Re: Chinese Mars flyby?

America had a mars flyby as a fall back in case the Russians landed on
the moon first. Basically, you have a skylab type module, with an
apollo capsule affixed, launched into a 24 month orbit launched to fly
by Mars and return to Earth without any further propulsive input.
You have an SIVB fully fueled on orbit, and boost to mars flyby speed
and when the SIVB is empty, use the hydrogen tank as a long duration
module. A module where the LEM would be has mars rovers, and a soil
analysis lab and so forth attached.

Except for duration, its actually easier to fly by Mars than land on
the moon..

In fact, you can fly by mars on the outbound leg, and fly by mars
again on the inbound leg.

With today's tele-robotic capabilities you can land asimo type robots
on Mars and drive them around, and then as you fly by on your inbound
leg, samples get returned to the vehicle for detailed analysis aboard
ship.

This system could use a Soyuz/Mir type setup as well, with a BRIS-M
module or two for boost.

Kepler's third law says the period of an orbit squared is equal to the
semi-major axis length cubed. So, the Earth orbits the sun once per
year and is one Astronomical Unit (AU) semi-major axis length. To
boost something into an orbit that circles the sun every TWO years
requires us to square two and take the curbe root of the result -
giving us.the cube root of four or 1.5874 AU - that's the semi-major
axis required to carry out this mission.

The orbit has a perihelion of 1 AU and to have a semi-major axis of
1.5874 then the apohelion must be 2.1748 AU.

This is not as far as Ceres, but is well inside the asteroid belt, so
there might be a telescope on board the spacecraft to allow it to take
photos of any asteroids that fly nearby during apohelion.

Using Kepler's first law - and simple geometry of an ellipse, this
gives us the eccentricity and semi-latus rectum of the ellipse that
describes this orbit.

The eccentricity is 0.5874 and
the semi-latus rectum is 1.5874

The Sun is at one focus of this ellipse, and the origin of the ellipse
is on a line of the major axis 0.5874 AU from the center of the sun.
All angles are measured from this point.

v = 0 degrees is perihelion, r = Rmin = 1.000
v = 90 degrees, r = p = 1.5874
v = 180 degrees r = Rmax = 2.1748

r = p / (1 + e*cos(v) )

Now Mars' orbits at a distance 1.5237 AU

So, v = 86.919 degrees - which explains why you can encounter the
planet on the way out, and then encounter it again on the way back.

This takes about 160 days to get there for flyby.
Then 205 days after flyby you're farther from the sun than any human
has flown
Then 205 days after apohelion you're back at mars
Then 160 days after the second flyby you're back at Earth.

We're starting at 1 AU from the sun, and we can use the Vis Viva
equation to figure out the speed at perihelion of 1 AU. requires 7.17
km/sec

Even though you've slowed from 36.92 km/sec relative to the Sun at
Earth to 34.48 km/sec relative to the Sun at Mars, you're flying by
Mars at about 10.38 km/sec. You're within 1 million km of Mars for
about 107 hours.

The S-IVB carried 73,280 liters (19,359 U.S. gallons) of LOX, massing
87,200 kg (192,243 lbs). It carried 252,750 liters (66,770 U.S.
gallons) of LH2, massing 18,000 kg (39,683 lbs), and massed 119,900 kg
total

SIVB

119.9 tonnes
87.2 tonnes LOX
18.0 tonnes LH2
14.7 tonnes structure

With an exhaust velocity of 4.5 km/sec and a delta vee requirement of
7.17 km/sec we have a propellant fraction of 0.7967 which means 132.0
tonnes total vehicle weight. Subtracting out the structure, we have,
12.13 tonnes of useful payload.

The Apollo Command Module was 5.2 tonnes in mass, this leaves 6.93
tonnes of equipment and consumables.

1 tonne per person per year consumables, for a crew of two for two
years totals 4 tonnes - allowing 2.93 tonnes for robotic landers and
perhaps even sample return systems.

As mentioned the SIVB would be equipped with a 'wet' stage that housed
propellants, and the crew would use it once it fired.

Werner vonBraun wanted to use the S-II upper stage on the Saturn V as
a 'wet workshop'

http://en.wikipedia.org/wiki/Wet_workshop

but later the SIVB was launched, dry as skylab. The Saturn V could
have launched a 'wet' SIVB into an escape trajectory, and the 'wet'
SIVB fired its engine (likely a RL10 engine operating at very high
expansion) draining the tanks,which would then be used in transit to
house a crew of 2. Solar panels would be deployed to power the space
station - and a reduced mass SM would carry supplies and so forth, for
the mission.

The Chinese have worked on their own version of a 'wet' workshop using
the upper stage of a Longmarch rocket project 921-2

http://en.wikipedia.org/wiki/Project_921-2

The second stage of the CZ-2E is given below;

Stage2: 1 x CZ-2E-2. Gross Mass: 91,500 kg (201,700 lb). Empty Mass:
5,500 kg (12,100 lb). Motor: 1 x YF-25/23. Thrust (vac): 831.005 kN
(186,817 lbf). Isp: 298 sec. Burn time: 295 sec. Length: 15.52 m
(50.91 ft). Diameter: 3.35 m (10.99 ft). Propellants: N2O4/UDMH.

The problem is exhaust velocity is 2.9 km/sec - not 4.5 km/sec - which
means lots more propellant is needed.

But the space station module size is very close to the Stage 2 size -

CZ-2E-2 length 15.52 m diam 3.35 m wgt 5.5 tonne
921-2 length 9.00 m diam 2.80 m wgt 8.0 tonne

Launching a CZ-2E-2 modified as a 'wet' space station - requires 11
launches to 'refuel' it.

The Shenzou spacecraft consists of 3 parts- an orbital module, a
service module and a re-entry module;

Orbital module - 1.5 tonne
Re-entry module - 3.3 tonne
Service module - 3.0 tonne

This is a total mass of 7.8 tonnes

Docking this system to a refueled 'wet' station means 13.3 tonnes of
payload and structure is propelled by 89 tonnes of fuel operating at
2.9 km/sec exhaust speed. This provides a final delta vee of 5.9 km/
sec from LEO - which has an excess velocity of about half that needed
for the mission we're discussing.

Refueling TWO CZ-2E-2s one 'wet' the other not, with 22 fueling
flights, and firing them in sequence. We have 18.8 tonnes of payload
plus 89 tonnes of propellant being accelerated by another 89 tonnes of
propellant with an exhaust speed of 2.9 km/sec.adds another 1.74 km/
sec to the delta vee - which when added to the first calculation
totals 6.64 km/sec from LEO

Adding TWO MORE CZ-2E2s - with 44 fueling flights, and firing the
first two then the sequence just described - same propellants and
performance - adds another 1.79 km/sec - bringing the total to 8.43 km/
sec from LEO.

So, if they had one flight every month - it would take four years to
accumulate the fuel on orbit. They could also accumulate 'empties'
for additional missions - or to expand a space station on orbit.

This is quite a program - and requires a substantial commitment by the
Chinese.

It is also a gradual development over time - providing clear
milestones along the way - and provides a means to work on lowering
costs by steady continual regular production schedules.

Orbiting a crew,
Orbiting a space station module,
Refilling a space station module
Lunar flyby - with lunar orbit - and drop robot probes on moon
Mars flyby - with robot probes on mars
Lunar lander and return
Mars lander and return

Which is pretty much the program outlined in vonBraun's CONQUEST OF
SPACE - which has been supressed since vonBraun was marginalized after
the assasination of JFK.

Of course, that doesn't stop knowledgeable engineers from other
nations reading the book and planning their own programs.








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