sevenperforce

If it's farther than GEO then it will still be able to see all of Earth over the course of its orbit. Yes, some amateur astronomer might point a scope at it, but it would be difficult to pick up again if it was in an orbit like that. Also limited reflectivity would make it virtually invisible. I don't see how it could be responsible for anything on Earth but then again I haven't read the book.

My understanding of New START was that MIRVs are fine, but each warhead counts as an individual missile even when clustered on a single missile for the purposes of deployment limits. But maybe I am incorrect.

It is cooled internally by cryogenic fuel flow much like an engine bell from what I recall. That's how they are cooling the heat shield, but insulation is a different issue entirely -- keeping the propellant from boiling off due to ambient heat transferred through the skin. To @RyanRising's question -- my guess is that these early hopper prototypes have simply dispensed with insulation altogether. The low fineness of the stage minimizes surface area, anyway. For short hops and even suborbital hops, the amount of propellant lost to boiloff is going to be relatively negligible. Once they have an actual orbital vehicle under construction, we'll see whether they do some sort of insulation. It may be that the additional weight of insulation would be greater than the weight of propellant lost to boiloff anyway, and so eschewing insulation entirely is the best approach.

No, that doesn't have anything to do with inefficiencies; it's a fundamental aspect of rocket engines. A rocket engine has to push something out of the back end in order to produce forward thrust. Well, some batteries lose weight during discharge and others gain weight during discharge, but most remain about the same. However, that's not quite the right comparison. You can use up a battery in any number of ways on board a spacecraft, but unless you're yeeting some sort of exhaust out of the back of your spacecraft, your spacecraft isn't going to move. Electric ion thrusters operate by using electrical energy to accelerate tiny amounts of ionized gas at high velocity, but even they are subject to changes in acceleration as the gas is used up and the vehicle becomes lighter. The Dawn spacecraft, for example, carried a total of 425 kilograms of xenon gas, about 35% of its total weight at launch. If fuel use is constant then yes, you have a constant change in acceleration and you can determine the distance covered accordingly. You just have to go to a third-order kinematic equation instead of the more familiar second-order kinematic equations. Velocity is the rate at which position changes; an object moving at 10 m/s is changing its position by 10 meters every second. Acceleration is the rate at which velocity changes; an object accelerating at 10 (m/s)/s or 10 m/s2 is changing its velocity by 10 m/s every second. The rate of change in acceleration is called "jerk" (or sometimes "jolt") and it has units of ((m/s)/s)/s or m/s3. The third-order kinematic equation for distance can be readily derived by integration, but I'll skip that step and just give it to you: x = x0 + v0t + 0.5*a0t2 + 1/6jt3 In order to solve this equation, you'll need to know each of the starting values (x0 , v0, and a0) as well as your value for jerk (j). In your case, the starting position x0 seems like it is going to be zero, and the same appears to be true for the starting velocity v0 as well. So you need to define a0 and j. If your initial acceleration is 7.3392 m/s2, then that's your a0. How do we calculate j? Well, with constant propellant use, j will be a constant, so we can use this equation (should be self-evident; let me know if it's not): j = (af - a0)/t In order to accelerate a starting mass of 86,400 tonnes at 7.3392 m/s2, you'll need an engine producing 634 meganewtons. By the end of your burn, your vehicle mass has dropped to 21,600 tonnes but you're still thrusting at 634 meganewtons, giving you a final acceleration of 29.36 m/s2. Using our equation above, we get jerk equal to 29.36 m/s2 minus 7.3392 m/s2 divided by 600 seconds, or 0.0367 m/s3. Now that we have that, we plug all the rest in. x = x0 + v0t + 0.5*a0t2 + 1/6jt3 x = 0 + 0*600 sec + 0.5*7.3392 m/s2 * (600 sec)2 + 1/6*(0.0367 m/s3)*(600 sec)3 x = 0 + 0*600 sec + 3.6696 m/s2 * 360,000 s2 + 0.00612 m/s3 * 216,000,000 s3 x = 1,321,056 m/s2 * s2 + 1,321,920 m/s3 * s3 x = 2,642,976 meters So there's your answer -- which you will note is about double the answer you would have gotten without accounting for change in acceleration. Nope, it is not. Kinematic equations assume no air resistance. If you're talking about something moving through an atmosphere then we're gonna have a whole other set of issues. The standard kinematic formulae are unit-agnostic.

I've been out of pocket. Wenhop? Fairly certain that everything after about T-30 seconds is almost completely automated. They program in all of the parameters and tell it what to do with all contingencies and give it abort modes for anything out of spec.

As usual @RCgothic is entirely correct but I will also point out, just in case OP missed it, that the constant acceleration equation will not work properly for a reaction engine. If you are shoving exhaust out of a nozzle, then you also have to account for the decrease in your total vehicle mass across the period of the burn. You'd need to use the Tsiokolvsky rocket equation for that. The 600 is most definitely the result of the 600 seconds. You're just doing iteratively what s=0.5*a*t^2 does non-iteratively.

I don't believe so.

The small size is just incredible.

Lol, logged in at T+14 seconds. Staging and second engine start are good. Aaaand stage 1 landed. Waaaay off to one side though.

I can't remember if we've covered it -- are these propulsive vents?

The hiring laws prohibit unlawful employment discrimination (on the basis of race, color, religion, sex, or national origin) but the ITAR laws trump them, because refusing to hire someone based on ITAR is not unlawful. In other words, you aren't refusing based on national origin; you're refusing based on ITAR status. So if you are a foreign national who applies to work for ULA in an ITAR-sensitive role, ULA can honestly say "Sorry, we can't consider your application, at least not without a specific ITAR exemption from the State Department, because ITAR will not allow us to share information with you." Where SpaceX screwed up (intentionally or unintentionally) was saying the same thing to refugees and asylees with legal status even though ITAR does not prohibit them from hiring refugees and asylees. To add to my prior examples: it would be like a hospital telling a gay man, "We can't hire you because the FDA prohibits gay men from coming into contact with blood" when in fact the FDA rule (which has now been rescinded, finally) prohibited gay men from donating blood but said nothing about healthcare employment generally. If it was JUST a job posting, that might be chalked up to a mistake, but if the hospital was also refusing to consider applications from gay men who applied for maintenance work or billing jobs (e.g., no physical patient contact) then it would look much more like a deliberate lie to disguise homophobic discrimination.

Oh, so they need to change the specific laws cited on their job postings? Thanks for that analysis, makes more sense now. Yeah, basically. But not just the laws they cite -- it's also the terms they use, and the terms they don't use. It's like if the admissions office at a university had a posting that said "Federal law prohibits us from considering student applications without proof of Selective Service registration or proof of exemption." Since only males can register or be exempted from Selective Service, this would imply that federal law prohibits the university from accepting female students, which obviously is not true. Such a posting might just be a mistake/misstatement, but if there was evidence that the contractor refused to accept female students on this basis (e.g., if it refused applications from women on the basis that they did not provide proof of Selective Service registration) then the DOJ could take enforcement action under Title IX. Or imagine a job posting for a teacher at an elementary school which said "Bright or exotic dyed hair can be a distraction to students: acceptable hair colors for applicants include blonde, strawberry blonde, platinum, light brown, and auburn." This suggests that individuals with naturally black hair (e.g., people of a particular marginalized ethnicity) need not apply. Again, this could just be a miscommunication or mistake, but if they actually turned away applicants based on hair color then they could get in trouble.

The DOJ lawsuit sets out some pretty specific claims. Among these is the claim that SpaceX used its public announcements, job applications, and online recruiting communications to exclude asylees and refugees. This isn't hard to verify: the internet is forever, after all. Then I used WayBack to look up historical postings and found this: To conform to U.S. Government space technology export regulations, including the International Traffic in Arms Regulations (ITAR) you must be a U.S. citizen, lawful permanent resident of the U.S., protected individual as defined by 8 U.S.C. 1324b(a)(3), or eligible to obtain the required authorizations from the U.S. Department of State. Learn more about the ITAR here. And then I looked up a current job posting on the SpaceX recruiting website (emphasis shows additions): To conform to U.S. Government export regulations, applicant must be a (i) U.S. citizen or national, (ii) U.S. lawful, permanent resident (aka green card holder), (iii) Refugee under 8 U.S.C. § 1157, or (iv) Asylee under 8 U.S.C. § 1158, or be eligible to obtain the required authorizations from the U.S. Department of State. Learn more about the ITAR here. So that's the difference. It might not seem like a huge difference, but it's a meaningful one: as job postings go it plainly violates Title VII, by discouraging applicants based on national origin. An inaccurate job posting by itself isn't proof of discriminatory intent, but taken together with the other evidence proffered by DOJ's lawsuit it tends to support their allegations of broadly discriminatory practices (particularly if the DOJ notified them and they didn't correct it). Form I-589 (the paperwork required for application for asylum) can be filed before crossing the border or arriving at a point of entry to seek asylum, but it does not have to be filed in advance. Federal law gives asylum-seekers up to a year after they enter the country before they are required to file (18 USC § 1158(a)(2)(B)). In fact, if initial screening is conducted by USCIS, then USCIS will automatically treat the screening itself as a filing of the application. However, an "Asylee under 8 U.S.C. § 1158" would be a person who has not only filed Form I-589 but has actually been granted asylum under subsection (b)(1)(A) of the code section. So, not some rando who walked across the border and just claimed asylum.

If it's a working engine, yes. This is very common in Russian designs and other engines with Soviet legacy, but less common in Western designs (the Rocketdyne XLR-89-5 is a notable exception). However, it's important not to confuse the entire engine assembly with individual combustion chambers. Here's an example, shown by the venerable RD-180: This is one liquid propellant rocket engine with a single thrust structure, a single gimbal mount, a single turbopump, and a single propellant flow system. However, there are two combustion chambers and two nozzles. I suspect that when you are calling something an "engine" you are actually talking about the combustion chamber. You would never want to make a single combustion chamber feed multiple nozzles, because the pressure loss along the path from the chamber to the various nozzles would be too inefficient. The only exception would be for a very low-impulse system, like a set of attitude control thrusters or propulsive vents. Try to think about individual rocket engine combustion chambers like the individual pistons in an automobile engine and the entire rocket engine assembly like the completed automobile engine+fuel pump+crankshaft system. There is no real limit on the power from a single engine. When you're talking about rocket engines, you need to be conscious of both power input and power output. The power input is the amount of power that is generated by the turbopump in order to push the propellants into the combustion chamber(s). The turbopump is basically a jet turbine engine all on its own and can be almost arbitrarily large; the fuel turbopump for the Aerojet M-1 engine produced a 75,000 horsepower, and the RD-170's single-shaft turbopump produced a whopping 230,000 horsepower. That's up there with the horsepower of the largest diesel engines ever produced, although it is dwarfed by the effective horsepower of airliner engines (like the GE-90's 21 trillion horsepower). The drive to multiple combustion chambers was not an issue of power, but an issue of combustion stability. Once you end up with a gigantic combustion chamber (the volume of the F-1 engine combustion chamber was almost a full cubic meter), the flow of massive amounts of propellant starts to produce currents that can impede combustion. That's why the Russians used multiple combustion chambers, and it's one of the reasons that a company like SpaceX would rather have dozens of smaller Raptor engines than build a gigantic Super Raptor. Nope, not boom. Who told you that bigger nozzles are needed for higher chamber pressures? Or did you just assume that? That's not how it works. When exhaust comes yeeting out of the combustion chamber, it has a lot of heat and a lot of pressure. The nozzle allows the exhaust to expand, trading heat for velocity and providing a surface against which the expanding exhaust can push to produce thrust. Complete expansion is impossible without an infinitely long nozzle, so you have to truncate the nozzle somewhere. Pressure is the force that a fluid exerts on a containing surface (and on itself). Force is proportional to acceleration, so the higher your chamber pressure, the more rapidly it will accelerate as it travels down the nozzle. At very high combustion pressures, you can get away with a relatively short nozzle, because of how quickly the exhaust expands. At low combustion pressures, you need a longer nozzle to give the exhaust time to fully expand. So it's exactly the opposite of what you were told. If your nozzle is not large/long enough for your design pressure, it won't hurt anything; you're just wasting pressure that could otherwise be converted into thrust (the exhaust will spill out the edges of the nozzle before it has fully expanded). Higher chamber pressure equals more thrust for an equivalent sized nozzle. If you increase the pressure, the forces acting on the nozzle are greater, and so even though you still have spillover at the edge of the nozzle, you will have gotten more out of the exhaust flow by the time it reaches the edge of the nozzle. If you want to increase thrust directly, you have to increase the amount of propellant going into your engine, usually by spinning the turbopumps faster. There is a limit to how fast you can spin your turbopumps before you just kinda have to build a bigger turbopump. There is no real limit on how high the chamber pressure can be, except for the fact that a higher chamber pressure requires a heavier chamber (to contain the pressure) and a beefier turbopump (to push the propellant into the chamber at a higher pressure) which means your engine will be much heavier overall. So there's a tradeoff between chamber pressure and engine weight. There's also a tradeoff between chamber pressure and turbopump power consumption. The turbopump typically burns a portion of your propellant to provide the power to push the rest of the propellant into the chamber, but if you ramp up your chamber pressure too high, then you end up using up too much of your propellant inside the turbopump. It's perfectly fine to have multiple combustion chambers feeding a single nozzle -- it's common in aerospike designs. Certainly no problem with melting the nozzle, if the nozzle is regeneratively cooled. I'm not sure why you're always afraid of things melting. Melting is certainly a common challenge in the world of rocket science, but it's a technical/efficiency challenge, not a basic limitation of physics. Melting is usually NOT the limiting factor on making an engine bigger or more powerful.

SpaceX does not conduct "full flight duration" static fire tests of any of its launch vehicles. When SpaceX announces that it has completed a "full duration" static fire, it means that the static fire went the full duration it was intended to go. This is rarely more than a few seconds; the longest Falcon Heavy static fire went for a dozen seconds. AFAIK, no launch vehicle in history has ever conducted a full thrust, full flight duration static fire test. The closest analogue is probably the Green Run testing campaign for SLS which sometimes goes full flight duration but of course is only operating at a little over 20% of full liftoff thrust. ...something that has never been done on any rocket, ever... What, precisely, is the advantage of this imagined full-thrust full-flight-duration static fire over a test launch? If your goal is to line the pockets of whatever company is supplying parts for ultra beefy hold-down clamps, I suppose this makes sense. No other reason though. How do you figure? A booster with a third of the liftoff thrust of Superheavy cannot possibly deliver 100 tonnes to LEO with full reusability when Superheavy+Starship can only do 120-150 tonnes to LEO with reusability. If you mean flying smaller payloads with full reusability, then it becomes really questionable whether you're reaching any advantages over Falcon 9 and Falcon Heavy, since the reuse of your "mini-Starship" becomes challenging. How do you EDL the upper stage? You can't just scale down the current Starship design, because even the "skydiver" entry works the same way, your landing won't: Raptor Vacuum is vastly overpowered for a hoverslam landing. If you mean smaller payloads with only first-stage reuse, then you really have no advantage over the Falcon family. SpaceX already has the market cornered with Falcon 9 alone. Why do you think they aren't flying Falcon Heavy more often? The answer is simple: when Falcon 9 can deliver virtually every commercial payload with first stage reuse, there's really no need to add more cores. Going from Falcon 9 to Falcon Heavy was a nightmarish dev path. The Falcon Heavy core is a completely new rocket compared to the Falcon 9 booster. If you want a Raptor-based vehicle with more than 9 engines, it's a better plan to just go big from the start. A scaled-up Falcon 9 upper stage powered by a single Raptor Vacuum would suffer from high dry mass, making it much less useful for single-launch missions BLEO. That's what they thought the first time.... Unless I misremember, SpaceX was very much anticipating that substantial pad damage would be a potential outcome. They thought there was a chance it might survive, but they knew there was a chance that it wouldn't, and figured that as long as they got good launch data, it would be worth it. They didn't anticipate the degree of pad destruction, but clearly it wasn't catastrophic given that they have already gotten it functional again. Do you have any data to support your dismay? Do you have any data that suggest the early shutdowns were an engine reliability issue? I don't think anyone outside of SpaceX and perhaps certain people at the FAA know whether there were actually any engine failures in the first flight or not. There were engine shutdowns, yes. There are no data that suggest one way or another whether these engine shutdowns were commanded. While I think we are generally in agreement, I will note that RS-25 failures have been responsible for a total of seven launch aborts: six with the Shuttle and one with SLS. In one case the RS-25 failure was seconds away from causing LOCV. This just goes for the proposition that engine-out capability is generally a very good thing.

Oh, I certainly agree with that. SLS is far too expensive for cargo. Far too expensive for anything, really, but definitely for cargo. If it's "Centaur-like" then you're not getting 45 tonnes of hydrolox inside even the Falcon Heavy extended fairing, as I explained above (and re-explain in more detail below). The RL10-C-2-1 has never been paired with a Centaur, but is still in production and achieves your desired specific impulse. It's a whopping 163.5" long but fortunately 77.2" of that is a deployable nozzle extension, leaving the 64.1" engine and the 22.2" fixed nozzle extension to give a total stowed height of 86.3". Here are the internal dimensions of that extended fairing, lifted straight from the F9 user guide. The numbers on the left are height above the payload adapter; the numbers on the right are diameter: I'm not sure what cargo you're thinking of, but for the sake of simplicity let's cut the hydrolox stage off at ST = 477.976", right where the ogive starts. That means your payload, whatever it is, needs to fit inside a right truncated cone with a height of 175", a base with diameter 180", and a top with diameter 49". Let's allow 4" of clearance all the way around the stage for downcomers, vents, and the like (remember that this thing has to perform some significant maneuvering), giving us a stage diameter of 172". Tank wall thickness is of course negligible. Rule of thumb is that the most efficient dome shape is an ellipsoid of height R/2(1/2), so this gives an ellipsoidal cap height of 60.8". The double-walled intermediate bulkhead has a thickness of 0.3". Neglecting any volume occupied by stringers and the like, the total available propellant volume is represented as the sum of the volumes of an oblate spheroid of height 121.6" and a cylinder of height 269.8", both with a diameter of 175": The ellipsoid has a volume of 31.95 cubic meters and the cylinder has a volume of 106.34 cubic meters, giving this "Centaur Plus" frankenstage a total available internal propellant volume of 138.29 cubic meters. At hydrolox's bulk density of 0.28 g/cc, this gives 38.7 tonnes of propellant. Now, the RL10-C-2 has a propellant mass flow rate of 24.07 kg/s, giving it a total burn time of 1,608 seconds. That's far, far too long -- a full third of an LEO orbit (although obviously it wouldn't happen all at once; I'm just illustrating) -- so the Oberth losses would be immense. You're going to need two engines. It will be a tight squeeze under the payload fairing but it can probably be done. If we're going "Centaur-like" then let's apply the Centaur mass ratio closely. Centaur has an empty mass of 2,247 kg, which drops to 2,032 kg when you subtract the mass of the RL10-C-1. It carries a total of 20,830 kg of hydrolox, giving it an engineless mass ratio of 10.25:1. Thus to hold 38.7 tonnes of propellant, our frankenstage needs a dry mass of 3,775 kg. Add the increased weight of two RL10-C-2s and the stage mass comes up to 4,377 kg. With 12 tonnes payload, the combined stack develops 5,540 m/s of Δv. Falcon Heavy can deliver 63.8 tonnes of payload to LEO, but that necessarily includes its own residuals. The PAF can't handle a ~55 tonne payload so it would need a special stage adapter to brace against the bulkhead or something. Estimating 1.8 tonnes for that structure, Falcon Heavy's "payload" is 56.9 tonnes and it reaches LEO with 6.9 tonnes of its own propellant to spare. Estimating FHUS at 4.6 tonnes, this means FH can burn the last of its props to give the upper stage stack thingy a boost of around 363 m/s past LEO, which is at least something helpful. Thus the total Δv available to the payload from LEO is 5.9 km/s. TLI is 3.2 km/s, LOI is 0.9 km/s, and descent to the lunar surface from orbit is 1.87 km/s, for a total requirement of 5.97 km/s. Just shy. However, you do have to actually LAND your payload, and our Centaur Plus frankenstage isn't going to be useful for that. If you go with a crasher stage architecture, then we can easily give the actual payload some small pressure-fed hypergolic thrusters to perform the last 70 m/s of the landing burn, plus whatever is necessary for the actual landing maneuvers. Of course, a 12-tonne payload is enough for a lunar ascent vehicle. On the other hand, this would require building a completely new intermediate Centaur frankenstage. IIRC, Centaur V was announced ca. 2012 and they still haven't ironed out the kinks, so we can assume a similar timeframe for this sort of retrofit. If you want to just play rocket legos with the existing Centaur SEC, that's much more doable. With the same intended 12-tonne payload, the existing Centaur SEC develops 3,982 m/s of dV and has a stack mass of 35.1 tonnes. It fits easily in the extended fairing with significantly more room for payload. Launched on Falcon Heavy (using a notional 1.3-tonne payload adapter frame), the Falcon upper stage would reach LEO with an impressive 27.4 tonnes of propellant residuals, allowing it to deliver a heft 1,747 m/s to the Centaur and its payload. This means the whole stack boasts 5.7 km/s of Δv, just 200 m/s less than the first design despite using only about half as much hydrolox. AND that's with the lower specific impulse of the RL10-C-1. Just goes to show that specific impulse ain't everything. Once you are going beyond LEO, mass ratio becomes more important than specific impulse, which is why the Merlin 1DV and the single RL10-C-1 can beat a pair or RL10-C-2s. The high specific impulse of hydrolox is more useful for lifting large monolithic payloads into LEO in the first place (think: Saturn V second stage).

F9 is already close to the maximum fineness ratio for a rocket. Wind conditions at altitude are a problem with high fineness because the bending moment increases dramatically. A significant tank stretch isn't really on the table. I’ve seen numbers for Falcon Heavy to TLI in the range of ~20 tons Oh, absolutely. FH can absolutely send around that much to TLI. Which is why replacing the upper stage with hydrolox for only ~9-10 tonnes to TLI wouldn't be a good idea. I'm confused -- are you now talking about a four stage vehicle? Or 4.5 stage, counting the FH side boosters? Look at pages 83 and 88 of the Falcon 9 User Guide. The extended fairing has a cylindrical payload area 478" high and 181" in diameter, and there's an additional 175" of conical payload volume above that cylindrical plane. Let's imagine that your notional TLI hydrolox stage used a cluster of BE-7 engines to maximize utilization of volume and needed about 4" for tank walls, external fittings, and clearance from the fairing. The BE-7 engines are 80" tall. Let's imagine a lunar lander slightly squattier than the Apollo LM, at 5 meters height. That cuts 23" into the cylindrical region. Finally, let's assume a 1" insulated common bulkhead and shave off another 18" equivalent of vertical volume to account for the ellipsoidal caps. So that leaves a total available tank volume of 137 cubic meters, which gives us space for 38 tonnes of hydrolox. That's best-case-scenario assumptions with a single-stage architecture. If you try to stack two hydrolox stages on top of each other, your total combined propellant load drops to 28 tonnes of hydrolox. I'm still not quite sure what exactly you're proposing in terms of the mission profile, but from the overall context I think you are envisioning a lander+stage combo being launched on Falcon Heavy, meeting up with separately-launched crew somewhere in cislunar space, and then taking the crew down to the lunar surface and back up. For safety and simplicity, I'll consider a crasher stage architecture, where a zero-boiloff hydrolox stage performs both the braking burn into lunar orbit as well as the descent burn, and drops off to allow the lander to perform the final hovering landing as well as the ultimate ascent. Let's use the 9-12 tonne ascent vehicle concept from the Artemis initial studies; I'll go with 10 tonnes to make sure we have enough extra mass for landing legs and hovering propellant. The crasher stage needs 2.77 km/s of Δv to brake into cislunar space and then take the lander down to the surface. Assuming 450 seconds of specific impulse, the stack propellant fraction needs to be on the order of 47%. Assuming a relatively decent stage mass ratio of 10:1 including engine(s) and insulation system, the stage needs to be 52% of the total stack mass, giving us a total stack mass of 20.8 tonnes. But here we see there's no need for stacking up multiple hydrolox stages. Falcon Heavy can already deliver ~20 tonnes to TLI, so adding an additional stage underneath is completely unnecessary.

With a hydrolox upper stage, a fully expendable Falcon Heavy can only send 9.3 tonnes of payload to TLI, let alone to the lunar surface. Hydrolox is fluffy; a hydrolox Falcon upper stage would only be able to carry around 30 tonnes of propellant.

Sorry, but this is incorrect. Shockwaves or any kind of wave can be reflected any direction. Shockwaves cannot travel upstream in a supersonic flow. If they could, then supersonic jets would be destroyed by shockwave reflections. But more to the point, it's now become unclear what you're talking about here, exactly. Are you saying that the exhaust plumes could be reflected off of the water deluge and impinge on the booster? Or are you saying that the sonic shockwaves produced by engine sound pressure could be reflected off of the water deluge and cause more harm to the booster than sonic reflection off bare concrete? If the former, the answer is simply no. Exhaust plumes don't bounce. If the latter, what you're suggesting is at least physically possible, but since the water is not in a steady-state flow that is both planar and laminar, it's not possible here. The fact that rocket engines inevitably produce reverberating sound waves -- something we have known for a long long time -- does not lead to the unsupported conclusion that this specific water deluge system would cause damage to the engines. There is neither a mechanism of action for such damage nor evidence that such a mechanism is in play. Which Apollo test taught the "lesson" of using a flame trench? The Apollo launch site was built from the ground up with a flame trench. So you're not really advocating that we learn from "the lessons of Apollo"; you're just saying "they should do it like Apollo did it" without any reference to lessons. The pads that launched the Saturn V never had a sound suppression deluge at all; their meager water spray was for cooling and fire prevention. A sound suppression deluge was not added until after the first few Shuttle flights, to reduce sonic shockwave tile damage. And SpaceX does have a flame trench. It has six of them, in fact. The Saturn V's flame trench was 13 meters deep; the Orbital Launch Platform for Superheavy is 25 meters deep.

Exhaust flow is supersonic, so pressure waves can't travel backwards. Explosion shockwaves for example can bounce off of atmosphere layers and bounce back to Earth known as atmospheric focusing: Atmospheric focusing is a type of wave interaction causing shock waves to affect areas at a greater distance than otherwise expected. Variations in the atmosphere create distortions in the wavefront by refracting a segment, allowing it to converge at certain points and constructively interfere. In the case of destructive shock waves, this may result in areas of damage far beyond the theoretical extent of its blast effect. Examples of this are seen during supersonic booms, large extraterrestrial impacts from objects like meteors, and nuclear explosions. Atmospheric focusing can take place when a shockwave interacts with a discrete atmospheric boundary later, similar to the way that light will be partially refracted and partially reflected at the boundary layer between water and air or between glass and water. This requires, however, that the boundary layer exist in the same medium through which the wave is propagating. Shock waves are pressure waves that travel through the air, and since the air is the thing that has the boundary layer, the boundary layer can refract the shock wave. The exhaust flow from the business end of a Raptor engine is not a pressure wave in a static medium; it's a supersonic flow of exhaust. Supersonic flows are not shock waves, and the mass of chaotically-moving water spray is not a propagation medium. Even if this was an apt analogy, which it isn't, a pressure wave experiencing atmospheric focusing is still going to be traveling away from the source, not back toward it. When a near-laminar supersonic flow impinges on a surface, it doesn't bounce. Rather, the collimation is destroyed and the flow experiences a transition normal to the original direction of travel, carrying the energy away in the normal plane. Sound pressure is an issue, of course, but the water significantly damps the sound pressure. There's no reflection off the water spray.

Water, like everything else, travels in a parabolic arc. So if it is angled outward at the base, it will be traveling at a shallower angle once it reaches the level of the engines, meaning that it will miss the engines. Exhaust flow is supersonic, so pressure waves can't travel backwards. Purely speculative and highly implausible. Also, a commanded shutdown is not the same as a failure. CRS-1 had an engine failure on ascent. A commanded shutdown during a static fire doesn't tell us anything.

I don't know what the RC setup is like, but if you add a gyro to the vehicle so that you can see the pitch and yaw in real time then perhaps you'd have better luck controlling the TVC manually.

Someone actually did it. I had an idea for a much simpler self-landing rocket a while back. The idea was to do a two-stage solid-fueled rocket, but use a cluster of motors on the first stage. Using clustered motors can be dicey with a hobby rocket because of inconsistent startup transients, but it can be remedied as long as you have a sufficiently large central motor to provide the bulk of liftoff thrust. Basically you'd use a short, stubby, high-powered central motor ringed with four additional smaller, longer motors to act as sustainers, each angled slightly to point through the center of mass and thus reduce the impact of off-axis thrust during startup and shutdown transients. The second stage would pop off and fire normally. At separation, the first stage would have "grid fins" that pop out, releasing long transparent streamers to provide drag and keep the vehicle oriented tail-first. The landing would be achieved using a CO2 cartridge leading to a pair of cold-gas nozzles tucked in with the sustainers. A cold-gas blow-down thruster has the advantage of starting with high thrust that drops off quickly, which is just what you need. The idea would be for the grid fins and streamers to provide pointing and give it a reasonable terminal velocity, then trigger the CO2 burst disc at about 30 feet above the ground. The blow-down would kill maybe 80% of terminal velocity, at which point the shock-absorbing landing legs would handle the rest. Any remaining pressure in the CO2 cartridge would be lost quickly enough that the residual T/W ratio would be less than 1. Cold gas also has the important advantage of not starting gas fires, which is helpful on a test range.

The engine exhaust gases are in collimated, laminar flow; the water jets are not in laminar flow. Thus the pressure drop is exponential.

Just glancing through this thread but I spotted right out of the gate that you're adding 10 tonnes of propellant to the Orion service module, but you're not adding any dry mass or different geometry to the Orion service module. Where would you propose these 10 tonnes of hypergolics hang out? Ziplock bags?