sevenperforce
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Posts posted by sevenperforce
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2 hours ago, tater said:
Cool. We need NTP.
Surprising exactly…nobody.
But I agree, we absolutely need it.
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17 minutes ago, Royalswissarmyknife said:
I wonder how Nasa deals with all the water after a launch.
The same way SpaceX deals with all the water after a launch of Falcon 9 or Falcon Heavy.
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On 7/19/2023 at 10:01 AM, kerbiloid said:
The Einstein's postulates. Just given without any explanation.
Not without explanation, just without derivation. But then proven to the most rigorous of standards by every possible test (and by some tests thought to be impossible at the time the postulates were postulated).
On 7/19/2023 at 10:01 AM, kerbiloid said:The empirical Hubble's law, to which the expansional model was pulled by ears.
The very idea of the Universe expansion, a perpetuum mobile.
You're confusing observations with conclusions.
On 7/19/2023 at 10:01 AM, kerbiloid said:The Michelson-Morley experiment. A search for aethereal flow, made by the tools consisting of the same aether, so moving with the aether speed. It's like measure the wind speed, sitting on a balloon.
(I don't say, that the aether exists, but the experiment looks... strange.)You realize the Michelson-Morley experiment wasn't only done once, right?
The laser interferometers at LIGO are the new and improved versions of Michelson-Morley, and they do quite a good job at detecting motion in the "aether" that is spacetime. Confirmed optically, no less.
23 hours ago, kerbiloid said:How can the age stay same, when it was calculated before the Dark Matter had entered the room?
Dark matter is directly observable via objects like the Bullet Cluster. It would exist regardless of overall cosmology.
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17 hours ago, darthgently said:
I can't even visualize the effects of differential throttling of the scarfed design engines. It seems if you throttle down the port side , the starboard side engines would push the central pressure point toward port, which would somewhat cancel the throttling down on the port side, making it more complex.
I don't think the scarfing will cause overmuch problems for differential throttling. It feels like you're imagining secondary plume interactions, but in a supersonic flow there is no upstream propagation of plume interactions, so any change in the wake pressure across the heat shield will be very very minor compared to the impact of the throttling itself.
Thinking about it in the context of equal and opposite forces:
The equal and opposite forces from each thruster are split between the forces on the chamber/injector, which have a significant medial-radial component (green) but balance each other out, and the forces on the scarfed expansion surface, which have a lower medial-radial component and progressively greater superior component (blue):
By downthrottling on one side, the medial-radial component of the chamber force decreases on that side, causing a direct torque on the opposite side; it also causes the center of vertical force to shift away from that side, which also creates a torque.
If anything, differential throttling of a design like this may result in too much authority and lead to oscillations that are hard to damp, like putting a Vector engine underneath too lightweight of a rocket in KSP. It will be less of an issue in space but it might cause issues with landing.
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4 hours ago, darthgently said:
And on the BO patent side the nozzle scarfing seems promising and something Stoke may be interested in if it pans out. I'm picturing it as maybe working to keep the flow both near the shield on the inboard side while still allowing the flow to expand aft on the outboard side. The inward flow against the shield would create a high pressure peak at the center of the shield also resulting in aft directed thrust. Or something like that. Any ideas on what I'm missing?
You're absolutely right.
It looks like Stoke really started off with the idea of the actively-cooled heat shield and sort of stumbled onto an aerospike design, while BO appears to have been designing this as an aerospike from the ground up.
Stoke may have better luck with its landing burns than BO, since they can just accept flow detachment.
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16 hours ago, tater said:
@sevenperforce, so in all the cases it seems like stoke holds the patent priority? Does this result in some ugly legal wrangling you think?
My patent law experience is not extensive, but IMHO the levels of potentially infringing overlap aren't very significant. You can't patent physics, after all, and just because a patent might describe certain attributes in conjunction with its core aspects doesn't mean all of those attributes are protected.
But yes, Stoke holds patent priority on everything here.
The core of the Stoke patent is the dual use of the expander cycle with the heat shield cooling manifold to pump the coolant during re-entry. It's a truly novel idea and I think it will stick. The question is whether BO tries to copy that approach or not.
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The latest patent from Stoke includes a depiction of a non-axisymmetric heat shield to provide lift during re-entry with 0° AOA:
It's not immediately clear from the patent whether this orientation is imagined as fixed or as moveable. The patent doesn't discuss it moving, so it's a little tough to figure out exactly how this would interact with the plume coming off the engines themselves, if it's "stuck" like this.
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JARVIS has emerged!
I should note that the original Actively-Cooled Heat Shield System and Vehicle Including the Same patent by Stoke, establishing the concept of an engine which uses heat from an actively-cooled shield to circulate the coolant, was filed in August 2020 and claims priority to a prior filing from December 2019, and Stoke's Augmented Aerospike Nozzle, Engine Including the Augmented Aerospike Nozzle, and Vehicle Including the Engine patent was filed August 2021 and claims priority to November 2019. In contrast, Blue Origin's patent above was filed in July 2022 and claims priority to December 2021. So Blue Origin seems to be solidly behind the patenting curve here. Even Stoke's most recent Annular aerospike nozzle with widely-spaced thrust chambers, engine including the annular aerospike nozzle, and vehicle including the engine patent, despite being filed in April 2022, claims priority to April 2021, predating all priority by Blue Origin.
I'll post more over in the Stoke thread, but I did note that their latest patent includes a depiction of a non-axisymmetric heat shield which would provide lift during re-entry at 0° AOA, although it's unclear how that would interact with the aerospike expansion in vacuum.
One of the nifty things about BO's patent is the "plurality of scarfed nozzles" depicted around the heat shield:
As Elon Musk has pointed out, one of the reasons that aerospikes suffer is that the primary challenge in rocket engine nozzles is getting the exhaust to go DOWN, and aerospikes aren't great at that. Actually angling and "scarfing" the nozzles into the heat shield like this will make the recessed nozzle surface present a more consistent surface against which the exhaust can expand, reducing intramolecular cosine losses, and it also protects the engines more directly than Stoke's design seems to.
An element very similar to Stoke's design (and potentially grounds for a patent infringement battle) is that "The heat shield may be actively cooled [and] The heat shield may include a cooling circuit configured to dissipate heat encountered during reentry of the upper stage." They also suggest "secondary fluid injectors" which may eject fluid (presumably from an open/bleed expander cycle exhaust) along with the scarfed thrust nozzles to help shape the plume, which seems to be the configuration depicted here:
Later on they contemplate that the heat shield would "be constructed of thin face sheets separated" by spacers and that "turbine exhaust gas is fed directly into the space between the face sheets and exhausted into the space . . . inside the annular engine exhaust stream through orifices in the outer sheet." Using some sort of base bleed in an aerospike design is a well-known way to improve efficiency (for example, here).
It looks like they are considering both a closed expander design (using numerous BE-7 turbine units) and an open/bleed expander design (using one or more BE-3U turbine units), depending on whether they use the base bleed or not. The patent explicitly states that their design saves development time by "repurposing turbomachinery (e.g., powerpacks) and thrust chambers designed for other engines" and references designs being created "for use in other space vehicles." Later on, it gives the non-limiting example of using "two BE-3U powerpacks" to operate the nozzles but notes that as many as five or more powerpacks could be used. BO contemplates that they could "power down some number of powerpacks and/or nozzles to meet thrust requirements" for certain aspects of a mission; this would require that the various thrust nozzles be plumbed to alternating turbopumps so as to keep the thrust vector consistent.
They suggest a 23-foot diameter vehicle and a 21-foot-diameter ring of engines which creates, effectively, a single 21-foot-diameter nozzle. They anticipate a specific impulse of 395-425 seconds in one configuration, 400-420 seconds in another configuration, and 405-415 seconds in a third configuration; these configurations are not clearly differentiated. They anticipate that for the two-BE-3U-powerpack configuration, only a single powerpack would be used for vertical landing, with sea level thrust of "about 100 klbf" and throttling capability down to as low as 20%. They talk about each thruster producing 2000 lbf in some configuration but it's not clear how many thrusters they are envisioning with that thrust level. Given that the BE-3U is expected to produce as high as 160 klbf in vacuum, this would imply probably thirty thrusters plumbed to each of the two powerpacks.
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16 hours ago, tater said:
I have to wonder if the use of electric pumps paradoxically makes the engines more amenable to being dunked in seawater, simply because there is no complex turbine plumbing which could be impacted by incursions.
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Ships in science fiction have such high specific impulse that fuel reserves are really never an issue unless they suddenly need to be a plot point.
If you want them to be a plot point, make them so.
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1 hour ago, RCgothic said:
Nice!
That might just do it.
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2 hours ago, tater said:2 hours ago, CatastrophicFailure said:
Wait, I thought only three engines were plumbed for relights?
Might be, I recall hearing that—but watch the vid, and the exhaust profile looks more like launch, not like the 3 engine entry burn which seems to read more linear to me.
Can confirm that only three engines are plumbed for relights. The other engines (and those three, at launch) get their TEA-TEB from GSE.
Also, I'm not sure even Falcon 9 could handle running all nine engines on a nearly-empty booster. Think about it -- 9x941 kN on a ~30 tonne stage is almost thirty gees. That's really getting into ridiculous materials stress.
2 hours ago, tater said:Interesting that residuals for a 20 second, 3 engine entry burn results in a ~20% payload loss, but a total of ~70 seconds is only a 40-50% loss.
The upper stage is really doing most of the work to get to orbit.
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Fictional spaceships depicted in film typically don't show any RCS thrusters at all -- the ships turn magically with engines running at full blast but without any forward impulse.
Part of this was the combination of a lack of knowledge (film directors don't know how spaceships work) and rule of cool (it felt more "advanced" to have spaceships that turn without any visible means of rotation). The other part was the cost of CGI. CGI used to be incredibly expensive, and so any added animations drove up the budget. It is only in recent years that the cost of high-quality animation has dropped to the point that animating things like RCS pods is no longer a big-budget item.
The decrease in CGI cost has combined with an additional desire to show space as grittier, with more realism and dynamic elements. The CGI ships in The Mandalorian and the rest of the Disney expansion of Star Wars are a good example of this. Prior to the Disney expansion, CGI ships in Star Wars basically just floated around without any visible propulsion or mechanisms, but now you see animations of stuff like RCS and vertical thrusters. For example, vertical thrusters on the Razor Crest here:
Vertical thrusters on the N1 Starfighter here:
And the use of active gimbal on the main engines, to control the ramming action of the Hammerhead Corvette:
An astute observer will of course note that the engine plumes from the corvette appear to have Mach diamonds showing overexpansion, which is de facto impossible in the vacuum of space. Perhaps these aren't actually Mach diamonds but are instead some sort of magnetohydrodynamic vortices that just look like Mach diamonds? /s
One possible in-universe reason we don't often see RCS in science fiction is that the inertial control mechanisms are just massively more powerful, or they have used some sort of magnetic field coupling as the primary orientation control. In real life, magnetic field coupling can be used to desaturate reaction wheels (the Starlink satellites use an electromagnetic rod to push against Earth's magnetic field for wheel desaturation to since they don't have any attitude thrusters), but you'd need immensely more power to do that for your primary mechanism.
Science fiction, of course, is famous for huge power sources. The TIE fighters in Star Wars are said to be extraordinarily maneuverable, and the X-Wing pilots are always saying "Lock S-foils in attack position." Why? The X-Wings have their weapons mounted on the wingtips, so opening the wings would increase their field of fire...but that might not be the only reason.
The X-Wings do have sings that look vaguely like they could provide aerodynamic lift, but the TIE fighter's wings certainly do not:
What are those wings useful for, other than looking cool? They could be solar panels (at least, that's one in-universe answer offered in some of the extended universe stuff) but their utility would be meager given the enormous power requirements of Star Wars vehicles. They could also be thermal radiators of some kind, although we don't see them glow like radiators. My head-canon is that any flat, thin surfaces on Star Wars vehicles are high-energy magnetically-coupled reaction control planes, capable of pushing against any ambient planetary or stellar magnetic field in order to rapidly change orientation. That's why the X-Wings appear to maneuver in space just like ordinary aircraft: their "wings" are acting like real control surfaces. That's also why the TIE fighters are so maneuverable: they are basically just two giant RCS surfaces with a pilot and engine and blasters in a pod at the center. Even the Millennium Falcon's long, flat external surfaces are acting this way.
This allows some aspects of basic physics to still apply. The moment of torque around the center of mass of a ship is going to depend on the physical orientation of those reaction planes. X-Wings and Naboo Starfighters have good roll and pitch authority ("let's try spinning, that's a good trick") but poor yaw authority and will probably rely on differential thrust for yaw. The TIE fighters have ridiculously good yaw and roll authority but very little pitch authority other than what phased gimbaling of their ion thruster can provide; this explains the strange, otherworldly way in which they seem to zip around through the sky. The Falcon has some yaw and roll control but FANTASTIC pitch control, which allows it to simply pitch its nose up and use its sublight thrusters to blast its way into space like a KSP spaceplane loaded down with oversized RCS wheels.
To summarize, Star Wars ships basically use the following engines and maneuvering systems:
- Repulsorlifts: a "gravity brake" which "locks" to a particular point in the gravitational gradient and allows a vehicle to hover with little or no power consumption; similar technology is used for inertial dampening to keep the crew from getting smeared
- Reaction control planes: a magnetically-coupled "wing" which pushes against planetary or stellar magnetic fields to provide roll, pitch, and yaw authority as necessary
- Vertical turbine-based thrusters: a turbine bypass which sucks in atmospheric gases, diverts them through the main engine system, and then expels them downward at high velocity in order to provide vertical translation for liftoff when you need to ascend or descend vertically
- Sublight thrusters: extremely powerful reaction engines with ridiculously high specific impulse which somehow do not torch everything behind them...perhaps using a exhaust made of something which decays into non-interacting particles shortly after leaving the nozzle
- Hyperdrives: whatevermajiggy that lets you go to hyperspace
It's a good head-canon anyway.
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First methane-powered rocket to orbit...on my birthday!
Not SpaceX, but not half bad either.
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23 hours ago, Beccab said:
The clear disadvantages in this design are (a) no obvious abort mode and (b) overweight engines carried all the way to orbit. Every kg that can be moved off the orbiter is an extra kg of payload, so if there was a way to offload two of those three engines it would be a significant benefit.
Obviously reducing dead weight carried by the first stage isn't a 1-to-1 benefit to payload, but it does something. If you can do crossfeed, it's absolutely worth doing a sustainer architecture IF it doesn't result in a less-efficient upper stage engine and IF it doesn't force other suboptimal design choices.
A three-core crossfed design with the side boosters returning Falcon Heavy style and the center core going to orbit is the straightforward choice, but then EDL of the orbital stage (for the sake of reusability) becomes the challenge. With crossfeed and a potential tripropellant sustainer, the disadvantages of hydrolox are ameliorated because the nice thrusty boosters carry the fluffy hydrolox tanks out of the atmosphere where drag is no longer an issue, and drop them off full. Then the fluffy tanks help reduce heat load on EDL.
I wonder how large and draggy a hydrogen tank would need to be in order to reduce heat load to the "hot structure" regime.
Of course you still need descent control and a landing mode. You can do a standard biconic entry with split flaps but that doesn't provide a straightforward landing mode.
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22 minutes ago, Exoscientist said:
By the way, there might be a simpler way of achieving the same thing as the thrust-augmented increased oxygen flow. That is if the Vulcain can be made to be of variable mixture ratio. On take-off you would increase the oxygen rate to increase thrust. This though would decrease Isp so you would only use it for a few seconds during take-off for the added thrust. Thereafter when the propellant burn has made the rocket lighter you would switch back to the usual mixture ratio to get the high Isp.
You wouldn't want to only use it for a few seconds. You'd want a smooth transition as you climbed, balancing the need for higher specific impulse against the need to avoid gravity drag.
Gravity drag is a REALLY big deal.
22 minutes ago, Exoscientist said:Altitude compensating nozzles, or adaptive nozzles, is one of those methods that have been known about for years to be able to increase payload but hasn’t been used.
Usually because altitude-compensating nozzles won't actually increase payload.
22 minutes ago, Exoscientist said:I mentioned that an all-liquid Vulcain at 7.5% payload fraction would mean other launch companies such as SpaceX would be forced to catch up to this
There's no reason that other launch companies would need to catch up to any particular payload fraction. Launch companies try to catch up to cost; they don't care about catching up to a particular payload fraction.
22 minutes ago, Exoscientist said:I thought maybe the Falcon 9 could do it by using the same ultra efficient lightweighting used on the Ariane 5 core. Since kerolox is 3 times denser than hydrolox, the 16.3 to 1 mass ratio of the Ariane 5 core would correspond to a mass ratio for the Falcon 9 first stage of 50 to 1(!).
The mass ratio of the Ariane 5 core is based on a tank which is partially in tension rather than compression for the highest-stress portions of the flight and thus can be very lightweight. So translating to the Falcon 9 tanks won't work for that reason. It also is not a 16.3:1 mass ratio. Additionally, the LOX tank is not significantly different in weight than a typical LOX tank; the difference is exclusively in the hydrogen tank. You can't just take the (incorrect) 16.3:1 mass ratio and multiply by the difference in density.
22 minutes ago, Exoscientist said:But after running the numbers this still would not be enough for the Falcon 9 to get to 7.5% payload fraction.
Which it doesn't have any reason to do anyway.
22 minutes ago, Exoscientist said:What might work though is if you also used altitude compensation on the first stage of the Falcon 9.
The Merlin 1D already uses a nozzle that is designed for optimal performance at all altitudes in its trajectory.
22 minutes ago, Exoscientist said:the Merlin on the first stage has such poor vacuum Isp at ca. 312s. But note with altitude compensation its vacuum Isp could be raised to > 360 s.
It absolutely could not. The Merlin 1D Vacuum has a specific impulse of 348 seconds; there is no nozzle in existence that could raise the vacuum specific impulse of ANY Merlin engine to 360 seconds, let alone a nozzle which could be fired at sea level.
22 minutes ago, Exoscientist said:the vacuum Isp of the Vulcain could be raised to > 470s(!)
It absolutely could not. The Vulcain engine already has near its maximum vacuum specific impulse for its engine cycle. A gas generator hydrolox engine is not going to be able to get above 445-450 seconds. We have never ever had a hydrolox engine with over 470 seconds of specific impulse.
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1 hour ago, Beccab said:
Nobody actually built anything with that architecture, but it was somewhat heavily proposed during the early shuttle program for drop tank designs, such as Starclipper. The relative engine is the XLR-129
The XLR-129 would have been a variable-mixture-ratio fuel-rich staged combustion engine, similar to the RS-25 but with only a single preburner feeding two pumps rather than two separate preburners.
I'm reminded of the RD-701 which also would have been used for an SSTO-like or SSTO-lite application, although it was an ORSC architecture that used only kerosene for the preburners but switched between kerosene and liquid hydrogen in the main chamber.
Sustainer architectures lend themselves to inventive concepts for increasing thrust at liftoff and increasing specific impulse at altitude, and so they really trigger my weakness for clever engineering. It would be cool to have a tripropellant engine with an ORSC kerosene-based LOX pump and a split hydrogen-based expander cycle fuel pump, for example...especially one with a nozzle which would be sea-level expanded in mostly-kerosene-mode and vacuum-expanded in mostly-hydrogen-mode, so that changing the mixture ratio and propellant during ascent would automatically provide altitude compensation. Cooler still if you could use crossfeed from a strap-on booster assembly (maybe even a fly-back booster) to provide the kerosene, minimizing tankage on the main stage.
Of course if you're going for full reusability this doesn't help EDL.
1 hour ago, wumpus said:But you need a lot less thrust on a second stage than you'd put in a sustainer, and that weight comes out of your cargo capacity (much worse than forcing the first stage to lift the dead weight of the second stage engine).
Then again, the weight of a second-stage engine is typically much larger than the weight of a similar-flow-rate first stage engine (e.g. Merlin, Rutherford, Raptor), so if you can use a larger engine bell with some sort of thrust augmentation (especially crossfed) then you're winning.
1 hour ago, wumpus said:And it looks like the highest Chinese launch site gets .8 bar at launch. Helps a little, but might not be worth redesigning a nozzle.
I believe most sea-level engines are actually optimized for around 2-3 km altitude: a little overexpanded at launch in order to be less underexpanded at altitude. The optimal expansion ratio depends not only on the integration of specific impulse across the flight regime to maximize actual delta-v but also on reducing delta-v lost to gravity drag from low thrust. I suppose they might be designed for a slightly higher altitude if they are launching from a higher elevation, but I don't think it would be worth a redesign unless those rockets are ONLY going to be launching from high elevation and never from closer to sea level.
1 hour ago, wumpus said:You'd be surprised how thin air can be and still be able to breathe (a lot higher than that).
Fair point.
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13 minutes ago, wumpus said:
The gotcha is that 3 launcher platforms that solved altitude compensation all failed.
Pegasus (+XL) is basically retired
Stratolaunch isn't interested in orbital launches
Virgin Galactic (the air-launch to orbit company) is bankrupt and nobody is interested in the rocketYeah, air-launch is not my favorite. Not unless the separation is exoatmospheric.
13 minutes ago, wumpus said:Not to mention, using a 2 stage rocket gives you ideal vacuum Isp for most of the delta-v needed.
The (perhaps deceptively) attractive element of parallel staging is that you can get better T/W at liftoff without needing more engine mass. Of course, crossfeed makes it much more advantageous.
16 minutes ago, wumpus said:Didn't somebody slap a moveable extension onto a sustainer already, or was the proposal tabled before construction/launch?
I don't believe any vehicle has ever launched with a first-stage moveable nozzle extension, except maybe some of the earliest solid rocket ICBM designs.
16 minutes ago, wumpus said:I think China has a launch facility at altitude, but no idea how that will affect nozzles and pressure design.
If the air pressure is high enough to support breathing, it's essentially sea level for the purposes of nozzle design.
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1 minute ago, tater said:
I would love to see some really out of the box ideas. Maybe at some point one of these other rocket outfits will do this to try and make themselves relevant. Stoke springs to mind as trying something really novel.
I like altitude-compensating nozzles, so I really want to see a better sustainer architecture.
So far we have three main sustainer architectures:
- Wimpy core hydrolox stage with beefy boosters to get T/W over 1 (Ariane 5, SLS, STS, LM-5)
- Beefier core stage with large upper stage and optional small solid boosters (Atlas V, Delta IV M, Ariane 4)
- Two-stage architecture with multi-stick first stage (Delta IV Heavy, Falcon Heavy, Soyuz)
I want something more interesting.
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22 minutes ago, tater said:
Maybe a handful of Arianespace employees can work some of these issues—it will give them something to do for the next year/whatever while they are not launching rockets.
Course the ones who would actually do this work are working on a slow-follower version of the F9, and no such changes to Ariane 6 will ever happen or even be considered—just as frankenrocket changes to SLS will never happen.
If you wanted to REALLY frankenrocket things up, you could use the Vulcain 2 engine to get a turbineless scramjet-based HTOL vehicle off the ground and up to ram compression speed, let the airbreathers get you moving high and fast without sucking up too much of your hydrogen, then switch the Vulcain 2 back on to provide the rest of the push to a high-suborbital trajectory, like 4 km/s. Then release a cryogenic upper stage from inside a fairing to deliver a comsat payload to GTO. The HTOL vehicle would be able to handle the re-entry with minimal shielding and coast back home.
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On 7/8/2023 at 9:20 AM, Exoscientist said:On 7/6/2023 at 3:58 PM, sevenperforce said:
How will it get off the ground? Its thrust capabilities are based on vacuum thrust off a second stage with a payload on top, not sea level thrust off the ground with both an upper stage and a payload.
I’m going by Lapsa’s statement in the video where he literally says this stage can go to 400,000 feet, which is suborbital space. So he must mean taking into account the fact the thrust and Isp levels of the engines are reduced at sea level at takeoff.
400,000 feet is ~124 km, which would require about 1.45 km/s. Add a three-tonne upper stage and payload and I'm unsure that's enough umph to get high enough for the upper stage to achieve orbit.
On 7/8/2023 at 9:20 AM, Exoscientist said:There are several European start-ups on the horizon planning to take advantage of the small launch market of < 1,000 kg to LEO. With the miniaturization of satellites, this market is expected to be sizable. With it’s level of development Stoke Space could probably beat these European start-ups to market joining Rocket Lab as the only companys offering such launches.
I'm skeptical about such an architecture being able to get ANY satellites larger than cubesat size to LEO.
On 7/8/2023 at 9:20 AM, Exoscientist said:The examples shown there including the famous DC-X show a ground-launch conical stage is not prohibitive against a stage reaching suborbital space or even being a SSTO.
Well, the DC-X certainly was not going to be an SSTO.
Any SSTO which can reach orbit, deliver a payload, and return to the ground in one piece can deliver MORE payload with an upper stage.
On 7/8/2023 at 9:20 AM, Exoscientist said:Quite notable as well is Bono wanted to use the efficiency of the aeroplug/aerospike for the launch, not disparage it.
Getting an SSTO or sustainer to orbit more or less requires an altitude-compensating nozzle system, yes. But you'll also note that none of Bono's designs proposed a using an aerospike on the first stage of a two-stage architecture.
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18 hours ago, magnemoe said:20 hours ago, AckSed said:
Science fight! Science fight!
Edit: Say we have a thrust-augmented nozzle that injects extra oxygen to increase thrust. Not looking for anything special, just a bit more mass-flow at the start for a O/F ratio of say 13:1 for 30 seconds. Does that do anything for the lift-off mass?
It should if TWR is low and therefore its more important to increase TWR than high ISP at liftoff.
Probably not burning oxygen rich but rather not burning hydrogen rich at liftoff, With SRB you would then switch to more hydrogen rich and then throttle down as TWR get high enough.Bumping up thrust at the beginning definitely helps. Liftoff thrust is a big problem on the first stage. Of course the more you increase thrust, the more you have to strengthen the LH2 tank, but the improvement is better.
The thrust improvements in the Vulcain 2 came from going to a less fuel-rich mixture ratio. Similarly (but in history), the Saturn V's increased lift margins to be able to send the rover to the lunar surface came from adjusting the F1 engine to be able to change the mixture ratio in flight, from high thrust and lower isp at liftoff to lower thrust and higher isp at altitude.
SpoilerSpecifically (for anyone interested) this increased capacity was the result of a LONG sequence. Higher thrust at liftoff meant lower gravity drag, and lower gravity drag meant a faster climb and a more rapid increase in F1 specific impulse. Then the increased specific impulse and subsequent mixture ratio change meant the first staging happened at a higher speed. This meant that the second stage got closer to orbit, which meant the third stage had less work to do to get into orbit and had more propellant left over for the lunar burn. This not only meant that the third stage could throw the added weight of the rover, but also that the third stage could place the whole stack in a more optimal higher-energy lunar trajectory, which meant fewer correction burns and a shorter insertion burn by the CSM. The CSM then had more propellant left to place the lunar module in a lower lunar orbit, reducing the amount of propellant that the lunar module would need to burn in order to reach the lunar surface.
Going up to a 13:1 O/F ratio would probably eat the engine -- that's way too oxygen-rich. However, the thrust-augmented nozzle could work with some sort of boundary layer film cooling.
I could calculate the actual change in thrust but it would depend on choices like injection pressure and set figures like turbopump capability.
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1 minute ago, tater said:
He was speculating about gimbal clearance...this could be fixed by using SL Raptors on the outside or going to 4 RVacs and 8 SLs.
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19 minutes ago, Exoscientist said:
Below are the input page I used and results page for the Silverbirdastronautics.com calculator. For the Vulcain engine, I took the vacuum thrust as 1,350 kN. So for three Vulcains I input 4050 kN in the thrust field for the first stage. I took the vacuum ISP as 434 s.
As discussed above, the Silverbird calculator will overestimate the result because it does not provide a sufficiently large sea level thrust penalty.
To get an engine that is properly expanded at sea level and also boasts a 434 second vacuum specific impulse, you would need something like the RS-25 powerhead with a smaller sea-level sized nozzle. And even then you'd probably still have to crank up the chamber pressure.
19 minutes ago, Exoscientist said:Since it had higher thrust than two Vulcains I used the later, larger version Ariane 5 “E” core at 170 ton propellant load and 14 ton dry mass.
Well now we have an entirely different first stage. I can do the math for that stage, but different inputs mean different outputs.
19 minutes ago, Exoscientist said:I added ~4 tons for two additional Vulcains and another ~2 tons for strengthening the tank for the higher thrust, but also subtracted ~2 tons for removing the JAVE.
The JAVE only has a mass of 1.7 tonnes, and you're going to need a lot more than 2 additional tonnes to strengthen the hydrogen tank.
19 minutes ago, Exoscientist said:For the second stage, the higher thrust enabled a larger upper stage. So I took it comparable to the Centaur V at a 50 ton propellant load and 5 ton dry mass.
So we also have an entirely different second stage too.
19 minutes ago, Exoscientist said:For the engines I used three Vinci engines at 180 kN vacuum thrust each for a total of 540 kN, but I mentioned I assumed a larger nozzle to reach the highest 465.5 s Isp of the RL10.
The Vinci engines have a nozzle extension that is 2.15 meters in diameter and makes the entire engine 4.2 meters long. How are you going to fit those -- with room to gimbal -- inside a 5.4-meter interstage? And the RL10B2's 465.5 second specific impulse already comes from a 2.15 meter wide nozzle, so what makes you think you can squeeze out more specific impulse from the Vinci? Just use the Vinci's actual specific impulse. And you can really only fit two of them on the stage anyway.
19 minutes ago, Exoscientist said:The result was 19.6 tons to LEO. This is nearly 3 times your result of 7 tons.
That's to be expected because you used a different first stage and a different second stage.
The military wants nuclear propulsion to the Moon???
in Science & Spaceflight
Posted
Wow, they're building an FTL engine!?!?
In all seriousness, this is great.
Of course this is a pure marketing mockup and we shouldn't assume any engineering details from it. I wonder what kind of engine cycle they will go for. The NTR that was intended to serve as a Earth-Lunar ferry as part of the original STS architecture would have been a dual-mode version of NERVA, balancing LOX injection for high thrust at the start of the burn against the pure-LH2 low-thrust end burn for high efficiency. Nuclear engines have specific impulse to spare and so the usual challenge is trying to get high enough thrust to avoid Oberth losses.
NERVA was a thrust chamber tap-off (or "hot bleed" since you can't say combustion tap-off) design, which is probably why the flight-configured engines in the above mockup and in the BWXT image below are depicted with the turbine exhaust injection manifold wrapped halfway around the nozzle like an F-1 or MVac:
It's unclear why there are two separate flows from the tank and apparently two (or four?) separate turbopump assemblies as well. Maybe it's supposed to have two modes, one with a higher tap-off ratio and one without? Or are the turbopumps running in series?
For ground testing the nozzle extension was not used and so the turbine exhaust had a separate nozzle:
You can see that the tap-off comes from the very bottom of the reactor chamber right before entering the throat.
An alternative approach for a flight-configured engine would have had dual turbine exhaust nozzles, presumably for roll control (or even all control authority if the main engine was fixed:
With the thrust levels that you need from a NTR, a closed expander cycle really isn't possible. Granted, you make better use of the Carnot cycle power because you don't have to pump a bunch of heavy oxygen like closed expander hydrolox, but all that heavy oxygen is what's giving you most of your thrust, so you're right back where you started. Plus, the high mass of the nuclear reactor means that you're really not as concerned about squeezing out extra Isp as you are about getting more thrust.
The LockMart promotional materials in that video depict four gold foil tanks surrounding the reactor. Even though we obviously shouldn't assume anything gratuitously from the pure promotional stuff, it makes me wonder if they're considering a dual-mode design like the STS ferry, or even something more creative like using a hydrolox gas-generator as an independent pump cycle for the main pumps. There are definitely some unique considerations for a nuclear engine, beyond just the obvious (like safety issues). In a conventional liquid bipropellant combustion engine, you have a limited amount of total thermal energy you can extract from a given mass of propellant, but maximizing the heat energy increases the average molecular weight in the exhaust, lowering the total specific impulse. As a result, a bipropellant combustion engine has to balance maximal thermal output against an optimally efficient propellant mix (which is why engines don't typically run stoichiometric). With a nuke, on the other hand, the total amount of thermal energy available is not a function of the amount of propellant you have: instead, total available energy is a function of reactor mass and propellant dwell time. So the baseline assumptions don't always work out quite the same way.