Something on another thread got me to wondering. What are the fundamental limits to throttling a kero-lox turbopump engine like the Merlin. I believe that pressure fed pintle engines can throttle much deeper. Is the 70% envelope on the Merlin-1D a fairly hard limit or just a design trade off? Anyone have any insight?
Someone mentioned the possibility of flow separation if the M1D is throttled too low. From my limited understanding that among other things depends on the shape and size of the nozzle extension. So maybe it could be possible to throttle a M1d more deeply with a different nozzle design. That might mean less efficiency in other parts of the flight envelope though.The Merlin 1D has the whole nozzle made of a structure that circulates RP-1 to cool it (it's called regeneratively cooled nozzle). It would look, from the extremely clean look of it and amazing T/W, that they have the inner and oute linings and use a corrugated separator in the middle. Then they would braze it. Like the NK-33. You can add an extension, that will have to be cooled differently (either ablatively, or film-cooled+radiatively, like the Vacuum version of the Merlin 1D). But you can't cut it easily. You'd have to modify or even get new tooling, and the pressure loss on the nozzle cooling circuit will be reduced, which I don't know how would impact the power balance on the engine. Thus, is not easy to reduce the thrust because of flow separation.
Merlin 1D | .. | Merlin 1D Vac | .. | Merlin 1D+ | .. | Merlin 1D+ Vac | |
Chamber Pressure, MPa | 9.7 | 9.7 | 10.8 | 10.8 | |||
Pressure exhaust, bar | 0.43 | 0.03 | 0.475 | 0.034 | |||
Throat area, m^2 | 0.042 | 0.042 | 0.042 | 0.042 | |||
Nozzle area, m^2 | 0.9 | 7.22 | 0.9 | 7.22 | |||
Nozzle diameter, m | 1.07 | 3.03 | 1.07 | 3.03 | |||
Mixture ratio | 2.34 | 2.36 | 2.36 | 2.375 | |||
Mass flow chamber, kg | 229 | 229 | 253.5 | 253.5 | |||
Mass flow gg, +% | 3.25 | 3.25 | 3.25 | 3.25 | |||
Net mass flow, kg | 236 | 236 | 262 | 262 | |||
SL Thrust, klbf | 147 klbf | 165 klbf | |||||
SL Isp, s | 282 s | 286 s | |||||
Vac Thrust, klbf | 167 klbf | 181 klbf | 185 klbf | 200 klbf | |||
Vac Isp, s | 320 s | 347 s | 321 s | 347 s |
Someone mentioned the possibility of flow separation if the M1D is throttled too low. From my limited understanding that among other things depends on the shape and size of the nozzle extension. So maybe it could be possible to throttle a M1d more deeply with a different nozzle design. That might mean less efficiency in other parts of the flight envelope though.The Merlin 1D has the whole nozzle made of a structure that circulates RP-1 to cool it (it's called regeneratively cooled nozzle). It would look, from the extremely clean look of it and amazing T/W, that they have the inner and oute linings and use a corrugated separator in the middle. Then they would braze it. Like the NK-33. You can add an extension, that will have to be cooled differently (either ablatively, or film-cooled+radiatively, like the Vacuum version of the Merlin 1D). But you can't cut it easily. You'd have to modify or even get new tooling, and the pressure loss on the nozzle cooling circuit will be reduced, which I don't know how would impact the power balance on the engine. Thus, is not easy to reduce the thrust because of flow separation.
Someone mentioned the possibility of flow separation if the M1D is throttled too low. From my limited understanding that among other things depends on the shape and size of the nozzle extension. So maybe it could be possible to throttle a M1d more deeply with a different nozzle design. That might mean less efficiency in other parts of the flight envelope though.The Merlin 1D has the whole nozzle made of a structure that circulates RP-1 to cool it (it's called regeneratively cooled nozzle). It would look, from the extremely clean look of it and amazing T/W, that they have the inner and oute linings and use a corrugated separator in the middle. Then they would braze it. Like the NK-33. You can add an extension, that will have to be cooled differently (either ablatively, or film-cooled+radiatively, like the Vacuum version of the Merlin 1D). But you can't cut it easily. You'd have to modify or even get new tooling, and the pressure loss on the nozzle cooling circuit will be reduced, which I don't know how would impact the power balance on the engine. Thus, is not easy to reduce the thrust because of flow separation.
The extension can be cut very easy. Water jet can cut it without a problem.
Two things about throttling:
-thanks to aero for this interesting link
http://www.ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20100033271_2010034521.pdf
the post
http://forum.nasaspaceflight.com/index.php?topic=33142.msg1130796#msg1130796
-in last GH1 flight, it's likely that throttling of merlin 1D went to about 60%.
While trying to understand Merlin 1D and in particular "Merlin 1D+"* in depth, I've iterated my calculations (http://tinyurl.com/merlin1d) a few times and have reached internal coherence and good balance with reality using the following characteristics/specs.
Merlin 1D .. Merlin 1D Vac .. Merlin 1D+ .. Merlin 1D+ Vac Chamber Pressure, MPa 9.7 9.7 10.8 10.8 Pressure exhaust, bar 0.43 0.03 0.475 0.034 Throat area, m^2 0.042 0.042 0.042 0.042 Nozzle area, m^2 0.9 7.22 0.9 7.22 Nozzle diameter, m 1.07 3.03 1.07 3.03 Mixture ratio 2.34 2.36 2.36 2.375 Mass flow chamber, kg 229 229 253.5 253.5 Mass flow gg, +% 3.25 3.25 3.25 3.25 Net mass flow, kg 236 236 262 262 SL Thrust, klbf 147 klbf 165 klbf SL Isp, s 282 s 286 s Vac Thrust, klbf 167 klbf 181 klbf 185 klbf 200 klbf Vac Isp, s 320 s 347 s 321 s 347 s
*) Merlin 1D+ is a nick-name for the engine Elon hinted during the SES8 teleconference; a 165 klbf SL "Merlin 1D"
Disclaimer; IANARS & many constants are read from analogue charts (http://www.braeunig.us/space/comb-OK.htm), so there are some wiggle room.
Nozzle diameter should probably be a few cm less.
I have a question for you. If SpaceX were to try the D+, what would they do for propellant? Mass flow goes from 236 to 262 so the burns would be shorter on both stages. Sure, the Isp is up a little but in the wash, it really doesn't make much difference to payload capability. With more powerful engines, will they extend the length of the tanks by 10 % (a guess) to keep the lift-off T/W the same or will they just drive the engines harder for little benefit?If they did not need the center engine for breaking, the extra thrust of the 1D+ should allow them to just get rid of it and save the weight. I don't think an 8 engine layout would work... hmmm...
And by the way, does anyone know the actual propellant mass for the stages of the existing F 9? Simulating the F 9 is a trick because each time the Isp changes, the lift-off mass changes, or the dry weight of the stages changes to compensate.Not sure how accurate this is, but there are some numbers here:
I know the mixture ratio so if I knew the length of the tank I could estimate the propellant mass probably more accurately than using the fuel flow rate and burn time since fuel flow rate is a mystery anyway.
If SpaceX were to try the D+, what would they do for propellant? Mass flow goes from 236 to 262 so the burns would be shorter on both stages. Sure, the Isp is up a little but in the wash, it really doesn't make much difference to payload capability.
Yes, it does make a difference. The reason is that with higher thrust you get a quicker ascent and reduced gravity losses. So the answer is that they could increase the maximum payload even without changing the amount of propellant.But wouldn't that also increase the g forces on the payload?
I have a question for you. If SpaceX were to try the D+, what would they do for propellant? Mass flow goes from 236 to 262 so the burns would be shorter on both stages. Sure, the Isp is up a little but in the wash, it really doesn't make much difference to payload capability. With more powerful engines, will they extend the length of the tanks by 10 % (a guess) to keep the lift-off T/W the same or will they just drive the engines harder for little benefit?
Yes, it does make a difference. The reason is that with higher thrust you get a quicker ascent and reduced gravity losses. So the answer is that they could increase the maximum payload even without changing the amount of propellant.But wouldn't that also increase the g forces on the payload?
Hi everyone,
this question maybe obvious for many of you - please excuse me if this is the case but I'm just in learning mode.
As well known, the Merlin 1D engine can be throttled down from 100% to 60%.
The vacuum thrust of a rocket engine is proportional to the exhaust velocity and to the mass flow rate of the gas exhaust: dm/dt * V_exh.
In vacuum and if also the gravity drag is negligible, when a Merlin 1D engine is throttled down, is dm/dt only changing, OR also V_exh (aka Isp_vacuum) is affected someway, with a consequent reduction in the engine efficiency in using the available propellant?
If the second hypothesis is right (as I would say by looking at the physical principles the engine is based on), does someone know a reasonable way to write the relation Thrust(%) = f(%,dm/dt, V_exh)?
Thanks for any hint,
Thrust is proportional to chamber pressure (to first order), so you're right. Good old Sutton.Hi everyone,
this question maybe obvious for many of you - please excuse me if this is the case but I'm just in learning mode.
As well known, the Merlin 1D engine can be throttled down from 100% to 60%.
The vacuum thrust of a rocket engine is proportional to the exhaust velocity and to the mass flow rate of the gas exhaust: dm/dt * V_exh.
In vacuum and if also the gravity drag is negligible, when a Merlin 1D engine is throttled down, is dm/dt only changing, OR also V_exh (aka Isp_vacuum) is affected someway, with a consequent reduction in the engine efficiency in using the available propellant?
If the second hypothesis is right (as I would say by looking at the physical principles the engine is based on), does someone know a reasonable way to write the relation Thrust(%) = f(%,dm/dt, V_exh)?
Thanks for any hint,
V_exh is not affected by thrust directly. There is a second order effect because you slightly lower the chamber pressure. This is my impression at least. I may be wrong.
Yes, it does make a difference. The reason is that with higher thrust you get a quicker ascent and reduced gravity losses. So the answer is that they could increase the maximum payload even without changing the amount of propellant.But wouldn't that also increase the g forces on the payload?
I have a question for you. If SpaceX were to try the D+, what would they do for propellant? Mass flow goes from 236 to 262 so the burns would be shorter on both stages. Sure, the Isp is up a little but in the wash, it really doesn't make much difference to payload capability. With more powerful engines, will they extend the length of the tanks by 10 % (a guess) to keep the lift-off T/W the same or will they just drive the engines harder for little benefit?
This additional loaded mass of 16,100 lb represented on average an 8.9-percent increase in onboard LO2 propellant. Test results also confirmed the presence of thermally stratified oxygen layers inside the tank. These layers varied in the vertical direction from 122 °R for the colder, denser fluid at the bottom to 166 °R for the warmer, less dense LO2 near the top outlet of the STA tank.
We could always subcool the kerosene...
I've heard of Rankine! :) It's like the imperial version of Kelvin. Also, if you're going to subcool the LOx... We could always subcool the kerosene...Flows like frozen molasses. Or were you making a joke? Please put a smiley. In the internet you can't say something ridiculous enough not to be taken at face value. ::)
V_exh is not affected by thrust directly. There is a second order effect because you slightly lower the chamber pressure. This is my impression at least. I may be wrong.
Do the first stage Merlin D engines, as built, throttle?
On photos of various Merlin 1Ds, I've noticed two different turbopump exhausts......(snip)
They are throttled down during a typical launch as the first stage approaches empty in order to limit acceleration.
Hmm, SpaceX Falcon9 main page reports this:They are throttled down during a typical launch as the first stage approaches empty in order to limit acceleration.
Do you have any reference for this?
thanks
I'm working on a 3D printable model of the F9 v1.1...
Hmm, SpaceX Falcon9 main page reports this:
"The first stage engines are gradually throttled near the end of first-stage flight to limit launch vehicle acceleration as the rocket’s mass decreases with the burning of fuel."
http://www.spacex.com/falcon9
Available data from first v1.1 flight on altitude and velocity fit quite well with a full throttle simulation with a reasonable trajectory.
Available data from first v1.1 flight on altitude and velocity fit quite well with a full throttle simulation with a reasonable trajectory.
Because MECO-1 happened earlier than on an expendable flight, roughly at the time it would normally start throttling down due to G limits. I would expect the SES-8 flight did involve throttling to some degree.
43 sec after throttling down started if the 2 min marker is taken for granted. However, a fit with the first 2 mins is enough to get some indication on the trajectory to be checked with available imaging.
Hi everyone,
this question maybe obvious for many of you - please excuse me if this is the case but I'm just in learning mode.
As well known, the Merlin 1D engine can be throttled down from 100% to 60%.
The vacuum thrust of a rocket engine is proportional to the exhaust velocity and to the mass flow rate of the gas exhaust: dm/dt * V_exh.
In vacuum and if also the gravity drag is negligible, when a Merlin 1D engine is throttled down, is dm/dt only changing, OR also V_exh (aka Isp_vacuum) is affected someway, with a consequent reduction in the engine efficiency in using the available propellant?
If the second hypothesis is right (as I would say by looking at the physical principles the engine is based on), does someone know a reasonable way to write the relation Thrust(%) = f(%,dm/dt, V_exh)?
Thanks for any hint,
[Changing thrust settings will change the chamber pressure, the chamber temperature, the characteristic velocity and correspondingly the specific impulse. And in the process also the mass flow.
It will leave key characteristics unchanged, such as the exit area, throat area, the oxidizer-to-fuel ratio and thus the ratio of specific heats , the vandenkerkhove function, the exit to chamber pressure ratio and therefore the thrust coefficient.
Since the throat area and the thrust coefficient stay constant, a change in thrust leads to a linearly corresponding change in chamber pressure and from this pressure ratio, a change in chamber temperature can be found using Poisson’s equation
This leads in turn to a change in characteristic velocity and specific impulse.
At 50% throttle a decrease of specific impulse from 340s to 317s is to be expected
On top you can expect some losses due to combustion efficiency that gets lower with lower throttle in a pintle engine.
Why that? Maybe just because progressively throttling is more efficient, at the end of the day, because it allows to follow much more carefully the thrust demand than just switching off engines?
IMO a big reason would be if another engine unexpectedly shutdown, as has happened with STS and Falcon9, then they just throttle up the remaining engines, with much shorter delay and avoiding the effects of restarting an engine in flight.
While trying to understand Merlin 1D and in particular "Merlin 1D+"* in depth, I've iterated my calculations (http://tinyurl.com/merlin1d) a few times and have reached internal coherence and good balance with reality using the following characteristics/specs.
Merlin 1D .. Merlin 1D Vac .. Merlin 1D+ .. Merlin 1D+ Vac Nozzle diameter, m 1.07 3.03 1.07 3.03
While trying to understand Merlin 1D and in particular "Merlin 1D+"* in depth, I've iterated my calculations (http://tinyurl.com/merlin1d) a few times and have reached internal coherence and good balance with reality using the following characteristics/specs.
Merlin 1D .. Merlin 1D Vac .. Merlin 1D+ .. Merlin 1D+ Vac Nozzle diameter, m 1.07 3.03 1.07 3.03
I've tried to draw the Falcon 9 first stage with your diameter, but I can't get it to fit. A circle of 8 engines with a diameter of 1,07 m each is going to have an outer diameter of at least 3,8 m as drawn in my CAD program, and that's with the engine nozzles touching each other. But if you look at images of the launch, the engines do not protrude outside the first stage diameter. And there's a gap between the engines.
If I limit the outer diameter of the 8 engines to 3,66 m and allow some spacing between them, the nozzle diameter is around 96,5 cm (my drawing was in 1:144).
I've also tried measuring the diameter from the second photo. Ignoring the distortion, the space between the center engine and the outer engines is 0.452 times the diameter of the center engine. So the total diameter of the ring of 8 engines is (3 + (2*0.452)) times the diameter of one engine nozzle. If the total ring diameter is 366 cm, then one engine must be 93,5 cm in diameter.
I might just not be looking in the right place, but has the Isp of Merlin-1D-VAC been confirmed or even reliably estimated yet?
I fully agree, nozzle diameter should be something like ~93-94 cm. I'm not fully there with the model. I have a similar issue with the RD-0162, not sure yet if those are related or there is a more trivial problem with the 1D model (such as adjusting the chamber pressure a bit, since that info might be old).I fully agree, nozzle diameter should be something like ~93-94 cm. I'm not fully there with the model. I have a similar issue with the RD-0162, not sure yet if those are related or there is a more trivial problem with the 1D model (such as adjusting the chamber pressure a bit, since that info might be old).
There is a picture up the nozzles of the Merlin C posted here.
http://forum.nasaspaceflight.com/index.php?topic=33598.msg1161838#msg1161838 (http://forum.nasaspaceflight.com/index.php?topic=33598.msg1161838#msg1161838)
It looks to me as though SpaceX has done something to suppress flow separation at sea level. I would call it a duel bell nozzle design except it looks like quadruple bell design would be more accurate.
I can see that the nozzles are uniformly divergent on the outside but on the inside it is different. What do you all think?
There is a picture up the nozzles of the Merlin C posted here.
http://forum.nasaspaceflight.com/index.php?topic=33598.msg1161838#msg1161838 (http://forum.nasaspaceflight.com/index.php?topic=33598.msg1161838#msg1161838)
It looks to me as though SpaceX has done something to suppress flow separation at sea level. I would call it a duel bell nozzle design except it looks like quadruple bell design would be more accurate.
I can see that the nozzles are uniformly divergent on the outside but on the inside it is different. What do you all think?
What are you drawing *that* conclusion from? That is just an engineering drawing with different sub-component coloured differently. I think you are reading more into those colours that you should.
There is a picture up the nozzles of the Merlin C posted here.
http://forum.nasaspaceflight.com/index.php?topic=33598.msg1161838#msg1161838 (http://forum.nasaspaceflight.com/index.php?topic=33598.msg1161838#msg1161838)
It looks to me as though SpaceX has done something to suppress flow separation at sea level. I would call it a duel bell nozzle design except it looks like quadruple bell design would be more accurate.
I can see that the nozzles are uniformly divergent on the outside but on the inside it is different. What do you all think?
What are you drawing *that* conclusion from? That is just an engineering drawing with different sub-component coloured differently. I think you are reading more into those colours that you should.
You're right. That's not a picture, my bad. Does anyone have a picture of the inside of a SpaceX rocket engine nozzle?
In particular I'm wondering; when an engine is throttled I have assumed the Mach number will stay constant, i.e., this number characterise a particular nozzle/throat/chamber design - and using this number one can calculate the exhaust pressure if given the chamber pressure (see spreadsheet for 90% throttle, for example).
In particular I'm wondering; when an engine is throttled I have assumed the Mach number will stay constant, i.e., this number characterise a particular nozzle/throat/chamber design - and using this number one can calculate the exhaust pressure if given the chamber pressure (see spreadsheet for 90% throttle, for example).
That assumption isn't quite correct, because the Mach number will change as the ratio of specific heats changes, and that ratio, of course, changes with chamber pressure.
kremlinology on the available Merlin data
A "staged combustion Merlin" would be a whole new engine. Not much would be in common.
Would be two completely different engines with just (some) construction methods in common. For starters, they would have to use an oxygen rich preburner. So, the would go from burning 3% of the RP-1 to 100% of the Oxigen, which means the preburner will have to process 90 times more mass, in a highly corrosive oxygen environment. Then, the turbine will have to pass 90 times more mass, so it will have a lot more power. Since the pump has more power, but have a lot more pressure losses, the pump outlet pressure should be 2 to 3 times as high. And then the injector will have to mix gas (O2) with liquid (RP-1), instead of both liquid. And the main combustion chamber would work at least at 70% more pressure. Not to mention that they wouldn't have the GG exhaust.A "staged combustion Merlin" would be a whole new engine. Not much would be in common.
Well, the rocket would be in common. Seems like a lot of commonality to me.
I know that the prop pumps are a very expensive part of the engine and they would need to change the turbine/pumps and routing, but would they need to change the feeds and engine attachments, nozzle? Would an SC cycle engine even fit into the space available?
I think an important consideration is their long term plan for retaining the Falcon 9 in their operational inventory which I don't know.
A "staged combustion Merlin" would be a whole new engine. Not much would be in common.
Well, the rocket would be in common. Seems like a lot of commonality to me.
Has anyone else considered that SpaceX may modify the Merlin D in order to achieve Elon's projected 15% increase in thrust (12 to 17 % depending on how you figure)?
I have been wondering if maybe, just maybe they might convert the Merlin D engine fueling system to a staged combustion cycle. That would give them the performance and it is not totally unreasonable now that they are gaining experience in staged combustion by working on the Raptor engine. Building a smaller staged combustion engine than the Raptor makes some sense, and building one for LOX/RP-1 seems logical since they are familiar with and have available everything that goes into building and testing LOX/RP-1 engines.
What do you think about a staged combustion Merlin D+?
Would it be possible/practical/acceptable to test a SC Merlin D + on the Falcon 9 while leaving 8 of the engines as is?
They may yet produce a M1E version, being engineering driven company I'd be surprised if they didn't tweek the M1D. Reducing the manufacturing cost of engine is one reason to make a new version, 3D printing may add significant cost savings if they are not already using it.Tom Mueller said a couple of years ago that there won't be a M1E. However the M1D is supposed to get a thrust upgrade to 165klbf at SL which has been discussed widely. SpaceX could have changed their mind over a possible M1E but I think it is still unlikely since SpaceX are moving on to their next gen. engine.
Nonsensical argument. We are talking about comparing engines alone - Merlin 1D and mythical Merlin 1SC - and about commonality between them.A "staged combustion Merlin" would be a whole new engine. Not much would be in common.Well, the rocket would be in common. Seems like a lot of commonality to me.
Merlin 1D may be too large to easily print using a metal 3d printer. I believe they rely on other techniques, like explosive forming.Maybe not the large parts, but I could see them print many of the smaller items.
While looking for Merlin data I came across this interesting little chalk talk by Tom Mueller on the Merlin 1C. Have your note pad ready, he gives numbers for the C model. I wonder how much they carry over to the D?
http://www.livescribe.com/cgi-bin/WebObjects/LDApp.woa/wa/MLSOverviewPage?sid=n7SMWV6PdG7r (http://www.livescribe.com/cgi-bin/WebObjects/LDApp.woa/wa/MLSOverviewPage?sid=n7SMWV6PdG7r)
The article where I found the link is here:
http://blog.linuxacademy.com/space/under-the-hood-with-the-spacex-merlin-engine/ (http://blog.linuxacademy.com/space/under-the-hood-with-the-spacex-merlin-engine/)
JMO but I expect the numbers carry over reasonably well to the D model, better than the numbers that we don't have on the Merlin D.
Ok, I searched. The article has been linked before but I didn't see anything about the fuel and LOX tank pressures or the turbo pump output pressures. Things I found interesting are:
Tank pressure = 50 psi for both tanks,
turbo pump output pressure, LOX 1400 psi, RP-1 1500 psi
gg combustion temperature 1033 deg. F
Combustion chamber temperature 6000 deg. F
Mdot 350 lbm/s
And the description of how the gas generator works.
Yes, it was for the Merlin 1 C, but interesting anyway.
Does anyone (or everyone but me) know what the area ratio is for the Merlin 1 D Vac engine. That is the upper stage engine?
I did not see this posted here. I apologize if I missed it somewhere, but it's taken by Steve J last thursday
http://www.flickr.com/photos/jurvetson/13160622333/in/photostream/
I did not see this posted here. I apologize if I missed it somewhere, but it's taken by Steve J last thursday
http://www.flickr.com/photos/jurvetson/13160622333/in/photostream/
What are the things in the background (under and to the right)?
What are the things in the background (under and to the right)?
It's a display they have hanging up. I want to say it's a 1A due to the thrust structure and GG exhaust, but I could be wrong.
It's a display they have hanging up. I want to say it's a 1A due to the thrust structure and GG exhaust, but I could be wrong.
Its not a 1A. Its up there purely to look pretty - its an MVacC thrust structure, chamber, nozzle, and nozzle extension with M1C (ie boost phase) gas generator and exhaust, likely just made up of spares they no longer have a use for other than decoration.
The video from the latest F9R-Dev1 test using a Merlin 1D also seems to demonstrate the new and improved throttling capabilities (seems like it achieves almost perfectly sustained hover for several seconds).
http://forum.nasaspaceflight.com/index.php?topic=33892.msg1184601#msg1184601 (http://forum.nasaspaceflight.com/index.php?topic=33892.msg1184601#msg1184601)
Far better than the Merlin 1C.
Far better than the Merlin 1C.
How is that far better, given the current statistically insignificant sample size?
Far better than the Merlin 1C.
How is that far better, given the current statistically insignificant sample size?
ugordon has a valid question for you, RDMM2081.
"Engineering is done with numbers. Everything else is just opinion."
Please make some attempt at calculating confidence intervals if you want to make an assertion we can agree to.
I'm not sure it would benefit. 3d printing generally makes poorer quality parts (ie grain structure, higher porosity and void content, rougher surface finish, poorer as-produced precision, etc) than traditional manufacturing and is severely size limited.
Nope. Not relevant to anything I said. Of course you can 3d print a gun out of metal, doesn't make it better quality (precision, porosity, etc) than traditional manufacturing.I'm not sure it would benefit. 3d printing generally makes poorer quality parts (ie grain structure, higher porosity and void content, rougher surface finish, poorer as-produced precision, etc) than traditional manufacturing and is severely size limited.
Technology may have already outstripped your generalization, late last year.
First 3D-Printed Metal Gun Fires 50 Rounds and Counting (http://mashable.com/2013/11/11/3d-printed-metal-gun/)
d. The multi engine scheme is just a marketing ploy. More than a 2/3's of a Falcon 9 flight time has the same consequence from a engine failure as the EELV's. It uses a single engine (which is also significantly different than the first stage engines) for second stage flight.
I'm not sure it would benefit. 3d printing generally makes poorer quality parts (ie grain structure, higher porosity and void content, rougher surface finish, poorer as-produced precision, etc) than traditional manufacturing and is severely size limited. It's nice because of complicated geometries being easier to fabricate and design-to-build cycle is shorter. A good fit for SuperDraco, not so much Merlin. No doubt some parts would make sense to 3d print, though.
Compared with a traditionally cast part, a printed valve body has superior strength, ductility, and fracture resistance, with a lower variability in materials properties. The MOV body was printed in less than two days, compared with a typical castings cycle measured in months.
The benefits of re engineering for 3D printing may not be worth costs and risk ie if it ain't broke don't fix it. With recent announcements of plans to make most flights have recoverable boosters.
Engine production rate should drop off in long run, even though flight rates increase.
well yeah, if you compare it to castings! Castings usually have inferior material properties to forged or rolled stock which is then machined or stamped or otherwise worked to shape.I'm not sure it would benefit. 3d printing generally makes poorer quality parts (ie grain structure, higher porosity and void content, rougher surface finish, poorer as-produced precision, etc) than traditional manufacturing and is severely size limited. It's nice because of complicated geometries being easier to fabricate and design-to-build cycle is shorter. A good fit for SuperDraco, not so much Merlin. No doubt some parts would make sense to 3d print, though.
Sounds like SpaceX have data to the contrary.QuoteCompared with a traditionally cast part, a printed valve body has superior strength, ductility, and fracture resistance, with a lower variability in materials properties. The MOV body was printed in less than two days, compared with a typical castings cycle measured in months.
M-1D was already a production optimized engine... why change it unless the benefit is significant.
I suspect that 3-D printing will play a substantial role from here out... not just boutique part making here and there. A majority 3-D printed power pack anyone?
SpaceX update, "SPACEX COMPLETES 100TH MERLIN 1D ENGINE":
http://www.spacex.com/news/2014/10/16/spacex-completes-100th-merlin-1d-engine
High-res photos attached
SpaceX update, "SPACEX COMPLETES 100TH MERLIN 1D ENGINE":
http://www.spacex.com/news/2014/10/16/spacex-completes-100th-merlin-1d-engine
High-res photos attached
This is where this shoud be posted.
Quite an achievement. Not clear to me why SpaceX need a production rate of 5 engines per week this soon. Surely pad and payload availability mean they don't need more than 200 engines next year?15 F9s and 3 Falcon Heavies and an in-flight abort F9. Though much of that will probably be for ramping up for 2016 (wouldn't be terribly surprised to see FH slip to 2016, but they'll need to start doing lots of tests in Texas starting in 2015). That's not completely unreasonable as a production goal. They doubled the number of launches from last year to this year and nearly doubled that from the year before. Wouldn't be impossible for them to hit 18 launches in 2016.
Quite an achievement. Not clear to me why SpaceX need a production rate of 5 engines per week this soon. Surely pad and payload availability mean they don't need more than 200 engines next year?15 F9s and 3 Falcon Heavies and an in-flight abort F9. Though much of that will probably be for ramping up for 2016 (wouldn't be terribly surprised to see FH slip to 2016, but they'll need to start doing lots of tests in Texas starting in 2015). That's not completely unreasonable as a production goal. They doubled the number of launches from last year to this year and nearly doubled that from the year before. Wouldn't be impossible for them to hit 18 launches in 2016.
Consider they'll probably do 7 launches this year and a pad abort for a total of nearly 8 all from one launch site. By 2016, they'll probably be using 3 launch sites and possibly 4. 3 times 7 is 21. 4 times 8 is 32.
That's a lot of launches, so the bottleneck will likely not be the pads. Texas may be the bottleneck. But anything less than 5 engines a week COULD be nearly a bottleneck by the end of 2015.
Interesting to see that the launch of SES-9 may use the higher thrust Merlin 1D engines.
Peter B. de Selding @pbdes
SES: We may skip spring SpaceX launch slot & wait till mid-year to let someone else be 1st using Falcon 9 main engine in full-thrust regime.
In light of the below quoted tweet, I think this thread needs a bump. Especially as I've seen a potential uprating of the M1D discussed in various places as a result.Interesting to see that the launch of SES-9 may use the higher thrust Merlin 1D engines.
Peter B. de Selding @pbdes
SES: We may skip spring SpaceX launch slot & wait till mid-year to let someone else be 1st using Falcon 9 main engine in full-thrust regime.
In light of the below quoted tweet, I think this thread needs a bump. Especially as I've seen a potential uprating of the M1D discussed in various places as a result.Interesting to see that the launch of SES-9 may use the higher thrust Merlin 1D engines.
Peter B. de Selding @pbdes
SES: We may skip spring SpaceX launch slot & wait till mid-year to let someone else be 1st using Falcon 9 main engine in full-thrust regime.
My main question is whether (as I've speculated before), this also requires sub-cooling of the propellants?
The extra thrust will be a requirement to lift more prop, but are the two joined at the hip? If not, liftoff T:W would be over 1.3.
Antares flew with subcooled lox, so this is an area where SpaceX don't need to blaze a trail. Still, I look forward to hearing more about this.
Cheers, Martin
In light of the below quoted tweet, I think this thread needs a bump. Especially as I've seen a potential uprating of the M1D discussed in various places as a result.Interesting to see that the launch of SES-9 may use the higher thrust Merlin 1D engines.
Peter B. de Selding @pbdes
SES: We may skip spring SpaceX launch slot & wait till mid-year to let someone else be 1st using Falcon 9 main engine in full-thrust regime.
My main question is whether (as I've speculated before), this also requires sub-cooling of the propellants?
The extra thrust will be a requirement to lift more prop, but are the two joined at the hip? If not, liftoff T:W would be over 1.3.
Antares flew with subcooled lox, so this is an area where SpaceX don't need to blaze a trail. Still, I look forward to hearing more about this.
Cheers, Martin
More thrust with same amount of fuel means less gravity losses. So it will give some benefit. But also adding more fuel (by either bigger tanks of subcooling the fuel) easily gives about 3 times bigger benefit than just increasing the thrust. My factor 3 is based on stettson-harrison method.
ULA did increase the trust of Delta IV by changing from RS-68 to RS-68A and not increase amount of fuel. This upgrade was needed to launch some big NRO payload about a year ago.
Makes me wonder if F9H will fly from the start with uprated Merlins and subcooled prop? Better to validate with final combination from the start.If the Heavy boosters are longer, they will weigh more. Perhaps full-thrust 1D was created for that reason.
If the Heavy boosters are longer, they will weigh more. Perhaps full-thrust 1D was created for that reason.
Makes me wonder if F9H will fly from the start with uprated Merlins and subcooled prop? Better to validate with final combination from the start.
If the Heavy boosters are longer, they will weigh more. Perhaps full-thrust 1D was created for that reason.Getting more thrust from the same engines helps you pretty much no matter what you're doing, no?
If the Heavy boosters are longer, they will weigh more. Perhaps full-thrust 1D was created for that reason.Getting more thrust from the same engines helps you pretty much no matter what you're doing, no?
Makes me wonder if F9H will fly from the start with uprated Merlins and subcooled prop? Better to validate with final combination from the start.
It is my strong belief that FH will never fly any other way. I suspect it's one of the things that has been holding up the maiden flight.
cheers, Martin
Makes me wonder if F9H will fly from the start with uprated Merlins and subcooled prop? Better to validate with final combination from the start.
It is my strong belief that FH will never fly any other way. I suspect it's one of the things that has been holding up the maiden flight.
cheers, Martin
Huh? What? As ugordan wrote, FH should still have a better T/W ratio at liftoff than F9v1.1...
Not necessarily; not if it costs you ISP, or results in exceeding structural limits on your stage or g-limits on your payload.Fair enough, it's helpful provided you don't intentionally blow up the rocket. I took that to be implicit.
Not necessarily; not if it costs you ISP, or results in exceeding structural limits on your stage or g-limits on your payload.Fair enough, it's helpful provided you don't intentionally blow up the rocket. I took that to be implicit.
It seems like the impact of lower ISP would depend on the the impact on GLOW.
In light of the below quoted tweet, I think this thread needs a bump. Especially as I've seen a potential uprating of the M1D discussed in various places as a result.Is there \ could there also be a lower thrust version of M1-D ?Interesting to see that the launch of SES-9 may use the higher thrust Merlin 1D engines.
Peter B. de Selding @pbdes
SES: We may skip spring SpaceX launch slot & wait till mid-year to let someone else be 1st using Falcon 9 main engine in full-thrust regime.
In light of the below quoted tweet, I think this thread needs a bump. Especially as I've seen a potential uprating of the M1D discussed in various places as a result.Is there \ could there also be a lower thrust version of M1-D ?Interesting to see that the launch of SES-9 may use the higher thrust Merlin 1D engines.
Peter B. de Selding @pbdes
SES: We may skip spring SpaceX launch slot & wait till mid-year to let someone else be 1st using Falcon 9 main engine in full-thrust regime.
I was thinking maybe, if they upgrade the thrust levels of the outer 8 engines, and seriously downgrade the thrust level of the inner engine, they can get the same total thrust at lift of but with better chances of landing the core.
This goes against basic rocket engineering logic of maximizing performance , but can possibly serve spaceX better.
In light of the below quoted tweet, I think this thread needs a bump. Especially as I've seen a potential uprating of the M1D discussed in various places as a result.Is there \ could there also be a lower thrust version of M1-D ?Interesting to see that the launch of SES-9 may use the higher thrust Merlin 1D engines.
Peter B. de Selding @pbdes
SES: We may skip spring SpaceX launch slot & wait till mid-year to let someone else be 1st using Falcon 9 main engine in full-thrust regime.
I was thinking maybe, if they upgrade the thrust levels of the outer 8 engines, and seriously downgrade the thrust level of the inner engine, they can get the same total thrust at lift of but with better chances of landing the core.
This goes against basic rocket engineering logic of maximizing performance , but can possibly serve spaceX better.
In light of the below quoted tweet, I think this thread needs a bump. Especially as I've seen a potential uprating of the M1D discussed in various places as a result.Is there \ could there also be a lower thrust version of M1-D ?Interesting to see that the launch of SES-9 may use the higher thrust Merlin 1D engines.
Peter B. de Selding @pbdes
SES: We may skip spring SpaceX launch slot & wait till mid-year to let someone else be 1st using Falcon 9 main engine in full-thrust regime.
I was thinking maybe, if they upgrade the thrust levels of the outer 8 engines, and seriously downgrade the thrust level of the inner engine, they can get the same total thrust at lift of but with better chances of landing the core.
This goes against basic rocket engineering logic of maximizing performance , but can possibly serve spaceX better.
What makes you think their chances of landing the core are too low now? They seem to be very close to success with surprisingly little learning curve so far.
That makes sense only if reusability turns out to be impossible with full size M1Ds due to T\W >>1 .
No, that doesn't make sense. Why throw away the advantage of a single engine mass production and create a second line? If all the engines are uprated (above 100% thrust) and somehow also raise their lowest thrust setting, they'll most likely eat that difference and instead tweak the final landing burn.
How much is that as a percentage of the RP-1? About 4%?
NASA did tests densifying lox, and found they could get an 8.9% increase in lox load at 122 Rankin. ;-)
Making either change without the other will change the Merlin mixture ratio.
The only guesses I can find for M1D mixture ratio are somewhat fuel rich, which will partly be the effect of the gas generator. It looks like sub-cooling both would make Merlin run less fuel rich, which would increase combustion temperatures. Could that be coped with by the greater regen cooling of the RP-1? Note, also, that the colder lox would absorb some of that extra energy.
This assumes they don't shift the bulkhead between the tanks for whatever is the ideal mixture ratio for M1D @ 112%. I guess that would only be a minor reconfiguration of the TEL? (And a fair bit of work to re-validate the stage structure.)
Cheers, Martin
Well, you know, I wish them well and good luck and I hope that won't be needed.
But I guess that if they can have a T/W closer to 1 then they will have better chances for successful landings.
We can hope that that won't be necessary but we haven't seen that work yet. All landings and hoverings must have been done with additional weight and at much lower velocities, IMO.
Now, let's suppose the booster's terminal velocity is 150m/s. That means for each clock cycle of this VersaLogic Falcon, a falling Falcon will have moved 0.0937 microns. The average thickness of a human hair is about a thousand times larger than this, and the Atom chip can execute up to two instructions per clock cycle (http://en.wikipedia.org/wiki/Intel_Atom_%28CPU%29#Bonnell_microarchitecture). And if you offload calculations to the GPU, the power available to you during those 0.0937 microns of travel is multiplied even further.
So while a booster falling through the atmosphere like a hypersonic spear looks overwhelmingly daunting to our plodding human perceptions, it is quite a leisurely pace as far as the guidance computer is concerned.
...
(And this is my 500th post on NSF. Yay! ;D)
So while a booster falling through the atmosphere like a hypersonic spear looks overwhelmingly daunting to our plodding human perceptions, it is quite a leisurely pace as far as the guidance computer is concerned. And while the work that went in to devising the software for guiding a booster from near space down to a tiny platform in the middle of the ocean is a breathtaking work of mathematical and software art, it is just another day at the office for the computer systems tasked with following the instructions of that software.The bottleneck is not the computer guidance calculation speed, it is the sensors integration time, which can be millions of times slower. Worst even are the actuators, in this case grid fins, engine gymbal, engine thruster and gas thrusters. These act very slowly to compensate.
Forgive me if this has been discussed. I searched but didn't find my question.
With an updated 1D+ and its increased thrust, when in the flight profile would the engines throttle back to limit G loads?
My idea that the engines are working harder, but that the burn time of the first stage would be shortened. So there would be a partial trade off regarding engine wear.
I thought as well that nearly 1.7 million pounds of thrust that the booster wouldn't be at full throttle the whole time. Again reducing the impacts of the additional thrust.
So while a booster falling through the atmosphere like a hypersonic spear looks overwhelmingly daunting to our plodding human perceptions, it is quite a leisurely pace as far as the guidance computer is concerned. And while the work that went in to devising the software for guiding a booster from near space down to a tiny platform in the middle of the ocean is a breathtaking work of mathematical and software art, it is just another day at the office for the computer systems tasked with following the instructions of that software.The bottleneck is not the computer guidance calculation speed, it is the sensors integration time, which can be millions of times slower. Worst even are the actuators, in this case grid fins, engine gymbal, engine thruster and gas thrusters. These act very slowly to compensate.
So sensor + actuators speed, accuracy and precision eventually determines the landing accuracy, unrelated to CPU speed.
Still, I agree with the general principle you described, I just would not go that far on boasting CPU performance when that clearly is not the bottleneck.
So while a booster falling through the atmosphere like a hypersonic spear looks overwhelmingly daunting to our plodding human perceptions, it is quite a leisurely pace as far as the guidance computer is concerned. And while the work that went in to devising the software for guiding a booster from near space down to a tiny platform in the middle of the ocean is a breathtaking work of mathematical and software art, it is just another day at the office for the computer systems tasked with following the instructions of that software.The bottleneck is not the computer guidance calculation speed, it is the sensors integration time, which can be millions of times slower. Worst even are the actuators, in this case grid fins, engine gymbal, engine thruster and gas thrusters. These act very slowly to compensate.
So sensor + actuators speed, accuracy and precision eventually determines the landing accuracy, unrelated to CPU speed.
Still, I agree with the general principle you described, I just would not go that far on boasting CPU performance when that clearly is not the bottleneck.
Would this also increase Dragon payload mass to ISS or is Dragon payload constrained by other factors?
Well, you know, I wish them well and good luck and I hope that won't be needed.
But I guess that if they can have a T/W closer to 1 then they will have better chances for successful landings.
We can hope that that won't be necessary but we haven't seen that work yet. All landings and hoverings must have been done with additional weight and at much lower velocities, IMO.
It's important to remember that the computer controlling the landing doesn't really care what the T/W ratio is, per se. It's just another item in a list of factors which govern the software's decisions about the landing operation.
Take, for example, the VersaLogic "Falcon" embedded system:
(http://www.versalogic.com/Products/Photos/EPU-2610-A-DS.jpg) (http://www.versalogic.com/products/ds.asp?productid=230)
It has a 1.6-gigahertz Intel Atom E6x0T processor (http://www.intel.com/content/www/us/en/intelligent-systems/medical-applications/atom-e6xx-series-datasheet.html#) and gigabit Ethernet, with up to two gigabytes of RAM, as well as integrated 3D graphics capabilities.
Now, let's suppose the booster's terminal velocity is 150m/s. That means for each clock cycle of this VersaLogic Falcon, a falling Falcon will have moved 0.0937 microns. The average thickness of a human hair is about a thousand times larger than this, and the Atom chip can execute up to two instructions per clock cycle (http://en.wikipedia.org/wiki/Intel_Atom_%28CPU%29#Bonnell_microarchitecture). And if you offload calculations to the GPU, the power available to you during those 0.0937 microns of travel is multiplied even further.
So while a booster falling through the atmosphere like a hypersonic spear looks overwhelmingly daunting to our plodding human perceptions, it is quite a leisurely pace as far as the guidance computer is concerned. And while the work that went in to devising the software for guiding a booster from near space down to a tiny platform in the middle of the ocean is a breathtaking work of mathematical and software art, it is just another day at the office for the computer systems tasked with following the instructions of that software.
The computer is not a human helicopter pilot who needs an eternity - in computer time - to figure out the landing maneuver, and it need not concern itself with the physical comfort of itself or any passengers. If the T/W is higher, it will just pull higher G's by triggering the landing burn later, and then twiddle its thumbs for a bit while it waits patiently for the Merlin to spin up, the same as before. (At launch, ignition is begun at T-3 seconds (http://www.spaceflight101.com/falcon-9-generic-countdown-timeline.html) - that's 4.8 billion clock cycles for a VersaLogic Falcon, and up to 9.6 billion instructions.) This persistent notion that the ability to hover, or a lower T/W, will make a difference to the landing operations arises from thinking about the maneuver in human terms, rather than machine terms.
And on the contrary, we have seen it work repeatedly, on each of the multiple occasions that the booster has made a soft touchdown in the Atlantic. The fact that it fell over and popped like a balloon afterward is irrelevant to the landing sequence. If there had been a barge underneath CRS-3 or OG-2, they'd have a booster on display outside Hawthorne today.
(And this is my 500th post on NSF. Yay! ;D)
So while a booster falling through the atmosphere like a hypersonic spear looks overwhelmingly daunting to our plodding human perceptions, it is quite a leisurely pace as far as the guidance computer is concerned. And while the work that went in to devising the software for guiding a booster from near space down to a tiny platform in the middle of the ocean is a breathtaking work of mathematical and software art, it is just another day at the office for the computer systems tasked with following the instructions of that software.The bottleneck is not the computer guidance calculation speed, it is the sensors integration time, which can be millions of times slower. Worst even are the actuators, in this case grid fins, engine gymbal, engine thruster and gas thrusters. These act very slowly to compensate.
So sensor + actuators speed, accuracy and precision eventually determines the landing accuracy, unrelated to CPU speed.
Still, I agree with the general principle you described, I just would not go that far on boasting CPU performance when that clearly is not the bottleneck.
Clearly landing with T/W >1 is not impossible, hovering is for cowards, computers are awesome.
With that logic only, nothing stops from landing a single engine core stage like atlas, falcon 1 or falcon 9 2nd stage.
Do you think that that is possible with good enogh computers, actuators and sensors?
Really, do you think its possible?
OTOH, using 9 engines gave spacex the ability to throttle deep enough and therefore to approach the ground slow enough and hopfully land.
So there must be some golden zone for landing T/W, depending on the electromechanic abilities of the LV,
And probably a smaller engine won't be necessary.
It's a lot simpler than that. If you try to land an empty stage with 9 engines (or land an Atlas, or any of your other examples) then you'll exceed the allowed acceleration limit of the stage, and it will fail mechanically. Having a 1:9 throttle really helps when the stage empty weight is only a few % of the takeoff weight.
It is true that with higher T/W, the avionics has to "work harder", but as has been pointed out above, it's really not an issue. You can land at T/W=2 (a=1g), or T/W=3 (a=2g) and for the computers and actuators, this is slow-mo.
So while a booster falling through the atmosphere like a hypersonic spear looks overwhelmingly daunting to our plodding human perceptions, it is quite a leisurely pace as far as the guidance computer is concerned. And while the work that went in to devising the software for guiding a booster from near space down to a tiny platform in the middle of the ocean is a breathtaking work of mathematical and software art, it is just another day at the office for the computer systems tasked with following the instructions of that software.The bottleneck is not the computer guidance calculation speed, it is the sensors integration time, which can be millions of times slower. Worst even are the actuators, in this case grid fins, engine gymbal, engine thruster and gas thrusters. These act very slowly to compensate.
So sensor + actuators speed, accuracy and precision eventually determines the landing accuracy, unrelated to CPU speed.
Still, I agree with the general principle you described, I just would not go that far on boasting CPU performance when that clearly is not the bottleneck.
This. And as of right now ,do we know if the throttle response of the engines is fast, accurate and repeatable? That's the bit that I think is the slowest and most indeterminate. All the other stuff can be accurately timed (even if slow). If it takes 3s seconds +- 0.5s for the engine to respond to a throttle setting, you can use a 6502 processor and still be fast enough.
It's a lot simpler than that. If you try to land an empty stage with 9 engines (or land an Atlas, or any of your other examples) then you'll exceed the allowed acceleration limit of the stage, and it will fail mechanically. Having a 1:9 throttle really helps when the stage empty weight is only a few % of the takeoff weight.
It is true that with higher T/W, the avionics has to "work harder", but as has been pointed out above, it's really not an issue. You can land at T/W=2 (a=1g), or T/W=3 (a=2g) and for the computers and actuators, this is slow-mo.
I completely agree with your point, but I have a minor quibble about how you present it here. While T/W=2 does give 1g acceleration with respect to the surface of the Earth, the load on the structure of the rocket is 2g, and at T/W=3 the load is 3g. When T/W=1, the rocket hovers and the load is 1g. Since we're talking about the structural capacity of the rocket, I think that's the more useful number.
It seems like there is a lot of unfounded speculation that a uprated Merlin 1D is physically different than the model currently flying. I don't think that is the case. Isn't it just a case of running more propellant through the engine? Nothing more than some simple code changes to the engine controller.
Now I do imagine that there are just about 100% for sure some minor tweaks to the M1D since it started flying, but I just don't think that there are any substantial changes for this soon to fly uprated engine.
Is there any info on what is being considered. In the past there have been numbers up to 117%. I have a feeling that they will be starting at 105% or maybe even 110%.
“You know SpaceX is introducing into their manifest ... a modification of the current engine, with about a 20 percent increase in thrust," said Martin Halliwell, SES’s chief technical officer. "We’re making a decision internally as to whether we want to be the first to fly it.”
SES: We may skip spring SpaceX launch slot & wait till mid-year to let someone else be 1st using Falcon 9 main engine in full-thrust regime.
It seems like there is a lot of unfounded speculation that a uprated Merlin 1D is physically different than the model currently flying. I don't think that is the case. Isn't it just a case of running more propellant through the engine? Nothing more than some simple code changes to the engine controller.
Now I do imagine that there are just about 100% for sure some minor tweaks to the M1D since it started flying, but I just don't think that there are any substantial changes for this soon to fly uprated engine.
Is there any info on what is being considered. In the past there have been numbers up to 117%. I have a feeling that they will be starting at 105% or maybe even 110%.
With respect to the future potential for the rocket, I do think we've got.. I'm really happy with this rocket design. It's an incredibly capable vehicle. It's actually one of the biggest rockets in the world, it's worth noting, at about 1.3 million pounds of thrust, and we're only actually operating the engines at about 85% of their potential. Down the road, in future missions, we anticipate being able to crank them up to their full thrust capability, which would give about 165,000 pounds of sea-level thrust per engine. Anyway, it really is something that is, I think, going to serve really well for the commercial launch market, for government satellites and for Dragon, both crew and cargo. I believe its inherent reliability potential is better than any other rocket in the world. It will be up to us to show that it lives up to that reliability potential.
165klb is 112% of today's published SL thrust level (147 klb).
If it was really 117%, that would be 173 klb, and it seems unlikely Elon would quote 165 klb if the true number was 173 klb.
I think these are the only hard numbers we have on the subject:-
165klb is 112% of today's published SL thrust level (147 klb).
If it was really 117%, that would be 173 klb, and it seems unlikely Elon would quote 165 klb if the true number was 173 klb.
I think these are the only hard numbers we have on the subject:-
Those aren't the only hard numbers we have. Those are numbers from 2013. In 2015, SES said a 20% increase in thrust, and they are a customer who is considering delaying a launch because of the changes, so they should know exactly what the real numbers are.
In 2013, Elon was likely less sure of what the numbers would be, and it's possible he was quoting the lower end of the range they expected. If it was really a 12% increase, I really doubt SES would be going around saying 20%.
If it was really a 12% increase, I really doubt SES would be going around saying 20%.
Is it possible that today's F9 actually has a sub-optimal T:W, so that the upgraded version is optimal?
cheers, Martin
It's my suspicion that the higher thrust is dependent on propellant sub-cooling.
Granted what I'm about to say doesn't hold for whole vehicle issues like prop loading or a higher Max Q.
But if customers are worried about higher stresses on the engines at higher power, and no one wants to go first....
Well there are 9 engines, and engine out capability, so do all 9 engines need to be run at the same power level? If two engines opposite each other were run at the new full thrust while the others stayed put, then if an engine did happen to fail, the payload would still make orbit by sacrificing the fuel set aside for landing. If everything works as planned, they have a validation and customers may feel better about going first on a full 9 engine upgrade.
I wonder if they're considering that?
I'm no turbo pump engineer, but I will also speculate that increased viscosity also reduces cavitation margins unless tank pressure is increased.
My guess (just that) is that SpaceX was planning on running the first uprated launch on a Dragon mission. With the loss of Antares late last year, it wouldn't surprise me if NASA expressed discomfort with such a plan and asked SpaceX to do it with someone else. SpaceX is now looking for someone else to go first and SES would rather have it be someone else, too...
I doubt they will run a mission with a subset of engines operating in the higher thrust regime, but I suppose if nobody will agree to go first, and that assuages anyone's concerns, it is possible.
If push comes to shove, they could put a mass simulator on top of the first stage and fly a suborbital flight, with an orbital mission profile, to test the new engines. The mass sim would stage at the proper time, in the usual manner, but simply fall into the ocean. The first stage would make a "routine" landing on Just Follow the Instructions.
If push comes to shove, they could put a mass simulator on top of the first stage and fly a suborbital flight, with an orbital mission profile, to test the new engines. The mass sim would stage at the proper time, in the usual manner, but simply fall into the ocean. The first stage would make a "routine" landing on Just Follow the Instructions.
Hm... interesting idea... and I think you just described the DragonV2 in-flight abort test...
<snip>
The first stage would make a "routine" landing on Just Follow the Instructions.
<snip>
<snip>
The first stage would make a "routine" landing on Just Follow the Instructions.
<snip>
A small nit... the autonomous spaceport drone ship is named Just Read the Instructions nothing was actually said about following instructions.
;)
Is it possible that today's F9 actually has a sub-optimal T:W, so that the upgraded version is optimal?
cheers, Martin
That's my bet, I believe (without any proof) that this "higher thrust" was planned from the beginning for the 1D and has been throttled down for first set of flights just to be conservative and also waiting for tweeks like sub cooled fuel.
Sub cooling also causes viscosity increase as well as density increase. The increased viscosity means more pressure loss in the piping. Not too much of an issue for the lox but it is more significant for the rp1 through all the small cooling channels. I'll venture a guess that sub cooling means that more pumping power is needed.
Do we know if the new Merlin 1D+ Engine would be mounted in the future on all Falcon 9, or will there be two versions falcon falcon 1D and 1D+ ?
I agree. My understanding was that they over built the engine to have extra margins. Now that they have lots of firings and data on the stresses they are releasing those margins because they are more comfortable.Do we know if the new Merlin 1D+ Engine would be mounted in the future on all Falcon 9, or will there be two versions falcon falcon 1D and 1D+ ?
As I understand it from what I have read, Merlin 1D+ is the Merlin 1D with 165 klb instead of 147 klb of thrust. It is not a new engine and it is debated of how much change has actually been made on the Merlin 1D to get the new thrust.
I agree. My understanding was that they over built the engine to have extra margins. Now that they have lots of firings and data on the stresses they are releasing those margins because they are more comfortable.Do we know if the new Merlin 1D+ Engine would be mounted in the future on all Falcon 9, or will there be two versions falcon falcon 1D and 1D+ ?
As I understand it from what I have read, Merlin 1D+ is the Merlin 1D with 165 klb instead of 147 klb of thrust. It is not a new engine and it is debated of how much change has actually been made on the Merlin 1D to get the new thrust.
The increased viscosity means more pressure loss in the piping.
As far as everything else, your moving such a huge amount of fuel in such a short time, you'd have to be inputing a crazy amount of heat from somewhere to warm the fuel load up a single degree.
That's rather high, I think.I agree. My understanding was that they over built the engine to have extra margins. Now that they have lots of firings and data on the stresses they are releasing those margins because they are more comfortable.Do we know if the new Merlin 1D+ Engine would be mounted in the future on all Falcon 9, or will there be two versions falcon falcon 1D and 1D+ ?
As I understand it from what I have read, Merlin 1D+ is the Merlin 1D with 165 klb instead of 147 klb of thrust. It is not a new engine and it is debated of how much change has actually been made on the Merlin 1D to get the new thrust.
The real question is there an ISP increase. A 5 point increase would result in a 15% or more increase in payload capability.
QuoteThe increased viscosity means more pressure loss in the piping.
Forgive this plebeian question. At first glance, I would think that increased viscosity would decrease flow rate. I can't see how this becomes a decrease in pressure.
I agree. My understanding was that they over built the engine to have extra margins. Now that they have lots of firings and data on the stresses they are releasing those margins because they are more comfortable.Do we know if the new Merlin 1D+ Engine would be mounted in the future on all Falcon 9, or will there be two versions falcon falcon 1D and 1D+ ?
As I understand it from what I have read, Merlin 1D+ is the Merlin 1D with 165 klb instead of 147 klb of thrust. It is not a new engine and it is debated of how much change has actually been made on the Merlin 1D to get the new thrust.
The real question is there an ISP increase. A 5 point increase would result in a 15% or more increase in payload capability.
The real question is there an ISP increase. A 5 point increase would result in a 15% or more increase in payload capability.
If the first stage averages 300s Isp, and the second stage averages 340s Isp, theoretical dV at 13,150kg payload is 10,084m/s. Take that number, and find the payload to match it at 305s Isp for the first stage and 345s Isp for the second, and you get 14,077kg, a 7.5% increase to LEO. Applying the same analysis to GTO using quoted payload figures of 4850kg, raises payload to 5423kg, an 11.8% increase to GTO.
If you also add the overall prop increase due to super cooling of ~1.5% to the 1st stage ISP increase of 2 points at sea level only which would be only an average of just 1 for atmospheric a vac operation of 1st stage, I wonder what you get for a payload increase with your model.If the first stage averages 300s Isp, and the second stage averages 340s Isp, theoretical dV at 13,150kg payload is 10,084m/s. Take that number, and find the payload to match it at 305s Isp for the first stage and 345s Isp for the second, and you get 14,077kg, a 7.5% increase to LEO. Applying the same analysis to GTO using quoted payload figures of 4850kg, raises payload to 5423kg, an 11.8% increase to GTO.
If the table on the first page can be trusted, the average increase in Isp will only be about 2s as there isn't much increase in vacuum Isp. I suspect the Isp increase comes from the higher chamber pressure resulting in the exhaust being less under-expanded at sea level. There was another post later on the first page where someone saw a similar relationship when they modeled throttling at sea level. There is no increase in Isp for the M1D-vac. So you're looking at 300 vs 302 for the first stage and no change for second stage.
If you also add the overall prop increase due to super cooling of ~1.5% to the 1st stage ISP increase of 2 points at sea level only which would be only an average of just 1 for atmospheric a vac operation of 1st stage, I wonder what you get for a payload increase with your model.
The second item is that if there is a sea level ISP increase then there should also be about the same or proportionately large vac ISP increase. This due to the fact that ISP is dependent on TC pressure. Higher pressure higher ISP. This implies that ISP will increase for vac operation as well.
I calculated the vac ISP from the thrust and flow rate which gives an ISP of 358s (200000/(2.2*253.5)) a gain of 11. Something is not correct in his numbers.If you also add the overall prop increase due to super cooling of ~1.5% to the 1st stage ISP increase of 2 points at sea level only which would be only an average of just 1 for atmospheric a vac operation of 1st stage, I wonder what you get for a payload increase with your model.
The second item is that if there is a sea level ISP increase then there should also be about the same or proportionately large vac ISP increase. This due to the fact that ISP is dependent on TC pressure. Higher pressure higher ISP. This implies that ISP will increase for vac operation as well.
The 2s increase is alread averaging 4s at sea level --> 1s vacuum from malu5531's post (http://forum.nasaspaceflight.com/index.php?topic=32983.msg1131598#msg1131598 (http://forum.nasaspaceflight.com/index.php?topic=32983.msg1131598#msg1131598)).
The other post I was referring to was this one from deltaV: http://forum.nasaspaceflight.com/index.php?topic=32983.msg1136160#msg1136160 (http://forum.nasaspaceflight.com/index.php?topic=32983.msg1136160#msg1136160)
He modeled 100% vs 50% throttling for the M1D. It seems like the same relationships should apply for 112% vs 100%, i.e. much larger Isp increase at sea level than vacuum.
Depends on where you measure pressure. It's like the idiots solution to low engine oil pressure, which is to add thicker oil. That might increase pressure at the pump and sensor, but it decreases oil flow and pressure at the outlets or the injectors in the rocket engine case.QuoteThe increased viscosity means more pressure loss in the piping.
Forgive this plebeian question. At first glance, I would think that increased viscosity would decrease flow rate. I can't see how this becomes a decrease in pressure.
Each engine is shielded from the adjacent engines by Kevlar blankets.
Do we know if the new Merlin 1D+ Engine would be mounted in the future on all Falcon 9, or will there be two versions falcon falcon 1D and 1D+ ?
Do we know if the new Merlin 1D+ Engine would be mounted in the future on all Falcon 9, or will there be two versions falcon falcon 1D and 1D+ ?
maybe 3 vers. 9FR being the 3rd.
This brings into question the whole expression of disgust by SX management regarding certification. This thread becomes meaningful.... http://forum.nasaspaceflight.com/index.php?topic=34818.0. The USAF can't be expected to certify a moving target can they?
ULA makes small tweaks to their LVs quite frequently. Think the M1D thrust upgrade as RS-68 vs RS-68A, just to name an example.Do we know if the new Merlin 1D+ Engine would be mounted in the future on all Falcon 9, or will there be two versions falcon falcon 1D and 1D+ ?
maybe 3 vers. 9FR being the 3rd.
This brings into question the whole expression of disgust by SX management regarding certification. This thread becomes meaningful.... http://forum.nasaspaceflight.com/index.php?topic=34818.0. The USAF can't be expected to certify a moving target can they?
ULA makes small tweaks to their LVs quite frequently. Think the M1D thrust upgrade as RS-68 vs RS-68A, just to name an example.Do we know if the new Merlin 1D+ Engine would be mounted in the future on all Falcon 9, or will there be two versions falcon falcon 1D and 1D+ ?
maybe 3 vers. 9FR being the 3rd.
This brings into question the whole expression of disgust by SX management regarding certification. This thread becomes meaningful.... http://forum.nasaspaceflight.com/index.php?topic=34818.0. The USAF can't be expected to certify a moving target can they?
How big of a change needs to be made before it is considered a new rocket and require re-certification? Could ULA argue that their BE-4 powered rocket is just an evolution of the Atlas V a not require re-certification?
"While the new system will require certification, we're confident this new system will enable us to further reduce costs while continuing to provide the most affordable and reliable launch services to our customers," ULA spokeswoman Jessica Rye wrote in an email.
Space News: SES Rethinking Being First To Fly a Full-throttle Falcon 9 (http://spacenews.com/ses-rethinking-being-first-to-fly-on-a-full-throttle-falcon-9/)Quote“You know SpaceX is introducing into their manifest ... a modification of the current engine, with about a 20 percent increase in thrust," said Martin Halliwell, SES’s chief technical officer. "We’re making a decision internally as to whether we want to be the first to fly it.”
120%? That's 175klb. I don't think we've seen anything over 112% before for the upgrade have we? This is along with densification would allow enough margin for boost back even with com SATs? How they get to 53t Without cross feed for FH?
"at about 1.3 million pounds of thrust, and we're only actually operating the engines at about 85% of their potential. Down the road, in future missions, we anticipate being able to crank them up to their full thrust capability, which would give about 165,000 pounds of sea-level thrust per engine."
Space News: SES Rethinking Being First To Fly a Full-throttle Falcon 9 (http://spacenews.com/ses-rethinking-being-first-to-fly-on-a-full-throttle-falcon-9/)Quote“You know SpaceX is introducing into their manifest ... a modification of the current engine, with about a 20 percent increase in thrust," said Martin Halliwell, SES’s chief technical officer. "We’re making a decision internally as to whether we want to be the first to fly it.”
120%? That's 175klb. I don't think we've seen anything over 112% before for the upgrade have we? This is along with densification would allow enough margin for boost back even with com SATs? How they get to 53t Without cross feed for FH?
The various percentages mentioned for this upgrade/uprate are all mixed. If you want to know where SES's quoted 20% comes from, take Elon's statement that previous launches have all been running the M1D @ 85%, a 20% increase over that is actually 102% (0.85x + 0.2*0.85x=1.02x) or close enough to 100% to make no difference in PR-speak. The confusion is whether the "20%" you're talking about is of the nominal 100% or just of the originally quoted 85%.
The bigger problem is that none of the calculated numbers are very consistent. Some are based on math using quoted percentages, some are based on using quoted thrust levels, some are a mixture of the two, etc. The best thing to do is just to read what Elon actually said about the M1D+ and ignore all the percentages:Quote from: Elon Musk"at about 1.3 million pounds of thrust, and we're only actually operating the engines at about 85% of their potential. Down the road, in future missions, we anticipate being able to crank them up to their full thrust capability, which would give about 165,000 pounds of sea-level thrust per engine."
165klb of SL thrust per M1D+.
Interesting page claiming to show Merlin progression
http://www.b14643.de/Spacerockets_2/United_States_2/Falcon-IX/Merlin/index.htm
165klb @10.8MPa and SL ISP of 286
"About 85%" from what I can tell implies no more precision than 82.5-87.5%.
Interesting page claiming to show Merlin progression
http://www.b14643.de/Spacerockets_2/United_States_2/Falcon-IX/Merlin/index.htm
165klb @10.8MPa and SL ISP of 286
Interesting page claiming to show Merlin progression
http://www.b14643.de/Spacerockets_2/United_States_2/Falcon-IX/Merlin/index.htm
165klb @10.8MPa and SL ISP of 286
Neat - and M1D has now doubled the M1A thrust, while staying the same size. (Actually is effectively more compact, since the thrust structure and gimbaling mechanism is much more compact)
Interesting page claiming to show Merlin progression
http://www.b14643.de/Spacerockets_2/United_States_2/Falcon-IX/Merlin/index.htm
165klb @10.8MPa and SL ISP of 286
Neat - and M1D has now doubled the M1A thrust, while staying the same size. (Actually is effectively more compact, since the thrust structure and gimbaling mechanism is much more compact)
If this can be taken at face value it is really interesting, because it's not been clear to me whether the extra thrust came from opening the throat or increasing chamber pressure and Ae.
Yeah, greater chamber pressure is more likely than changing throat area to me.
Interesting page claiming to show Merlin progression
http://www.b14643.de/Spacerockets_2/United_States_2/Falcon-IX/Merlin/index.htm
165klb @10.8MPa and SL ISP of 286
Neat - and M1D has now doubled the M1A thrust, while staying the same size. (Actually is effectively more compact, since the thrust structure and gimbaling mechanism is much more compact)
If this can be taken at face value it is really interesting, because it's not been clear to me whether the extra thrust came from opening the throat or increasing chamber pressure and Ae.
The extra thrust can be gained by upping chamber pressure and mdot, so there is no need to change physical dimensions of the throat and nozzle.
http://forum.nasaspaceflight.com/index.php?topic=32983.msg1131598#msg1131598
Should SpaceX call this higher-thrust engine the Merlin 1E?
Should SpaceX call this higher-thrust engine the Merlin 1E?
How about just M1D+, as Norbert Brügge had it in his Merlin table (http://www.b14643.de/Spacerockets_2/United_States_2/Falcon-IX/Merlin/index.htm)? More than a D, less than an E. Just for our discussion purposes.
Should SpaceX call this higher-thrust engine the Merlin 1E?
I would not think so. It appears that the higher thrust was intended from the beginning. They are setting the 'standard' throttle from ~85% of planned thrust to ~100%. The upgrade might be as simple as a change to the engine firmware.
How about just M1D+, as Norbert Brügge had it in his Merlin table (http://www.b14643.de/Spacerockets_2/United_States_2/Falcon-IX/Merlin/index.htm)? More than a D, less than an E. Just for our discussion purposes.
I would not rely on designations from the Brügge website. He is unfortunately very creative in inventing his own designations ("Voskhod-U", "PSLV-G+" etc.).
I would not introduce a new engine designation, unless there is confirmation of it. AFAIK, the uprated M-1D is pretty identical to the current M-1D, just certified for higher thrust.
I'm curious and don't have time to run numbers on volume lost due to additional
Tank shrinkage if they do sub cool? It would be a function of delta Temp * thermal expansion coefficient to the 3rd power.. Correct?
I'm curious and don't have time to run numbers on volume lost due to additional
Tank shrinkage if they do sub cool? It would be a function of delta Temp * thermal expansion coefficient to the 3rd power.. Correct?
I think not, but almost.
It should be Delta-Temp*Coeff_of_Thermal_exp times 3
The volume of a cylinder is Pi*R^2*L*(1+dT*CTE)^3.
If the last term is expanded you get 1+ 3*dT*CTE +3*(dt*CTE)^2 +(dt*CTE)^3
The terms with higher powers of dt*CTE are insignificant.
The derivative of the constant term is zero.
The derivative of the second term with respect to temperature is 3*Pi*R^2*L*CTE. To get the volume change multiply by dt.
Should SpaceX call this higher-thrust engine the Merlin 1E?It appears to be the same engine with increased prop flow.
Should SpaceX call this higher-thrust engine the Merlin 1E?
I would not think so. It appears that the higher thrust was intended from the beginning. They are setting the 'standard' throttle from ~85% of planned thrust to ~100%. The upgrade might be as simple as a change to the engine firmware.
And don't forget prop densification by lowering its temperature.
Should SpaceX call this higher-thrust engine the Merlin 1E?
I would not think so. It appears that the higher thrust was intended from the beginning. They are setting the 'standard' throttle from ~85% of planned thrust to ~100%. The upgrade might be as simple as a change to the engine firmware.
And don't forget prop densification by lowering its temperature.
There is no thrust increase strictly from increased prop density, just stage mass ratio increases.
Should SpaceX call this higher-thrust engine the Merlin 1E?
I would not think so. It appears that the higher thrust was intended from the beginning. They are setting the 'standard' throttle from ~85% of planned thrust to ~100%. The upgrade might be as simple as a change to the engine firmware.
And don't forget prop densification by lowering its temperature.
There is no thrust increase strictly from increased prop density, just stage mass ratio increases.
Centrifugal pumps are volume-pumping machines at constant rpm so density increases do increase flow rate and thus thrust. Stage mass fraction has no effect on thrust. The turbopump power also has to be increased to accommodate the increased mass flow at constant volume.
And the T/W ratio should be over 170 :o
Sorry if this is an obvious question. After the 1D+ has been tested and validated, would they continue to use the new higher thrust for typical missions? I.e. low-mass payload, enough fuel for RTLS. In other words, if the goal is to maximize engine lifespan, does logic dictate running the engines slower when possible?
Upgrades in the works to allow landing for geo missions: thrust +15%, deep cryo oxygen, upper stage tank vol +10%
And the T/W ratio should be over 170 :o
Which means that the engines make up about 0.7% of the total mass of the stage, whereas the tanks are on the order of 3-4%. While high T/W ratios are great (and I'd love to see a 200 there), it really doesn't matter all that much. Reducing tank mass and residual fuel mass is much more important in terms of performance - and so is any increase in ISP, even if it means a lower T/W ratio.
AviationWeek posted an interesting article on updated Falcon 9/Merlin 1-D
http://aviationweek.com/space/upgraded-falcon-9-may-need-additional-certification
Is there any possibility that SpaceX has been slowly ramping up max throttle on the M1D at each launch ? Like adding 1% per launch. The weight of those GEO payloads vs DeltaV to GEO, any data suggesting that ? Of course, increasing throttle without any of the other changes would result in just reduced gravity losses, or a relatively small performance gain.
That doesn't really sound like very good logic. RPM and amount of propellant moved are both functions of power to the pump. Increased density isn't what's causing more flow. Power is.There is no thrust increase strictly from increased prop density, just stage mass ratio increases.Should SpaceX call this higher-thrust engine the Merlin 1E?
I would not think so. It appears that the higher thrust was intended from the beginning. They are setting the 'standard' throttle from ~85% of planned thrust to ~100%. The upgrade might be as simple as a change to the engine firmware.
And don't forget prop densification by lowering its temperature.
Centrifugal pumps are volume-pumping machines at constant rpm so density increases do increase flow rate and thus thrust. Stage mass fraction has no effect on thrust. The turbopump power also has to be increased to accommodate the increased mass flow at constant volume.
Pressure differential multiplied volume flow rate divided efficiency.
Engine thrust is related to mass flow rate.
Increasing density you have more mass flow (more thrust) with same volume flow (same pumping power).
Pressure differential multiplied volume flow rate divided efficiency.
Engine thrust is related to mass flow rate.
Increasing density you have more mass flow (more thrust) with same volume flow (same pumping power).
It would seem that unless you changed the geometry of the combustion/thrust chambers, that it's going to take higher pressures and temperatures to increase the thrust, and so pumping power.
The same orifice, the same (or close) mach 1?, higher pressure...
Centrifugal pumps are not volume machines, but pressure machines with the power needed being that to maintain a pressure at a given flow. The pressure is generated by the centripetal? acceleration of the mass and then the force outward.
I wonder how all this relates to the "sweet spot" Musk mentioned long ago.
edit: drifted a little.
In March 16 and 17 appearances at the Satellite 2015 conference here, Shotwell said the new-version Falcon 9, which has yet to be named, will be about 30 percent more powerful than the rocket’s current version. -
Shotwell said the company stopped full qualification of the Merlin 1D engine’s capabilities to keep the first Falcon 9 v1.1 flights on schedule. But the qualification work has continued.
“We’ve gone back and gotten that performance on the engine to place it on the vehicle,” Shotwell said. “So we’ve got a higher-thrust engine. We’ve finished development on that and are going into qual [qualification testing].
I am not expert, but does not reducing the thrust of an engine to below design optimum also reduce the Isp? So if they reduced the throttled the engine down to 85% thrust, they would also reduce the Isp a bit and throttling it up to 100% would improve Isp as well. Densification will do the rest of the 30% increase in performance. That is at least my idea of what is happening.It's about the chamber thrust and the nozzle expansion ratio. In existing engines, usually if you increase the thrust, you do so by increasing the chamber pressure (there are other ways to do it, but usually it means you're essentially producing a new engine).
30% more powerful is meaningless without knowing what exactly they're referring to with the word 'power'. There is no reason to think she's countermanding Elon.
She's clearly talking about payload capability of the rocket... and probably payload to LEO.
Aerospace 101 style rocket equation of new numbers confirms ~30% more payload to orbit with 285ISP for first stage, 345ISP for second stage, +10% second stage tank volume and +9% propellant mass (approximated the weighted average of 10% LOX density and 5% RP-1 density). New GLOW is 525613kg. T/W at lift off is ~1.16:1.
Elon Musk @elonmusk · 2h 2 hours ago
Falcon 9 lifting off with 1.3 mmmillion pounds of force
A rather cryptic tweet from Elon
Tweet (https://twitter.com/elonmusk/status/579861658337574912)QuoteElon Musk @elonmusk · 2h 2 hours ago
Falcon 9 lifting off with 1.3 mmmillion pounds of force
Is he rounding up the thrust of the F9 core from 1.26M lbf or did SpaceX manage to squeeze out another 40000 lbf of thrust with maybe a gain of a few seconds in impulse?
Right. The base of the first few v1.1 rockets was white, but then changed to black.Looks like they mildly changed octoweb. Perhaps lacking white 'covers'?
Credit: SpaceX, via parabolicarc.
http://www.parabolicarc.com/wp-content/uploads/2014/10/SpaceX_Merlin_1D_octoweb.jpg
Where? I don't see a difference.
The 30% improved performance (payload wise) would increase the F9 to [approximations payload values to show scale of increases]:
LEO GTO
Reusable 13.3 -> 17 4.5 -> 5.8
Fully Expendable 16 -> 20 5.2 -> 6.7
To put it in competitive payload capability with an Atlas V:
GTO AV-521 vs F9R, AV-531 vs F9E
LEO AV -531 vs F9R, AV-551 vs F9E
The LEO capability may be even great enough to launch a BA330 on a fully expendable F9 flight vs having to use a FH.
The basic thing is that the new performance numbers means that any payload services sold that required the use of an F9E (no legs) can now be performed by a F9R (with legs).
Could it be "unpainted" TPS?Right. The base of the first few v1.1 rockets was white, but then changed to black.Looks like they mildly changed octoweb. Perhaps lacking white 'covers'?
Credit: SpaceX, via parabolicarc.
http://www.parabolicarc.com/wp-content/uploads/2014/10/SpaceX_Merlin_1D_octoweb.jpg
Where? I don't see a difference.
- Ed Kyle
Right. The base of the first few v1.1 rockets was white, but then changed to black.Looks like they mildly changed octoweb. Perhaps lacking white 'covers'?
Credit: SpaceX, via parabolicarc.
http://www.parabolicarc.com/wp-content/uploads/2014/10/SpaceX_Merlin_1D_octoweb.jpg
Where? I don't see a difference.
The 30% improved performance (payload wise) would increase the F9 to [approximations payload values to show scale of increases]:
LEO GTO
Reusable 13.3 -> 17 4.5 -> 5.8
Fully Expendable 16 -> 20 5.2 -> 6.7
To put it in competitive payload capability with an Atlas V:
GTO AV-521 vs F9R, AV-531 vs F9E
LEO AV -531 vs F9R, AV-551 vs F9E
The LEO capability may be even great enough to launch a BA330 on a fully expendable F9 flight vs having to use a FH.
The basic thing is that the new performance numbers means that any payload services sold that required the use of an F9E (no legs) can now be performed by a F9R (with legs).
I wish we would specify a little more clearly what we mean with things like 'GTO'. There are a lot of GTO orbits and they are not the same. One of which is SpaceX's GTO -1800m/s which is not comparable to, say, Ariane 5 GTO -1500m/s capability. I know you know the difference, but it would be nice, as a reader, to not have to interpret what you mean quite as heavily. :)
The 30% improved performance (payload wise) would increase the F9 to [approximations payload values to show scale of increases]:
LEO GTO
Reusable 13.3 -> 17 4.5 -> 5.8
Fully Expendable 16 -> 20 5.2 -> 6.7
To put it in competitive payload capability with an Atlas V:
GTO AV-521 vs F9R, AV-531 vs F9E
LEO AV -531 vs F9R, AV-551 vs F9E
The LEO capability may be even great enough to launch a BA330 on a fully expendable F9 flight vs having to use a FH.
The basic thing is that the new performance numbers means that any payload services sold that required the use of an F9E (no legs) can now be performed by a F9R (with legs).
I wish we would specify a little more clearly what we mean with things like 'GTO'. There are a lot of GTO orbits and they are not the same. One of which is SpaceX's GTO -1800m/s which is not comparable to, say, Ariane 5 GTO -1500m/s capability. I know you know the difference, but it would be nice, as a reader, to not have to interpret what you mean quite as heavily. :)
Which of the above underlying assumptions are people using as a baseline starting point for their numbers? I too would be happy to have some more explicitly defined conditions listed but I don't really see any practical way of doing that until SpaceX clears up some of the ambiguity surrounding their claimed capabilities. Until we get a new set of payload numbers from SpaceX to use as a baseline, one with the underlying assumptions explicitly laid out, I just don't see us being able to achieve much clarity on the subject. Just my $0.02
The 30% improved performance (payload wise) would increase the F9 to [approximations payload values to show scale of increases]:
LEO GTO
Reusable 13.3 -> 17 4.5 -> 5.8
Fully Expendable 16 -> 20 5.2 -> 6.7
To put it in competitive payload capability with an Atlas V:
GTO AV-521 vs F9R, AV-531 vs F9E
LEO AV -531 vs F9R, AV-551 vs F9E
The LEO capability may be even great enough to launch a BA330 on a fully expendable F9 flight vs having to use a FH.
The basic thing is that the new performance numbers means that any payload services sold that required the use of an F9E (no legs) can now be performed by a F9R (with legs).
I wish we would specify a little more clearly what we mean with things like 'GTO'. There are a lot of GTO orbits and they are not the same. One of which is SpaceX's GTO -1800m/s which is not comparable to, say, Ariane 5 GTO -1500m/s capability. I know you know the difference, but it would be nice, as a reader, to not have to interpret what you mean quite as heavily. :)
How exactly do you suggest we do that? SpaceX's numbers are either with or without 30% reserves for reuse already baked in; with or without the full thrust Merlin 1Ds already baked in; with or without the rest of the upgrades Elon mentioned--subchilling and +10% US volume-- already baked in; to a GTO that is 1800m/s short of GEO or one closer to GEO than that (ABS/Eutelsat was to about -1600m/s); realistic or somewhat wishful thinking; etc.
Which of the above underlying assumptions are people using as a baseline starting point for their numbers? I too would be happy to have some more explicitly defined conditions listed but I don't really see any practical way of doing that until SpaceX clears up some of the ambiguity surrounding their claimed capabilities. Until we get a new set of payload numbers from SpaceX to use as a baseline, one with the underlying assumptions explicitly laid out, I just don't see us being able to achieve much clarity on the subject. Just my $0.02
If the new engine pushes the Falcon heavy towards 60 tonnes to LEO, won't this make the $1 billion Per launch 70 tonne SLS look DOA. $150 mill for 60 tonnes.$1000 mill for 70 tonnes ummmSince cross feeding is no longer being used on the Falcon Heavy I suspect that the upgrades will be used to get it to its advertised lift capacity. SLS is not a good comparison to make. SLS will be capable of almost 90 tonnes from the start. The lift capacity is under reported. The program wanted room for mass growth should it happen, but it hasn't.
Does anyone know the figures for falcon heavy with updated engines to LEO
Since cross feeding is no longer being used on the Falcon Heavy I suspect that the upgrades will be used to get it to its advertised lift capacity.
It was mentioned in the thread about Gwynne Shotwell's comments at UC Berkeley. The LEO payload for Falcon Heavy is now 53 tonnes according to the website not 52 as it had been. So it looks like SpaceX is updating the official figures on the website and that they are at least somewhat current.
Since cross feeding is no longer being used on the Falcon Heavy I suspect that the upgrades will be used to get it to its advertised lift capacity.
I am such a newb but why is cross feeding no longer being used for FH? How do y'all know when it aint even launched yet, has spacex disclosed said design details so many months before it even debuts?
I am such a newb but why is cross feeding no longer being used for FH? How do y'all know when it aint even launched yet, has spacex disclosed said design details so many months before it even debuts?
13 tons to Mars is not bad. They could send several small ISRU and power production units prior to human landing at a much lower cost than SLS. Even two Falcon Heavy launches, docked, for 26 tons. One with a Mars booster, and the other with the lander ISRU equipment that could match SLS at a lower cost. Or they could launch a 50 ton unit and refuel it with Falcon 9 or another heavy for 50 tons unit to Mars.
It was mentioned in the thread about Gwynne Shotwell's comments at UC Berkeley. The LEO payload for Falcon Heavy is now 53 tonnes according to the website not 52 as it had been. So it looks like SpaceX is updating the official figures on the website and that they are at least somewhat current.
http://forum.nasaspaceflight.com/index.php?topic=36879.msg1337655#msg1337655
Correct me, if I am wrong, but deeply cooled LOX will also provide part of performace incrase for high energy orbits, because it will slow down LOX boil-off.
Correct me, if I am wrong, but deeply cooled LOX will also provide part of performace incrase for high energy orbits, because it will slow down LOX boil-off.
Actually no, as there is no boil-off during the ascent. However it does mean more mass in a smaller storage space to then be thrown out the tail pipe so to speak.
This is strange - SES-9 is supposed to be the mission for upgrade Merlins. Before they could launch Turkmenisat-1, Orbcomm og2 and what? Jason-3? In flight abort?
Also, the outer boosters could operate at full power (165,000 lb thrust) upgraded mode, while the core can operate at a lower thrust to burn longer thus higher staging for the second stage.
Also, the outer boosters could operate at full power (165,000 lb thrust) upgraded mode, while the core can operate at a lower thrust to burn longer thus higher staging for the second stage.
There is no performance benefit in running at anything less than maximum throttle. "Staging later" just means more gravity losses.
(Unless the engine has higher isp at lower throttle, but this is the opposite of what is usual)
(And of course there may be structural reasons for throttling down, but that's getting into a trade between vehicle dry mass and extracting full performance from the engines)
Also, the outer boosters could operate at full power (165,000 lb thrust) upgraded mode, while the core can operate at a lower thrust to burn longer thus higher staging for the second stage.
There is no performance benefit in running at anything less than maximum throttle. "Staging later" just means more gravity losses.
(Unless the engine has higher isp at lower throttle, but this is the opposite of what is usual)
(And of course there may be structural reasons for throttling down, but that's getting into a trade between vehicle dry mass and extracting full performance from the engines)
Also, the outer boosters could operate at full power (165,000 lb thrust) upgraded mode, while the core can operate at a lower thrust to burn longer thus higher staging for the second stage.
There is no performance benefit in running at anything less than maximum throttle. "Staging later" just means more gravity losses.
(Unless the engine has higher isp at lower throttle, but this is the opposite of what is usual)
(And of course there may be structural reasons for throttling down, but that's getting into a trade between vehicle dry mass and extracting full performance from the engines)
Of course there's a performance benefit from running the center booster at less than maximum throttle -- the benefit is that more of the propellant in the center core only has to lift the mass of the core, not the side boosters.
Now that we hope to be seeing the "full thrust" Merlin 1-D configuration fly soon I have to wonder how long they've been planning this. Back when the Merlin 1-D was first introduced, I believe it was Tom Mueller who said there would not be a 1-E, which generated some discussion here. I wonder if he had a hard time keeping a straight face. Given later statements about needing to qualify and fly even though they realized the engines had more to give I suspect it's been since day 1 and the stage 2 stretch has been sitting on the shelf waiting for quite a while. I do wonder how much tweaking it took to qualify the Merlin-1D at full thrust.
They didn't move thr red line... they just developed technology (fuel densification) that allowed operation at the original 100% red line instead of at 85% like v1.0.Highly speculative statement.
They didn't move thr red line... they just developed technology (fuel densification) that allowed operation at the original 100% red line instead of at 85% like v1.0.
They didn't move thr red line... they just developed technology (fuel densification) that allowed operation at the original 100% red line instead of at 85% like v1.0.Highly speculative statement.
Now that we hope to be seeing the "full thrust" Merlin 1-D configuration fly soon I have to wonder how long they've been planning this. Back when the Merlin 1-D was first introduced, I believe it was Tom Mueller who said there would not be a 1-E, which generated some discussion here. I wonder if he had a hard time keeping a straight face. Given later statements about needing to qualify and fly even though they realized the engines had more to give I suspect it's been since day 1 and the stage 2 stretch has been sitting on the shelf waiting for quite a while. I do wonder how much tweaking it took to qualify the Merlin-1D at full thrust.I believe that it was as related to the huge margin for thurst increase that Merlin 1D had, but also becuase of the Raptor development. Whatever comes after Falcon will be a thing out of science fiction books. If you have seen the simulations that Dimitry and Hyperion 5 did of a Falcon 9 M with Mini-Raptors, and you consider densification and stretched upper stage as a possibility, you're talking about 30tonnes to LEO and 10 to GTO from a single stick.
Now that we hope to be seeing the "full thrust" Merlin 1-D configuration fly soon I have to wonder how long they've been planning this. Back when the Merlin 1-D was first introduced, I believe it was Tom Mueller who said there would not be a 1-E, which generated some discussion here. I wonder if he had a hard time keeping a straight face. Given later statements about needing to qualify and fly even though they realized the engines had more to give I suspect it's been since day 1 and the stage 2 stretch has been sitting on the shelf waiting for quite a while. I do wonder how much tweaking it took to qualify the Merlin-1D at full thrust.I believe that it was as related to the huge margin for thurst increase that Merlin 1D had, but also becuase of the Raptor development. Whatever comes after Falcon will be a thing out of science fiction books. If you have seen the simulations that Dimitry and Hyperion 5 did of a Falcon 9 M with Mini-Raptors, and you consider densification and stretched upper stage as a possibility, you're talking about 30tonnes to LEO and 10 to GTO from a single stick.
In other words, Tom knew that anything that came after the Merlin 1D was going to be CH4/LOX full staged technology.
What market will be served by the fuller-thrust Merlin, with the other changes (densified propellant, stretched stage, etc)?
Doesn't F9R do everything it's supposed to right now, mission-wise?
Same market but with more possibilities for recovery of the 1st stage. Right now, that is not possible for GTO comsat missions. With the changes, it will be for sats up to X tonnes. "X" yet to be revealed/determined.
Same market but with more possibilities for recovery of the 1st stage. Right now, that is not possible for GTO comsat missions. With the changes, it will be for sats up to X tonnes. "X" yet to be revealed/determined.
Will these changes propagate to Falcon Heavy also? Once the new generation appears, does the previous generation go obsolete and get retired?
Will these changes propagate to Falcon Heavy also? Once the new generation appears, does the previous generation go obsolete and get retired?
The 30% improved performance (payload wise) would increase the F9 to [approximations payload values to show scale of increases]:I'm afraid the math doesn't work like that. The only figure for which they might achieve 30% increase is actual payload to GTO, and the only reason it's so large is that the upper stage dry mass is 4-5 tons, so a 15% increase in mass to GTO represents a ~30% increase in *payload* to GTO.
LEO GTO
Reusable 13.3 -> 17 4.5 -> 5.8
Fully Expendable 16 -> 20 5.2 -> 6.7
To put it in competitive payload capability with an Atlas V:
GTO AV-521 vs F9R, AV-531 vs F9E
LEO AV -531 vs F9R, AV-551 vs F9E
The LEO capability may be even great enough to launch a BA330 on a fully expendable F9 flight vs having to use a FH.
The basic thing is that the new performance numbers means that any payload services sold that required the use of an F9E (no legs) can now be performed by a F9R (with legs).
This got me to thinking, is the thrust being uprated on both the booster and vacuum engines? I don't remember seeing it specified and just improving the second stage would probably be the biggest help.
This got me to thinking, is the thrust being uprated on both the booster and vacuum engines? I don't remember seeing it specified and just improving the second stage would probably be the biggest help.
The second stage engine is overpowered already. No need to upgrade unless it involves a sizable ISP increase too. The second stage upgrade is in the tank stretch.
But isn't an ISP increase implicit? How do you get more thrust without increasing the chamber pressure?What if the preburner requires proportionally greater propellant?
I tried to calculate the isp of the uprated Merlin 1D Vac, but I found out that the gas generator (this is not staged combustion) dumped mass basically cancelled out the increase in Pc. A bigger expansion ratio would otherwise help, though.But isn't an ISP increase implicit? How do you get more thrust without increasing the chamber pressure?What if thepreburnergas generator requires proportionally greater propellant?
Hmmm, how big a change would it be to adjust the relative s1 tanks to optimize for chilled prop?
Discussion there was that it would only require repositioning the common bulkhead.Hmmm, how big a change would it be to adjust the relative s1 tanks to optimize for chilled prop?
There is discussion about this in other threads, I just can't find where right now. Probably the General Falcon 9 thread somewhere.
Discussion there was that it would only require repositioning the common bulkhead.Hmmm, how big a change would it be to adjust the relative s1 tanks to optimize for chilled prop?
There is discussion about this in other threads, I just can't find where right now. Probably the General Falcon 9 thread somewhere.
Great photo--thanks deruch. The nozzle expansion looks huge--is this a Merlin Vac or is this just early in the process and they will "bring it in" to a smaller diameter? (I realize they will "bring it in" regardless, but ...)
I don't think so. Explosively formed is inner and outer layer of combustion chamber with integrated cooling channels. Nozzle shouldn't be produced that way sice its made out of one layer of quite thin sheet of metal.Merlin 1D nozzles are explosive formed. This is a 1D-Vac extension.
Great photo--thanks deruch. The nozzle expansion looks huge--is this a Merlin Vac or is this just early in the process and they will "bring it in" to a smaller diameter? (I realize they will "bring it in" regardless, but ...)
Why does it read "Former launch engineer at SpaceX"? Is Tom Mueller not with SpaceX any more?
"The Merlin 1D weighs 1030 pounds, including the hydraulic steering (TVC) actuators. It makes 162,500 pounds of thrust in vacuum. that is nearly 158 thrust/weight. The new full thrust variant weighs the same and makes about 185,500 lbs force in vacuum. You can do the math! BTW, I believe most other engines don't include the thrust vector control actuators in their F/W numbers."
http://www.quora.com/Is-SpaceXs-Merlin-1Ds-thrust-to-weight-ratio-of-150+-believable
Why does it read "Former launch engineer at SpaceX"? Is Tom Mueller not with SpaceX any more?
That is weird. According to his LinkedIn profile (https://www.linkedin.com/pub/thomas-mueller/3b/451/209) his job title has changed but he's still at SpaceX.
Why does it read "Former launch engineer at SpaceX"? Is Tom Mueller not with SpaceX any more?
That is weird. According to his LinkedIn profile (https://www.linkedin.com/pub/thomas-mueller/3b/451/209) his job title has changed but he's still at SpaceX.
Looks like a focus on Mars......Raptor? EDIT: i.e. does the nozzle maybe become under expanded?
Why does it read "Former launch engineer at SpaceX"? Is Tom Mueller not with SpaceX any more?
That is weird. According to his LinkedIn profile (https://www.linkedin.com/pub/thomas-mueller/3b/451/209) his job title has changed but he's still at SpaceX.
Looks like a focus on Mars......Raptor? EDIT: i.e. does the nozzle maybe become under expanded?
Just speculation? Or did I miss something on his Linkedin?
And I wonder, is the 180:1 T/W ratio accompanied with a drop in performance?
180:1 thrust to weight... Absolutely incredible.
I guess that's what happens when you push the limits of RP-1 without switching to staged combustion.
I look forward to future progress with Raptor!
"The Merlin 1D weighs 1030 pounds, including the hydraulic steering (TVC) actuators. It makes 162,500 pounds of thrust in vacuum. that is nearly 158 thrust/weight. The new full thrust variant weighs the same and makes about 185,500 lbs force in vacuum. You can do the math! BTW, I believe most other engines don't include the thrust vector control actuators in their F/W numbers."
http://www.quora.com/Is-SpaceXs-Merlin-1Ds-thrust-to-weight-ratio-of-150+-believable
Doing the math... That is a 180:1 thrust/weight ratio. That is insane! This is certainly more of a testament to improved manufacturing techniques since there are so few american liquid rocket engines developed in the last 2 decades, but whatever the reason the results are impressive.
"The Merlin 1D weighs 1030 pounds, including the hydraulic steering (TVC) actuators. It makes 162,500 pounds of thrust in vacuum. that is nearly 158 thrust/weight. The new full thrust variant weighs the same and makes about 185,500 lbs force in vacuum. You can do the math! BTW, I believe most other engines don't include the thrust vector control actuators in their F/W numbers."
http://www.quora.com/Is-SpaceXs-Merlin-1Ds-thrust-to-weight-ratio-of-150+-believable
Doing the math... That is a 180:1 thrust/weight ratio. That is insane! This is certainly more of a testament to improved manufacturing techniques since there are so few american liquid rocket engines developed in the last 2 decades, but whatever the reason the results are impressive.
Question: Might SpaceX have changed manufacturing technique for the "full-thrust" Merlin 1D? Perhaps 3D-printing the thrust chamber to reduce weight? Sounds a bit crazy, but read on.
Musk said this week at the ISSRDC that they are 3D-printing LOTS of engine parts for Raptor, and we know they are doing so for Super Draco. He emphasized lower weight as a result. Link to the source, and summary of what he said, is in another NSF thread, here (http://forum.nasaspaceflight.com/index.php?topic=37839.msg1402002#msg1402002).
If SpaceX is perhaps simply changing the manufacturing technique of some parts of the Merlin 1D as part of the full-thrust engine, that would result in a crazy high thrust-to-weight ratio.
Edited to clarify.
180:1 thrust to weight... Absolutely incredible.
I guess that's what happens when you push the limits of RP-1 without switching to staged combustion.
I look forward to future progress with Raptor!
FFSC will likely have much lower T/W due to requirement for an additional preburner, turbine, and gas flow piping that raises mass considerably. Somewhat compensated by higher combustion pressure and higher ISP
180:1 thrust to weight... Absolutely incredible.
I guess that's what happens when you push the limits of RP-1 without switching to staged combustion.
I look forward to future progress with Raptor!
FFSC will likely have much lower T/W due to requirement for an additional preburner, turbine, and gas flow piping that raises mass considerably. Somewhat compensated by higher combustion pressure and higher ISP
Two questions:
1. What was the previous record holder for gas gen T/W?
2. Couldn't SpaceX apply same techniques/technology to shatter the records for SC engines?
I found this exchange with Tom Mueller amusing. I'm not a Quora expert, but it appears to be one of only two things he has posted so he must really have wanted to set the record straight!I've been out of the loop for a while, but those numbers (at least at the vac stage of the burn) are basically 114% thrust vs current.
"The Merlin 1D weighs 1030 pounds, including the hydraulic steering (TVC) actuators. It makes 162,500 pounds of thrust in vacuum. that is nearly 158 thrust/weight. The new full thrust variant weighs the same and makes about 185,500 lbs force in vacuum. You can do the math! BTW, I believe most other engines don't include the thrust vector control actuators in their F/W numbers."
http://www.quora.com/Is-SpaceXs-Merlin-1Ds-thrust-to-weight-ratio-of-150+-believable
I've often wondered which engines include TVC actuators in their reported masses as we now know M1D does, and which engines consider them part of the stage systems? RD-180 presumably includes them since they appear to be part of the engine/thrust structure assembly, but SSME presumably does not? The frequently quoted 137:1 thrust-to-weight ratio of the NK-33, against which M1D is often compared, presumably also does not include TVC actuators as it had none in the configuration the NK-15 flew on the N1 vehicle, and which NK-33 was intended to replace. Anyone know the story for LR-87-5 which is also claimed to have a very high thrust-to-weight ratio (although strangely does not usually make it into lists of high thrust-to-weight engines)?
180:1 thrust to weight... Absolutely incredible.
I guess that's what happens when you push the limits of RP-1 without switching to staged combustion.
I look forward to future progress with Raptor!
I found this exchange with Tom Mueller amusing. I'm not a Quora expert, but it appears to be one of only two things he has posted so he must really have wanted to set the record straight!I've been out of the loop for a while, but those numbers (at least at the vac stage of the burn) are basically 114% thrust vs current.
"The Merlin 1D weighs 1030 pounds, including the hydraulic steering (TVC) actuators. It makes 162,500 pounds of thrust in vacuum. that is nearly 158 thrust/weight. The new full thrust variant weighs the same and makes about 185,500 lbs force in vacuum. You can do the math! BTW, I believe most other engines don't include the thrust vector control actuators in their F/W numbers."
http://www.quora.com/Is-SpaceXs-Merlin-1Ds-thrust-to-weight-ratio-of-150+-believable
I've often wondered which engines include TVC actuators in their reported masses as we now know M1D does, and which engines consider them part of the stage systems? RD-180 presumably includes them since they appear to be part of the engine/thrust structure assembly, but SSME presumably does not? The frequently quoted 137:1 thrust-to-weight ratio of the NK-33, against which M1D is often compared, presumably also does not include TVC actuators as it had none in the configuration the NK-15 flew on the N1 vehicle, and which NK-33 was intended to replace. Anyone know the story for LR-87-5 which is also claimed to have a very high thrust-to-weight ratio (although strangely does not usually make it into lists of high thrust-to-weight engines)?
Cheers, Martin
The russians dont even bother with that type because they pushed SC kerolox technology about as far as it can be pushed already in the 60s/70s.
The russians dont even bother with that type because they pushed SC kerolox technology about as far as it can be pushed already in the 60s/70s.
70s/80s, don't give them too much credit. What they accomplished was amazing but from the mid 60's onward they were 10 years behind the American program. For example their shuttle program. Also, how many liquid hydrogen engines have come out of USSR/Russia?
The excellence in Soviet technology, although impressive has not been unsurpassed because others simply have gone down a different path.
Edit: Spelling
The russians dont even bother with that type because they pushed SC kerolox technology about as far as it can be pushed already in the 60s/70s.
70s/80s, don't give them too much credit. What they accomplished was amazing but from the mid 60's onward they were 10 years behind the American program. For example their shuttle program. Also, how many liquid hydrogen engines have come out of USSR/Russia?
The excellence in Soviet technology, although impressive has not been unsurpassed because others simply have gone down a different path.
Edit: Spelling
But the Russians seems to have gone down the RIGHT path (hydrocarbon rather then hydrogen and solids), being much further along on a wrong track doesn't put you ahead, if anything it's a road block (due to vested interests) to actually getting to the right solution. The Russians mistake was to try to do copy our path rather then sticking with their own.
I believe the choice of rocket fuel is simply a design decision based on the intended market and application of the launch vehicle.
The russians dont even bother with that type because they pushed SC kerolox technology about as far as it can be pushed already in the 60s/70s.
70s/80s, don't give them too much credit. What they accomplished was amazing but from the mid 60's onward they were 10 years behind the American program. For example their shuttle program. Also, how many liquid hydrogen engines have come out of USSR/Russia?
The excellence in Soviet technology, although impressive has not been unsurpassed because others simply have gone down a different path.
Edit: Spelling
Engine | Merlin 1D (Design) | Merlin 1D (Test article) | Merlin 1D (Production) | Merlin 1D (Plus) |
SL Thrust | 140klbf | ? | 147klbf | 168klbf? |
Vac Thrust | 155klbf | ?(better than expected) | 161klbf | 184klbf? |
SL isp | 280s? | ? | 282s | 282s? |
Vac isp | 310s | 309s? | 311s | 311s? |
T/W | 160 | ? | 150+ | 157? |
Chamber Pressure | 1410psi | ? | 1410psi? | 1410psi? |
Expansion | 16 | 16? | 16?(this shouldn’t vary much from design) | 16 |
Throttle range | 70-100% | 70%-100%? | 70%-100% | 70%-100%? |
Engine weight | 440kg? | ? | 485kg? | 485kg? |
Costs | <M1C? | - | <M1C | =M1D(Production) |
The following is a table summary of the M1D through its many forms. Please update and correct any errors in the values of the table.
Engine Merlin 1D (Design) Merlin 1D (Test article) Merlin 1D (Production) Merlin 1D (Plus) SL Thrust 140klbf ? 147klbf 168klbf? Vac Thrust 155klbf ?(better than expected) 161klbf 184klbf? SL isp 280s? ? 282s 282s? Vac isp 310s 309s? 311s 311s? T/W 160 ? 150+ 157? Chamber Pressure 1410psi ? 1410psi? 1410psi? Expansion 16 16? 16?(this shouldn’t vary much from design) 16 Throttle range 70-100% 70%-100%? 70%-100% 70%-100%? Engine weight 440kg? ? 485kg? 485kg? Costs <M1C? - <M1C =M1D(Production)
That "Center Pusher" stage separation system is something I'd never heard of. Is that the current technique?I don't think so. They currently use 4 pushers spaced around the perimeter of the interstage. I've never seen an indication of a center pusher, either in assembly photos or 2nd stage sep videos from mission launches.
In the slides posted above, thanks for that btw, it shows a "Center Pusher" under the Mvac during stage sep.
I don't recall ever seeing that. I wonder if that was needed because of the new thrust profile and/or 2nd stage stretch. Was the bell extended along with the interstage requiring an additional guide or is it because the pushers along the interstage are no longer able to push the stretched 2nd stage without additional help. Or all of the above.
In the slides posted above, thanks for that btw, it shows a "Center Pusher" under the Mvac during stage sep.
I don't recall ever seeing that. I wonder if that was needed because of the new thrust profile and/or 2nd stage stretch. Was the bell extended along with the interstage requiring an additional guide or is it because the pushers along the interstage are no longer able to push the stretched 2nd stage without additional help. Or all of the above.
Has that ever been done before...this center line, internal pusher?
Some interesting new info
170klbf thrust at sea level and 210klbf in vacuum :)
I would be -very- surprised if the new thrust was achieved without a commensurate increase in ISP since they presumably are not redoing the thrust chamber to open the throat. (I.e. they modestly increased chamber pressure.) But that's just speculation.
Some interesting new info
170klbf thrust at sea level and 210klbf in vacuum :)
It will be interesting to see what happens to the specific impulse for the vacuum version. 210,000 is a 26,000 more than the expected number listed in the table.
Anyway have a guess?
70s/80s, don't give them too much credit. What they accomplished was amazing but from the mid 60's onward they were 10 years behind the American program.Were that the case losing access to 1980s Soviet engine technology would not be a crisis for American assured access to space.
70s/80s, don't give them too much credit. What they accomplished was amazing but from the mid 60's onward they were 10 years behind the American program.Were that the case losing access to 1980s Soviet engine technology would not be a crisis for American assured access to space.
No, that doesn't logically follow.Perhaps, but we can also see the outcome of attempts to do hydrogen boost and the technology gap for American attempts to do hydrocarbon boost, which rules out the other possibilities. These aren't equally valid technology tracks. Hydrocarbon boost is really better. The US finally getting credible ORSC hydrocarbon boost is probably the biggest advance in American spaceflight in generations, but even then it's not quite up to RD-180 standards. AR-1 and BE-4 look to be a regression in ISP.
Were that the case losing access to 1980s Soviet engine technology would not be a crisis for American assured access to space.
It's only a "crisis" because the rocket was designed for the engine and slotting in a new engine that is not designed identically is a big effort. And developing a new engine takes time.Granted, but AFAICT AR-1 still has reduced ISP. More akin to an uprated NK-33. That's nothing to complain about but when the US resets and picks up where the Soviets were in the 1960s/70s and it's an improvement, it's really hard to believe the two technology bases are equal but different.
Regardless, that window is closing, and the not-so-distant future will see only one US LV provider (and a marginal one at that) relying on Russian engine technology.Yes, changing tracks and resetting to 1960s/70s Soviet technology is an improvement, and more than enough for viable launch capability. It may never be economical to replicate RD-180/191.
Ballast would cut down the max G without new levels of throttling.Ballast on an upper stage sufficiently to reduce max g-loading makes zero sense. Payload is king, especially in high-energy missions like commercial GTO.
Payload is king, especially in high-energy missions like commercial GTO.
Why reengineering? They'd know the specs MVac 1D+ needed to hit from the outset. It's just engineering.Payload is king, especially in high-energy missions like commercial GTO.Yes, the payload is king. What counts is its safe and sound delivery. What is the problem with ballast when doing this for a lightweight payload requires either a reengineering or ballast?
Why reengineering? They'd know the specs MVac 1D+ needed to hit from the outset. It's just engineering.
Payload is king, especially in high-energy missions like commercial GTO.
Yes, the payload is king. What counts is its safe and sound delivery. What is the problem with ballast when doing this for a lightweight payload requires either a reengineering or ballast?
So the answer to your "what is the problem" question is that it shortens the lifespan of the satellite once it reaches orbit. Which reduces the value of the launch, and therefore the price that can be charged.
Cheers, Martin
the 1D+ is still a Merlin 1D basically. To lower throttle they would need to do a lot of reengineering. It is completely unnecessary if ballast can do the same.Don't think that follows.
I'm sure ballast would be considered if necessary, but not seeing the obvious necessity.
So the answer to your "what is the problem" question is that it shortens the lifespan of the satellite once it reaches orbit. Which reduces the value of the launch, and therefore the price that can be charged.
Cheers, Martin
No it doesn't. We are talking about very light satellites that get accelerated too much. That would be satellites that are well below F9 capabilities. Heavier satellites won't need ballast.
The M1D+VAC engine with 210klbf would mean that the engines needs to be able to throttle down to 50% in order to keep the max g's for the lighter GEO sat missions (3mt payload) to remain under 7g's.
Bumped into this enjoy!
Servo Motors Survive Space X Launch Conditions
Robust Servo Motors Survive Launch Conditions to Optimize Rocket Fuel Burn
http://www.micromo.com/applications/aerospace-defense/space-x-shuttle-launch
Upgraded Merlin rocket engine with special black coating for thermal testing
Is this a significantly updated version of the Merlin engine (like a Merlin 1E), or is this just the Merlin 1D with full thrust. In the latter case I would have to wonder why they need to test its thermal properties. I would have assumed that they already tested all this a long time ago?
Is this a significantly updated version of the Merlin engine (like a Merlin 1E), or is this just the Merlin 1D with full thrust. In the latter case I would have to wonder why they need to test its thermal properties. I would have assumed that they already tested all this a long time ago?
We have no idea how old this picture is... Could easily be many months ago.
Is this a significantly updated version of the Merlin engine (like a Merlin 1E), or is this just the Merlin 1D with full thrust. In the latter case I would have to wonder why they need to test its thermal properties. I would have assumed that they already tested all this a long time ago?
According to their Facebook page, this is "special black coating for improved thermal testing"
Whatever that means...
According to their Facebook page, this is "special black coating for improved thermal testing"
Whatever that means...
Maybe because the shiny chrome gets all sooty on the trip up, and they need to test it dirty for improved accuracy.
center engine already had a partiall coating full is for all enginesQuote from: SpaceXUpgraded Merlin rocket engine with special black coating for thermal testing
Attached
The 'stand' of sorts the engine bell is in, is the same type of hold-down/support used for the Merlin 1d vac? Is this method (at least the main) reason the M1D-vac needs the 'stiffener' ring that comes off in after the first few seconds of stage 2 firing?
EDIT:picture added
The 'stand' of sorts the engine bell is in, is the same type of hold-down/support used for the Merlin 1d vac? Is this method (at least the main) reason the M1D-vac needs the 'stiffener' ring that comes off in after the first few seconds of stage 2 firing?
EDIT:picture added
MVAC nozzle is a soft Niobium nozzle that needs some help to prevent deformation.
M1D is not a soft metal. The stand is just to keep the engine upright and secure.
The 'stand' of sorts the engine bell is in, is the same type of hold-down/support used for the Merlin 1d vac? Is this method (at least the main) reason the M1D-vac needs the 'stiffener' ring that comes off in after the first few seconds of stage 2 firing?
EDIT:picture added
MVAC nozzle is a soft Niobium nozzle that needs some help to prevent deformation.
M1D is not a soft metal. The stand is just to keep the engine upright and secure.
I was under the impression that the M1D was the metal niobium bell and that the Vac had a carbon-carbon nozzle extension.
My understanding is that Merlin Vacuum is not just a Merlin 1D with a nozzle extension. I think it has a different base nozzle as well. I'm not sure about the thrust chamber.The 'stand' of sorts the engine bell is in, is the same type of hold-down/support used for the Merlin 1d vac? Is this method (at least the main) reason the M1D-vac needs the 'stiffener' ring that comes off in after the first few seconds of stage 2 firing?
EDIT:picture added
MVAC nozzle is a soft Niobium nozzle that needs some help to prevent deformation.
M1D is not a soft metal. The stand is just to keep the engine upright and secure.
I was under the impression that the M1D was the metal niobium bell and that the Vac had a carbon-carbon nozzle extension.
Is this a significantly updated version of the Merlin engine (like a Merlin 1E), or is this just the Merlin 1D with full thrust. In the latter case I would have to wonder why they need to test its thermal properties. I would have assumed that they already tested all this a long time ago?Full thrust Merlin 1D. They may have been awaiting a turbopump upgrade, or something similar, before being able to put full-thrust into development testing. Also, this coating may be for a cluster firing on the first "v1.2" stage, or something. Maybe. Or maybe not.
My understanding is that Merlin Vacuum is not just a Merlin 1D with a nozzle extension. I think it has a different base nozzle as well. I'm not sure about the thrust chamber.The 'stand' of sorts the engine bell is in, is the same type of hold-down/support used for the Merlin 1d vac? Is this method (at least the main) reason the M1D-vac needs the 'stiffener' ring that comes off in after the first few seconds of stage 2 firing?
EDIT:picture added
MVAC nozzle is a soft Niobium nozzle that needs some help to prevent deformation.
M1D is not a soft metal. The stand is just to keep the engine upright and secure.
I was under the impression that the M1D was the metal niobium bell and that the Vac had a carbon-carbon nozzle extension.
- Ed Kyle
Picture show the struts linking the TPA to the thrust structure (Merlin 1C).
Movement is allowed in the direction of the arrows.
Until Merlin 1C the TPA assembly was mounted partially on the thrust structure (following combustion chamber only on 1 DOF); movement was compensated with bellows.
On Merlin 1D TPA gimbals completely with engine and this is true specially for 1D Vac, where TPA exaust goes into nozzle.
T/W 165 (sl) 180 (vac) !!!!!!!!
From a first glance, it seems like the plan is to offload work from the first stage (18s less burning time) to make stage recovery easier.
Also, any idea what would the TWR be now?
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You need to look at more than the burn time. Look at the velocity of the stage at engine shutdown.
Yes SpaceX has increased density of the fuel and oxidizer, but that is probably not enough on it's own. The increase in thrust is coming from putting more fuel through the engine in the same amount of time.
Some "new" (changed 20 days ago) numbers on the SpaceX website (http://www.spacex.com/falcon9).Where did you find the ISP values? 348 was the earlier estimate for M1DVAC. Was there any actual ISP published so we can determine if the TC pressure has increased?- Merlin 1D vac thrust is now 934kN / 210.000lbf, isp is now 348s, burn time is 397s
- Merlin 1D FT thrust is now 756kN / 170.000lbf (SL) and 825kN / 185.500lbf (vac), burn time is 162s
From a first glance, it seems like the plan is to offload work from the first stage (18s less burning time) to make stage recovery easier.
Also, any idea what would the TWR be now?
Where did you find the ISP values? 348 was the earlier estimate for M1DVAC. Was there any actual ISP published so we can determine if the TC pressure has increased?
Where did you find the ISP values? 348 was the earlier estimate for M1DVAC. Was there any actual ISP published so we can determine if the TC pressure has increased?
Falcon 9's second stage is powered by a single Merlin vacuum engine nearly identical to the first-stage engines, but modified to operate in the vacuum of space. Like the main Merlin engines, the vacuum engine is designed and manufactured in-house by SpaceX. The engine is designed to burn for about six minutes, and can be shut down and restarted multiple times as needed to deliver different payloads into different orbits. SpaceX's Merlin vacuum engine has the highest vacuum specific impulse (isp)--a measure of engine efficiency--of any American liquid oxygen/kerosene engine with a vacuum isp of 348 seconds. The engine is housed inside the rocket's interstage.
Falcon 9's second stage is powered by a single Merlin vacuum engine nearly identical to the first-stage engines, but modified to operate in the vacuum of space. Like the main Merlin engines, the vacuum engine is designed and manufactured in-house by SpaceX. The engine is designed to burn for about six minutes, and can be shut down and restarted multiple times as needed to deliver different payloads into different orbits. SpaceX's Merlin vacuum engine has the highest vacuum specific impulse (isp)--a measure of engine efficiency--of any American liquid oxygen/kerosene engine with a vacuum isp of 340 seconds. The engine is housed inside the rocket's interstage.
...
Copying from the site.QuoteFalcon 9's second stage is powered by a single Merlin vacuum engine nearly identical to the first-stage engines, but modified to operate in the vacuum of space. Like the main Merlin engines, the vacuum engine is designed and manufactured in-house by SpaceX. The engine is designed to burn for about six minutes, and can be shut down and restarted multiple times as needed to deliver different payloads into different orbits. SpaceX's Merlin vacuum engine has the highest vacuum specific impulse (isp)--a measure of engine efficiency--of any American liquid oxygen/kerosene engine with a vacuum isp of 348 seconds. The engine is housed inside the rocket's interstage.
Regarding older isp numbers, I checked the site with the wayback machine. Here is a link from june 2015 (http://web.archive.org/web/20150519004028/http://www.spacex.com/falcon9). Quoting.QuoteFalcon 9's second stage is powered by a single Merlin vacuum engine nearly identical to the first-stage engines, but modified to operate in the vacuum of space. Like the main Merlin engines, the vacuum engine is designed and manufactured in-house by SpaceX. The engine is designed to burn for about six minutes, and can be shut down and restarted multiple times as needed to deliver different payloads into different orbits. SpaceX's Merlin vacuum engine has the highest vacuum specific impulse (isp)--a measure of engine efficiency--of any American liquid oxygen/kerosene engine with a vacuum isp of 340 seconds. The engine is housed inside the rocket's interstage.
Some "new" (changed 20 days ago) numbers on the SpaceX website (http://www.spacex.com/falcon9).- Merlin 1D vac thrust is now 934kN / 210.000lbf, isp is now 348s, burn time is 397s
- Merlin 1D FT thrust is now 756kN / 170.000lbf (SL) and 825kN / 185.500lbf (vac), burn time is 162s
From a first glance, it seems like the plan is to offload work from the first stage (18s less burning time) to make stage recovery easier.
Also, any idea what would the TWR be now?
I thought that decreasing the thrust of the engines (for the old Merlin 1D version) also decreased their ISP. So it would seem feasible that the full throttle version has a slightly higher ISP. But I might be remembering something wrong.
I thought that decreasing the thrust of the engines (for the old Merlin 1D version) also decreased their ISP. So it would seem feasible that the full throttle version has a slightly higher ISP. But I might be remembering something wrong.It's possible chamber pressure and therefore expansion ratio has increased.
Open cycle engines like gas generator have an issue that above a certain level of Pc, you need to dump more propellant to run the turbines, and thus you increase the efficiency at the Mcc but the lose propellant to the turbines. I would not be surprised that the increase in ISP is more due to the expansion ratio than the Pc.
Of course, the increased thrust would need the longer nozzle to keep the same nozzle exit pressure.
From 240 to 248 in ISP is like what? About ~12 % in fuel savings for the same Dv? According to spaceflight101.com they carried 92.670 kg of fuel on the second stage on the original F9 1.1.
In this case since they have made the tanks larger and added densification... That should do a lot for payload mass to orbit....
Copying from the site.QuoteFalcon 9's second stage is powered by a single Merlin vacuum engine nearly identical to the first-stage engines, but modified to operate in the vacuum of space. Like the main Merlin engines, the vacuum engine is designed and manufactured in-house by SpaceX. The engine is designed to burn for about six minutes, and can be shut down and restarted multiple times as needed to deliver different payloads into different orbits. SpaceX's Merlin vacuum engine has the highest vacuum specific impulse (isp)--a measure of engine efficiency--of any American liquid oxygen/kerosene engine with a vacuum isp of 348 seconds. The engine is housed inside the rocket's interstage.
Regarding older isp numbers, I checked the site with the wayback machine. Here is a link from june 2015 (http://web.archive.org/web/20150519004028/http://www.spacex.com/falcon9). Quoting.QuoteFalcon 9's second stage is powered by a single Merlin vacuum engine nearly identical to the first-stage engines, but modified to operate in the vacuum of space. Like the main Merlin engines, the vacuum engine is designed and manufactured in-house by SpaceX. The engine is designed to burn for about six minutes, and can be shut down and restarted multiple times as needed to deliver different payloads into different orbits. SpaceX's Merlin vacuum engine has the highest vacuum specific impulse (isp)--a measure of engine efficiency--of any American liquid oxygen/kerosene engine with a vacuum isp of 340 seconds. The engine is housed inside the rocket's interstage.
Just poking at this in RPA lite, it seems like you would get some gain from higher Pc at sea level with the same size nozzle. The exhaust is less overexpanded at sea level so you get increased Isp at low altitudes. The same effect reduces Isp when you throttle down at sea level -- exhaust becomes more overexpanded, flow separation makes the nozzle smaller, effectively.You surely meant underexpanded, right? I concur with your assesment for a first stage engine. If the Merlin 1D Full Thrust is exactly the same as the previous (MCC, throat and nozzle dimensions), then it should has a bigger isp bias towards sea level performance. It might increase isp overall, but should be higher at sea level.
Just poking at this in RPA lite, it seems like you would get some gain from higher Pc at sea level with the same size nozzle. The exhaust is less overexpanded at sea level so you get increased Isp at low altitudes. The same effect reduces Isp when you throttle down at sea level -- exhaust becomes more overexpanded, flow separation makes the nozzle smaller, effectively.You surely meant underexpanded, right? I concur with your assesment for a first stage engine. If the Merlin 1D Full Thrust is exactly the same as the previous (MCC, throat and nozzle dimensions), then it should has a bigger isp bias towards sea level performance. It might increase isp overall, but should be higher at sea level.
But I was talking about the vacuum version, which has the enlarged nozzle. There's basically no atmosphere for it and thus is a vacuum analysis.
He meant probably overexpanded nozzle (but wrote overexpanded exhaust).Just poking at this in RPA lite, it seems like you would get some gain from higher Pc at sea level with the same size nozzle. The exhaust is less overexpanded at sea level so you get increased Isp at low altitudes. The same effect reduces Isp when you throttle down at sea level -- exhaust becomes more overexpanded, flow separation makes the nozzle smaller, effectively.You surely meant underexpanded, right? I concur with your assesment for a first stage engine. If the Merlin 1D Full Thrust is exactly the same as the previous (MCC, throat and nozzle dimensions), then it should has a bigger isp bias towards sea level performance. It might increase isp overall, but should be higher at sea level.
But I was talking about the vacuum version, which has the enlarged nozzle. There's basically no atmosphere for it and thus is a vacuum analysis.
He meant probably overexpanded nozzle (but wrote overexpanded exhaust).
About the increase of thrust of merlin 1D, I believe it was obtained optimizing the flow, without increase of turbine power.
Optimizing the pumps, redesigning cooling channels, increasing diameter of pipes, you can minimize friction losses. The 15% gain seems consistent with this, and the good part is you need no extra propellant for turbine.
He meant probably overexpanded nozzle (but wrote overexpanded exhaust).
About the increase of thrust of merlin 1D, I believe it was obtained optimizing the flow, without increase of turbine power.
Optimizing the pumps, redesigning cooling channels, increasing diameter of pipes, you can minimize friction losses. The 15% gain seems consistent with this, and the good part is you need no extra propellant for turbine.
Yes, overexpanded nozzle was what I meant.
So the additional 15% thrust between M1D and "M1D FT" is due to reduced friction losses in the propellant flow? Does this imply a higher Pc at the same pump power? Or just larger mdot at the same pressure?
Optimizing the pumps, redesigning cooling channels, increasing diameter of pipes, you can minimize friction losses. The 15% gain seems consistent with this, and the good part is you need no extra propellant for turbine.SpaceX changing more than they let on again. :P
He meant probably overexpanded nozzle (but wrote overexpanded exhaust).
About the increase of thrust of merlin 1D, I believe it was obtained optimizing the flow, without increase of turbine power.
Optimizing the pumps, redesigning cooling channels, increasing diameter of pipes, you can minimize friction losses. The 15% gain seems consistent with this, and the good part is you need no extra propellant for turbine.
Yes, overexpanded nozzle was what I meant.
So the additional 15% thrust between M1D and "M1D FT" is due to reduced friction losses in the propellant flow? Does this imply a higher Pc at the same pump power? Or just larger mdot at the same pressure?
He meant probably overexpanded nozzle (but wrote overexpanded exhaust).
About the increase of thrust of merlin 1D, I believe it was obtained optimizing the flow, without increase of turbine power.
Optimizing the pumps, redesigning cooling channels, increasing diameter of pipes, you can minimize friction losses. The 15% gain seems consistent with this, and the good part is you need no extra propellant for turbine.
Yes, overexpanded nozzle was what I meant.
So the additional 15% thrust between M1D and "M1D FT" is due to reduced friction losses in the propellant flow? Does this imply a higher Pc at the same pump power? Or just larger mdot at the same pressure?
It's with looking at decreased viscosity at lower temperatures.
Not sure what percentage improvement it would be, but it should be less for both fuel and oxidizer.
It's been a long darn time since my undergrad classes in fluid mechanics, but it seems that I recall the viscosity of many (most?) liquids increases with lower temperature - admittedly, I haven't looked at either RP1 or LOX to see if that's the case.
It's been a long darn time since my undergrad classes in fluid mechanics, but it seems that I recall the viscosity of many (most?) liquids increases with lower temperature - admittedly, I haven't looked at either RP1 or LOX to see if that's the case.
This table of viscosity for oxygen shows LOX viscosity increases with decreasing temperature for all pressures in the relevant regime (0.1MPa to 10MPa). Dynamic viscosity nearly doubles between 90K and 70K while the increase in density is much more modest, so kinematic viscosity must also be increasing going from 90K to 70K.
http://www.nist.gov/srd/upload/jpcrd395.pdf (http://www.nist.gov/srd/upload/jpcrd395.pdf)
Open cycle engines like gas generator have an issue that above a certain level of Pc, you need to dump more propellant to run the turbines, and thus you increase the efficiency at the Mcc but the lose propellant to the turbines. I would not be surprised that the increase in ISP is more due to the expansion ratio than the Pc.
Of course, the increased thrust would need the longer nozzle to keep the same nozzle exit pressure.
Just poking at this in RPA lite, it seems like you would get some gain from higher Pc at sea level with the same size nozzle. The exhaust is less overexpanded at sea level so you get increased Isp at low altitudes. The same effect reduces Isp when you throttle down at sea level -- exhaust becomes more overexpanded, flow separation makes the nozzle smaller, effectively.
Open cycle engines like gas generator have an issue that above a certain level of Pc, you need to dump more propellant to run the turbines, and thus you increase the efficiency at the Mcc but the lose propellant to the turbines. I would not be surprised that the increase in ISP is more due to the expansion ratio than the Pc.
Of course, the increased thrust would need the longer nozzle to keep the same nozzle exit pressure.
Just poking at this in RPA lite, it seems like you would get some gain from higher Pc at sea level with the same size nozzle. The exhaust is less overexpanded at sea level so you get increased Isp at low altitudes. The same effect reduces Isp when you throttle down at sea level -- exhaust becomes more overexpanded, flow separation makes the nozzle smaller, effectively.
I wonder if it makes sense to throttle down as the atmosphere thins during ascent, once you get into regime where the acceleration is already relatively high and gravity losses much lower? Assumes there is a point where Isp becomes more important than gravity losses (and that GG is in the regime that baldusi describes).
Cheers, Martin
Do I understand correctly? Overexpanded first stage engines increase their ISP when throttled to be less overexpanded when the atmosphere gets thinner on altitude compared to not throttled?
Throttling on merlin 1D is done dosing propellants to the turbopump.
Likely at lower thrust not only the mass of propellants to TP is less, but also the percentage on total flow.
Therefore is possible to have a slight increase in ISP.
As ever IMHO.
Both M1-D+ and M1-D+ vac are turbopump-fed gas generator engines. In the vaccuum version, the pump exhaust is fed back to the nozzle, but in the S1 version it is not.Probably not. In fact the proposal for an updated F-1B engine would have done away with this as well, see http://arstechnica.com/science/2013/04/new-f-1b-rocket-engine-upgrades-apollo-era-deisgn-with-1-8m-lbs-of-thrust/
Would it make sense to design a future variant of the S1 engine to also do that (like the F1 engine did for example)? What are the possible advantages or disadvantages of a variation like that (in engine complexity/mass/reliability/performance and octaweb complexity/mass)?
Has SpaceX been clear on the question of whether all M1-D engines can be qualified for full thrust flight, or only M1-D engines produced since a certain date, or?
Has SpaceX been clear on the question of whether all M1-D engines can be qualified for full thrust flight, or only M1-D engines produced since a certain date, or?
Has SpaceX been clear on the question of whether all M1-D engines can be qualified for full thrust flight, or only M1-D engines produced since a certain date, or?
We know that Musk doesn't like type designations, but they have labeled the Merlin engines 1A, 1B, 1C, and 1D. To have two "flavors" of Merlin 1D with different thrusts would be very strange.
A reasonable assumption, but only an assumption, is that engines sharing a single designation are more or less interchangeable.
My guess, but its only a guess, would be that the Jason F9-1.1 engines could be run at full thrust
(snip)
Of course now we know the full potential (power) needed the colder RP-1 & LOX,
(snip)
My guess is that there are no major differences between the FT and non-FT versions of the M1D. If true, then the engines recovered from the Jason S1, if recovered, together with any spares could be repurposed for a Full Thrust stage with minor refurbishment. If the combustion chamber is different, that would not be possible.
(snip)
Of course now we know the full potential (power) needed the colder RP-1 & LOX,
(snip)
No. We don't really know that either.
Sub-cooling could be for just the increased density. SpaceX has stated the benefit of increased propellant density.
F9-FT needs sub-cooled P1 & LOX AND "Full Thrust" engines but we don't know that one leads to the other.
Coincidence is not causality.
We don't have that level of detail.
The question now that a set of M1DFT's have flow will we ever find out the real performance values for these engines?
There are some differences on the FT version. I wish I could define them here but I can't. Eventually I am betting SX will release a picture (like we have seen of legacy M1Ds in various stages of build on the factory floor) and the visual clues will be obvious to the keen eye.One obvious candidate would be slightly large plumbing to deal the the higher mass flow. Pictures when they happen will be interesting.
Can anyone give an update on information regarding the ablative coating of the M1D engines? I remember watching a video of the M1A a few years ago which was one of the first tests but I wondered if the technology has improved and what advancements in this area would aid to rapid reusability?
I know back on Jan 16th, 2014 that SpaceX fitted and launched an additive printed Main Oxidizer Valve on one of the nine booster Merlin-1D engines. I was wondering if there's been any update on this 'test' and if they are now using this version of the MOV (and any other printed components) on the Merlin-1D?
Source: http://www.spacex.com/news/2014/07/31/spacex-launches-3d-printed-part-space-creates-printed-engine-chamber-crewed (http://www.spacex.com/news/2014/07/31/spacex-launches-3d-printed-part-space-creates-printed-engine-chamber-crewed)
There are some differences on the FT version. I wish I could define them here but I can't. Eventually I am betting SX will release a picture (like we have seen of legacy M1Ds in various stages of build on the factory floor) and the visual clues will be obvious to the keen eye.
If the supercooled propellant is denser, then a greater weight of propellant can occupy the same volume. So why should anything be larger, which increases weight, however slightly? There may be changes in the engines that are less obvious--but if the engines were designed for supercooled propellants prior to this flight, perhaps not. I'm guessing more turbopump power is needed to stuff the additional mass of propellant into the combustion chamber, and the lower temperature of the propellants may have an effect on combustion efficiency (and may help cooling the engine, making the performance margin possible).There are some differences on the FT version. I wish I could define them here but I can't. Eventually I am betting SX will release a picture (like we have seen of legacy M1Ds in various stages of build on the factory floor) and the visual clues will be obvious to the keen eye.One obvious candidate would be slightly large plumbing to deal the the higher mass flow. Pictures when they happen will be interesting.
Landed M1D FT cluster up-close
https://twitter.com/SpaceX/status/684094378021801984
Landed M1D FT cluster up-close, plus a brightened version
Does anybody know what parts of the nozzles on the first stages engine are made of what?Best candidate IMHO is stainless steel.
Does anybody know what parts of the nozzles on the first stages engine are made of what?Best candidate IMHO is stainless steel.
You are right, could be chrome plated copper.Does anybody know what parts of the nozzles on the first stages engine are made of what?Best candidate IMHO is stainless steel.
The picture I just posted looks like the inner layer with the channels milled in it is copper. Could be that the inner and outer layers are some kind of nickel superalloy for heat resistance. Which makes the overall construction not too different from the stainless/copper/stainless layering used in some cooking pots ...
You just answered your own question. Denser=Thicker=More-Pumping-Loss. Either increase pump power or decrease pumping loss. If the turbopump is maxed out then slightly larger pipe are another answer.If the supercooled propellant is denser, then a greater weight of propellant can occupy the same volume. So why should anything be larger, which increases weight, however slightly? There may be changes in the engines that are less obvious--but if the engines were designed for supercooled propellants prior to this flight, perhaps not. I'm guessing more turbopump power is needed to stuff the additional mass of propellant into the combustion chamber, and the lower temperature of the propellants may have an effect on combustion efficiency (and may help cooling the engine, making the performance margin possible).There are some differences on the FT version. I wish I could define them here but I can't. Eventually I am betting SX will release a picture (like we have seen of legacy M1Ds in various stages of build on the factory floor) and the visual clues will be obvious to the keen eye.One obvious candidate would be slightly large plumbing to deal the the higher mass flow. Pictures when they happen will be interesting.
So there might be changes, yes, but I suspect they are subtle.
--Damon
- when the M1D or any similar engine throttles down, how is this throttling actually accomplished?
A few questions that perhaps the experts here can answer:For M1D we know for sure (after CRS-6) that a bipropellant valve regulates the flow to the preburner, therefore effectively controlling the RPM of the turbopump.
- when the M1D or any similar engine throttles down, how is this throttling actually accomplished? Is it by reducing the RPM of the turbo pumps, partially closing a valve or pintle, or combination?
- how does it ensure the proper combustion ratio is maintained, despite a possible change in the LOX temperature during a flight, and resulting expansion? Does it dynamically adjust the flow of LOX volume to account for different density?
The densification is about 10%, IIRC, which will have some effect on the energy used to pump it, but it seems the more significant effects of densification relate to controlling the combustion mixture.
Not that I'm an expert or anything.
only the center engine has a bi prop valve in order to throttle for landing ops...
only the center engine has a bi prop valve in order to throttle for landing ops...
^ "former SpaceX technician"
good enough for me
if only one engine throttles you still get a reduced 100 - 90% range eh
or
do we think that the typical 1d can do 100-70 and perhaps the middle engine has a greater range?
^ "former SpaceX technician"
good enough for me
if only one engine throttles you still get a reduced 100 - 90% range eh
or
do we think that the typical 1d can do 100-70 and perhaps the middle engine has a greater range?
We did see some deep throttling in Grasshopper. Made folks wonder if the M1D could throttle more than advertised.
wikipedia says the 1d-vac can do 39%?
We seem to be short on solid facts here.
That still doesn't answer whether it's ONLY the center engine that throttles, and if so, what is the real benefit of the Falcon Heavy configuration?
Maybe I was not so clear but I didn't state the engines as a group don't have the ability to throttle up and down to a degree...only that E9, the center positioned engine has a actual added valve (the bi prop valve) and sensor package for throttling. The other positions may or may not have the same or lesser degrees of throttling as well. ;)So you are saying there are _four_ flavors of 1D engine? One for vacuum, one for center engine, one for two outer engines used for SSRP and one for the rest?
Maybe I was not so clear but I didn't state the engines as a group don't have the ability to throttle up and down to a degree...only that E9, the center positioned engine has a actual added valve (the bi prop valve) and sensor package for throttling. The other positions may or may not have the same or lesser degrees of throttling as well. ;)So you are saying there are _four_ flavors of 1D engine? One for vacuum, one for center engine, one for two outer engines used for SSRP and one for the rest?
I believe Both Musk and Shotwell have publicly stated their desire to minimize variants - for both manufacturing and system architecture perspectives.
I suggest one and only one Marlin-1D SL flavor. It's restartable from an external (to engine) TEA/TEB system, and has the full throttle range afforded to a 1D.
Let me ask this - it's commonly understood that engine start during launch is via GSE, correct? What's that architecture? It certainly doesn't have nine pairs of lines running from GSE to individual engines, right? So if you accept that, then you must also accept that the engine start plumbing to each engine comes from a single point on the rocket. (Let's call it a distribution manifold - see where this is going?). Now it's also accepted that three engines are also restarted via onboard systems. So does it make sense to have a entire second system for the restarts designed into certain engines, with all that extra unneeded redundancy and complication, or would it make more sense to utilize the same plumbing and manifold structure already in place for the GSE fed start for the relights?
So the addition of valves to direct the relights to particular engine(s) provides a uniform architecture (only one flavor) and also provides opportunity to go to another outside opposing pair if the initial pair has issues relighting.
What has to be different between center and outside engines is not throttle, but gimbal.
What has to be different between center and outside engines is not throttle, but gimbal.
Gimbal has to be different.
(What is so difficult to understand in this?)
Hint: not hardware for gimbal has to be different; not no gimbal.
Gimbal should be limited:
-by software, the engine controller is fixed to the “thrust plate“, different between center and outboard engines.
-by hardware, stroke limiting clamps are visible on actuators of outboard engines.
stroke limiting clamps are visible on actuators of outboard engines.
stroke limiting clamps are visible on actuators of outboard engines.
Those appear to be electrical harness clamps, not mechanical stops, which are typically internal to TVC actuators and not externally visible.
Anyway, drifting off topic. The point is, there's no reason to doubt the statement that all F9 first stage engines are identical. The differences being discussed above (throttling, gimbal angle input limits) are all *external* to the engine itself.
That still doesn't answer whether it's ONLY the center engine that throttlesI don't see that being a question in the first place. Back in the day when Merlin 1C couldn't throttle they had to throttle by shutting off two engines to keep the acceleration within limits. Throttling down one engine cannot provide reduction in thrust similar to shutting off two engines. Therefore all engines can throttle.
That still doesn't answer whether it's ONLY the center engine that throttlesI don't see that being a question in the first place. Back in the day when Merlin 1C couldn't throttle they had to throttle by shutting off two engines to keep the acceleration within limits. Throttling down one engine cannot provide reduction in thrust similar to shutting off two engines. Therefore all engines can throttle.
Gimbal should be limited:
-by software, the engine controller is fixed to the “thrust plate“, different between center and outboard engines.
-by hardware, stroke limiting clamps are visible on actuators of outboard engines.
They all are able to throttle; I assume the center one is more precise or has a faster response due to the added biprop valve, and gives faster feedback about the actual performance due to added instrumentation. This does not necessarily change anything about the engine structure itself, but is more like an added package only installed for the center engine.That still doesn't answer whether it's ONLY the center engine that throttlesI don't see that being a question in the first place. Back in the day when Merlin 1C couldn't throttle they had to throttle by shutting off two engines to keep the acceleration within limits. Throttling down one engine cannot provide reduction in thrust similar to shutting off two engines. Therefore all engines can throttle.
Can we assume then that the center engine throttles more than the others?
Is there any known or expected difference between M1Dvac and M1D engines (except different nozzle)? In other words, would it be possible to take "worn out" engine from reusable S1, fit vacuum nozzle and mount it as S2 engine?
Is there any known or expected difference between M1Dvac and M1D engines (except different nozzle)? In other words, would it be possible to take "worn out" engine from reusable S1, fit vacuum nozzle and mount it as S2 engine?
Is there any known or expected difference between M1Dvac and M1D engines (except different nozzle)? In other words, would it be possible to take "worn out" engine from reusable S1, fit vacuum nozzle and mount it as S2 engine?
The thrust chamber / nozzle interface does not appear to be identical.
http://www.spacex.com/sites/spacex/files/shiny_merlin_edited.jpg
http://www.spacex.com/files/assets/img/merlinvac.jpg
Is there any known or expected difference between M1Dvac and M1D engines (except different nozzle)? In other words, would it be possible to take "worn out" engine from reusable S1, fit vacuum nozzle and mount it as S2 engine?
The thrust chamber / nozzle interface does not appear to be identical.
http://www.spacex.com/sites/spacex/files/shiny_merlin_edited.jpg
http://www.spacex.com/files/assets/img/merlinvac.jpg
Yes, those two photographs are very helpful in showing for the Merlin 1D and Merlin 1Dvac the same sorts of differences that were shown above in the excellent post by Lars-J.
Perhaps one of you, or another image-savvy editor, could make a side-by-side single image of those two, so that Lars-J's point could be well illustrated with the 1D technology.
Is there any known or expected difference between M1Dvac and M1D engines (except different nozzle)? In other words, would it be possible to take "worn out" engine from reusable S1, fit vacuum nozzle and mount it as S2 engine?
The thrust chamber / nozzle interface does not appear to be identical.
http://www.spacex.com/sites/spacex/files/shiny_merlin_edited.jpg
http://www.spacex.com/files/assets/img/merlinvac.jpg
Yes, those two photographs are very helpful in showing for the Merlin 1D and Merlin 1Dvac the same sorts of differences that were shown above in the excellent post by Lars-J.
Perhaps one of you, or another image-savvy editor, could make a side-by-side single image of those two, so that Lars-J's point could be well illustrated with the 1D technology.
Here's the best I could do. The angles are different, but I tried to size them such that the combustion chamber is the same size.
I had not even been thinking of having pics of engines being fired, just side-by-sides of the two jpg photos that dkovacic had inserted into the discussion a few posts ago (shiny engines on the factory floor):
http://www.spacex.com/sites/spacex/files/shiny_merlin_edited.jpg
http://www.spacex.com/files/assets/img/merlinvac.jpg
... which would still look great in a side-by-side comparison image.
I had not even been thinking of having pics of engines being fired, just side-by-sides of the two jpg photos that dkovacic had inserted into the discussion a few posts ago (shiny engines on the factory floor):
http://www.spacex.com/sites/spacex/files/shiny_merlin_edited.jpg
http://www.spacex.com/files/assets/img/merlinvac.jpg
... which would still look great in a side-by-side comparison image.
That last image is a M1C-Vac, not M1D-Vac. As you should be able to tell, the M1C-Vac is quite different than the M1D-Vac. That image of an M1DVac firing is the only public photo of the M1DVac that I have been able to find. (excluding rocket cam footage)
If you can find a shiny M1D-Vac photo from the factory floor I would be happy to use it. :)
I told a SpaceX engineer I met a few days ago I could build them a variable geometry nozzle out of an old Buick Roadmaster hood and two used garage door openers to help increase the throttle range of the M1D. I'm pretty sure he was serious when he said they'd think about it.
I had not even been thinking of having pics of engines being fired, just side-by-sides of the two jpg photos that dkovacic had inserted into the discussion a few posts ago (shiny engines on the factory floor):
http://www.spacex.com/sites/spacex/files/shiny_merlin_edited.jpg
http://www.spacex.com/files/assets/img/merlinvac.jpg
... which would still look great in a side-by-side comparison image.
That last image is a M1C-Vac, not M1D-Vac. As you should be able to tell, the M1C-Vac is quite different than the M1D-Vac. That image of an M1DVac firing is the only public photo of the M1DVac that I have been able to find. (excluding rocket cam footage)
If you can find a shiny M1D-Vac photo from the factory floor I would be happy to use it. :)
Okay, thanks for clarifying. I had understood that they were both M1D/M1Dvac engines from whomever originally posted the images.
Looking at the pictures of the ORBCOMM-2 engines from the returned core it seems that this coating might well have made it all the way to the final version of the engine, do we know if this is the case?
Hadn't seen that either and yes, that's a big nozzle. Like the sign they have up.
https://www.youtube.com/watch?v=6glAvN5APh4 (https://www.youtube.com/watch?v=6glAvN5APh4)
Hadn't seen that separation test around 0:36.
Damn, that new nozzle is big! You can also see the added center pusher (in black).
Try over here. (http://forum.nasaspaceflight.com/index.php?topic=39182.msg1483572#msg1483572)Hadn't seen that separation test around 0:36.
Damn, that new nozzle is big! You can also see the added center pusher (in black).
I get "This video has been removed by the user." Anyone get a rip? Or at the least, explain what was in the video?
Thanks.
Does anyone have the current price/cost of the Merlin 1D? DoD mentioned that the RD-180 costs ULA 30 million today and the last figure I heard about the 1D was 2 million each.
Maybe I was not so clear but I didn't state the engines as a group don't have the ability to throttle up and down to a degree...only that E9, the center positioned engine has a actual added valve (the bi prop valve) and sensor package for throttling. The other positions may or may not have the same or lesser degrees of throttling as well. ;)So you are saying there are _four_ flavors of 1D engine? One for vacuum, one for center engine, one for two outer engines used for SSRP and one for the rest?
I believe Both Musk and Shotwell have publicly stated their desire to minimize variants - for both manufacturing and system architecture perspectives.
I suggest one and only one Marlin-1D SL flavor. It's restartable from an external (to engine) TEA/TEB system, and has the full throttle range afforded to a 1D.
Let me ask this - it's commonly understood that engine start during launch is via GSE, correct? What's that architecture? It certainly doesn't have nine pairs of lines running from GSE to individual engines, right? So if you accept that, then you must also accept that the engine start plumbing to each engine comes from a single point on the rocket. (Let's call it a distribution manifold - see where this is going?). Now it's also accepted that three engines are also restarted via onboard systems. So does it make sense to have a entire second system for the restarts designed into certain engines, with all that extra unneeded redundancy and complication, or would it make more sense to utilize the same plumbing and manifold structure already in place for the GSE fed start for the relights?
So the addition of valves to direct the relights to particular engine(s) provides a uniform architecture (only one flavor) and also provides opportunity to go to another outside opposing pair if the initial pair has issues relighting.
Landed M1D FT cluster up-close
https://twitter.com/SpaceX/status/684094378021801984
higher res, using the :orig tag.
The new user's guide covers both F9 and FH: http://www.spacex.com/sites/spacex/files/falcon_9_users_guide_rev_2.0.pdf (http://www.spacex.com/sites/spacex/files/falcon_9_users_guide_rev_2.0.pdf)
Landed M1D FT cluster up-close
https://twitter.com/SpaceX/status/684094378021801984
higher res, using the :orig tag.
Lots of asymmetrical accretion of crusties* both inside and outside the nozzles, plus more at the mounting - that should give a lot of useful information by pattern and compositional analysis.
*First estimate is primarily mixed metal oxides, but I'd need to start with some FTIR and ICP-MS (maybe LIBZ, if they have calibration curves for their materials. Manufacturer curves tend to be good on steel and aluminum alloys) to get an outline of composition.
Landed M1D FT cluster up-close
https://twitter.com/SpaceX/status/684094378021801984
higher res, using the :orig tag.
Lots of asymmetrical accretion of crusties* both inside and outside the nozzles, plus more at the mounting - that should give a lot of useful information by pattern and compositional analysis.
*First estimate is primarily mixed metal oxides, but I'd need to start with some FTIR and ICP-MS (maybe LIBZ, if they have calibration curves for their materials. Manufacturer curves tend to be good on steel and aluminum alloys) to get an outline of composition.
To my highly untrained eyes it looks like residue/ash from very slow burning kerosene (hypoxic environment) that was dripping down from the injector after shutdown. (Notice that it's primarily on the two outboard engines probably of the three engine reentry burn.) My explanation for why it's not on the center engine is that the center engine refired again for landing and had plenty of atmosphere to more cleanly burn off the kerosene at sea level.
This has been discussed elsehwhere. It's likely the oxides of aluminum and boron, residue from the TEA/TEB ignition fluid. The TEA/TEB ignition fluid is probably injected into all three engines at every restart event for simplicity of plumbing design, even though the landing burn uses only the center engine.
The landing burn burns off the TEA/TEB residue in the center engine, but since the outer two engines are not used in the landing burn, the TEA/TEB residue from the landing burn restart remains in those two nozzles.
Elon tweeted this a few minutes ago:40% to 112% is a 65% range. (Can't remember if that 112% is current for the "full but not 1.7 mblf thrust" engine?)
https://twitter.com/elonmusk/status/728753234811060224
"@lukealization Max is just 3X Merlin thrust and min is ~40% of 1 Merlin. Two outer engines shut off before the center does."
There had been some disagreement about the throttle range of the M1D - did it have a 40% range or could it throttle down to 40%? But this confirms it - it can throttle down to 40%. (A 60% range)
Any new info on the further improvement of the Merlin 1D?
https://en.wikipedia.org/wiki/Merlin_%28rocket_engine_family%29#Merlin_1D
In May 2016, SpaceX announced plans to further uprate the Merlin 1D by increasing vacuum thrust to 914 kN and sea-level thrust to 845 kN.
Looking at the photos of aft end... it's amazing outer engine bells can work while so closely spaced. With all the vibration, I swear it looks like this should not be possible :D
During the Thaicom-8 launch coverage for regular folks, one of the announcers gave very specific thrust numbers for the Merlin - something like, "it can throttle from 80,000 lbf to 210,000 lbf". I believe he gave even more digits of precision, that's my rough memory.
I'm too lazy to rewatch it to get the exact quote.
I'm sure you all were watching the Technical Broadcast and missed it. ::)
During the Thaicom-8 launch coverage for regular folks, one of the announcers gave very specific thrust numbers for the Merlin - something like, "it can throttle from 80,000 lbf to 210,000 lbf". I believe he gave even more digits of precision, that's my rough memory.
I'm too lazy to rewatch it to get the exact quote.
I'm sure you all were watching the Technical Broadcast and missed it. ::)
I remember that, I believe that was for the M-Vac, which if I recall has a larger throttle range than the first stage Merlin
During the Thaicom-8 launch coverage for regular folks, one of the announcers gave very specific thrust numbers for the Merlin - something like, "it can throttle from 80,000 lbf to 210,000 lbf". I believe he gave even more digits of precision, that's my rough memory.
I'm too lazy to rewatch it to get the exact quote.
I'm sure you all were watching the Technical Broadcast and missed it. ::)
I remember that, I believe that was for the M-Vac, which if I recall has a larger throttle range than the first stage Merlin
You are both correct. It was 80,000 lbf - 210,000 lbf for the M-Vac. It is at 49:00 in the hosted webcast video which is T+27:32 on the countdown clock.
Elon Musk – Verified account @elonmusk
@lukealization Max is just 3X Merlin thrust and min is ~40% of 1 Merlin. Two outer engines shut off before the center does.
Not both correct. That range of throttling (down to 40%) can also be achieved by the M1D according to a recent Elon tweet in reply to a question about the 3-engine landing burn. He said center engine throttles down to 40% for landing.I wonder how much the back-pressure in the bells from the 300mph+ descent affects this?
Not both correct. That range of throttling (down to 40%) can also be achieved by the M1D according to a recent Elon tweet in reply to a question about the 3-engine landing burn. He said center engine throttles down to 40% for landing.I wonder how much the back-pressure in the bells from the 300mph+ descent affects this?
There isn't any "back pressure" after the engine restarts, so there's no aero effect on engine thrust or throttling.Thank you for the reminder of the implications of supersonic flow. My still slightly sleepy mind was thinking along the lines of back pressure increasing the chance of flow separation further down the nozzle, and thereby reducing the throttleability of the motors.
Exhaust flow after the nozzle throat is supersonic, which means no information about external pressure can flow upstream back into the nozzle. Conditions inside the engine/nozzle are exactly the same regardless of external atmospheric conditions (except for low ambient pressure that could cause flow separation inside the nozzle, but that's a different condition.)
If the stage is supersonic during retropropulsion, a bow shock will form outside the nozzle, but that's still downstream of the exit plane and so doesn't affect engine thrust. If subsonic, no bow shock, and same logic applies.
The supersonic flow is at the throat, not at the nozzle plane. So it might impact a bit and you could suffer from flow separation at the nozzle tip. Obviously not close to the throat.Not both correct. That range of throttling (down to 40%) can also be achieved by the M1D according to a recent Elon tweet in reply to a question about the 3-engine landing burn. He said center engine throttles down to 40% for landing.I wonder how much the back-pressure in the bells from the 300mph+ descent affects this?
There isn't any "back pressure" after the engine restarts, so there's no aero effect on engine thrust or throttling.
Exhaust flow after the nozzle throat is supersonic, which means no information about external pressure can flow upstream back into the nozzle. Conditions inside the engine/nozzle are exactly the same regardless of external atmospheric conditions (except for low ambient pressure that could cause flow separation inside the nozzle, but that's a different condition.)
If the stage is supersonic during retropropulsion, a bow shock will form outside the nozzle, but that's still downstream of the exit plane and so doesn't affect engine thrust. If subsonic, no bow shock, and same logic applies.
The back pressure actually probably improves things. It's easier to light an engine at pressure than in vacuum.Not both correct. That range of throttling (down to 40%) can also be achieved by the M1D according to a recent Elon tweet in reply to a question about the 3-engine landing burn. He said center engine throttles down to 40% for landing.I wonder how much the back-pressure in the bells from the 300mph+ descent affects this?
The supersonic flow is at the throat, not at the nozzle plane. So it might impact a bit and you could suffer from flow separation at the nozzle tip. Obviously not close to the throat.Not both correct. That range of throttling (down to 40%) can also be achieved by the M1D according to a recent Elon tweet in reply to a question about the 3-engine landing burn. He said center engine throttles down to 40% for landing.I wonder how much the back-pressure in the bells from the 300mph+ descent affects this?
There isn't any "back pressure" after the engine restarts, so there's no aero effect on engine thrust or throttling.
Exhaust flow after the nozzle throat is supersonic, which means no information about external pressure can flow upstream back into the nozzle. Conditions inside the engine/nozzle are exactly the same regardless of external atmospheric conditions (except for low ambient pressure that could cause flow separation inside the nozzle, but that's a different condition.)
If the stage is supersonic during retropropulsion, a bow shock will form outside the nozzle, but that's still downstream of the exit plane and so doesn't affect engine thrust. If subsonic, no bow shock, and same logic applies.
Well, separation occurs if Pa>Pe, (ambient pressure larger than exit pressure) not the opposite.
Well, separation occurs if Pa>Pe, (ambient pressure larger than exit pressure) not the opposite.True, though you can get away with a little Pa>Pe. I think Merlin 1D is slightly under-expanded on take-off.
You have to realize that the SL low throttle thrust if it was a M1C would be 77% of the M1C SL. The exit plane area of the M1D and M1C is the same. What is different is the the M1D has a smaller throat and higher TC pressures. So throttling down to just below the exit plane velocities that an M1C had should work fine without much problems in SL pressure conditions.
Almost all first stage engines are over-expanded at liftoff and under-expanded at MECO. At launch, Pa>Pe, and an oblique shock at the exit plane causes the flow to "neck down" as you always saw in the close-ups of SSMEs at ignition. (These oblique shocks also give rise to the classic "shock diamonds" in the exhaust.) Pa has to be more than twice Pe for the flow to actually be separated, which means the oblique shocks move up into the nozzle and the flow lines depart from the nozzle up inside, which is very inefficient and greatly to be avoided. First stage engines are ideally expanded (Pa=Pe) at only one altitude, and at that point the flow of the exhaust will perfectly match the exit angle of the nozzle. Above that altitude, the nozzle is under-expanded, and an expansion wave at the exhaust will cause the flow to bend outwards and form the large "fans" you see from about Max-Q to MECO. Upper stage engines are always under-expanded, because the nozzle would have to be infinitely long to match exit pressure to the vacuum ambient conditions. The nozzles are as long as practical to get the best performance in vacuum, though.Well, separation occurs if Pa>Pe, (ambient pressure larger than exit pressure) not the opposite.True, though you can get away with a little Pa>Pe. I think Merlin 1D is slightly under-expanded on take-off.
According to the info we have available.
M9
Thrust :.
SL 845 kN - (190,000 lbf)
Vac 914 kN - (205,500 lbf)
Weight :. 467kg (1030 lb) together with the actuators.
TWR :. SL 184.4, Vac 199.5
Throttle range :. 338 kN - 914 kN (76,000 lbf - 205,500 lbf)
MVac
Thrust :. Vac 934 kN (210,000 lbf)
Throttle range :. 356kN - 934 kN (80,000 lbf - 210,000 lbf)
And both sets of thrust values represent sligtly more than 60% reductions from max to min, showing throttling to about 37% and 38% for the M1 and M1Vac respectively.
Either these are excessive or Musk rounded off when stating throttling to "40%".
According to the info we have available.
M9
Thrust :.
SL 845 kN - (190,000 lbf)
Vac 914 kN - (205,500 lbf)
Weight :. 467kg (1030 lb) together with the actuators.
TWR :. SL 184.4, Vac 199.5
Throttle range :. 338 kN - 914 kN (76,000 lbf - 205,500 lbf)
MVac
Thrust :. Vac 934 kN (210,000 lbf)
Throttle range :. 356kN - 934 kN (80,000 lbf - 210,000 lbf)
There really isn't anything designated "M9", is there?
Perhaps that's a typo for M1?
And both sets of thrust values represent sligtly more than 60% reductions from max to min, showing throttling to about 37% and 38% for the M1 and M1Vac respectively.
Either these are excessive or Musk rounded off when stating throttling to "40%".
Questions...
Are they all in, all done at 190,000 sl-lbf... ??
Have we now seen the final top thrust numbers planned on SpaceX Merlin series rocket engines... ??
Will they run S1 at that 9 x 190K all the time off the pad... or only when needed... ??
Will this now get them to the published 5500kg to GTO-1800 with ASDS recovery and S1 in reusable condition target shown on their website... ??
???
Questions...
Are they all in, all done at 190,000 sl-lbf... ??
Have we now seen the final top thrust numbers planned on SpaceX Merlin series rocket engines... ??
Will they run S1 at that 9 x 190K all the time off the pad... or only when needed... ??
Will this now get them to the published 5500kg to GTO-1800 with ASDS recovery and S1 in reusable condition target shown on their website... ??
???
I'll try to answer some of that.
Are they all in, all done at 190,000 sl-lbf... ??
Have we now seen the final top thrust numbers planned on SpaceX Merlin series rocket engines... ??
No idea. They seem to keep upping performance as they get more data and knowledge on the engine and what it can do. It would not be absurd to think they will keep working on it with performance, cost and re-usability goals in mind.
Will they run S1 at that 9 x 190K all the time off the pad... or only when needed... ??
There is no reason to launch a vehicle @less than max thrust/throttle. Ever.
Will this now get them to the published 5500kg to GTO-1800 with ASDS recovery and S1 in reusable condition target shown on their website... ??
That is their official advertisement on their site, and what they are aiming for.
Will they run S1 at that 9 x 190K all the time off the pad... or only when needed... ??
There is no reason to launch a vehicle @less than max thrust/throttle. Ever.
Elon Musk @elonmusk Apr 30
F9 thrust at liftoff will be raised to 1.71M lbf later this year. It is capable of 1.9M lbf in flight.
Will they run S1 at that 9 x 190K all the time off the pad... or only when needed... ??
There is no reason to launch a vehicle @less than max thrust/throttle. Ever.
..... Musk stated that liftoff thrust will be less than "capable" in flight thrust. Whether that means there will be a throttle-up after liftoff isn't clear, but it seems possible.QuoteElon Musk @elonmusk Apr 30
F9 thrust at liftoff will be raised to 1.71M lbf later this year. It is capable of 1.9M lbf in flight.
That is just the difference between performance of 9 Merlin engines at sea level and in vacuum*That makes sense. FH will be throttling down the center core in flight, implying that it lifts off at or near open throttle, so a single F9 with 1/3 the thrust shouldn't have any pad issues at open throttle.
* based on published numbers at SpaceX website.
For the same exit area and a higher expansion ratio the throat has to be smaller. M1D has a significant larger expansion ratio over that of the M1C. Exit area/throat area = expansion ratio.You have to realize that the SL low throttle thrust if it was a M1C would be 77% of the M1C SL. The exit plane area of the M1D and M1C is the same. What is different is the the M1D has a smaller throat and higher TC pressures. So throttling down to just below the exit plane velocities that an M1C had should work fine without much problems in SL pressure conditions.
Is there evidence for this? I've not seen that particular assertion before.
Will they run S1 at that 9 x 190K all the time off the pad... or only when needed... ??
There is no reason to launch a vehicle @less than max thrust/throttle. Ever.
Will they run S1 at that 9 x 190K all the time off the pad... or only when needed... ??And yet, SSME max thrust was 109%, but it launched at 104.5% because that was the maximum they were comfortable with unless the situation was so desperate that this was the lesser risk.
There is no reason to launch a vehicle @less than max thrust/throttle. Ever.
Will they run S1 at that 9 x 190K all the time off the pad... or only when needed... ??
There is no reason to launch a vehicle @less than max thrust/throttle. Ever.
And yet, SSME max thrust was 109%, but it launched at 104.5% because that was the maximum they were comfortable with unless the situation was so desperate that this was the lesser risk.
Cheers, Martin
Why would NASA run the SSMEs at lower throttle? Was it because the thrust was angled, or because the SSMEs were way over expanded at SL?
Why would NASA run the SSMEs at lower throttle? Was it because the thrust was angled, or because the SSMEs were way over expanded at SL?
MP99 touched on it--mission and system managers determined the SSME throttle level was the best for safety and performance (and in that order).
...
Questions...
Are they all in, all done at 190,000 sl-lbf... ??
Have we now seen the final top thrust numbers planned on SpaceX Merlin series rocket engines... ??
When you throttle down an engine you reduce the chamber pressure and the efficiency of the engine quickly deteriorates. Correct? Much better shutting off one of the engines than throttling all engines down. Of course that raises the issue of restarting the engine...
When you throttle down an engine you reduce the chamber pressure and the efficiency of the engine quickly deteriorates. Correct?
When you throttle down an engine you reduce the chamber pressure and the efficiency of the engine quickly deteriorates. Correct?
Partly correct. The chamber pressure decreases, but engine efficiency does not quickly deteriorate.
The Merlin, for example, only loses about 15% of it's specific impulse when throttled to 40% thrust at sea level. When Merlin Vacuum is throttled to 30% of max thrust, it only loses about 1% of of it's specific impulse.
Source: model of known Merlin parameters in RPA lite.
Wow, so we've learned that SpaceX will be phasing out single engine testing at McGregor! (Gwynne, end of Q&A, Small Sat Conference)
We are likely to go away from single engine tests on Merlin, you know, once we have finalized the design and show great decrease variability, i say, greater non variability, decrease variabilityMy emphasis
Should be a nice decrease in costs and improvement in flow.
Talk of dropping the static fire another simplification that is discussed -- no timing known, but could flow from the same logic.
Did not see this here..
From SpaceX facebook by Mike HawesQuoteWow, so we've learned that SpaceX will be phasing out single engine testing at McGregor! (Gwynne, end of Q&A, Small Sat Conference)
Have not seen the video of the conference yet.
Edit: Just finished watching the video at about 1:01:46:QuoteWe are likely to go away from single engine tests on Merlin, you know, once we have finalized the design and show great decrease variability, i say, greater non variability, decrease variabilityMy emphasis
Should be a nice decrease in costs and improvement in flow.
Talk of dropping the static fire another simplification that is discussed -- no timing known, but could flow from the same logic.
So do they test all 9 at once on the stage at Macgregor? Or just in the static fire?
Interesting that even with all the accumulated manufacturing and testing knowledge they've built up making many hundreds of Merlin engines, they still have much variability at all. I guess the constant design iterations along the way are to blame but still interesting that (apparently) acceptance testing is used to characterize the engines (individually? by lot?).
It is difficult to imagine that they have a 'reject' rate much over zero.
Did not see this here..
From SpaceX facebook by Mike HawesQuoteWow, so we've learned that SpaceX will be phasing out single engine testing at McGregor! (Gwynne, end of Q&A, Small Sat Conference)
Have not seen the video of the conference yet.
Edit: Just finished watching the video at about 1:01:46:QuoteWe are likely to go away from single engine tests on Merlin, you know, once we have finalized the design and show great decrease variability, i say, greater non variability, decrease variabilityMy emphasis
Interesting that even with all the accumulated manufacturing and testing knowledge they've built up making many hundreds of Merlin engines, they still have much variability at all. I guess the constant design iterations along the way are to blame but still interesting that (apparently) acceptance testing is used to characterize the engines (individually? by lot?).
Did not see this here..
From SpaceX facebook by Mike HawesQuoteWow, so we've learned that SpaceX will be phasing out single engine testing at McGregor! (Gwynne, end of Q&A, Small Sat Conference)
Have not seen the video of the conference yet.
Edit: Just finished watching the video at about 1:01:46:QuoteWe are likely to go away from single engine tests on Merlin, you know, once we have finalized the design and show great decrease variability, i say, greater non variability, decrease variabilityMy emphasis
Interesting that even with all the accumulated manufacturing and testing knowledge they've built up making many hundreds of Merlin engines, they still have much variability at all. I guess the constant design iterations along the way are to blame but still interesting that (apparently) acceptance testing is used to characterize the engines (individually? by lot?).
Did not see this here..
From SpaceX facebook by Mike HawesQuoteWow, so we've learned that SpaceX will be phasing out single engine testing at McGregor! (Gwynne, end of Q&A, Small Sat Conference)
Have not seen the video of the conference yet.
Edit: Just finished watching the video at about 1:01:46:QuoteWe are likely to go away from single engine tests on Merlin, you know, once we have finalized the design and show great decrease variability, i say, greater non variability, decrease variabilityMy emphasis
Interesting that even with all the accumulated manufacturing and testing knowledge they've built up making many hundreds of Merlin engines, they still have much variability at all. I guess the constant design iterations along the way are to blame but still interesting that (apparently) acceptance testing is used to characterize the engines (individually? by lot?).
That's the *point*. Variability *has* dropped, so this is why they are getting to the point of skipping that step. I'm not sure why you read it and assume almost the opposite. In the meantime they are simply being thorough, something many people seem to assume they are not.
They still have a little while to go since the FT+ engines should be now showing up for single engine tests. But in 6 months the SL versions tested (due to the 18/yr launch rate) will be ~81 engines. So a year from now after >160 SL engines they could feel confident that the single engine test step is not required. Prior to that after a dozen flights of the FT+ versions they could do away with the on-pad hot fire test. Its a matter of like you say decreasing the variation distribution between engines to the point that the differences even on a 3sigma+ engine vs a 3sigma- engine is not very much. Add to that the likelihood of the 1st stage testing decreasing in occurrence due to stage reuse and the effort involved with M1D engine testing will be greatly reduced a year from now. M1DVAC single engine testing is likely to follow by sometime after the SL stop doing single engine testing. Making the most testing being that of US.
On another note is the statement of being able to cycle the test stand in <48hrs. So in a month they can test 15 engines. 9 engines in every 18 days. That supports a launch rate of 20 per year of new 1st stages.
They continue to head in the direction of lowering cost per launch and increasing launch rate.
I've asked this before, but does it look like pre-launch hotfires will continue forever? They would mean double the beach closures, which I don't think too many people here realize yet.
I've asked this before, but does it look like pre-launch hotfires will continue forever? They would mean double the beach closures, which I don't think too many people here realize yet.
I've asked this before, but does it look like pre-launch hotfires will continue forever? They would mean double the beach closures, which I don't think too many people here realize yet.
It might be that at some point, the static fire will be merged with the launch. Sort of like a live pre-flight check.
Changes will be required, but I don't see anything in principle against it, and it will reduce the number of cryo cycles on the tanks.
I've asked this before, but does it look like pre-launch hotfires will continue forever? They would mean double the beach closures, which I don't think too many people here realize yet.
It might be that at some point, the static fire will be merged with the launch. Sort of like a live pre-flight check.
Changes will be required, but I don't see anything in principle against it, and it will reduce the number of cryo cycles on the tanks.
The data review after static fire is done by humans, so are you both suggesting it is simple enough to be done by code? (Maybe it already is for the most part, I don't know.) If not then a simple AI will have to be developed and trained off prior data.Changes will be required, but I don't see anything in principle against itIn other words, normal launch sequence will subsume static fire.
The data review after static fire is done by humans, so are you both suggesting it is simple enough to be done by code? (Maybe it already is for the most part, I don't know.) If not then a simple AI will have to be developed and trained off prior data.Changes will be required, but I don't see anything in principle against itIn other words, normal launch sequence will subsume static fire.
I think the static fire is different from the first two seconds of launch.The data review after static fire is done by humans, so are you both suggesting it is simple enough to be done by code? (Maybe it already is for the most part, I don't know.) If not then a simple AI will have to be developed and trained off prior data.Changes will be required, but I don't see anything in principle against itIn other words, normal launch sequence will subsume static fire.
It's not like it hasn't already been done -- sort of.
I wish I could recall exactly which mission it was, but I recall quite well an occurrence where the Falcon 9 engines started up, then shut down just prior to the hold-downs coming off. After a short analysis and quick fix, the count was recycled and the launch proceeded that night (IIRC), late but still within the launch window.
It's not like it hasn't already been done -- sort of.
I wish I could recall exactly which mission it was, but I recall quite well an occurrence where the Falcon 9 engines started up, then shut down just prior to the hold-downs coming off. After a short analysis and quick fix, the count was recycled and the launch proceeded that night (IIRC), late but still within the launch window.
I believe the F9 sequence you are describing was prior to subcooled LOX...
Not sure if this was posted earlier.
Probably discussed many times but I can't find it: is there any likelihood that sea level Merlin will burn redirected turbo pump exhaust in future?IANAE, but it does not seem worth the trouble at this point in engine design cycle all for a couple of seconds of ISP on an SL engine. They probably want to be done with R&D on this engine.
Well it's already ducted into the after part of the nozzle to provide cooling.
Probably discussed many times but I can't find it: is there any likelihood that sea level Merlin will burn redirected turbo pump exhaust in future?
Probably discussed many times but I can't find it: is there any likelihood that sea level Merlin will burn redirected turbo pump exhaust in future?That was a planned feature for Merlin-2 LRE before Raptor Development became a thing and Merlin-2 work was shelved indefinitely. Any Merlin-2 improvements they decide to move over to the Merlin-1 family would likely become Merlin-1E although there are no present plans for Merlin-1E LRE development.
Probably discussed many times but I can't find it: is there any likelihood that sea level Merlin will burn redirected turbo pump exhaust in future?
I think I read somewhere that the actual turbopump exhausts pay a role (due to their position and use) on some aspects of the octaweb assembly performance.
It's not like it hasn't already been done -- sort of.
I wish I could recall exactly which mission it was, but I recall quite well an occurrence where the Falcon 9 engines started up, then shut down just prior to the hold-downs coming off. After a short analysis and quick fix, the count was recycled and the launch proceeded that night (IIRC), late but still within the launch window.
I believe the F9 sequence you are describing was prior to subcooled LOX...
Yes, I'm positive it was. I'm also pretty sure it was a Falcon 9 v1.1, prior to the current FT version that, we are told, should just be called Falcon 9... :)
Anyone remember which mission it was? I'm pretty certain it wasn't a CRS flight, they have very short launch windows. Maybe Cassiopeia? It was definitely pre-stage-recovery days.