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Old 05-23-2016, 02:31 AM   #16
RGClark
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OK, one more stab at the Mars moons proposal. We have discussed before the surprising speed that New Horizons was able to pass the orbit of Mars due to its high departure speed:

Math needed for 5-week flight from Earth to Mars.
http://orbiter-forum.com/showthread....3&postcount=78

The transit time of New Horizons to pass Mars' orbit was 78 days for a LEO departure delta-v of 8.4 km/s. But as discussed in that post, the Mars capture delta-v would be close to 12 km/s, prohibitive when added to the departure delta-v.

So instead I'm thinking of other ways of being captured at Mars such as aerocapture:

https://en.wikipedia.org/wiki/Aerocapture

This allows the spacecraft to be captured into Mars orbit with minimal propellant burn.

So using a limiting LEO departure delta-v of 8.4 km/s as with New Horizons and assuming the departure takes place in the time frame of the last quarter of this year to the first quarter of next year, what would be the travel time then?

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Old 05-23-2016, 05:13 AM   #17
Keithth G
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This is doable. But I'm travelling without computer for a couple of weeks. Will get back to this upon my return to HK.

Last edited by Keithth G; 03-05-2017 at 11:47 PM.
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Old 05-23-2016, 02:19 PM   #18
RGClark
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Thanks. I'll try also calculating it using the circular orbit, coplanar approximation discussed in post #9.

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Old 05-23-2016, 04:03 PM   #19
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Quote:
Originally Posted by Keithth G View Post
 Here, I have assumed a simple mission design in which the spacecraft escapes a 200 x 200 km Low Earth Orbit (LEO); makes a direct transfer to Mars; and then inserts into a 200 x 200 km Low Mars Orbit (LMO).
Don't most Martian missions inject into a highly elliptical orbit and then use successive aerobraking passes (sometimes over months) to pull the orbit in?

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  So instead I'm thinking of other ways of being captured at Mars such as aerocapture:
I'm fairly certain that aerocapture has never actually been used. It might not be worth it because of the extra weight required for heat shielding and stress mitigation.

OTOH, MSL, I believe didn't bother with establishing orbit at all. It came in from the transfer orbit directly to atmospheric entry.

Last edited by Shifty; 05-23-2016 at 04:13 PM.
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Old 05-23-2016, 05:14 PM   #20
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Quote:
Originally Posted by Shifty View Post
 Don't most Martian missions inject into a highly elliptical orbit and then use successive aerobraking passes (sometimes over months) to pull the orbit in?
I'm fairly certain that aerocapture has never actually been used. It might not be worth it because of the extra weight required for heat shielding and stress mitigation.
OTOH, MSL, I believe didn't bother with establishing orbit at all. It came in from the transfer orbit directly to atmospheric entry.
Yes, aerobraking has been used when the spacecraft has already been put in capture orbit at Mars. Aerobraking then reduces the size of the orbit. It may also have been used to circularize. I'm not sure about that.

Aerocapture is a more difficult proposition. It would require the spacecraft to plunge deep into Mars atmosphere, skimming the tree tops so to speak. That has not been tried yet. Since this is only a test flight anyway this may be a good chance to test it.

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Old 05-23-2016, 05:23 PM   #21
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According to Wikipedia atleast, the MSL entered the Martian atmosphere at 5.8 km/s and experienced a peak deceleration of 15g.

Last edited by Keithth G; 03-05-2017 at 11:46 PM.
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Old 06-06-2016, 04:22 AM   #22
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I've now had a chance to run the problem through PyKEP:

For possible transfer times of less than 500 days, the best day to leave is 11 October, 2016; the transfer time to Mars is 273 days; the departure dV is 8.4 km/s; and (at a 50 km periapsis altitude at Mars) the periapsis velocity would be 10.2 km/s.

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Old 04-09-2017, 04:57 PM   #23
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Originally Posted by RGClark View Post
 The launch window to Mars comes every two years where the delta-v to get there by a Hohmann trajectory is minimal. This year it was in March, when Exomars was launched. But you can still launch to Mars at different times. It just requires more delta-v.
SpaceX wants to decide on an initial mission for the Falcon Heavy, scheduled to launch at the end of this year:

SpaceX undecided on payload for first Falcon Heavy flight.
May 3, 2016 Stephen Clark
http://spaceflightnow.com/2016/05/03...-heavy-flight/

One possibility would be a lander mission to Phobos or Deimos. I calculated from a delta-v chart that it would be doable carrying a Dragon during the minimal delta-v window using an existing solid rocket upper stage for Earth escape. But that minimal delta-v launch window already passed in March.
How do you calculate the delta-v requirements during other times such as the end of the year?
The idea behind this post was that SpaceX was going to launch a Falcon Heavy flight to Mars outside the optimal launch window. However, now the plan will be to launch in 2018 which presumably will be within the optimal launch window.

So my proposed mission is still possible for this first FH launch, to do a fast flight to Mars, ca. 35 day duration, to demonstrate its feasibility for a fast manned mission. The delta-v needed for such a fast flight to Mars in 2018, a particularly close opposition, was discussed here:

Math needed for 5-week flight from Earth to Mars.
http://orbiter-forum.com/showthread....6&postcount=17

It's about 10 km/s. The updated specifications for the Falcon Heavy with the upgraded Merlin engines are on the SpaceX Falcon Heavy page.

I estimate we could get 2 to 3 metric tons as payload to Mars for the fast trip depending on whether we used for the in-space stage the small cryogenic Ariane 5 upper stage leaving from Trans Mars Insertion, or the larger Centaur leaving from geosynchronous transfer orbit.

We still have the problem of slowing down when the use such fast flight speeds which result in fast arrival speeds at Mars. Some preliminary calculations suggest it might work by plunging deep into the Martian atmosphere, skimming the tree-tops so to speak.

Bob Clark
Attached Thumbnails
Falcon Heavy specs.jpg  

Last edited by RGClark; 04-12-2017 at 01:33 PM.
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Old 04-20-2017, 07:37 PM   #24
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Actually, while the Red Dragon mission on the Falcon Heavy is set for 2018, SpaceX plans for two prior FH test flights for the latter part of this year.

Elon has discussed testing recovery of the upper stage on these missions which will reduce payload. He has also discussed putting a "fun" payload on one of them, like his cheese wheel on the first Dragon test flight.

Still, if low cost in-space stages could be used for a flight to the Martian moons perhaps Elon could be convinced to make one or both of these first FH flights be to the moons of Mars. Note that key to Elon's plan for manned flights to Mars is getting the fuel for the return trip from Mars. Taking the fuel from the Martian moons would have advantages such as low gravity for getting the fuel to an orbiting propellant depot. Then these first flights to the Martian moons could serve as scout missions for water ice deposits. Plus, it could resolve the mystery of Phobos' origin, whose low density led to much speculation about it.

In the table provided by Keithth G in post #14, the launch dates from Sept. to Dec. 2017 have travel times to Mars in the range of 270 days. But being outside the optimal launch windows, they have large delta-v requirements. So my plan to test short flight times by high departure speeds wouldn't be very useful for these flights. That would have to be reserved for the optimal departure windows.

In the blog post "Low Cost Europa Lander Missions", I discussed some small in-space stages for a possible Europa mission.

Two stages discussed were the storable propellant stage Delta K and the Integrated Apogee Boost Subsystem (IABS) stage. The Delta K has a 6 mT propellant load, 0.95 mT dry mass, and 319 s Isp. The Integrated Apogee Boost Subsystem (IABS) stage is a small kick-stage used to put geosynchronous satellites in their final orbits. It has a 1.6 mT gross mass and .3 mT dry mass, for a 1.3 mT propellant mass, with a 312 s Isp.

Then for a small 1.5 mT robotic rover it could get this to:

319*9.81ln(1 + 6/(.95 + 1.6 +1.5)) +312*9.81ln(1 + 1.3/(.3 + 1.5)) = 4,500 m/s.

The latest specs on the Falcon Heavy give it a 16.8 mT payload to Mars transfer insertion. This is about a 3,800 m/s delta-v.

The Dec. 23, 2017 departure according to Keithth G's table takes 4,836.1 m/s for Earth departure and 2,451.5 m/s to match Phobos orbit at Mars, for a total of 7,287.6 m/s. Then to land on Phobos requires an additional 500 m/s, so all together 7,787.6 m/s, call it 7,800 m/s.

Then since the Falcon Heavy will already provide 3,800 m/s for Earth departure, only 4,000 m/s would have to be provided by our two in-space stages, which is within their capability with a 1.5 mT payload.

This 1.5 mT payload that could be landed on Phobos is so large we might even be able to include a solid rocket stage to return a sample from Phobos to Earth.

In regards to the cost, NASA wants a mission to Phobos so they may be willing to pay for the cost of the in-space stages.

Bob Clark

Last edited by RGClark; 04-20-2017 at 07:40 PM.
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Old 04-22-2017, 04:56 PM   #25
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Quote:
Originally Posted by BrianJ View Post
 If you just want the dV values (rather than the fun of figuring out how to calculate it) then Piper's "Trajectory Planner" is very useful:
Trajectory Planner
I'm trying out the "Trajectory Planner" now. I assume the "Departure deltaV" means the dV from just the position of the Earth's solar orbit, i.e., no consideration of a spacecrafts orbit around Earth, and "Arrival deltaV" means the dV to match Mars in it's position in solar orbit, not what it takes to put it in orbit around Mars.


So how do you find the delta-V needed to make the Mars transfer injection assuming the spacecraft is already in LEO?

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Old 04-23-2017, 04:03 AM   #26
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Quote:
Originally Posted by RGClark View Post
 I'm trying out the "Trajectory Planner" now. I assume the "Departure deltaV" means the dV from just the position of the Earth's solar orbit, i.e., no consideration of a spacecrafts orbit around Earth, and "Arrival deltaV" means the dV to match Mars in it's position in solar orbit, not what it takes to put it in orbit around Mars.


So how do you find the delta-V needed to make the Mars transfer injection assuming the spacecraft is already in LEO?

Bob Clark
\Delta V = \sqrt{V_{\infty}^2 + V_{esc}^2} - V_{orb}

where V_{\infty} is the hyperbolic excess velocity (departure deltaV from trajectory planner).

V_{esc} is the local escape velocity, aka the escape velocity for the parking orbit altitude.

V_{esc} = \sqrt{\frac{2GM_{planet}}{R_{planet}+alt}}

where G is the gravitational constant, M_{planet} is the planet's mass, R_{planet} is the planet's radius and alt is the altitude of the parking orbit.

V_{orb} is the parking orbit velocity.

V_{orb} = \frac{V_{esc}}{\sqrt{2}}

Same applies for arrival. If you want to simply calculate the periapsis velocity and not the orbit insertion/injection dV, then don't use the V_{orb} term.

Source: ORBITAL MECHANICS

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Old 04-23-2017, 01:20 PM   #27
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Quote:
Originally Posted by dgatsoulis View Post
 \Delta V = \sqrt{V_{\infty}^2 + V_{esc}^2} - V_{orb}
where V_{\infty} is the hyperbolic excess velocity (departure deltaV from trajectory planner).
V_{esc} is the local escape velocity, aka the escape velocity for the parking orbit altitude.
V_{esc} = \sqrt{\frac{2GM_{planet}}{R_{planet}+alt}}
where G is the gravitational constant, M_{planet} is the planet's mass, R_{planet} is the planet's radius and alt is the altitude of the parking orbit.
V_{orb} is the parking orbit velocity.
V_{orb} = \frac{V_{esc}}{\sqrt{2}}
Same applies for arrival. If you want to simply calculate the periapsis velocity and not the orbit insertion/injection dV, then don't use the V_{orb} term.
Source: ORBITAL MECHANICS
Thanks for that. I was getting different numbers than in Keithth G's table in post #14. I suspected that was the answer.

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Old 05-07-2017, 01:59 PM   #28
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Some suggestions for the test flights of the Falcon heavy:

Test flights of the Falcon Heavy for missions to the moons of Earth and Mars, Page 1.
http://exoscientist.blogspot.com/201...heavy-for.html


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Old 05-19-2017, 06:05 PM   #29
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Quote:
Originally Posted by RGClark View Post
  One possibility would be a lander mission to Phobos or Deimos. I calculated from a delta-v chart that it would be doable carrying a Dragon during the minimal delta-v window using an existing solid rocket upper stage for Earth escape. But that minimal delta-v launch window already passed in March.
For ballpark Hohmann numbers you could use the spreadsheet I used to make my Cosmic Train Schedule page.

Here's a screen capture for arrival to a Deimos orbit:



You can see insertion to this orbit takes about a 2 km/s periapsis burn.

Here's a capture orbit with Deimos at apoapsis



A 1 km/s periapsis burn plus about a .7 apoapsis circuralize burn gives about 1.7 km/s

Quote:
Originally Posted by RGClark View Post
 How do you calculate the delta-v requirements during other times such as the end of the year?

Bob Clark
Here is another spreadsheet, NonHohmannEarthToMars.xlsx. Like the other spreadsheet I assume circular coplanar orbits.

The user can vary shape of transfer ellipse by inputing aphelion and aphelion. For example the typical Hohmann would have a 1 A.U. perihelion and a 1.52 A.U. aphelion.

Here's a screen capture when transfer orbit has a .9 A.U. perihelion and a 3 A.U. aphelion:



Underlined are numbers of interest: departure V infinity, arrival V infinity, and time of flight.

Off to the right on this spreadsheet the user can input the Mars orbit you wish to insert into and it will give the delta V.
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Old 05-20-2017, 03:04 AM   #30
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Quote:
Originally Posted by HopDavid View Post
 ...
Here is another spreadsheet, NonHohmannEarthToMars.xlsx. Like the other spreadsheet I assume circular coplanar orbits.
The user can vary shape of transfer ellipse by inputing aphelion and aphelion. For example the typical Hohmann would have a 1 A.U. perihelion and a 1.52 A.U. aphelion.
Here's a screen capture when transfer orbit has a .9 A.U. perihelion and a 3 A.U. aphelion:
{image}
Underlined are numbers of interest: departure V infinity, arrival V infinity, and time of flight.
Off to the right on this spreadsheet the user can input the Mars orbit you wish to insert into and it will give the delta V.
Thanks for that. Other flights I wanted to calculate were fast trips to Mars and also Jupiter. These would be non-Hohmann transfer orbits, so your spreadsheets would be very helpful here.


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