Titan Sample Return Concept

[RedKraken]

heh....first thing i did was go looking for ramjet engine configs in RO....couldn't find any.....but i have found forum references by well-known denizens. 

so we are looking for an N2-augmented rocket for titan.  

Edited October 24, 2016 by RedKraken

[KerikBalm]

24 minutes ago, RedKraken said:

Ok.... in RO 1.1.3 with my ITS lander, i only used 4000m/s to ascend from titans surface to 1000x1000km.

?? my dv map says 7600 ??

is this a bad estimate? or is RO out of wack?

Info about the map here https://www.reddit.com/r/space/comments/1ktjfi/deltav_map_of_the_solar_system/

MJ reads 1.35m/s2 gravity, 1.55 atmos, density 5.5 kg/m3 at surface.... all good.

3 SL raptors all the way....didnt turn on the vacs

Surface start at 150t/2100t dry/fueled no cargo. twr 3.2, 8100m/s (sl) 8730m/s (vac)

orbital vel is 1570m/s so i used about 2430m/s in losses.

It was a rough gravity turn (40%) starting a 10km ending at 250km.

Straight up to 1000km right angle turn might be more efficient here. Almost kerbal.

I had to burn an ap at 1500km cause titan was clawing it down, even out at 600km. 600km is low orbit.

i had mechjeb limit my q to 29kpa, so it was slow going off the surface 110 - 150 m/s in the first 20km.

Well, I was looking at this dV map:

Orbital height would have to be a little higher than Ganymede due to the atmosphere despite the solid surface radius being a bit smaller (2575 km vs 2634km).

Titan is almost the same size as ganymede.

In terms of mass, its less massive: 1.3452 vs 1.4819  (x10^23 kg) So its orbital velocity will be lower. If Ganymede takes ~2km/sec to get to orbit, thats more than what it would take to get to orbit on titan assuming no air resistance... but 5.6 km/s lost to air resistance?! that seems ridiculous. To get to Earth orbit, 7.8 km/s, one uses ~9.0 km/s... so 1.2 km/s to increase altitude by a couple hundred km + counter gravity drag + counter aerodynamic drag.

Titan's gravity drag is going to be nearly an order of magnitude less. Its atmosphere is only ~50% denser, and normally most losses are due to gravity drag instead of air drag anyway... 7.6 km/s seems absurd to me.

[RedKraken]

1 minute ago, KerikBalm said:

7.6 km/s seems absurd to me

agreed

[insert_name]

15 hours ago, RedKraken said:

heh....first thing i did was go looking for ramjet engine configs in RO....couldn't find any.....but i have found forum references by well-known denizens. 

so we are looking for an N2-augmented scram-jet for titan.  

nitrogen is one of the most inert gases out there that isn't a noble gas. good luck getting it to burn with something, much less convince someone to do this as a mission critical component for a multi billion dollar mission.

[RedKraken]

N2 will not burn and it doesnt need to for this engine.

The idea is to use N2 for extra mass...not combustion. We already have an oxidizer and a fuel.

Perhaps u could read back just a few posts. KerikBalm comments on ramjets and scramjets.

[insert_name]

10 minutes ago, RedKraken said:

N2 will not burn and it doesnt need to for this engine.

The idea is to use N2 for extra mass...not combustion. We already have an oxidizer and a fuel.

Perhaps u could read back just a few posts. KerikBalm comments on ramjets and scramjets.

 Pretty sure increasing the reaction mass tends to decrease exhaust velocity, decreasing ISP. Also scramjets are pointless if you are carrying your fuel and oxdizer with you

[RedKraken]

ok thanks for your comment.

Any engine enthusiasts want to comment on N2-augmented rockets for titan? 

Edited October 24, 2016 by RedKraken

[KerikBalm]

9 hours ago, insert_name said:

 Pretty sure increasing the reaction mass tends to decrease exhaust velocity, decreasing ISP. Also scramjets are pointless if you are carrying your fuel and oxdizer with you

Yes, it decreases real exhaust velocity. No its not pointless to do this

You know what else has really really poor exhaust velocity? The air from a propellor. The bypass air going through the fan of a turbofan... even the air coming out the back of a ramjet.... being 80% non-combustion products.

You really think a turbofan with ~10,000 effective Isp is pushing air out the back at ~100,000 m/s? If it were, then it would be one heck of a space drive if we just supply it with some Oxygen for something like 3,000 Isp (given the fuel to O2 ratios)

You can be limited by reaction mass, energy, or both.

Consider an ion drive (solar or nuclear powered): You aren't so limited by energy (rather energy/time), but reaction mass is limited so you try and expell your reaction mass with the most amount of energy possible to reduce use of reaction mass. The propellant (usually xenon) and energy source (solar or nuclear) are different things.

A chemical rocket: the energy supply is also the reaction mass supply. You want ot chose the reactants that have the most energy per unit mass. So that the exhaust velocity is highest

A jet/prop engine/air augmented rocket: Reaction mass is essentially unlimited, you're limited by energy. A normal jet engine also doesn't have to carry oxidizer... but this isn't the whole story. If you compared a (essentially)Kerosene burning J-57 on the SR-71 getting 2000+ Isp to a kerosene burning F-1 on the Saturn V getting 304 vac Isp... you'll notice a range that is too large to be explained by "free oxidizer". The oxidizer to fuel ratio is only 2.56:1 so lets look at the Isp you'd expect if the oxidizer just magically appeared from nowhere in the combustion chamber: 304 * (1+2.56) = 1082... hmmmm.... about half of what the J-57 turboramjet is getting burning essentially the same fuel? what is going on?

Its the N2! Its the kinetic energy equation KE=1/2 mv^2

For the energy it takes to expel 1 kg of propellant at 4,000 m/s, you could expel 4 kg or propellant at 2,000 m/s. Half the velocity takes 1/4 the energy. 4x the propellant at half the velocity is twice the momentum, and hence twice the thrust for the saem amount of fuel. ***

For the same amount of energy, you can also expel 16kg of propellant at 1,000 m/s, for 4x the thrust.

The key here is that the N2 is free propellant/reaction mass that you pick up and accelerate as you need it. Its not like the stored fuel that you have to accelerate with the rocket. You can use all the N2 you want as reaction mass and as long as it gives a thrust augmentation, it improves effective Isp.

You are using the thermal energy of combustion to expand the air and push it out the back. This is essentially the same concept as using electgric energy from combustion in a piston engine to turn a propeller and push air backward to propel a plane forwards.

When you are scooping up your reaction mass from the atmosphere, you don't care about exhaust velocity like you do for a rocket in the vacuum

*** Incidentally, if we assume the J-57 is consuming nearly all the oxygen that passes through it, we're pretty close to the 4x reaction mass, 1/2 exhaust velocity, double effective Isp case. In this case 1/4 of the reaction mass are the fuel's reaction produced, and 3/4 is inert gas.   25% and 75%. This of course assumes 100% efficiency in the heating of the inert gas for thermal expansion and production of thrust

In the atmosphere... 20% is O2 which will be burned, 80% inert N2 and Argon. O2 has a MW of 32, N2 is just 28... so the mass ratio is of course a bit different from the molar ratio... by mass its more like 22% O2, and 78% other gasses. Then there's the fuel mass we're injecting, in a 1:2.56 ratio... that 22% becomes roughly 29% O2 and kerosense reaction product, by mass, in the exhaust, if the engine was consuming all the O2 that passes through it.

Of course, it doesn't. Turbojets don't (otherwise there'd be no O2 available for afterburners to work with), turbofans certainly don't.

I don't know about the SR-71... that thing is designed to cruise with afterburners on, and is getting 2,000 effective Isp when it does.

 

Its really simple... there's a lot of energy from the combustion available to do useful work. An air augmented rocket uses that energy to heat the inert gas, which expands in such a way that it provides useful thrust.

Edited October 21, 2016 by KerikBalm

[RedKraken]

Physics, chemistry and math are so useful when you are trying to do engineering. Kelly Johnson is an interesting fellow. I will have to reread skunkworks.

[RedKraken]

Titan Sample Return Plan with methane engines

direct mission, no gravity assists, no isru

48000kg from LEO  gets 80kg back from titan after 16years
(scaled-down 4800kg mission could bring back 8kg)
lift to LEO with falcH 

all pretend methalox engines at 320s sl 360s vac 300s? titan surface 
7 years out, ~2years at titan, 7 years back ...... ****zero boiloff *****
aerobrake direct to surface (both ends)
titan landing cost 100 m/s and heatshield and legs and sampler
earth landing chutes and heatshield and crumple element(include in 80kg)

assume stage dry mass is ~10%
assume LEO to saturn intercept = 7300 m/s (vac)
assume titan ascent to LTO = 4000 m/s (atmos)
assume LTO to earth intercept = 4200 m/s (vac)
transfer window planner in RO ksp113 + solar system dv map (check)

mission mass progression
LEO                             48000kg (18670 dry)
-> TSI stage II                 14000kg (4380 dry)
-> titan intercept then surface 2800kg drops shield to 2600 (667 dry)
-> titan ascent to LTO          480kg (146 dry)
-> Earth intercept then surface 80kg

if isru (assume 300kg)then
LEO                             17000kg 
-> TSI stage II                 4800kg
-> titan intercept then surface 1050kg refuels to 2600....drops isru for ascent
-> titan ascent to LTO          480kg  (142kg dry)
-> Earth intercept then surface 80kg

Hypergolic direct stack is over 100t.  Use a superDraco for titan ascent. 

Edited October 26, 2016 by RedKraken