sevenperforce

Setting aside this microwave plasma jet, what about conventional props? Propellers can reach 92% efficiency at converting shaft torque into thrust. For an ICE or turboprop engine, torque at the shaft is only a fraction of the available chemical energy due to friction and Carnot inefficiencies, but a direct-drive electric motor can exceed 95% efficiency. I've only ever flown two planes, one of which was the ever-popular Cessna 172 Skyhawk. Specifications for the Cessna 172R: Fuel capacity: 163 kg Engine: Lycoming IO-360-L2A, four cylinders, 120 kW Engine weight: 117 kg Cruise speed: 63 m/s Fuel consumption at cruise: 35 kg per hour Maximum cruise duration: 4.6 hours Theoretical max cruising range: 1000 km A single Tesla Model S motor masses just 32 kg and pushes 270 kW shaft power. A linear downscale to 120 kW would require just 14 kg of engine. If we use that saved engine weight for more batteries then our new "fuel capacity" in the same plane becomes 266 kg. To account for 95% motor efficiency and 92% thrust efficiency, we can just act as though our battery weight budget is just 232 kg. That mass budget gives us only 41 kWh of battery power, which will give us a pitiful 21 minutes of cruise and a range of just 80 km. Of course, if you could use a lithium-air battery at an effective 18 MJ/kg (5 kWh/kg), that mass budget would get you 1,160 kWh which is enough for a theoretical 9.6 hours of cruise, more than double what a Cessna 172 can do. So that would be quite promising.

This is hands-down the best explanation I've ever heard of this. Thanks.

Blurry photo, but can we tell anything new from it?

They sent out another email to clarify that they intentionally didn't clarify: "For this phase of the screening process, you will be asked to shoot and submit a one minute video. Before shooting and submitting your video, make sure you follow the guidelines below. There is no specific topic for you to present in the video and the content is free at your own creativity."

Wouldn't be much of a difference -- jet exhaust is already 30-32% water vapor by weight.

Presser: As early as Friday, March 26, the SpaceX team will attempt a high-altitude flight test of Starship serial number 11 (SN11) – our fourth high-altitude flight test of a Starship prototype from Starbase in Texas. Similar to previous high-altitude flight tests of Starship, SN11 will be powered through ascent by three Raptor engines, each shutting down in sequence prior to the vehicle reaching apogee – approximately 10 km in altitude. SN11 will perform a propellant transition to the internal header tanks, which hold landing propellant, before reorienting itself for reentry and a controlled aerodynamic descent. The Starship prototype will descend under active aerodynamic control, accomplished by independent movement of two forward and two aft flaps on the vehicle. All four flaps are actuated by an onboard flight computer to control Starship’s attitude during flight and enable precise landing at the intended location. SN11’s Raptor engines will then reignite as the vehicle attempts a landing flip maneuver immediately before touching down on the landing pad adjacent to the launch mount. A controlled aerodynamic descent with body flaps and vertical landing capability, combined with in-space refilling, are critical to landing Starship at destinations across the solar system where prepared surfaces or runways do not exist, and returning to Earth. This capability will enable a fully reusable transportation system designed to carry both crew and cargo on long-duration, interplanetary flights and help humanity return to the Moon, and travel to Mars and beyond. There will be a live feed of the flight test available here that will start a few minutes prior to liftoff. Given the dynamic schedule of development testing, stay tuned to our social media channels for updates as we move toward SpaceX’s fourth high-altitude flight test of Starship!

I could be wrong here but I don't think that's accurate (and a reactor with energetic exhaust would be MORE subject to the Carnot cycle, not less). All hydrocarbon engines -- internal combustion engines, turboprops, turbofans, and so forth -- are fundamentally heat engines. They're using the heat of combustion and converting that into mechanical work using cycles of isothermal and adiabatic expansion and compression (whether reciprocating, as in a piston engine, or linear, as in a turbine). Even a magically frictionless heat engine can only convert a certain maximum amount of the heat into work. Carnot's Theorem says that the maximum possible efficiency of a heat engine is given by: where TC is the ambient temperature and TH is the engine operating temperature. Hydrocarbons burn at 700-1400 K and the ambient operating temperature is usually around 300 K, so the maximum possible thermodynamic efficiency of a hydrocarbon heat engine is 57-79%. Friction and other engineering inefficiencies mean that the most well-designed hydrocarbon heat engines convert only 30-50% of their heat into mechanical work, depending on their internal temperature. Electric motors use electromagnetic fields to convert energy directly into work without requiring a heat cycle, so they can go straight to 90% efficiency or higher. This thermodynamic efficiency is distinct from the subsequent thrust efficiency.

Electric engines are not limited by thermal efficiency. That's the critical distinction here.

Makes me wonder if you could use the unique properties of this microwave plasma to create a highly-efficient vortex to suck in and compress air without needing a fan in the intake. Or if you could spread the microwave plasma jets out across the wings/control surfaces to essentially use the entire body of the plane as an intake. You can't do that with a conventional turbojet because you're limited by the turbine rotation cross-section and bypass fan. It would be MUCH quieter, that's for sure. Not sure if it would achieve enough efficiency to overcome the energy density issues. A lithium-ion battery has an max specific energy of about 1 MJ/kg. Jet fuel has an energy density of 43 MJ/kg, assuming it is burned in air. So electric engines need a 1-2 order of magnitude increase in thrust-specific efficiency to compete with hydrocarbons. It's different with cars, of course, where curb weight is not as limiting an issue. Plus, an internal combustion engine is realistically limited to a Carnot efficiency of 20-35% while electric motors can easily exceed a shaft output efficiency of 90%. Because of their efficiency they're also much smaller and lighter, which offsets the weight of the batteries.

I wonder if a normal, planned re-entry is steeper and thus looks different.....

The question will be power and performance in comparison to conventional engines. We could easily build an "electric jet" now with ducted fans; the problem is how to get power. Batteries are heavy and they don't get lighter when you use them up, like fuel tanks. They claim linear extrapolation would give 8.5 kN with a 310 kW power source like the Tesla Model S battery pack. Where will that get us? Well, let's look at the cruising characteristics of a commuter airliner like the Dornier 328JET, powered by a pair of Pratt & Whitney PW306B turbofan engines. Cruise thrust is 5 kN at 236 m/s with a specific fuel consumption of 69.9 kg/kN/h (takeoff thrust is much higher but we're not worried about that right now). I'm not sure which of the various Model S battery packs they were discussing, but the 100-kWh Model S battery has a power output of 311-451 kW, so let's go with that. The 85-kWh battery has a mass of 540 kg, so a little allowable linear extrapolation would guesstimate the mass of the 100-kWh battery at 636 kg. At a power output of 310 kW, you'd need 1,972 kg worth of battery to get a single hour of continuous 8.5-kN thrust. That's a specific fuel "consumption" of 232 kg/kN/h. So this very rough BOE estimate suggests that such a microwave-air-plasma-powered jet would only have 30% the cruising range of a comparable hydrocarbon-based business jet. Plus, range would be even lower because a jet gets lighter as it burns fuel, but this would not. Also you have to deal with the increased thrust requirements for takeoff. Unless there's some way that the microwave plasma thruster could be VASTLY more efficient than a conventional engine in other, unique ways, there's simply no way to get the kind of energy density we need for electric planes right now.

I mean, that's always what it looks like, right? Usually it just does that over the Indian Ocean or something.

Playing around a little more with Starship legs. Selling points: Legs are fully stowed within the skirt for launch All vertical loads transfer through the current load paths Full three-point self-leveling capability for each leg Significant ground clearance Wide stance Tilt damping with good load path The trick is to have U-shaped feet so that the diagonal member can extend through the gap while stowed and during deployment.

Did RS-25 refurb cost more than new engines? Not for SLS (obviously yes) but during the Shuttle program.

Very possible they are going out staggered. I would just expect SOME sort of direction. Yuzaku says he wants artists and others who will create inspiration, right? So, what is the video supposed to highlight? "Tell us one thing in your life that excites you and why" or "Create a video that shows some of your past work" or "Tell us why this mission inspires you" or just "Introduce yourself and tell us what you do". Any of those would be great. But not having any direction means their submissions are going to be all over the place. I think the distinction is between limitation and direction. If the goal is to vet and compare applicants then you need to arrive at a coherent vision. It doesn't make sense to compare one person who made a video about their past art with another person who made a video introducing themselves to another person who talked about their ideas for the flyby. There shouldn't be any limitations on the output, but giving direction at this stage would just be common sense.

Well this is mildly confusing. A video.....about what?

Considering the shuttle landed the sustainers without the need to throttle this seems a bit odd. I mean, I stand by what I said -- it's still not a very efficient approach. I can't remember whether the refurb costs for the SSMEs was lower than build cost during the program or not, but either way, it's a lot of dead weight. At least if the engine is throttleable then you can use it on orbit and use it for a throttled landing burn. I wonder if they could have used a Zenit booster approach. Spray one side of the tank with ablative paint, give it cold-gas thrusters for control, and then add parachutes with retrorockets. Simple splashdown; tow it back.

Sustainer stages from sea level to orbit are stupid unless you are aiming for reusability and your sustainer engine can throttle low enough to perform a landing burn. Then, at least, you have some justification. But to make that work you have to solve EDL.

It would be easy enough to give SuperDraco a vacuum nozzle, but it would have far too much thrust -- around 93 kN, just slightly less than an RL-10C. Since the Kestrel was retired, SpaceX doesn't have an intermediate engine between the Draco and the SuperDraco. Their notional 10-tonne meth-gox thrusters for Starship will come in at about the same thrust as SuperDraco.

Wet workshop sounds good in theory but it is just very hard to implement in practice. Building living space inside a tank on orbit is...challenging. You can launch with the prefab living space attached to the top, then vent the propellant and slide the living space into the tank, but the living space would need to be inside a fairing anyway and so at that point just launch the whole dang thing as a monolith. One of Von Braun's ideas was to build "floors" into the interior of the terminal stage, with all of the floors being built out of an open grid which would be lightweight and would also allow the propellant to drain through. That is only a slight drop in propellant capacity and at least, gives you a rough sort of internal organization. But then -- what do you even do with the space?

Massive rainstorm here so I wouldn't have been able to see anything, but thanks!

I have a degree in physics, I'm a published scientist, and I will sit here and tell you quite clearly that it is not possible for this to work. There are a lot of things which make intuitive sense that simply don't work.

I did account for LAS mass, actually. I think that the Silverbird coding does a poor job of handling the sustainer stage architecture, generally. It says to enter the vacuum thrust and vacuum specific impulse and it will adjust down, but I don't think it adjusts down properly when dealing with the RS-25. It's fascinating to read through the design history of the orbiter and see how it was just a series of kludges being re-kludged. "This orbiter is too big, which makes our piloted reusable booster too big. What if we put the hydrogen in disposable tanks?" "Wow, putting the hydrogen in disposable tanks makes the orbiter so much smaller and gives us lots of space! But where do we put them? A pair of tanks would occlude the orbiter's wings. What about underneath?" "Hey let's just move the oxygen into the nose of the hydrogen tank and expand the cargo bay." "There's no way for our piloted reusable boosters to have abort modes now because the external tank is in the way. Let's use a pressure-fed liquid booster that splashes down." "Wait we want to use the Saturn V first stage as our booster! And we can even give it wings and fly it back!" "No, that's too much trouble. Let's use a cluster of 7 solid motors and split the external tanks up on either side." "We can use fewer solids if we stretch the external tank, put it back in the middle, and light all the engines at once!" And thus the Shuttle was born.

For a truly minimal ISS mission, you could even use the same core with two SSMEs and just add four of the AJ-60A solid motors used by Atlas V from 2003-2020. That will get you 22 tonnes to the ISS, which is plenty of margin for Orion Lite, especially if it uses its service module for orbital insertion. Liftoff TWR isn't great -- just 1.05 -- but you always have space to throw on a couple more solid motors if you need it. Then you could have been servicing the ISS with Orion Lite for substantially less than the cost of a Shuttle flight, even if operating Orion was exactly the same price as operating and refurbishing the Shuttle (which it would not have been). Well, I think the Silverbird Astronautics performance calculator may overestimate performance somewhat. For example, when I manually input each of the actual values for SLS Block 1, it tells me that I get 34 tonnes (27.5-42.4 tonnes, 95% CI) to TLI, where TLI is defined as 417 x 362600 km at 52 degrees (that would be a staging orbit reachable from the ISS). That's the same orbital parameter I used above. So if we take as gospel NASA's claim that SLS Block 1 can only send 27 tonnes to the moon, then perhaps Jupiter DIRECT 241 (with the DCSS as above rather than the larger six-engine Jupiter Upper Stage proposed by DIRECT v3.0) could only send 23 tonnes to the moon. But, again, that's with NO upgrades from Shuttle hardware. We could have been doing that a decade ago, with plenty of time to develop a proper upper stage. The decision to stretch the SLS core stage rather than work on a better upper stage makes no sense, to me. The Shuttle external tank was already oversized for a sustainer stage; making it bigger just made the problem bigger.

I decided to play around with some of the numbers using Silverbird Astronautics, and...dang. Even with just a single SSME on the bottom of a Space Shuttle external tank, and assuming that the upper and lower tank adapters nearly triple the dry mass of the external tank, this bare-bones "Jupiter 110" configuration could still deliver over 50 tonnes to the ISS...more than enough for Orion Lite and co-manifested cargo. Of course, only having one engine might make for a painfully long burn time so that's an issue. Two SRB, one SSME, no upper stage: ~58 tonnes to ISS Two SRB, two SSME, no upper stage: ~69 tonnes to ISS Two SRB, three SSME, no upper stage: ~73 tonnes to ISS Two SRB, four SSME, DCSS upper stage: ~109 tonnes to ISS Two SRB, four SSME, DCSS upper stage: ~29 tonnes to TLI via ISS staging orbit I guess I can see why opponents of DIRECT said it was overpowered for LEO. I wonder if the gimbal range on the Shuttle SRBs and on the SSMEs would have been enough to allow a single-SRB launch for an even more minimal mission profile to the ISS. With just a single SRB and two SSMEs mounted opposite, you could get over 40 tonnes to the ISS, provided gimbal range allowed it.