerobraking will work at either of those speeds.
Earth should be at apoapsis, because in this case we want to lose, rather than gain, velocity as Venus is in a lower orbit than Earth.
Let's begin. First, the velocity of Earth around the Sun at apoapsis. We don't really need to calculate this, as it is well known: 29290 m/s.
The transfer orbit is an elliptical orbit with its apoapsis at Earth, and its periapsis at Venus. This gives a semi-major axis of:
(Radius of Earth's orbit at departure + Radius of Venus' orbit at arrival)/2.
Or (152.1 * 10^9 m + 108.94 * 10^9 m)/2 = 130.52 * 10^9 m.
Velocity at apoapsis:
v^2 = 1.327 * 10^20 (2/152.1 * 10^9 - 1/130.52 * 10^9)
v = 26985 m/s
At periapsis:
v^2 = 1.327 * 10^20 (2/108.94 * 10^9 - 1/130.52 * 10^9)
v = 37676 m/s
The velocity of the spaceship in Earth orbit can be calculated from the altitude, and the altitude is as low as possible to take maximum advantage of the Oberth effect. I'm going to guess around 200 km.
v^2 = 3.986 * 10^14 (2/6578000 - 1/6578000)
v = 7784 m/s.
To figure out how much of an advantage the Oberth effect gives us, we use something called hyperbolic excess velocity. This is how much extra velocity the spaceship will have when it reaches infinite distance on a hyperbolic escape orbit. If you reach exactly escape velocity, HEV will be zero. The desired HEV is equal to the orbital velocity of the primary - the velocity of the transfer orbit at the point where it begins so when the spacecraft escapes the primary's gravity, the remaining velocity puts it on the transfer orbit.
The equation for HEV is HEV^2 = v^2 - ve^2
Our desired HEV is 29290 - 26985 = 2305.
Rearranging the equation:
v^2 =2305^2 + 11008^2.
(as escape velocity is equal to the velocity of a circular orbit at that altitude times √2)
So v = 11246 m/s.
11246 - 7784 = 3462.
Here you can see the clear advantage of this over a escape burn (3224) + a burn outside of Earth's sphere of influence to put the spacecraft on a transfer orbit (2305), which would be 2067 m/s more.
Venus is inclined 3.39 degrees off of Earth, however this amounts to at most a few m/s if it is corrected for at the very start of the transfer. For this kind of loose approximation, we can ignore this.
So, multiplying by 1.1, we have a Delta-V of 3808 m/s.
Next post, we will calculate transfers for other missions, including approximate landing Delta-V, and use the reverse rocket equation to figure out how much payload can be carried to them.
Sunday, April 30, 2017
Alternate ITS missions part 1
The Interplanetary Transport System is a spacecraft concept developed by SpaceX and designed to colonize Mars. While it is designed for that specific goal, it could feasibly be adapted to other missions. This post, part 1, talks about the performance of the ITS, and the second part will talk about what missions it is capable of.
The ITS has two configurations, the ship, and the tanker. The ship is designed to carry 450t or 100 people to Mars, while the tanker is designed to refuel the ship in Earth orbit.
Let's take a closer look at the performance.
First of all, the maximum delta-v.
Using the Tsiolkovsky rocket equation (DV = Ve*ln(minitial/mfinal)) like so:
DV = 3,747.42*ln(2100/150) = 9889 m/s
This is an impressively high delta-v, however it assumes no payload, with the standard ITS configuration (not the tanker).
With the tanker configuration:
DV = 3,747.42*ln(2590/90) = 12589 m/s
This also assumes no payload.
You can calculate how much delta-v you need for transfers between planets with a porkchop plot, or you could just use a handy delta-v map of the solar system that assumes reasonably efficient transfers.
(Chart from Atomic Rockets)
We can reverse the rocket equation to find out how much payload we can carry, given a specific required delta-v. Wolfram|alpha can do this for us easily, and the answer is mfinal = minitial*ℯ-DV/VE
ℯ = Euler's constant, or ~2.71.
Payload equals mfinal - empty mass
Empty mass differs between the tanker and the standard ITS configuration by 60t. It's hard to tell what of that is necessary for a much smaller crew, as might be used on these missions. A safe bet is probably around 120t, halfway between the two numbers.
Propellant mass also differs, with 1950t for the ship and 2500t for the tanker. In this case, to keep costs low, the ship value of 1950t would be used, as there is no space for crew in the tanker, and changing the size and shape of the composite tanks is expensive.
Note that while SpaceX uses a value of 6km/s from low earth orbit (LEO) to trans-Mars injection (TMI), more efficient (but slower) transfers use on the order of 3-4km/s. Presumably, this is done to minimize life support costs; both consumables and radiation shielding. However, if you're sending 5 people instead of 100, consumable costs per day only scale 1/20 as fast per unit time, so almost any other missions would sacrifice travel speed for payload.
The ITS has 3 sea level Raptor engines, and 6 vacuum Raptors. Basically, in a vacuum you could use the sea level engines in addition to the vacuum engines, but it would just hurt our delta-v. In an atmosphere, you can't use the vacuum engines due to flow separation, which happens when your engine is overexpanded. (flow separation is an extreme example of what causes the space shuttle main engines to wobble as they start up, as they have to be efficient throughout the entire atmosphere: https://www.youtube.com/watch?v=hDCCBgppG4s)
So the ITS has different thrust-to-weight ratios (TWR) for different situations.
3*SL Raptor = 9150kn = 933 metric tons-force = empty TWR of 7.775
= loaded TWR of 0.45
6*VAC Raptor = 21000kn = 2141 metric tons-force = empty TWR of 17.84
= loaded TWR of 1.03
Max possible thrust in VAC = 3074 metric tons-force = empty TWR of 25.62
= loaded TWR of 1.49
These are all assuming you're on Earth, with no payload.
The ITS can withstand 10-15gs, and a minimum of 1,700 degrees C.
In the next part we will go over what missions the ITS could perform now that we have determined its performance.
The ITS has two configurations, the ship, and the tanker. The ship is designed to carry 450t or 100 people to Mars, while the tanker is designed to refuel the ship in Earth orbit.
Let's take a closer look at the performance.
First of all, the maximum delta-v.
Using the Tsiolkovsky rocket equation (DV = Ve*ln(minitial/mfinal)) like so:
DV = 3,747.42*ln(2100/150) = 9889 m/s
This is an impressively high delta-v, however it assumes no payload, with the standard ITS configuration (not the tanker).
With the tanker configuration:
DV = 3,747.42*ln(2590/90) = 12589 m/s
This also assumes no payload.
You can calculate how much delta-v you need for transfers between planets with a porkchop plot, or you could just use a handy delta-v map of the solar system that assumes reasonably efficient transfers.
We can reverse the rocket equation to find out how much payload we can carry, given a specific required delta-v. Wolfram|alpha can do this for us easily, and the answer is mfinal = minitial*ℯ-DV/VE
ℯ = Euler's constant, or ~2.71.
Payload equals mfinal - empty mass
Empty mass differs between the tanker and the standard ITS configuration by 60t. It's hard to tell what of that is necessary for a much smaller crew, as might be used on these missions. A safe bet is probably around 120t, halfway between the two numbers.
Propellant mass also differs, with 1950t for the ship and 2500t for the tanker. In this case, to keep costs low, the ship value of 1950t would be used, as there is no space for crew in the tanker, and changing the size and shape of the composite tanks is expensive.
Note that while SpaceX uses a value of 6km/s from low earth orbit (LEO) to trans-Mars injection (TMI), more efficient (but slower) transfers use on the order of 3-4km/s. Presumably, this is done to minimize life support costs; both consumables and radiation shielding. However, if you're sending 5 people instead of 100, consumable costs per day only scale 1/20 as fast per unit time, so almost any other missions would sacrifice travel speed for payload.
The ITS has 3 sea level Raptor engines, and 6 vacuum Raptors. Basically, in a vacuum you could use the sea level engines in addition to the vacuum engines, but it would just hurt our delta-v. In an atmosphere, you can't use the vacuum engines due to flow separation, which happens when your engine is overexpanded. (flow separation is an extreme example of what causes the space shuttle main engines to wobble as they start up, as they have to be efficient throughout the entire atmosphere: https://www.youtube.com/watch?v=hDCCBgppG4s)
So the ITS has different thrust-to-weight ratios (TWR) for different situations.
3*SL Raptor = 9150kn = 933 metric tons-force = empty TWR of 7.775
= loaded TWR of 0.45
6*VAC Raptor = 21000kn = 2141 metric tons-force = empty TWR of 17.84
= loaded TWR of 1.03
Max possible thrust in VAC = 3074 metric tons-force = empty TWR of 25.62
= loaded TWR of 1.49
The ITS can withstand 10-15gs, and a minimum of 1,700 degrees C.
In the next part we will go over what missions the ITS could perform now that we have determined its performance.
Monday, December 28, 2015
Falcon 9 & SpaceX
From redditor tossha
The Falcon 9 is a two-stage rocket fully designed and manufactured by SpaceX. It is, most importantly, partially reusable. The first stage can land upright, autonomously. This has been compared to balancing a rubber broomstick in a windstorm. This is important because it means that all the cost of building a first stage can be recouped. In addition, SpaceX's crew capsule (crew dragon) can be reused.
We'll get back to the design of the Falcon 9, but first, a bit of background.
Rocket technology has not significantly advanced since the '60s, so it still costs as much as ever to put stuff in space (US$13,812 per kilogram to low earth orbit for an Atlas V), which is why President George H. W. Bush's plan (announced in July 20, 1989) for a manned Mars mission was quickly thrown out when NASA told him that it would cost 500 billion US dollars.
SpaceX was started when Musk tried to create interest in space by sending a greenhouse to Mars, but everywhere he went for look for his rocket, he found it was ridiculously expensive. So he started his own space exploration company. He took the cost of the raw materials required to make a rocket, and compared them to the cost of a typical rocket price, and got about 2%. The materials cost percentage of market price for a car is about 47%. Elon Musk discovered that major aerospace firms don't want to fly unflown equipment, which means that no new technology ever gets used. The way companies cut costs is by using engines actually manufactured in the '60s.
So the Falcon 9 can launch a kilogram to low earth orbit for US$4109, because that SpaceX makes almost all of the rocket in-house, so there's none of the levels of sub-contractors that all want a piece of the price tag pie. Plus, SpaceX uses modern methods of manufacturing and assembly that all cut costs. And, of course, reusability will make it far cheaper than before, possibly 100 times cheaper. Here Elon Musk sums it up nicely:
Imagine if you had to have a new plane for every flight. Very few people would fly. -Elon MuskAirplanes aren't any cheaper to build if you reuse them, but you can make the price for a flight cheaper if you can do many flights with them.
But, I hear you ask, why is Elon Musk doing all this? Considering that he was already rich from paypal, why did he decide to revolutionize space travel? Because he wants (in addition to running a profitable company) to put 1,000,000 people on Mars. And he wants their ticket to cost only 500,000 US dollars. He wants to do this for two reasons, to create a "backup" of the human race, so if a mass extinction event happens to a planet with humans on it, the human race can survive, and for this reason:
This seems like a good idea, but we should be careful not to think of this as a solution for humans' impact on Earth. Unsustainable fuels are unsustainable on Mars, and a terraformed Mars' ecosystem would be just as fragile as Earth's. It is, however, a partial solution for overpopulation. Ideally, it could open up far more space for humans. Nonetheless, that space is not infinite. Humans would have to continuously colonize new planets to support the growth of population.
Now, if you look almost all the way to the top of the post, you'll see that this was actually a digression. Back to the Falcon 9!
The Falcon rocket family is named after the Millennium Falcon from the Star Wars franchise. The first version of it was called the Falcon 1. The Falcon 1 was designed to test construction techniques for the Falcon 1, while also launching satellites for profit. SpaceX was at this being funded directly from Elon Musk's pocket, and it had money for 3 or 4 launches when it started. The first launch was a failure. So was the second and third. Finally, on the fourth launch, the Falcon 1 made it into orbit, saving SpaceX from closing. After that, NASA awarded a US$ 1.6 billion contract to SpaceX.
Then SpaceX built the Falcon 9. Here we're going to be talking about the Falcon 9 v1.1 full thrust, the latest version. The Falcon 9 has two stages, with 9 Merlin 1D engines on the first stage and one Merlin Vacuum 1D on the second stage. First let's look at the Merlin engine:
The Merlin is developed fully in-house by SpaceX, and is the highest thrust to weight ratio (TWR) rocket ever built. The 1D version has an absurd TWR of 165.9. For comparison, the space shuttle main engine (SSME) gets only 73.1. TWR is important in rocket engines because that the higher the TWR, the more mass the rocket engine can lift compared to it's mass, so you need less engine mass to lift the same amount of payload. i.e., it's an incredible rocket engine.
Here's a Merlin 1D firing:
Now, the rocket:
The entire rocket uses many smart construction techniques, such as 3D printing to make their rockets cheaper. Look at this first, for an idea of how it looks: http://www.spacex.com/falcon9
The first stage has one very important feature: reusability. The first stage lands on four carbon fiber and aluminum honeycomb landing legs, which fold down for landing.
From SpaceX
The first stage also has four deployable grid fins, at the top of the first stage, which help to keep the rocket on the right course. The honeycomb structures in the photo below are the fins:
From SpaceX
In addition to these, there are small cold gas thrusters to keep the first stage on course during landing.
The second stage is like a small version on the first stage, with only one Merlin Vacuum 1D engine and a smaller tank. It is not reusable.
Subscribe to:
Posts (Atom)