Moon mission today - Dragon & Mars hab

kbd512 · 2016-12-10 17:14:34

GW,

I'm of the opinion that capsules make poor landers in thin to nonexistent atmospheres, but Orion consumes all development funding not allocated to SLS and something is still better than nothing. Red Dragon can be made to work, but the compromises are severe and the benefits limited to merely providing a capability we wouldn't otherwise have for contrived reasons. When I think of a Mars lander for humans, I think of something with lift fans that can do more than land and take off once per mission or something that's so small, light, and simple (inexpensive / disposable) that retro-propulsion is not required.

Assuming NASA is willing to either use an unpressurized capsule or accept a very small and functionally limited pressurized capsule, a Red Dragon variant could feasibly deliver four people to the surface of the moon or Mars. As you've already indicated, Red Dragon would require a complete redesign to provide both descent and ascent capabilities for lunar missions. I find little merit in that undertaking unless we're dead set on building a lunar base. If Red Dragon could be refueled on the surface using a second capsule, then the problem becomes more manageable.

Apart from the battery technology, what other relevant technical challenges would there be in designing an electric lift fan-powered lander? MIRAGE is about the same vehicle mass minimally required to deliver two people to the surface and it was designed to take off and land vertically at gross weight using its 45kW electric lift fan.

Let's look at a table of mass figures from the MIRAGE design:

Gross with Fuel and Scientific Payload: 490kg; 186kg on Mars
Empty Vehicle (8090-T3): 283kg
Empty Vehicle (Composite): 199kg
LOX: 256kg
LH2: 31kg

Graphene Polymer batteries can store 1kWh/kg. Let's presume that a controlled descent would require 15 minutes of total flight time (11.25kWh), but for utility purposes let's say we want a vehicle capable of two hours (90kWh) of hover to explore the area around the Mars Surface Habitat. A 90kg battery corresponds with a 200kg payload, in a vehicle with a similar structural mass as MIRAGE, and that is roughly equivalent to what two average men would weigh wearing MCP suits. A pair of bungee suspended fabric seats and flight controls would certainly weigh no more than what is already allocated for fuel storage and scientific instruments in MIRAGE.

The point I'm trying to make here is that every other Mars lander design is basically a single use vehicle that provides no additional utility to the astronauts if it's not a full surface stay duration habitat sitting packaged with a substantial quantity of fuel to make a propulsive landing.

What if we skipped separate crew and cargo lander designs altogether and simply landed our astronauts in a habitat module attached to an electric lift fan so that the entire habitat could swiftly travel from location to location?

Are there any reasons why the deep space habitat, surface habitat, and crewed portion of the ascent vehicle couldn't be the same physical piece of hardware?

The electric motor power requirements for such a vehicle to hover works out to roughly 100kWe/t, or 1MWe for a 10t vehicle that contains enough food to feed two astronauts for 500 days. If the vehicle is designed to fly for a half hour per flight, that translates to a 500kg battery.

Hmm...

If the astronauts didn't recycle anything at all on the surface of Mars, a 500 day nominal surface stay translates into 5,030kg worth of consumables. We would certainly obtain oxygen from Mars using MOXIE, so some significant portion of that 840kg of oxygen required would be provided by that system. Although still experimental, it's a safe bet that some portion of the 2,420kg of water required would be baked out of the Martian regolith.

1,770kg - food
2,420kg - water
840kg - oxygen
500kg - graphene polymer flight batteries
200kg - 4 Siemens 260kW electric aircraft motors

That's 5,730kg for the consumables, flight batteries, and electric motors, leaving 4,270kg for the pressure vessel, flight-specific structures, power plant, and scientific instruments. An actual LENR power plant was independently measured as producing 20kWt/kg, so 50kg for 100kWe. We'd have to run a CO2 coolant pump to cool the reactors. Those figures apply to first generation LENR technology. The second generation units may be capable of producing 1MWe on their own without the requirement for a battery, but either way you have to dissipate the heat that the reaction produces.

In any event, flying habitat modules make the most sense to me for the Mars missions. After ADEPT slows the habitat vehicles to Mach 2 or thereabouts, you separate the heat shield and turn on the electric lift fans to soft land. Surface exploration involves flying to whatever sites look interesting. Return involves flying the habitat modules back to the LOX/LCH4 ascent stages, mating, ascent to orbit, and then mating with Earth return stages for return to Earth. Detachable heat shields have been used by every Mars mission to date, so nothing new there. Habitat modules capable of atmospheric flight shouldn't be overly complicated since the Martian atmosphere is so thin. Robotic propulsively landed ascent stages are required, but SpaceX has already proven that that technology works. The only really new technology required is the use of LENR in long duration power production for the LOX/LCH4 plant and to recharge the flight batteries in the habitat module. No radioactive materials are involved, nor any delicate and expensive solar panels.

SpaceNut · 2016-12-10 21:34:46

Thanks for the food for thought and response to use of Red Dragon for both moon or mars....
Back to the drawing board.....

RobertDyck · 2016-12-11 11:39:41

I have to point out my initial post.

Mars Direct ERV will not work on the Moon. The reason is Moon does not have a CO2 atmosphere, so ISPP will not work. NASA initially tried to design Apollo to direct launch the CSM from surface of Earth to surface of Moon and back, but found that even a Saturn V was not large enough to launch it. Why would you think a Mars Direct ERV could do the job when Apollo couldn't?
SLS block 1 can launch an Apollo style mission to the Moon. Several catches:
- Dragon capsule is much lighter than Apollo capsule
- LCH4/LOX instead of hypergolics for the service module to reduce mass
- carbon fibre composite propellant tanks instead of aluminum alloy
- Soviet LK architecture, again to reduce mass
- the stage used for LOI and to de-orbit the lunar module will also use LCH4/LOX and carbon fibre composite
- the lunar module will carry all 4 astronauts, but no rover, no science instruments, and only enough life support to go from lunar orbit to surface and back, as well as one pressure cycle. All time on the surface, life support will be spacesuits only.
- the lunar module will have one propulsion stage only, just like the Soviet LK. That propulsion stage will also use LCH4/LOX, and carbon fibre composite propellant tanks.
This is not intended to be used on its own. It's a taxi for astronauts. It must be accompanied by a Mars Direct habitat on the Moon. That habitat will have the lunar rover and science instruments. It will also have recycling life support capable of multiple months, or a couple years. It will be a lunar base. The habitat will require SLS block 2B.
Mars Direct habitat has an upper floor alone with as much floor area as a 60-foot class A motor home with slide-outs. Such a motor home is so large that in most states it's categorized as a mobile home, not an RV. The lower floor of a real Mars Direct habitat will have an air lock, storage compartment for the rover and surface science instruments, life support, landing rockets, propellant tanks for landing rockets, RCS thrusters, and propellant tanks for RCS thrusters. The storage compartment will also hold the inflatable greenhouse. During transit that storage compartment will be packed full of stuff, but once on the Moon (or Mars) that storage compartment will be empty, available for use as a laboratory or workshop. But the storage compartment will only be as large as a single car garage.
The inflatable greenhouse, once set up, will be the same width as a double car garage, and twice the length of a double car garage, so twice the floor area of a double car garage. Yes, twice double.
This does require launching a Dragon capsule on SLS. If SpaceX doesn't want to do that, we can do it with CST-100 Starliner. Boeing builds CST-100 and are involved with SLS, so they should be willing to do so. CST-100 Starliner is heavier than Dragon, but lighter than Orion. You can't do this with CST-100 on SLS block 1, but could with SLS block 1B. The only difference is replacing the Interim Cryogenic Upper Stage with the Exploration Upper Stage. But CST-100 would also need a new service module, also with LCH4/LOX and also with carbon fibre composite propellant tanks.
You can't do this with Orion. SLS block 1B is barely capable of launching Orion alone into lunar orbit and back. No room/launch mass for any sort of lunar module.
Constellation required 2 launches for each lunar mission: Ares V and Ares 1. You could do it with SLS block 2 and SLS block 1. But that would be all expendable equipment. And none of that equipment can be used for a Mars mission. My mission architecture is strongly based on Mars Direct; it required SLS block 2B to launch a Mars Direct habitat. Once established, that habitat is a permanent Moon base. Not only does it prepare for Mars by using Mars equipment on the Moon, but it establishes a permanent base on the Moon. Further missions can be done to that same base with a single launch of SLS block 1 only. Or SLS block 1B with CST-100.

Last edited by RobertDyck (2016-12-12 09:38:45)

Tom Kalbfus · 2016-12-11 12:04:58

Orion is what we got now, perhaps it can be lifted by a falcon 9 rocket, then the falcon can land, get more fuel and launch again, add fuel to the Orion or a rocket stage so it can make a TLI with lander.

kbd512 · 2016-12-11 15:20:11

Rob,

What I proposed is basically Mars Semi-Direct, as it requires three Falcon Heavy rockets for each pair of astronauts. For each additional pair of astronauts, add three more Falcon Heavy launches.

Opportunity #1:
Launch #1: Earth Return Stage (NTO / MMH; aerocapture at Mars and Earth)
Launch #2: Mars Ascent Stage (NTO / MMH; aerobraking reentry; rocket landing)

Opportunity #2:
Launch #3: Mars Deep Space / Surface Habitat (tethers to FH upper stage for AG; aerobraking reentry; lift fan landing)

Consumables should be split between the pieces of this mission architecture. The habitat contains consumables to get to Mars and half of the consumables for the surface stay. The ascent stage contains the other half. The Earth return stage contains consumables for Earth return. If you can't reach LMO, then you don't need food or water for a return to Earth. The habitat is lighter at descent to Mars and significantly lighter at ascent to LMO.

We need to figure out some scheme to roughly equalize the masses of all components in the architecture so there are not disparate delivery solution requirements.

RobertDyck · 2016-12-11 17:00:45

Tom Kalbfus wrote:

Orion is what we got now, perhaps it can be lifted by a falcon 9 rocket, then the falcon can land, get more fuel and launch again, add fuel to the Orion or a rocket stage so it can make a TLI with lander.

Falcon 9 can lift 22,800kg to LEO. That means 185km altitude @ 28.5° inclination. Orion total mass is 25,848kg not including launch abort system. That mass is just the capsule plus service module. I don't think that total mass even includes the fairing around the service module. So no, you can't launch Orion on Falcon 9.

Orion was launched by Delta IV Heavy into LEO. Notice that's Delta IV Heavy, even Atlas V medium with the maximum number of solid rockets isn't big enough. It could be launched by Atlas V Heavy, but that configuration has never been built, and United Launch Alliance (a 50/50 joint venture of Boeing and Lockheed-Martin) has no plans to build Atlas V Heavy. Current plans are to launch Orion into Earth orbit (any altitude) on SLS block 1. Exploration Mission 1 will launch Orion on a fly-by around the Moon and back on SLS block 1. To launch Orion into Lunar orbit requires SLS block 1B.

SpaceNut · 2016-12-12 22:07:34

Thanks for the numbers on Orion and Falcon 9 launch compatibility which its not possible to do...Originally it was to launch on the advance SRB but due to being so heavy and lots of vibrations that would not be much better to use as a taxi launch system and while the Delta IV is capable its not enought to make a lunar mission possible without lots of add ons in orbit.
lets stick to the Space x products and the Orbital ATK to create any chance for a mission to the moon without Nasa....all that we seem to need besides the earth departure stage, service module, lunar lander is a lunar departure unit to return the team home. So what do we do for the lunar lander?

GW Johnson · 2016-12-13 11:31:39

Answering kdb512 in post 76 above, if all things were otherwise equal, I would agree with you that an entry capsule is a poor way to deliver payload to an airless body (Mars is not airless).

All things are not otherwise equal. NASA is hogtied with a congressionally-mandated rocket and capsule that are so big, heavy, and expensive, as to eventually prove fairly useless. It means one has to use whatever is actually available, whether it is optimal or not.

Mars has too thin an atmosphere for large objects to have useful terminal velocities hanging on a parachute. But it's quite enough to provide hypersonic aero-deceleration. One comes out of that at about Mach 3 speeds (on Mars, near 0.7 km/s), and lower altitudes the heavier you are. You need some propellant margin for maneuver and terminal precision, maybe a delta vee near 1+ km/s, for a retropropulsive touchdown from end-of-hypersonics.

I have no good weight statements for any of Spacex's Dragon configurations. The best I have is nothing but guesses. But I do show gross delta-vee near 1 km/s for about 2000 kg of payload in my best guess for a Red Dragon. I show about 3/4 km/s at 4000 kg payload, and 1.2 km/s at 1000 kg. What I calculate is not correct, but it's not very far from what Spacex hints at.

Because of the hypersonic aero-deceleration, the retropropulsion delta-vee is about the same, whether you enter from orbit (at 3.7 km/s), or from direct entry (above 5 km/s).

Things are not the same at the airless moon, where all the deceleration must be propulsive. Lunar orbit velocity is 1.7 km/s, escape speed 2.4 km/s. Delta-vees like that are way out-of-reach for any conceivable version of Dragon using the common inner pressure shell all versions have shared so far. It's similar but even worse at Mercury.

But for the outer gas giant moons, versions of Dragon could easily serve. Seems like I saw some estimates of delta vees being bandied about that were at or under 0.4 km/s. So Dragon could serve at Mars, Jupiter/moons, and Saturn/moons, at least. For one-way landings.

We more-or-less have them available already, so we ought to think about using them where they can serve.

GW

Last edited by GW Johnson (2016-12-13 11:36:29)

RobertDyck · 2016-12-13 20:57:13

What would happen if you build a modified Falcon Heavy? Core stage with 9 Raptor engines instead of Merlin 1D. Of course that means LCH4/LOX. Side boosters would still be Falcon 9 core stages with Merlin 1D engines and RP1/LOX. And since the core stage would not be the same as boosters, that means no cross-feed. Upper stage would also have to be modified, it would use a single Raptor engine, modified for vacuum. That means a bell extension similar the Merlin vacuum engine currently used. And of course that means the upper stage will also use LCH4/LOX. Would that increase lift capacity sufficiently to throw the stack I described for the Moon?

Last edited by RobertDyck (2016-12-14 06:14:33)

Oldfart1939 · 2016-12-13 22:51:21

Robert; here's a repost from the Interplanetary Transportation thread:

Oldfart1939 wrote:

If a standard engineering approach is taken, then a switch to liquefied methane as the fuel will allow the new Raptor engines to replace the Merlins currently using RP-1 with a significant improvement in thrust generated with a higher ISP. The thus improved Falcon Heavy should be able to throw more and bigger chunks of hardware skyward. Maybe this is what Musk has in mind for his tentative 2018 Red Dragon to Mars launch? What about a Falcon Super Heavy with 4 detachable booster stages instead of 2? Maybe a bigger core stage with 4 Falcon 9 size boosters all using Methylox propulsion?

Last edited by Oldfart1939 (2016-12-13 22:53:37)

RobertDyck · 2016-12-14 06:14:48

Ok, now the hard work. Exactly how much could it lift?

Does anyone know velocity at the time of second stage separation? And how much propellant is reserved for fly-back? For a Dragon launch.

Last edited by RobertDyck (2016-12-14 06:18:20)

kbd512 · 2016-12-14 10:55:28

GW,

If you know of a better way to land a 10t payload using less than 1t of propulsion hardware, I'd love to know about it. Even if the battery is single-use, lift fans surpass all other alternatives that I know of on delivered tonnage. LENR and graphene polymer batteries are real technologies that will be ready for prime time in the next couple of years. Electric lift fans and electric aircraft motors are ready right now. At this point, electric lift fans are an aerospace engineering solution waiting to happen. Mars has an atmosphere that lift fans can produce thrust within, so now that our power production and storage technologies make electric lift fans practical there's little reason not to use them.

We're talking about mounting four 350hp electric aircraft engines to the habitat module. This is real hardware that Siemens makes and no more experimental than RS-25's at this point. Tadiran makes primary batteries with a .71kWh/kg. There is no need to think / believe that this is anything more sophisticated than a scaled-up quad-copter protected by a heat shield during reentry.

GW Johnson · 2016-12-14 12:01:37

What is needed to verify whether a lift fan is even feasible in "air" that thin is a test. Your quadcopter idea needs the same weight/area loading as it will have to support on Mars, which means 38% of the Earth weight for the system. It has to support the reduced weight in Earth's atmosphere at an altitude near 110,000 feet to have the same density and pressure as Mars surface, and higher still to simulate flight above Mars's surface. That is where they tested the supersonic ribbon chutes used at Mars on Viking, with a rocket-propelled item moving about Mach 2.5 (opening speed).

This quadcopter thing is not that expensive a test to run at subscale to demonstrate feasibility. You use a balloon to take a small one up to 150,000 feet or more, cut it loose, and attempt stable flight at selected steps of altitude on the way down. It does not have to "look like" any particular Mars payload, it just needs the correct weight/area loading. If it demonstrates effectiveness at power levels and blade speeds that are reasonable, then we know it can be scaled up and used.

Honestly, I have my doubts about fans, propellers, and wings as useful items for Mars, because at "ordinary" speeds, that "air" is first cousin to the vacuum of space. At near-orbital speeds, there is enough friction to slow you down handily, and enough shocked density to generate large forces. But at "ordinary" speeds, these forces are but a whisper. Surface density is about 0.7% that of Earth normal, while weight is 38%. That ratio is off from Earth by a factor just over 50:1. Very hard to design around that.

Yet, the test should be run. Fans and wings don't have the same constraints on weight loading-to-open that parachutes do. This would be a very interesting concept to apply at Mars, if it can be made to work. I doubt you can deploy these things supersonically, though. Some retropropulsion is likely required to get you down subsonic before you attempt a fan-supported landing. Just an old-timer's hunch.

GW

elderflower · 2016-12-14 12:58:39

My examination of a helicopter for use as crew/ stuff transport was enough to convince me that it isn't going to work on Mars using present materials for blades and structures, and subsonic operation.
If you can devise a way of making lift fans effective I should be very interested. otherwise we are going to need a rocket powered hopper.

kbd512 · 2016-12-14 18:56:01

elderflower wrote:

My examination of a helicopter for use as crew/ stuff transport was enough to convince me that it isn't going to work on Mars using present materials for blades and structures, and subsonic operation.
If you can devise a way of making lift fans effective I should be very interested. otherwise we are going to need a rocket powered hopper.

Elderflower,

A lift fan is not the same type of device as a propeller or rotor. It's more akin to a turbofan without combustion. We have to spin the turbine blades faster on Mars, just as we have to spin propeller or rotor blades faster on Mars, but the rotational velocity is well within the limitations of aerospace alloys or composites and electric motors. There is nothing technologically infeasible about the motors, electrical power supply (even if it's just a single use primary battery limited to 15 minutes of flight time), nor the control systems required.

For 40kWe input, an electric lift fan .9m in diameter can move 36.45kg/s when it's rotating at 50,000rpm and produce 174kgf at a .013kg/m^3 atmospheric density. I want to scale up the electric lift fan system MIRAGE was designed to use by a factor of 6.25 per fan.

A notional 10t habitat weighs 3.8t on Mars. If the habitat is only loaded with half the consumables for the 500 day surface stay, it would weigh near to 6t (a base heavy inflatable BEAM-derived habitat for two crew members suspended below the lift device and above fixed landing gear attached to ADEPT). This means the thrust of each fan has to be slightly more than tripled. We're increasing the electric motor power input by a factor of 6.25 per fan and using 4 fans, so input power should be sufficient for stable flight, with performance margin, given a lift requirement of 570kg per fan.

A Tadiran 710Wh/kg primary battery would weigh 352kg for 15 minutes of flight. The electric motors add 200kg. The associated structures add 100kg to 150kg. If there's a supersonic parachute + retro-rockets solution that weighs 700kg or less, you have my attention. Please share with the rest of us.

Even if graphene polymer batteries and LENR never materialize, which will certainly not be the case for the batteries, we come out far ahead of all other competing technologies on delivered tonnage as a function of propulsion system mass. The numbers aren't even close. The technology also scales up or down by adding lift fans and batteries. Double the battery and lift fan mass and you double the propulsive force for delivery of a 20t habitat for 4 crew members.

To recap, lift fans are not rotors or propellers and lift fans do not have the rigid Mach number limitations of propellers and rotor blades. The power storage and electric motor performance requirements are quite reasonable. This is much closer to economic and technological feasibility than LOX/LCH4 plant, which is not required for exploration missions when electric lift fans and NTO / MMH rockets are available.

SpaceNut · 2016-12-14 20:23:41

Blade lift is a function of atmospheric density, surface area of the blade and the RPM to which it moves, more bleeds with smaller surface area are possible for the same RPM to produce the same lift capability. It comes back to lift value to mass which occurs for what amount of available power which is consumed by running it.

kbd512 · 2016-12-14 21:21:13

GW,

There's no combustion process to initiate using electric motors, so I don't understand why they couldn't start in a supersonic flow. My surmise is that ADEPT brings our pair of explorers in from interplanetary space, slows them to Mach 2, the crew starts the lift fans, ADEPT is forcibly separated as the engines spool up, and the plunging descent is arrested within the first few kilometers of downward travel. The crew has a couple minutes of powered flight to visually locate their ascent stage using external high resolution cameras and radio beacons and to pick out a suitable landing spot near their ascent stage.

The ascent stage could also be landed using the same battery and more electric motors. The ascent stage won't move until the astronauts ascend to orbit, so its only EDL task is to successfully land on a relatively flat surface that has been pre-selected using satellite imagery.

Half way through the surface stay, the ascent stage lift fan hardware detaches and delivers the second batch of consumables to the habitat module. There are three batches of consumables split between the habitat module, ascent stage, and Earth return stage.

Batch 1 - outbound transit (in habitat module)
Batch 2 - Initial surface stay increment (in habitat module storage container)
Batch 3 - Final surface stay increment (in ascent stage storage container)
Batch 4 - inbound transit (in Earth return stage storage container)

The crew will use the habitat lift fans or ascent stage lift fans to attach their habitat to the ascent stage prior to departure. Prior to liftoff all unnecessary weight (mostly consumables), including lift fan hardware, is removed from the ascent stage and habitat module. The habitat module mass will be between less than 3000kg at liftoff. The ascent stage dV increment is something like 4.4km/s or thereabouts- I know it's 4.1km/s to LMO, but I can't recall what the dV increment is for the circularization burn. 4 pump-fed AJ-10-190's with extended nozzles can provide that dV increment. The ascent stage is just a modernized reprise of MIT's Scott Alan Geels Mars Ascent Vehicle using detachable EDL hardware.

At any rate, the Earth return stage is mated to the habitat module, consumables are transferred, and the vehicle is reconfigured for Earth reentry by attaching another ADEPT to the landing gear and lift fan to the top of the habitat module. The Earth return stage also uses pump fed AJ-10-190's and additional dV capability is used to slow the habitat module prior to Earth reentry using ADEPT and electric lift fans.

The idea here is to use as much of the same hardware as is feasible and to split the delivered payloads into roughly equal masses for Falcon Heavy or Vulcan Heavy to deliver. I could be wrong, but I think I've achieved that. 3 Falcon Heavy or Vulcan Heavy launches per pair of astronauts, mission crew head count limited by launch cadence rather than allocated funding, and a total mission hardware package cost somewhere in the $500M range per pair of astronauts. It's expensive, but workable using the sort of things we already have or can easily develop.

elderflower · 2016-12-15 05:00:02

Do you know of any demonstration of the kind of lift required at the Mach nos and densities you are considering, using electrically powered fans?

GW Johnson · 2016-12-15 16:29:06

I think you'll find that the fan tips must stay under the local speed of sound inside the fan shroud. This is true in all aircraft gas turbine fan/compressor assemblies. The oncoming stream is decelerated ahead of, and inside, the inlet to make that component about 0.3 Mach at the fan/compressor face.

The tip speed must be low enough so that the vector combination is about 0.95 Mach max, for a noisy, short life unit, and under 0.85 Mach for a quiet, long-life unit. The range of feasible pressure rise ratios for one fan/compressor stage fall in the 1.05 to 1.15 range, lower for fans, higher for axial compressors. Overall pressure rise as a difference above atmospheric, times total disk area, is one measure of thrust appropriate to a lift fan. It would be numerically equal to fan stream massflow times the appropriate fan ejection velocity.

Supersonic propellers, fan blades, and compressor blading have been tried many times over about the last 70 years. They have never been successful.

GW

Last edited by GW Johnson (2016-12-15 16:31:39)

kbd512 · 2016-12-15 18:18:28

GW,

The Tu-95 has supersonic propellers. More than half a century after the Tu-95's went into service, there still seem to be a number of them flying around today. The fan tips of modern jet engines used in airliners are also, briefly, mildly supersonic. If the blade tips are thin enough, this shouldn't be a show stopper. Wouldn't Cd be substantially reduced in the very low atmospheric pressures on Mars?

Are the low airspeeds a function of what works best for combustion or what is required to reduce drag on the fan blades?

By the very nature of the type of flying we're doing, the lift fans are started facing away from the oncoming supersonic flow. The habitat module is moving rear (heavy) end first with the faces of the lift fans oriented slightly upward and opposite of the direction of travel. Is there any way to model the flow immediately behind ADEPT since the heat shield is facing into a supersonic flow and, presumably, creating a fairly substantial pressure wave ahead of the stack?

Is there any reason why fans not used in a combustion process can't suck air in a supersonic flow with the compressor faces facing away from the oncoming flow?

Oldfart1939 · 2016-12-15 18:59:19

Robert Zubrin's book "Entering Space" had some examples of rotary wing type rockets mentioned in one of the earlier chapters. I don't have the book handy now, as I loaned it to a friend. He concluded that it wasn't going to work.

GW Johnson · 2016-12-16 10:52:15

Well, the propeller tips (tips only!!) go slightly supersonic on a Tu-95 when throttled-up to full power at max rpm. That is why the craft is so incredibly loud that it can be heard passing overhead from a submerged submarine, if flying at max speed. That is energy diverted to waste noise instead of propulsive power, by the way. You'll note that no other propeller-driven craft designed after that time deliberately spins net vector tip speeds above Mach 1. Accidentally perhaps, but not deliberately.

What you are talking about with a lift fan deployed at the end of entry in supersonic flight is quite different. All of the oncoming airflow would be supersonic. Even at the hub, where the tip speed due to rotation is near zero, the axial component would be very significantly supersonic.

We do very mild supersonic flow on turbine blades, but we restrict it to mild, because the shock wave energy losses get truly enormous. Most of that is taken on the fixed stator vanes behind each turbine stage. At least the pressure gradient is favorable, so there's not so much flow separation loss. We do not do that supersonic stuff with compressors at all, because the pressure gradient is adverse, adding pressure-induced flow separation to shock-induced flow separation as horrendous loss mechanisms. Plus, if the tips do go mildly supersonic, you have a much noisier engine.

Drag coefficients in general, including those of propeller blade airfoil sections, show a very strong transonic drag rise. They roughly double (or more) near Mach 0.9 through 1.1 or so. Lift does not rise (sometimes even drops), so L/D (and propulsive efficiency as a propeller) drop sharply in that transonic condition.

And no, drag coefficient CD does not reduce in thin density conditions, density figures into dynamic pressure. CD (and CL) is a function only of shape, Mach, and attitude angles. Dynamic pressure is 0.5 density velocity-squared. That's the only place density appears, but it affects everything from there, except weight.

Your propeller or fan blade drag and lift are coefficients times dynamic pressure time a reference area. Your thrust depends upon a component of that lift. The other lift component, and both components of drag, contribute to the torque required to drive the propeller. Make the blade L/D low, and you just reduced thrust while greatly increasing the power required to obtain that thrust. Rotational speed at cruise is usually reduced below takeoff max, just to get a more efficient combination in cruise.

And THAT is why no one today uses vector-addition tip velocities that exceed about 0.9 Mach in new propeller designs. Much less all-supersonic oncoming flow. Propeller, compressor, doesn't matter (only the details are different). Has nothing to do with combustion in the burner cans of a gas turbine. Has everything to do with the aerodynamics of rotating wings.

As for the "rotary rocket" devices, I think the jury is still out on that one, since none have flown. But most of us think its a long shot, except for subsonic flight. There are exceptions: a lot of the rotary rocket crowd were the founders of XCOR. Some of them would like to see the rotary rocket attempted.

GW

Last edited by GW Johnson (2016-12-16 10:59:46)

kbd512 · 2016-12-16 15:54:31

GW,

Is there any way to use the supersonic centrifugal CO2 compressor technology that Dresser-Rand and Ramgen are working on for waste heat recovery and apply it to aviation? The test facility is in Olean, NY. It's a DoE-funded project, so I don't see why NASA couldn't get their hands on the test data.

It's worth noting that we're not trying to obtain an efficient cruise here and the minimum design life is measured in minutes. All we need to know is whether or not we can generate enough mass flow with a 1MWe input power.

GW Johnson · 2016-12-17 19:09:10

kbd512:

I looked and found a few references and 3 patents for a "supersonic centrifugal compressor" on-line. All these documents related to supersonic flow speeds from the impeller into the diffuser. Some of them dealt with sonic to transonic speeds in the inlet and inducer, which is more like what you are contemplating.

All of them seemed to indicate such a thing (supersonic impeller-into-diffuser velocities) was possible. All of them seemed to indicate industry has found such designs very undesirable because of a very narrow range of conditions in which such compressors work at all, which severely restricts applications. Surge instability seems to be one limit.

I'm not at all sure how a centrifugal device might serve to produce lift forces, other than as direct jet thrust out of some nozzle. So I'm not sure how supersonic centrifugal technology might help with your lift fan concept. That seems more closely aligned with axial compressor technology, and those folks avoid supersonic flow like the plague.

GW

kbd512 · 2016-12-17 21:42:15

GW,

Can a centrifugal supersonic compressor be combined with an axial compressor in some rough facsimile of the stage layout of the Lycoming T53 (axial / centrifugal / axial), except with the first stage or stages being a supersonic centrifugal compressor to first compress the gas to something more usable (subsonic and higher pressure), followed by an axial compressor / lift fan (centrifugal / centrifugal / axial)?

Ramgen Power Systems for Military Engine Applications

A Numerical Study for Flow Excitation and Performance of Rampressor Inlet considering Rotor Motion

Ramgen Power Systems Workshop on Future Large CO2 Compression Systems

Ramgen Supersonic Shock Wave Compression Technology

Ramgen Supersonic Shock Wave Compression and Engine Technology

Does this stop working efficiently when the oncoming flow is subsonic?

Alternatively, would it be better from a mass and complexity standpoint to simply devise an inlet that slows the oncoming flow to subsonic speeds and then feeds it to a multi-stage axial compressor / lift fan?

All I want to know is whether or not it is feasible to suck in supersonic and subsonic CO2 and either sufficiently compress it to produce aerodynamic lift by combining centrifugal compression with axial lift fans or simply expanding the CO2 in an expansion nozzle to produce thrust from high velocity, compressed CO2.

Which system would or should have the lowest mass and hp requirements?

Can we do this with 1MWe or less?

Last edited by kbd512 (2016-12-17 21:43:37)

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