sevenperforce

I'll start once you post the mission requirements.

I wish there was some way to incorporate "orbital" or "orbit" into the name, to underscore that we are targeting LEO and not just suborbital flight. Hybrid rockets downthrottle somewhat more easily than bipropellant liquid rockets, but you're right, this is one of the factors that needs to be evaluated. Downthrottling too far may reduce chamber pressure and decrease specific impulse, even if it doesn't choke at the nozzle. This is one of the independent variables that can be altered to a point, but we need to watch carefully. If we say the minimum throttle is 20% rather than 10%, then total dV (on that initial spreadsheet) drops from 7,130 m/s to 7,118 m/s. So it's not a huge difference. A lot of this is fine-tuning. If we run into some more advantageous numbers along the way, you can get up over 8 km/s easily even without increasing stage radius. For example, if we use average isp rather than initial isp, if total aerodynamic drag is only 580 m/s, volumetric efficiency goes up to 0.7, White Lightning's propellant fraction goes up to 85%, and we kick naked-stage TWR up to 9, then we end up at 8,028 m/s on a standard 22-cm stage. In contrast, if the fixed independent variables turn out poorly, it gets difficult. If aerodrag is closer to 700 m/s, volumetric efficiency is no better than HEROS-3, White Lightning prop fraction is only 70%, the hybrid's isp is 300/240, and we can only manage a naked-stage TWR of 6, then we'd need each stage to be a full 85% wider to hit 7.8 km/s. Pyrotechnic bolts are probably easy enough to self-manufacture. Seems viable to me. All right, anyone who wants it -- here you go! Your mileage may vary. Any improvements welcome, and if you find any errors in my formulas, more power to you!

Still pining for the next mission.

Okay, so here's the spreadsheet I've been working on. The values in red are independent variables, the values we can play with at will. We can increase base thrust, or alter throttle for a given ascent phase, etc., and see where things come out. Of course, we still have to be careful. We can't go below the minimum throttle rating, and cranking up the TWR too high can cause problems, and so on. The values in orange are fixed independent variables. These are the numbers that are determined by systems outside our control, but which may change as we acquire more information. For example, we may only be able to push 300 seconds of vacuum specific impulse instead of 313, or we may only need to factor in 575 m/s of aerodrag rather than 650 m/s. The values in blue are the output of the equations. Obviously, the number at the bottom right in blue-bold is the total dV; that's the number we want to maximize. But we have to keep an eye on the rest of them, as well; we don't want TWR to go too low or too high at any point. For aerodrag, I used the simplifying assumption that the percentage of total aerodrag for a given ascent phase will be proportional to the fraction of ascent time spent in the atmosphere. So I split total aerodrag (a fixed independent variable) up based on the duration of each of the first three ascent phases. Aerodrag after that point should be negligible. For conservatism, I have calculated the dV for each phase based on the starting isp, rather than the average isp over that phase. This should give us a significant buffer. The nature of the spreadsheet is such that changing any one of the independent variables produces an entirely new optimization curve for the other variables, so it takes some work. Here's what it looks like when I increase the size of the booster and then tweak all the numbers to their optimal values: So it definitely looks doable! If anyone would like to play around with the spreadsheet, let me know and I'll try to upload it.

I think what he means is that the RCS thruster is placed to aim into a tube of some sort and used to push it away instead of a normal seperation motor. What I don't understand is why not use a tiny-ish hobby motor (like C or D size) as a separation motor. It's because we need an actual separation mechanism, and the RCS thruster can provide both mechanism and reaction. An explosive bolt disintegrates when it receives the electrical separation signal, but if we were using any kind of clamping mechanism, then you need a servo or solenoid or some other electromechanical system to actually push the clamp open via electrical current. With an RCS thruster in an airtight tube, the normal action of the RCS thruster can build up enough pressure to force open the clamps without needing any electrical signal, and then automatically push the booster away. The trouble with a hobby motor is twofold: first, the plume would tend to damage the stage; second, ignition is not instant and so the whole function is potentially compromised. All the cores will be more-or-less identical for the sake of commonality. Differential throttling is more efficient and more powerful than RCS, but each of the cores will be plumbed for RCS, so adding a single RCS thruster in the existing port shouldn't be difficult. And we'll need RCS on a standard single core when we are testing, as well. Lots of experience, haha! I make the selection box transparent and do a lot of copying and flipping and manipulating with the arrow keys, if that helps.

@Bottle Rocketeer 500 Bump?

I'm gonna stop you right there. Solid-fueled rockets do not "rely on the air around the vehicle to provide oxygen for ignition". Otherwise, everything @qzgy said. Not to rain on your parade, but this is rocket science. Nothing wrong with giving your input, of course. I'll set it at 650 m/s but make it one of the independent variables in the spreadsheet so we can tune it down a little later on if needed. If you watch the HEROS-3 launch, you can see that they packed the chutes into an interstage between the N2O tank and the payload module. That should work well enough. Do we necessarily need both drogues and mains? It would seem to increase the number of possible failure points. What about a reefed main that opens fully at lower altitudes? I don't think we need dedicated sep motors; I think we can repurpose an RCS thruster to act as both pneumatic decoupler and sep motor. The thruster would be fitted inside an airtight cylinder attached to the core. At the separation signal, the pressure valve opens and pushes residual HTP through the catalyst bed and RCS thruster, filling the cylinder with high-pressure gas until the force overcomes the spring on the clamps that hold the stages together. After separation, the RCS thruster continues to fire, pushing the nose of the booster away from the core just as a typical separation motor would. Not sure about attachments on the base, though. We cannot very well just strut them together.

According to page 2 of this study, vacuum isp for peroxide+kerosene ranges from 304 seconds at 2 bars chamber pressure to 313 seconds at 30 bars chamber pressure. If you look, you'll notice that the O/F fuel ratio changes based on chamber pressure as well. One advantage of a pressure-fed engine is that highest combustion chamber pressure is at launch, where pressure losses are highest; as it ascends and pressure losses decrease, tank pressure drops. OTRAG's pressure-fed design, which we are aping, calls for 600 psi ullage pressure or around 41 bars. So I think it is safe to aim for the high end of chamber pressures and O/F ratios. Higher chamber pressure helps tremendously with pressure losses at launch. I think I'll set launch isp at 250 seconds, just under the SL isp of the 47-bar Gamma 8 engines on the Black Arrow kerosene+peroxide LV, increasing stepwise to 310 seconds by core burnout. This requires an O/F mass ratio of 7.4:1, which I'll factor back into my equations. The "simulated" gravity drag and air drag for the Lambda 4S don't really make sense at all, because we determined earlier that there were 1.2 km/s of losses in aerodynamic and gravity drag. Gravity drag isn't too hard to calculate; the time from launch to orbital insertion was 104.8 seconds, 28.2 seconds of which was the terminal stage and 19.4 seconds of which was the second-to-last stage. Operating under the assumption that gravity drag pretty much dies out over the course of the second-to-last-stage burn, I'll use 67 seconds of gravity drag for a total of 656.6 m/s of gravity drag losses, leaving 543 m/s of aerodynamic drag losses. That's a little more reasonable. It's a tossup whether the Lambda 4S would have higher or lower aerodynamic drag than our launch vehicle; higher acceleration means greater speed and lower isp means a lighter stage, both resulting in more severe initial drag losses, but our vehicle will have more parallel boosters and will spend more time in the atmosphere, which means more drag losses for us. I'll just set total aerodynamic drag at 550 m/s for conservatism, and split it among the first few acceleration phases. Notably, I estimated 600 m/s of aerodynamic drag on the HEROS-3, which is very close. Recall the layout (only two boosters shown, but four planned total): A screengrab from the video of the HEROS-3 launch shows that the recovery interstage (containing chutes) is 0.48 meters long and the nose cone is 0.97 meters long, meaning that the actual tankage length is 6.05 meters. This means the volume of the tankage portion of the vehicle is 236 L, but based on earlier calculations (using N2O and paraffin density, stated propellant mass, and known O/F ratios for N2O+paraffin hybrid rockets), total propellant volume is 128 L. This means volumetric efficiency is just 54%. We should be able to squeeze slightly better volumetric efficiency out of our design...perhaps around 65%. Once I get all these numbers into the spreadsheet, I'll post it.

HEROS-3 is 7.5 meters long and 223 mm in diameter. I'm going to go ahead and set up a cleaner version of the same spreadsheet I used yesterday, but factor in stepwise drag losses and allow for some deeper throttling, and tie it all to a single diameter multiplier. This will allow me to baseline total dV at a HEROS-3 equivalent size and then increase diameter gradually until we have enough dV to reach orbit. If I can't do it with a single-core I'll try again with a quad-core. Interesting. Hydra would also be a neat name, since the base will have "many heads". Or we could call it the ARDYH since it's an upside-down Hydra...maybe backroynm that.

Oh, while we're at it...any ideas on a name? A few possibilities: Legion (since it uses clusters and parallel staging) Jebediah (in honor of KSP) EELOO (in honor of KSP, could backronym to something) Kilgore (character in Apocalypse Now who delivers the famous line, "I love the smell of napalm in the morning")

It should be easy enough to have a videocamera that stores high-quality video locally while simultaneously capturing still frames every few seconds and sending them back via VHF. This gives us nearly-live visuals but also preserves a video feed for later recovery. To clarify my earlier proposition: I was already working under the assumption that the second stage was a full-size single-stick core, so there's no extra dV to be gained there. To a point, there are really no significant disadvantages in making the stages slightly wider; it doesn't make them significantly harder to handle, and the increase in drag is more than counterbalanced by the linear increase in propellant fraction. The maths, in case anyone is curious: There's a slight increase in skin thickness in order to maintain structural integrity, but this is such a small value that it's negligible for relatively small increases in cylinder radius. And the increase in drag is also counterbalanced by the increase in overall vehicle weight. Another reason hobby rockets need high fineness ratios is that their low isp means their mass drops rapidly, but with our higher isp we hang on to our propellant for longer and so drag has less of an impact. I think it will be much simpler to make a slightly wider stage than it would be to double or triple the number of first-stage cores. Yes, this is definitely getting complex. I'd like to be able to model it with aerodrag, gravity drag, and pressure drag losses included, but the only thing that can be directly calculated is gravity drag, and even there it only holds for the initial part of the ascent (the kick stage should have no gravity drag, as it will already be in free-fall, but I'm not sure about the hybrid upper stage). I suppose I could estimate aerodrag from this analysis of the Lambda 4S, but 66 m/s seems ridiculously low, and it only calculates gravity drag for the 1.5 stages, out of 4.5 stages total. Then again, with its very high thrust maybe it really does clear the atmosphere that fast. If anyone can find or come up with an isp curve for a pressure-fed HTP+petrol+gel engine, I'd deeply appreciate it.

Somewhere between 8.32 km/s and 8.65 km/s, though symmetry becomes more complicated at that point. I'm unsure whether even 8.65 km/s is enough to get into orbit once gravity drag, pressure drag, and aerodynamic drag are pulled out. For a small launcher, aerodrag is high. If we could push to around 9.2 km/s, I'd breathe easier. Increasing the diameter of the rocket grows propellant mass quadratically while growing stage mass linearly, so I can look into exactly how wide it would need to be in order to achieve 9.2 km/s, either with a single-stick or a four-core. And all this is dependent on the assumption that I can use the propellant fraction of the HEROS-3 rocket as a viable starting point. Finally, I really wish I had a better grasp of what isp I can reasonably expect. Peroxide + kerosene ranges from 230 seconds to 319 seconds, depending on pressure and expansion ratio, which is really broad. For all I know, the aluminum salts used to gel the gasoline could push specific impulse considerably higher.

I hate not having a camera. Maybe one on the first-stage core, if we can reasonably assume that it will survive recovery at least partially intact? That wouldn't unduly cut into payload budget but it would give a nice view of ascent. A camera on the terminal stage probably wouldn't work anyway, due to spin-stabilization and the weight of a live video transmitter.

Okay, so I completely screwed up the numbers on the first go-round because I only included half the propellant of the parallel boosters. Fixed that. Correct numbers follow. T-0:02. Valves open at low throttle for hypergolic ignition. T=0:00. Ignition times five confirmed, valves to full throttle. TWR is 3.5:1, GLOW is 1.5 tonnes T+0:12. +443 m/s. Core and one booster pair throttle down to 50%. Throttledown drops TWR from 4.11 to 2.88. T+0:48. +1,229 m/s. First booster pair jettison, second booster pair throttles back up to 100%. TWR drops from 4.37 to 3.79. T+1:05. +748 m/s. Second booster pair jettison, core throttles back up to 100%. TWR drops from 4.88 to 2.69. T+1:14. +245 m/s. Core burnout; TWR has climbed up to 2.93. Separation and S2 ignition. T+2:02. +2,570 m/s. S2 burnout; TWR climbed from 3.68 to 8.83. Spin-up, separation, kick stage ignition and orbital insertion. Kick stage is +3,085 m/s. Cumulative dV is 8.32 km/s, which is definitely in the neighborhood of LEO. We can squeeze out more dV by deeper throttling, depending on the tradeoff between TWR, gravity drag, and aerodynamic drag. If you use a bolted-together four-core sustainer cluster rather than a single-stick core sustainer, you get 8.65 km/s cumulative dV and slightly lower gravity drag losses.

GLOW is Gross Lift Off Weight. All the numbers are determined by the rocket equation: dV = ve*ln( m0/mf ). At each step, you take the total inert mass (payload + structure) plus the remaining propellant in each stage, and that's your m0; then you subtract the propellant you burn, and the difference is your mf. The natural log of this ratio, times the exhaust velocity (specific impulse times 9.8) gives you dV. In the above design, the second stage, first stage, and boosters would share a common core. The second stage would admittedly be slightly less massive due to not having recovery gear, but I didn't factor that into my calculations. EDIT: fixing my numbers.

On to the first/parallel stage(s). Let's assume we use @TheEpicSquared's 1+2+2 configuration, with a core sustainer and two pairs of side boosters. GLOW is 1.2 tonnes. Determining when to throttle down and how far to throttle down is itself a major optimization problem, but let's start by assuming the following sequence, just to get ballpark performance: Launch: core at 100%, booster pair A at 100%, booster pair B at 100% 25% propellant consumption: core at 50%, booster pair A at 50%, booster pair B at 100% Booster pair B burnout: separation, core at 50%, booster pair A at 100% Booster pair A burnout: separation, core at 100% At this rate, we should expect to get 336 m/s from launch to throttledown, 817 m/s from throttledown to booster pair A burnout, 471 m/s from booster pair A burnout to booster pair B burnout, and 245 m/s remaining on the core, for a total of 1.87 km/s off the first stage. Now, the sooner we can downthrottle and the deeper we can downthrottle, the better, but TWR will become an issue if we downthrottle too low or too early. Combustion instability is also a reason not to downthrottle too aggressively. With 1.87 km/s for the parallel/staggered first stage and 5.655 km/s for the second stage and kick stage, we're already up to 7.52 km/s, which is DEFINITELY in the neighborhood of orbital flight. Of course, we have to subtract out gravity, pressure, and drag losses. But it's certainly in the right ballpark. And there's nothing saying we can't make a slightly wider stage than HELOS-3 and pack more propellant in, especially because we don't need the ultra-extreme fineness ratio typical of low-isp, high-thrust amateur rockets (amateur rockets make up in part for their poor isp by their very high thrust, cutting down on gravity drag but necessitating an ultra-streamlined profile to minimize more significant aerodrag losses). If our final design is generally in this ballpark, we can also get an idea of the minimum thrust rating we'll need in order to make it all work. In this configuration, the lowest TWR condition is at secondary booster burnout, when the core is lifting its own partially-fueled mass plus the mass of the upper stage, kick stage, and payload. In order to ensure sufficient TWR at this point in the ascent, I would suggest that an entirely full core ought to be able to lift an upper stage, a kick stage, and the payload with no less than a 2:1 TWR. Crunching the numbers, this means a naked stage TWR at least 4.4 and a single-motor thrust of at least 10.5 kN, about the same as the HELOS-3. Launch TWR would be 4.56:1, and hybrid-stage mass flow at 100% is 3.57 kg/s with a burn time of 47.3 seconds at full throttle. EDIT: I added in fuel incorrectly; correction below.

One nice thing about hybrids is that you can increase thrust to pretty much whatever you want by varying chamber length and throat diameter, without the mass penalty you get from increasing a liquid-rocket engine design. Let's say we use the Lambda 4S dV budget as a design target. I want to see if we are anywhere in the right ballpark. The Lambda 4S essentially had five stages. With the 1+2+2 first stage, a hybrid second stage, and a solid kick stage, as proposed by @TheEpicSquared, we're looking at a similar dV layout for our LV. If we want a total of around 6 km/s of dV on the upper hybrid stage plus our kick stage and we are guesstimating 4 kg payload (avionics + aeroshell + cubesat), then let's look at what we can do with White Lightning. Let's assume a propellant fraction of 80% on a hand-cast SRB and a vacuum specific impulse of 220 seconds. A little iterative modeling in Excel and I've got a sweet spot at 35 kg of solid propellant (any more than that, and too much impulse is lost on the hybrid stage lifting the weight of the SRB casing). The 300-s hybrid stage gives 2.67 km/s and the kick stage gives 3.08 km/s, for a total dV of 5.655 km/s. Just with these two stages, we'd already be dealing with a high-suborbital spaceflight. There are Android phones with accelerometers and gyroscopes; I wonder if someone could write an app to calculate telemetry in real-time based solely on that. Fair point.

As a random note, I just found out that GIRD-9, the first liquid rocket ever launched by the USSR, used LOX and jellied gasoline. Here's a photo of the unique combustion chamber (GIRD-9 at left; GIRD-10, a true liquid bipropellant rocket, at right): Additional thought: if we pressurize with N2O rather than atmospheric air, we could squeeze a little extra dV out of everything, but I dunno if it's worth the extra trouble. Anyway, back to the drawing board. I took a look at the HEROS-3 rocket and its propellant fraction is just so poor that it would be virtually impossible to get it into orbit no matter how many clusters we used. Thankfully, we have some improvements. The optimal O/F ratio for nitrous+paraffin (from page 18 here) is around 9.5, so if we go back to the HEROS-3 rocket, we can estimate that the 88 kg of propellant was 8.4 kg of paraffin and 79.6 kg of N2O. HTP is 2.2x as dense as N2O and jellied gasoline about the same as paraffin (0.9 kg/L). So the HEROS-3 carried around 9.33 L of paraffin and 119 L of pressurized N2O, for a total propellant volume of 128 L. HTP+kerosene has a mass O/F ratio of 7, corresponding to a volume O/F ratio of 3.23 (VERY nice), so I estimate we could pack 30.2 L of jellied petrol and 97.8 L of HTP into the same rocket, for a total propellant mass of 169 kg. This increases our propellant fraction from a damning 54% on HEROS-3 to a much more respectable 69.2%. Other improvements can be made, since we don't need avionics packages on each booster, but given the likelihood of including chutes, etc. we can ignore that. A 69% propellant fraction at an average specific impulse of 300 seconds gives us a whopping 3.47 km/s of dV on a single-stick. NOW we're cooking with gas (pun intended)! Repeated static fires will give us a very good projection of stage burnout time, and a pressure sensor in the main chamber above the fuel load (near the head pressure valve in the earlier diagrams) will show an instant pressure drop at fuel burn-through. This pressure drop will immediately send a signal to cut the oxidizer flow and trigger the staging sequence. I'm all for pneumatic deployment using residual HTP. RCS thrusters can be repurposed as open-cycle pneumatic pushers using a simple airtight seal. Tumbling terminal velocity shouldn't be too terribly high but this is definitely an issue to consider.

Absolutely. You won't hurt my feelings, haha. From my numbers, I just don't think clustering is avoidable. Without six-or-seven-figure budgets, hobbyists/amateurs/whatever simply can't manufacture solid-fueled or hybrid-fueled stages large enough to launch an orbital rocket serially. Copenhagen Suborbitals has built and tested some of the largest hybrid rockets ever, and they encounter serious combustion instability once they get up to 100 kN or so. Of course they also struggle with using nitrous oxide, since its low gaseous density gives the stages a pitiful propellant fraction of 40%. Amateurs can, however, construct peroxide-based hybrid rockets with much higher propellant fractions as long as they are kept reasonably small. And that's where clustering comes in. I don't like the computational requirements either, but I don't think there's another way. Yes, checkout is difficult. However, combustion instability won't be a problem; the injection is behind the catalyst bed, so as soon as the HTP hits the catalyst bed it will decompose anyway. Anyway, this is a smaller factor. Yeah, most likely not. Of course, given that decomposed HTP is hypergolic with petrol, we can potentially use a second hybrid stage in serial, in which case we would be able to recover the core but not the second stage: Although not shown, you could have anywhere from 2 to 6 parallel boosters, and they could separate simultaneously or pairwise. That's not too terribly complicated, really. A very kerbal solution to be sure. There's definitely a limit to asparagus staging; unless you have crossfeed, there's only so much you can do with throttle control. If you go much beyond 1+2+2, you'll end up with viciously diminishing returns. As a very basic initial approach, what if we use the height, diameter, and mass of the successful HEROS-3 sounding rocket, and simply swap in our propellants? I'll try to do the math on that and see what kind of performance we would have. A hybrid pressure-fed rocket is going to be MUCH sturdier than the wispy flying tin can that is the Falcon 9 first stage. Side boosters (and possibly the first-stage core, if serial staging was used) would chute down and land on their sides.

The advantage of clustering over a single-stick approach is that amateur (read: low-cost) solid and hybrid rocket manufacturing does not scale well. With differential throttling, some underperformance can be compensated for. Besides, these would be designed for reuse, so they would all be static-fired ahead of time.

Mass fraction will depend on tank and intertank wall thickness, which in turn will depend on combustion pressure. So if anyone can dig up something instructive regarding the range of possible combustion pressures for decomposed HTP burning with jellied petrol in a hybrid rocket engine, I'd appreciate it. Just a general range is enough to get things going. The higher the pressure, the stronger (and heavier) the tank and intertank will have to be in order to contain it. I know it's easy enough to find tank pressures for pressure-fed peroxide-kerosene rockets, but I don't know how hybrid rocket combustion pressures differ.

And now that this semantic argument has been beaten to death... You can really just replace "amateur" with "minimum cost" since professional-grade turbopumps, SRBs, and the like are going to eliminate amateurs from a cost basis before anything else. Once we have a good first-order approximation of stage mass fraction, I can start some basic optimization...if nothing else, to see how many OTRAG clusters you'd need if we took a purely serial approach. After I have the calculations down for a serial approach, we can start moving from serial to parallel to see how small we can make the overall vehicle. Then start optimizing.

My assumption was that a high-pressure valve's flow rate scales roughly with its weight, so the number of valves required would be less about throttling range and more about the flow requirements. If the max flow into the combustion chamber is 4 times as high as flow into each individual RCS thruster, then you'd want 1 valve per RCS thruster and 4 valves in the main chamber. If flow is 8 times higher, then you'd use 8 valves in the main chamber, and so forth. Yeah, it's fast. I can do more complicated stuff but for this kind of thing, it's hard to beat MS Paint for speed. Ehh, I don't know. Reaction wheels are expensive, heavy, and require rather complex electronic controls. Better to use the last puffs of HTP in the terminal hybrid stage to spin up just before separation. Spin-stabilization works well enough. I'm not confident that a 4-core design would be enough to get to orbit, but I'll see. Fineness ratio is really important for small launch vehicles. We want it to be as narrow as possible without running into bending-moment problems. Just how narrow we can get is going to depend on how thick the HTP tank walls will need to be. That, in turn, will depend on what kind of combustion pressures we are dealing with. Recall that the head pressure in the HTP tank needs to be higher than combustion pressure all the way to burnout; that's the challenge with pressure-fed rockets. By PV=nRT, the ratio of pressurant volume to total tank volume must be equal to the ratio of liftoff pressure to combustion pressure, since the pressurant will need to expand to fill the whole tank. So this means the HTP tank walls need to be thick enough to contain that kind of pressure. You want a large diameter so that you maximize tank volume for a given tank wall area, but not so large that your fineness ratio drops too low. (I wish there was a way to decompose a small flow of the HTP inside the tank in order to maintain tank pressure, as this would mean the tank would only need to hold slightly higher pressures than the combustion region rather than 5-10x greater, but I can't think of a way to do it. HTP decomposition is heat-catalyzed; if one gram of HTP decomposes inside the tank, the whole tank instantly decomposes. I suppose you could try to do something with liquid nitrogen to keep the tank temperature low enough to prevent runaway decomposition but that seems ridiculously complicated and rather heavy.)

Actually 9,414 m/s net dV. The final orbit was a 322 x 2414 km eccentric orbit with perigee velocity of 8,217 m/s, so that's 1,197 m/s in aerodynamic and gravity drag losses.

Okay, the Lambda 4S had 509 m/s of net dV from launch to side-booster burnout, 932 m/s of net dV from side-booster jettison to first-stage burnout, 1,849 m/s of net dV on the second stage, 2,209 m/s of net dV on the third stage, and 3,915 m/s of net dV on the fourth stage. It also had a launch TWR of around 8.