Falcon 1 and Upper Stages

Status
Not open for further replies.
B

barrykirk

Guest
Warning: Long winded post with a lot of questions.<br /><br />I've noticed in the literature that I've been able to dig up the following facts about the Falcon 1.<br /><br />First Stage Burn Time is 169 seconds.<br /><br />Second Stage Burn Time is 378 seconds.<br /><br />So, the second stage burns for slightly longer than twice as long as the first stage...<br /><br />Having said that, I found a chart that showed the second stage developed 0.65 g's at ignition and took around 130 seconds to work up to 1 g. Although it slowly builds accel over its burn to peak at a little over 4 g's by burnout.<br /><br />Does that make any sense that they would go that much under 1 g and for that long? Would they do better starting the Kestrel with less fuel so she starts at 1 g accel to reduce gravity losses?<br /><br />I found the rocket total launch mass is 60,000 pounds and the Merlin engine develops 77,000 pounds.<br /><br />I also found that the Kestrel engine develops around 7,000 pounds force. Around 9.1% the power of the Merlin. But with a higher ISP because it's entire flight is in a vacuum or close to it.<br /><br />Obviously, the Kestrel burns fuel at a much lower rate than the Merlin.<br /><br />No where did I find the breakdown of how much fuel is on the first stage and how much fuel is on the second stage.<br /><br />Also, the number I have, 7000 LBs for the second stage, does that include the payload?<br /><br />Another question I have is what is the velocity at first stage seperation? Does most of the velocity come from the second stage?<br /><br />I know that with the Saturn vehicles like the Saturn IB which was two stage to orbit. The velocity at first stage burnout was less than half the final orbital velocity. Is the falcon similar?<br /><br />Here is another one that I've been pondering.<br /><br />When the BFG comes out. If SpaceX makes it into a five engine cluster like Saturn V and puts a full blown Falcon IX on top of it, what would be the total mass to orbit capability?<br />
 
B

barrykirk

Guest
The five stage cluster I was talking about was their new BFG or Merlin 2 engine which has been announced as maybe being in R & D, or maybe not even that far developed.<br /><br />Elon Musk described an engine with about the same thrust as the old F-1 engine.<br /><br />He never mentioned what upper stages to put on top of that engine, but he did mention he wanted at least 125 tons to LEO.
 
B

barrykirk

Guest
Thank you for point out something I already knew, but was too lazy to think of myself.... Doh!!! I did know but forgot I knew.. Gee,,, definition of ISP is....<br /><br />Senility is setting in.<br /><br />OK...<br /><br />Launch mass of rocket is 27,200 Kg<br /><br />So, fuel for first stage is<br /><br />343KN * 169 sec / ( 255 sec * 9.8 m/sec/sec )<br /><br />fuel and oxidizer is roughly 23,200 Kg.<br /><br />Assuming that the fuel usage rate is a constant irrespective of external air pressure. That leaves 4000Kg for second stage and first stage engines plus tanks plus everything else.<br /><br />Second stage fuel is<br /><br />31 KN * 378 sec / ( 327 sec * 9.8 m/sec/sec )<br /><br />fuel and oxidizer is roughly 3657 Kg.<br /><br />for a total of 26857 Kg.<br /><br />Some of these number must be off a little.... It wouldn't take much.<br /><br />The rocket is virtually all fuel and oxidizer and nothing left over for engines, payload, tanks and everything else. I would have assumed no more than a total of 25,000Kg for fuel and oxidizer.<br /><br /><br />
 
B

barrykirk

Guest
Most of the second stage burn is tangent to the Earth's radius.<br /><br />Yes, that makes sense, and as was pointed out, the first stage goes almost straight vertical for a long time.<br /><br />I don't know what the max altitude of the first stage is. Obviously, it acheives that long after seperation.<br /><br />If that altitude is close to the final orbital altitude, then the second stage burn time of 378 seconds will probably be complete about the time that the second stage reaches orbital altitude. Even if she goes perfectly tangential the whole time.
 
P

propforce

Guest
<font color="yellow">OK... <br /><br />Launch mass of rocket is 27,200 Kg ...........<br /><br />So, fuel for first stage is ...... roughly 23,200 Kg. ...............<br /><br />Second stage fuel is .............. roughly 3657 Kg. ...........<br /><br />for a total of 26857 Kg. </font><br /><br /><br />Sounds a little high. This would put the mass fraction of propellant at 98.74%, that is WITHOUT the weight of payload and fairings.<br /><br />I think the error is in using the burn time. The total duration, e.g., burn time, may be correct but this does not account for the engine throttle, particularly for the 1st stage to limit the max axial G-loads.<br /> <br />It's too bad one can not find the propellant mass fraction on the Falcon 1. <br /><br />Here's Falcon 1 nominal trajectory and mission timeline.<br /> <div class="Discussion_UserSignature"> </div>
 
S

spacester

Guest
Sigh. I drop in to see if I can learn anything about Falcon . . .<br /><br /><font color="yellow">Most of the second stage burn is on a tangent to the Earths radius which means there are no gravity losses. </font><br /><br />Incorrect.<br /><br />Any time you are thrusting in a gravity field there are gravity losses. Period.<br /><br />Thrusting tangent to the <i> flight path </i> (not the Earth's radius) will <i>help to minimize</i> the gravity losses, not eliminate them: off-axis thrust is wasted on holding the thing up in the gravity field instead of accelerating the vehicle.<br /><br />In the case of a vehicle which is already in a circular orbit, this flight path will be tangent to Earth's surface and thus thrusting along the line of flight will bring the losses to the minimum possible. But not to zero.<br /><br />Any time you are thrusting in a gravity field there are gravity losses.<br /><br />(I will continue to look elsewhere for new knowledge and insight, sdc sure ain't the place any more.) <div class="Discussion_UserSignature"> </div>
 
S

spacester

Guest
Aah, so I'm a 'star' now, eh? My post count has decreased by about 1200 or so. Good ol' Infopoop strikes again, is that it?<br /><br />Whatever. At the rate I plan to post here in the future, I should get my count back up to where it was in about, oh, the year 2067 or so. <div class="Discussion_UserSignature"> </div>
 
S

scottb50

Guest
Manic Monday? <div class="Discussion_UserSignature"> </div>
 
P

propforce

Guest
<font color="yellow">Any time you are thrusting in a gravity field there are gravity losses. Period. </font><br /><br />True. However; most of the gravity losses (~ 80%) take place during the 1st stage burn, so are the aero & base drag (~100%). Take a look at the Falcon 1 mission timeline right above your post. The vehicle reaches 266,000 feet at the end of the first stage burn. <br /><br /><br /><font color="yellow">Thrusting tangent to the flight path (not the Earth's radius) will help to minimize the gravity losses, not eliminate them:</font><br /><br />True. What the vehicle wants to do is a ballistic flight path to minimize energy (or optimize payload) to orbit, so that means sometime during the 1st stage burn the vehicle tilt over to ~ 40 to 45 degree angle (gamma) relative to Earth's radius and starts a parabolic ballistic trajectory upward. For the 2nd stage part of trajectory, it's axis of thrust is very close to its flgiht path angle in order to minimize energy spent (thrust vector loss), usually within 5 degree (alpha). <br /><br />However; Shuttle_Guy is also correct in saying that the thrust axis is also fairly tangential to Earth's radius (gamma), if you look at a typical trajectory profiles. It's usually within 10 dgrees during the 2nd stage 1st burn, and for the Shuttle it's probably closer to zero degree, and practically equals to zero degree during the 2nd burn for circularization. In any case, the gravity losses in this segment of flgiht is fairly small, though not technically zero, as compared to the 1st stage performance. I think that's what the Shuttle_Guy was referring to.<br /><br /> <br /><font color="yellow">In the case of a vehicle which is already in a circular orbit, this flight path will be tangent to Earth's surface and thus thrusting along the line of flight will bring the losses to the minimum possible. But not to zero</font><br /><br />In the circularization burn, the gravity losses is less than 1% <div class="Discussion_UserSignature"> </div>
 
J

josh_simonson

Guest
Perhaps a better way to look at is is that the fuel expended to avoid hitting the earth approaches zero as the craft approaches orbital velocity. <br /><br />Viewed as two thrust vectors, one provides acceleration to orbital velocity and is tangential to the earth, the other points towards the earth to maintain altitude while accelerating. The vertical component must be at least m(V-Vorbital)**2/r(orbit) - so this drops off as the square of your delta from orbital velocity. The tangential force can be whatever you want, but you want it to be big.<br /><br />Once in orbit, a spacecraft doesn't lose anything to gravity, thrust is directly converted into potential and kinetic energy. <br /><br />
 
P

propforce

Guest
<font color="yellow">I found a chart that showed the second stage developed 0.65 g's at ignition and took around 130 seconds to work up to 1 g. Although it slowly builds accel over its burn to peak at a little over 4 g's by burnout. <br /><br />Does that make any sense that they would go that much under 1 g and for that long? Would they do better starting the Kestrel with less fuel so she starts at 1 g accel to reduce gravity losses? </font><br /><br /> <br />It is pretty common for 2nd stage start with less than 1 G of acceleration. As a matter of fact, many starts way less than 1 G and do not end up with 1 G even at the end of burn.<br /><br />If you look at both Atlas and Delta 2nd stages, both starts with 0.2 ~ 0.3G burn and ends with a 0.5~0.9 G<br /> <div class="Discussion_UserSignature"> </div>
 
B

barrykirk

Guest
Sorry for getting your hopes up and dashing them Spacester. I too feel starved for information from SpaceX.<br /><br />I would like to get more technical information about their rockets... More specs that are sadly lacking. Such as the individual masses of the stages.....<br /><br />Does anybody know where I can download a free software package, preferably a JAVA applet or an excel spreadsheet that would simulate a rocket. I'm not talking about a model rocket, but an orbital rocket. It would be real nice if this software allows the following parameters.<br /><br />1) Number of Stages.<br />2) ISP of individual stages, maybe even a variable ISP for the first stage as it changes altitude.<br />3) Different trajectory profiles.<br />4) Mass ratios of each stage.<br />5) Payload mass.<br />6) Frontal Area of rocket.<br />7) Height of rocket.<br />8) Fuel density.<br />9) Thrust to weight ratio of various engines<br /><br />It would be nice if you put in the masses and the fuels and the program looked up the density of the fuel and then allowed for rocket height and frontal area.<br /><br />I doubt whether a program that does all that exists. Actually, I'm sure it exists, but probably not as freeware.
 
B

barrykirk

Guest
Josh, I was unaware that the equation took that form.<br /><br />Vertical Accleration Component is ( V - Vorbital )^2 / r ( orbit )<br /><br />That means that most by the time you get to half orbital velocity. The rocket is accelerating towards the earth at 0.25 G unless offset by thrust.<br /><br />If that is the case than it makes the much lower acceleration of the upper stages much more understandable. It means that you can start with much more fuel for the same size engine.<br /><br />I had assumed that since centripetal acceleration is of the form V^2/ R, that the vertical acceleration required would be of the form.<br /><br />[ (Vorbital)^2 - V^2 ] / R<br /><br />The difference is key.<br /><br />With the equation you provided, the biggest force reduction occurs at the lowest velocities.<br /><br />With the equation I came up with the biggest force reduction occurs as one approaches orbital velocity.<br /><br />Sadly, my equation makes it much tougher for the rocket designer.<br /><br />Or am I missing something or is there an error in my thought processes?
 
P

propforce

Guest
<font color="yellow">"...If you look at both Atlas and Delta 2nd stages, both starts with 0.2 ~ 0.3G burn and ends with a 0.5~0.9 G..." <br /><br />What does Ariane V's look like? Its upper stage burn seems huge! </font><br /><br />The Ariane V has a very low G for the upper stage, it starts with ~ 0.1 G and ends not much higher. The reason is that it uses a very low-power storable propellant, pressure-fed, Aestus engine (F=6,500 lbf). Most of the velocity was delivered by the SRM and the Vulcain engine. That's why it can only deliver ~40K lbm to LEO while both Atlas V540/ V550 and the Delta IV-H can match or exceed its payload capability. <div class="Discussion_UserSignature"> </div>
 
P

propforce

Guest
<font color="yellow">Does anybody know where I can download a free software package.....</font><br /><br />Write your own program, especially on EXCEL spreadsheet. This way, it will force you to learn the equations and think about what the number really means. <br /><br /> <div class="Discussion_UserSignature"> </div>
 
E

edkyle98

Guest
Try this. <br /><br />http://www.geocities.com/launchreport/falcon.html<br /><br />I referenced most of these numbers from the Falcon 1 <br />user's guide that SpaceX used to have on-line. The <br />design has changed since then, so the real numbers <br />may differ slightly.<br /><br />In summary, my numbers were:<br /><br />Stage 1:<br /><br />Dry: 1.296 tonnes<br />Propellant: 21.092 tonnes<br />Total: 22.388 tonnes<br /><br />Stage 2: <br /><br />Dry: 0.36 tonnes<br />Propellant: 3.385 tonnes<br />Total: 3.745 tonnes<br /><br />Payload Fairing: 0.44 tonnes<br /><br />Using these numbers, assuming a 0.5% propellant <br />residual, and with a 0.5 tonne payload, the rocket <br />equation has the first stage providing about 45% <br />of the total ideal delta-v. The rocket just barely <br />makes it to orbit - very little margin.<br /><br /> - Ed Kyle
 
B

barrykirk

Guest
Well that makes sense that the rocket just barely makes it to orbit. If it had excess delta V to make orbit, they would increase the payload capacity until it just barely capable of making orbit.
 
P

propforce

Guest
Thanks for the link. Very informative indeed.<br /><br />FYI. I also got the Falcon 1 engine data update, it's 'better' than what's shown on in your link above.<br /><br />Merlin<br />Thrust (sl) = 72,000 lbf<br />Thrust (vac) = 85,000 lbf<br />Ivac = 325 sec.<br /><br />Kestrel<br />Thrust (vac) = 7,500 lbf<br />Ivac = 378 sec.<br /><br />Based on these numbers, I agree that it's payload capability is far less than the 1,400 lbm to LEO that it claims. Very little margin indeed. It will not meet the "assured access" payload criteria.<br /> <div class="Discussion_UserSignature"> </div>
 
P

propforce

Guest
<font color="yellow">Is there a definition of "LEO" that allows one breed to be compared to another, or is a lot of this PR smoke and mirrors anyway? </font><br /><br />There's always lots of PR smokes and mirrors when it comes to "marketing" type of payload performance chart as shown in each company's payload planner's guide. That's why one would have to ask more detailed questions such as velocity reserves, % propellant (fuel or oxidizer) bias, % boil-off, % undrainable, margin/ "knock-down" factor on engine Isp, etc. <br /><br />In this case, SpaceX's number shows 100% of fuel & oxidizer loaded = 100% used for performance, e.g., no margin, no reserves, no undrainables, no boil-offs, etc. what it says is this is a "marketing" type of payload number. Kinda like when engine companies put out the "theoretical" engine Isp numbers. There are ways to check these numbers for those who do this type of calculation. You develop a "rule of thumb" for each parameters and get a "feel" of when the numbers are not in the ball park.<br /><br />Typical "LEO" (other than a law enforcement officer) means a 185km (100 nmi) altitude with a 28.5 deg inclination for a "due east" launch. That number is often used as a comparison, and can translate a payload weight from there to other orbits. Strictly speaking, however; LEO covers anything under 1,000 nmi altitude circular orbit, MEO covers from 1,000 to 10,000 nmi altitude, and HEO covers anything above 10,000 nmi altitude. Other more well-known orbits are:<br /><br />Space Station orbit at 407km (220 nmi) and 51.6 deg, <br />the Sun Synch. orbit at 800 km (432 nmi) and 98.6 deg, <br />or the Geo Transfer Orbit (GTO) at 167 x 35,788km (90 x 19,323 nmi) and 27 deg. <br />Geo Synchronous Orbit (GEO) is a circular obit at 19,323 nmi. <br /><br />There are some more interesting orbit such as the Molniya orbit, a highly elliptical orbit developed by the Russian communication satellites, that will spend upto 10 hours cover the c <div class="Discussion_UserSignature"> </div>
 
P

propforce

Guest
Here's a view from space on the Molniya orbit. Notice the coverage on Russia.... <br /> <div class="Discussion_UserSignature"> </div>
 
E

edkyle98

Guest
Your specific impulse numbers are higher than the numbers that SpaceX has posted on its web site.<br /><br />The SpaceX web site says:<br /><br />"SpaceX Merlin Engine<br /><br /> Sea Level Thrust 77,000 lb<br /> Vacuum Thrust 92,000 lb<br /> Sea Level Isp 255s<br /> Vacuum Isp 304s<br /> Thrust to weight (fully accounted) 96<br /><br />SpaceX Kestrel Engine<br /><br /> Vacuum Thrust 7,000 lb<br /> Vacuum Isp 327s<br /> Thrust to weight 42 "<br /><br />As I recall, SpaceX found that Merlin was less <br />efficient than planned, so the company had <br />to increase thrust to offset the lower than planned <br />specific impulse.<br /><br /> - Ed Kyle
 
P

propforce

Guest
<font color="yellow">Your specific impulse numbers are higher than the numbers that SpaceX has posted on its web site. </font><br /><br />That very well could be. Those numbers are a couple of years old from SpaceX, I'd imagine they've updated them since. But I just check SpaceX website and they don't post any of these performance numbers anymore (or I just don't know where to look?).<br /><br />Now with the 'downgrade' of their engine Isp would make their payload claim even less likely. <div class="Discussion_UserSignature"> </div>
 
Status
Not open for further replies.

Latest posts