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6.1990-2894

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PROPOSED NEAR SUN PROBE (NSP)
MISSION OPTIONS
Q
W. Vaning
1309-5th Ave.,
San Rafael, CA
JGA
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ABSTRACT
Due to the intrinsic interest in a
Near Sun P r o b e , a variety of mission
options were evaluated, including several
J G A t s , E J G A ' s , V G A , M G A and VHGA. T h e optimum
trajectory depends on the spacecraft
instrumentation, but is narrowed t o three
choices: VGA,VHGA or IV EJEGA. For experiments requiring continuous solar monitoring att'safe" distances, VHGA is superior.
For experiments requiring brief, periodic
visits extremely close to the s u n , the
dV EJEGA is best. Flight times are fairly
long (3 to 8 years or more before the
spacecraft i s on station).Propellant
requirements are achievable, with VGA&VHGA
requiring little more propellant then a
trip to Venus, and the /.V EJEGA needing
less propellant than a trip to Jupiter.
Solar sails may be a viable technique to
extend the usefulness of these trajectories.
T h e JGA uses a simple flight out to
Jupiter, with a gravity turn, and a return
close to the sun. It requires just about
3 years from launch to solar perihelion.
However, Total mission velocity is second
greatest of all the cases investigated. A
typical flight path is presented in Fig. 1
JGA - NSP Trajectory Specimen
h. Jupiter
Prograde
\
Jupiter Gravity Assist
Figure 1.
1
'
. --
Earth
-.
Launch
I
EARTH DIRECT
---
INTRODUCTION
Interest in a Near Sun P r o b e (NSP) is
increasing. T h e only probe which has been
launched close to the Sun w a s Helios, and
it only came within .25AU of the S u n , 16
years ago! Interest in sending spacecraft
close to the sun i s reviving, in part due
to the concept of a JGA plus a A V burn
close to the sun to launch vehicles out of
our solar system at unprecedented speeds.
Other reasons for closely approaching our
sun include relativity experiments, Helioseismology, verifying the missing solar
neutrinos, and particles and fields
experiments.
Currently, the recommended technique
to get close to the sun is a JGA , quite
similar to the U lysses mission. However,
since I had eval ated alternate trajectories 15 years ago , a study was performed
to determine the best trajectories for
NSP's. This study revealed that the VnGA,
VHGA and A V EJEGA trajectories ( Venus(Hg)Mercury Gravity Assist; Delta V
Earth-Jupiter-Earth Gravity Assist) need
less propellant, and support sustained
exploration close to the sun.
T h i s paper starts by expounding some
considerations about basic NSP trajectories such a s VGA,MGA and JGA. Then the V n G 4
VHGA and C V EJEGA are evaluated, including
options for extended missions. Criteria
for using Solar Sails a r e discussed. New
concepts, like pseudoheliostationary and
heliostationary orbits are covered. T h e
conclusion recommends using
4 spacecraft,
one for each NSP case, over time.
NSP
T h e direct trajectory i s the fastest
method for getting the spacecraft on
station, but requires an excessive amount
of propellant. Helios i s the only spacecraft to use this technique. As an example,
Helios only took 3 months from launch to
perihelion. O n the other hand, Helios
required a s much propellant a s a flight to
Jupiter. Helios only approached within .25
AU (Astronomical Units) of the s u n , and
alternate techniques do much better.Closer
solar approaches with this technique use
rather excessive amounts of fuel. Helios'
flightpath is depicted in figure 2.
Figure 2.
Direct
Y
* Member AIAA
Copyright (c) 1 9 9 0 by Walker S.Vaning
All Rights Reserved,Publ7d by AIAA with permission.
NSP OPTION
,,
-
.
NSP
Trajectory
-
Earth
/
/
\
\
\
_
---._I-.-
VENUS GRAVITY ASSIST (VGA)
By utilizing o n e or more swingbys of
the Planet Venus, the propellant requirement is substantially reduced, over either
the direct, or JGA method. However, it is
hard to get real close to the sun. Spacecraft perihelia may run between.15 and .25
AU. Since the spacecraft Aphelia must be
near Venus (.7233 AU), the resulting tra-
jectories are within .3AU of the sun only
20% of the time. However, the final orbital inclination can vary f15O or more!
Mission time i s quite variable, depending
on the number of Venus swingbys required.
Flighttime to final perihelion may vary
between 4 months and 5 years. A typical,
3 swingby flightpath is depicted in Fig.3.
-.
Launch oportunities for either of these 2
cases occur at 13 month .intervals. A
typical flightpath is shown in Figure 5 ,
with both the 2 and 3 year orbits.
Figure 5. A V EJGA Specimen T r a j e ~ t o r y
Plus:
AV EJEGA
Figure 3. Venus G_ravLty Assist (VGA) NSP
.
V~GA
#
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Launch
\
'.--__-
'
\
\
\,
,
Venus
*,
/'.~arth orbit
LV
EJEGA
T'
-----*,
MARS GRAVITY ASSIST (MGA)
MGA is included mostly for theoretical reasons. Martian gravity is so weak,
that decades, or centuries would be required to perform the many, repeated
swingbys past Mars, which are needed. It
is interesting, that at low launch velocities Venus requires less fuel, but at
high velocities Mars is more efficient at
getting really close to the sun. Note that
Mars' orbit is rather elliptical, and
that it is more efficient to swingby Mars
a t aphelion, than perihelion on this mission. Comments regarding the VGA apply
here, also, but MGA takes much longer.
A specimen flightpath is shown in Fig.4.
-
Figure 4. Mars Gravity AssisLJMGA)
-.-.- M ~ G Ashown
.
NSP
NSP OPTION
--
The Delta V,Earth-Jupiter-Earth Grav-+
ity assist adds another Earth swingby
after the Jupiter swingby. Earth Swingby
may occur just before, or just after
spacecraft perihelion. The purpose of the
second Earth swingby is to slightly lower
the perihelion, and (most important),
inject the spacecraft into a multiple of
the Earth's orbit, by lowering the aphelion. A single Earth swingby is sufficient to produce a 3 year orbit.Further
swingbys may produce 2 4 , 2 , 1 3 and 1 year
orbits.Earth swingbys may also change the
orbital inclination, to separate out
relativity effects from the sun's 5 2 and
higher components. The frequent returns
to Earth allow for precise orbit determinations.Mission time is quite long. If
the initial Earth-Earth orbit is just 2
years, than subsequent perihelions are
5,8,13,16,17,
years from launch.3 year
Flight path is shown in
Figure 5 , with
Earth return occuring a t either intersection of Earth orbit with spacecraft.
Launch velocities are close to, or slightly less than the previous A V EJGA case.
...
VHGA NSP OPTION
---
Typical: 3 0 years
to .21 AU
.
1
0
The proper name for this technique is
Delta-V, Earth-Jupiter Gravity Assist.It
starts with either a 2 or 3 year orbit
that returns to Earth. Near aphelion on
this spacecraft orbit, a Delta-V rocket
burn is performed, which efficiently increases the velocity upon Earth return. An
Earth swingby then sends the spacecraft
zooming out to Jupiter, which can deflect
it arbitrarily close to the sun. Since
the preliminary orbit is merely tacked on
to the JGA method, trip times are 5 to 6
years to perihelion. In return for this
longer journey, the total mission velocity
is reduced by . 5 miles per second,or more.
The Venus-Mercury Gravity Assist
option employs multiple Venus and Mercury
Gravity assists. Notice that Mercury is
denoted with an H (for chemists Hg), to
avoid confusion with Mars and MGA. Usually two Venus swingbys are used a t the
start Useful launch opportanities are
knowni, a s they are identical to Mercury
orbiter mission times. VHGA trajectories
are catagorized by having the lowest
launch velocities known, being just a
little more than a trip to Venus.Trip
times are over 1 0 years to final orbit,
however useful observations can start
about the 6th year. There are two types
of final orbit, depending on whether the
spacecraft aphelion matches the aphelion
or perihelion of Mercury. If the match
occurs at Mercury perihelion, the trip
time is longer, and the orbit more
circular. If the match occurs at aphelion
of Mercury, the final orbit is more
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oblate, and approaches the sun more
closely. If the spacecraft orbit perihelion is near .18 AU, the spacecraft will
hover over one solar longitude, for about
a week at a time. I call this a pseudoHeliostationary orbit. This orbit could
be quite sensitive to solar mascons.
Subsequent orbits occur at roughly 90' of
solar longitude from previous orbit. This
orbit should be excellant for Helioseismology a s well. The transfer from
Venus to Mercury must occur near Mercury
perihelion, for minimum total trip time.
If the final spacecraft orbit i s to match
Mercury Aphelion, a 180' transfer from
perihelion to aphelion must occur about
halfway through the pumpdown. A Specimen
trajectory is presented in Figure 6.
Figure 6. VHGA Specimen
- - - Trajectory
V ~ H G Ashown , , , : 2 - ~ , L a u n c h , Earth
n m o
Q2Zu-d
.- .- - _ - -
NSP EVALUATION
-
*
.,
CRITERIA
Several criteria are of importance in
designing NSP missions. Of primary importance is a low launch velocity.Parergon is
a short total mission time.Communications
may experience interference when the ship
is close to, in front or behind the sun.
In these cases, it is desireable for the
spacecraft aphelion to exceed .3 AU, so
taped communications can retransmit results. In light of the communications
problem, orbits that periodically return
to Earth are extremely desireable.
Certain mission experiments may
impose additional restrictions. Helioseismology requires long, uninterupted
observations of the sun, preferably over
a fixed solar longitude, over part of an
orbit which is neither approaching nor
receding from the sun. A good solution t o
this requirement is a perfectly circular
orbit with a radius equal to Mercury
perihelion, or .3075 AU. Particles and
fields experiments may also be performed
from this "safe" orbit.
On the other hand, relativity experiments and measurements of the solar 5 2 ;
J3,etc or solar Neutrino experiments need
to get a s close to the sun a s possible.
With contemporary technology, I selected
this distance to be about .025 AU, or 6
solar radii.It is safe to approach Jupiter
to 6 Jupiter Radaii, and the sun may have
similar radiation fields. Even at .025 AU
heavy shielding will be essential for
the intense thermal and radiation environments. Eventually spacecraft will achieve
even closer distances, but . 0 2 5 AU is
about the current limit for a heavily
shielded vehicle. Since the magnitude of
the solar radiation field is uncertain, it
is desireable to ease down into the final
orbit, gradually. If it is discovered that
Solar radiation is more intense than expected, the spacecraft can remain further
from the sun. T h i s i s especially relevant
to the VHGA and &J EJEGA options. Thus the
current design studies selected a minimum
perihelion of .025 AU.
Notice that solar radiation should be
1600 times more intense at .025 AU than at
Earth. A kilogram of isotope of a neutrino
detector at this distance from the sun may
be more sensitive than 2 torsof the bulk
material on Earth, and cheaper t o o . T h i s
could be used to determine if neutrinos
change variety on the trip from sun to
Earth. A few milligrams of Tritium may
constitute an excellant neutrino detector,
if i t s natural radioactivity is subtracted.
Some of the launch trajectories are
capable of reaching modest solar inclinations. This may be useful for measuring
latitude dependent Solar properties, like
cosmic r a y s , s u n s p o t s , m i c r o m e t e o r s , s o l a r
corona,rotation and solar wind.Some of
these properties may vary a s much between
Earth and .025 AU, a s they do between here
and the Voyager spacecraft at 4 0 AU.
Experiments to measure the zodiacallight, and the gegenshien may also be
informative.
Since total mission velocity is the
important determinator of mission cost, &
feasibility, it is presented in Figure7 a s
it varies with closest solar approach
distance. These curves were computed
assuming circular coplanar orbits. actual
results may differ slightly. A total mission velocity of 2.00, corresponds to a C 3
of zero, plus no man0uvers.A distance equal
to the solar radius is drawn a s a dotted
horizontal line. Total mission velocity is
a reliable indicator of mission cost because the ISP of the IUS is comparable to
the ISP of common postboost propulsion
systems.Figure 7 is derived from Ref.1,
Figure 7
.3
AU
.2
AU
Closest Approach vs Total
AV
LEO
-
.1 AU
Ro
-Solar impacL- - O (C3.0)
2:5
3(~3=38);
4(~3=83)
2.0 Total Velocity (Miles/second)
iS5
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Figure 7 reveals several interesting facts.
First, the VHGA requires the lowest total
mission velocities, for distances under .3
AU from the sun. Direct Earth launches are
least efficient. Notice that there are two
VHGA curves, one for Mercury Aphelion, the
other for Mercury perihelion. Perihelion
is least efficient of the two. The &V EJGA
is quite efficient using a two year precurser orbit,but for solar approaches less
than .1 AU, the 3 year precurser is superior. Both trajectories present a 40% to 502
payload increase over a straight JGA. Only
the trajectories using a Jupiter swingby
appear capable of coming closer to the sun
than . 1 AU with existing technology.
The
total mission velocities in Figure 7 were
calculated using the circular (except for
Mercury),coplanar
approximations, of the
planetary orbits.
SAMPLE V ~ + G A NSP
-One member of the VGA class appears
especially interesting. This trajectory i s
described here, even though it i s less
efficient than the VHGA, because of the
shorter mission time, and the possibility
of varying the spacecraft inclination.
This sample mission starts with a
launch C 3 of 3 8 km2/sec2 and uses 3 V e n u s
swingbys to pump the spacecraft down to t k
2/1 resonance. Intermediate orbits a t the
312 and 5 / 3 harmonics a r e required. While
the first Venus swingby may need to be
powered (merely .13 km/s), the second and
third swingbys are not critical. N O W this
211 harmonic orbit is interesting. Returns
to Venus occur every Venus year (.615 yrs),
with enough excess energy to vary the inclination up to f14.5'.
Spacecraft perihelion is .18 AU for this orbit, quite
close to the .17 AU final orbit. T h i s allows the spacecraft to sample, s a y , 5 different solar inclinations in just 3 years.
By flying a variety of inclinations, it
should be possible to distinguish orbital
perturbations due to relativity, from perturbations due to the J 2 , J 3 terms, and
other effects. It takes a little over 3
years to reach the 211 harmonic, and launch
opportunities occur twice every 1 9 months
(type I O0-180' transfer; type I 1 18O0-360°
transfer). Since there exists substantial
turn angle capability on the 3rd swingby,I
suggest that the inclinations relative to
the Venus ecliptic on the subsequent orbits
follow a pattern like t 5 ° , - 5 0 , t 9 0 , t 1 4 0 , t 3 0 ,
-1l0,-14",0°, and final pumpdown. T h i s sequence requires about 8 years. If this m a q
Venus swingbys a r e used, then the spacecraft will need at least .8 k m / s in manouvering propellant. If a slightly greater
launch C3 i s used, a wider variety of
inclinations is possible. Two spacecraft
perihelia
are obtained on each orbit, ard
perihelia occur a t about 120" apart in
solar longitude. Figure 3 shows the flight
path for this case. Figure 8 illustrates
the launch C 3 requirement, vs the final
range of inclinations at Venus. There is
also a limit on the change of inclination
per Venus swingby (see fig.10).
N S AP C 3 vs Venus Inclination
Figure 8. V ~ ~ G
,0
,
5
50- Launch
~ m ~ / ~ e c
4G7
Maximum available
+Venus Inclination(')
While this trajectory i s unique, it
does possess several disadvantages, a s
compared to the VHGA. F i r s t , the final
orbits are fairly elliptical, rendering
orbital calculations tedious. Second, it
does not stay over the same solar hemisphere for more than 2 w e e k s , near perihelion, and the large radial velocities
make it useless for helioseismology work.
T h i r d , it i s near to the sun (within .3 AU)
only 20% of the time. VHGA trajectories are
close to the sun m o r e of the time. Fourth,
the launch velocity i s nearly.3 km/second
greater than a similar VHGA, reducing payload.
T h e best advantages to this trajectory
are the relatively short 3 year "boost"
time, and the larger inclination changes.
Figure 9 depicts the launch window for
this mission, for a typical launch. C 3 in
Km/s i s the ordinate a x i s , and time i s the
abcissa.
Figure 9. V ~ ~ G
Launch
A
Window for NSP
43
-1
\
\
Launch C 3 vs Time
\
\
38
1
T-10 days
T-5d
'.- - .
<
T
/
Tt5d
Time
I
T t 1 0 days
Figure 9 illustrates that the launch
window i s about 2 weeks wide for this mis2.
sion, if the maximum C 3 is about 41 ~m'/s
Actual requirements will vary slightly, depending on Earth-Venus geometry.
While figure 8 depicted the maximum
possible heliocentric orbital inclinations
relative to Venus, the maximum inclination
change per Venus swingby i s also a function
of the excess energy, and consequently the
launch C3. T h i s is shown in Figure 10, on
the next page. Notice that for maximum inclinations under 14.5',
the spacecraft may
change the inclination from 0 " to the maximum, in just o n e swingby. If the maximum
inclination exceeds 21' then 3 or more
Venus swingbys a r e required to traverse
from 0 " to the maximum. T h i s added time
can greatly extend the total mission time
Hence, it is recommended that the maximum
inclination be between 15" and 20'.
This
can also keep the manouvering propellant
to a minimum, a s each added Venus swingby
uses about .025 k m / s of fuel.
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Figure 10. VnGA,NSP @Incl
O0 :
s
u
$
--
--- 211 Harmonic'
----+-
vs Max. Incl
'
-
-
---___
'--
7-1___1_
0"
+lo0
+20°
+30°
Maximum I n c l i n a t i o n ( V e n u s , d e g r e e s )
Notice that this mission requires that
the inclination start from a small value,
swing to o n e maximum (+incl), then to the
other maximum (-incl) and finally return
to a value near z e r o , so the spacecraft
perihelion can achieve its ultimate value,
a few million miles closer to the sun.
Figure 11 shows how these orbits that
are 2/1 harmonic with Venus, appear from
the sun. Unlike the other planets, our sun
has a rotation that varies with latitude.
Near the solar equator a day lasts 2 5
days. Near the Sun's poles, a solar day
lasts 3 0 days. For this plot, we assume an
intermediate spin rate, equal to a solar
day lasting 2 6 days. T h e heliocentric orbit is then close to heliostationary for
a brief period, near perihelion. Actually
the spacecraft stays within Z0 of solar
longitude, for the two day period surrounding spacecraft perihelion. Outside this
brief period, synchronization i s quickly
lost. Of course, closer to the solar equator, synchronization might last longer.
Figure 11. 2/1 P l o t viewed from the Sun
With tick marks a t 5 day intervals and
assuming 2 6 day solar day.
I t is recommended that this mission
be launched a s soon a s possible. There are
about 4 launch opportunities in the 19931995 period, which may be practical. Some
variation on the old Helios spacecraft may
be used, or a different design could be
used. Solar radiation at .17 AU is about
3 6 times a s intense a s at Earth. T h i s is
twice a s much a s Helios endured, but NASA
testing facilities can duplicate much higher values, up to 8 1 times the intensity a t
Earth. Such a mission is a vital precurser
to closer NSP's. Numerous particles and
fields experiments can be performed. Other
studies,of Solar Mascons associated with
sunspots, relativistic effects, Solar 5 2 ,
5 3 and higher coefficients, or the effects
of erosion on surfaces and solar sail
materials, etc., would be essential for
more advanced probes. If this mission were
launched near 1 9 9 4 , with a 3 year pumpdown
and 3 more years of orbital inclination
changing, then the results would be just
in time for more advanced NSP's at the turn
of the century.
T h i s mission i s essential for closer
missions. T h e difference between .25 AU
(Helios) and .025 AU (JGA type NSP's) is a
factor of 10 in distance. Radiation and
micrometeorites may fluctuate dramatically
over this range. Since this mission will
extensively explore the region from .16 AU
to - 1 8 AU, we will then only have to extrapolate over a range of 6 in distance.
This will make predictions of the necessary
shielding from radiation and particles more
reliable. This will allow lighter, more
accurate NSP's in the future. This may make
it possible to design Solar Sail powered
NSP's.
SAMPLE -VHGA NSP
-
A wide variety of trajectories of the
general type VXHYGA are feasible. V a r i a b l e
include launch velocity, initial transfer
orbit type, use of powered, reverse C\V G A ,
powered gravity assists, inplane or out-ofplane pumpdowns, and whether the perihelion
or aphelion of Mercury is used for the
final pumpdowns. While theoretically, this
type of trajectory can produce the closest
solar approach distances for a given launch
velocity (see Fig.7), there are problems.
F i r s t , the large number of Mercury flybys,
required due to Mercury turn angle limits,
demands long total flight times, and much
manouvering propellant.
Although one or more Venus swingbys
can be used, reference 2 suggests that 2 or
3 Venus swinqbys i s nearly optimum, for
this type of mission. Hence two sample
trajectories were generated, both using 2
Venus gravity assists. T h e difference between these two samples, is in whether the
final pumpdown uses Mercury Aphelion or
Perihelion (see Fig.6 and Fig.12). Since
the Aphelion pumpdown case fails to get
closer to the S u n than the previous VnGA,
discussion is brief. 13 HGA reach a 6 / 5
Mercury Harmonic. T h e final orbits
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are .4667 AU by .2189 AU, with the spacecraft aphelion intersecting Mercury's
ap helion, Since this type of mission is
restricted by long flighttimes, very little excess velocity can be used. Thus only
small variations from Mercury's 7' inclination, say about f 2 to f3 degrees, can be
acc.omplished, on subsequent orbits.
The pumpdown sequence is rather long.
First, two Venus Swingbys are used, with
an intermediate orbit resonant (1/1) with
Venus. Next, 6 Mercury swingbys occur at
Mercury perihelion, each time increasing
the spacecraft inclination, relative to
the Mercury ecliptic. The resonances are:
2/3t(to reach Mercury perihelion), 2/3,
3/4, 5/6, 8 / 9 , 1/1. Since this last orbit
is intergrally harmonic with Mercury, it
will return to Mercury at Aphelion, which
will subsequently be used for further p u m p
down. The resonances continue:l8/17, 9 / 8
and finally 6/5. Further Mercury swingbys
can be used to change the spacecraft inclination by a few degrees. Launch C 3 is
low, typically 2 0 ~ m 2 / ~ e cTotal
~ .
fuel for
maneuvering is about .6 Kmlsec. Total time
from launch to orbit insertion is 15.5
years. This is quite long, but may be cut
in half, by using about 1 km/sec more fuel.
An alternate sequence is possible, by
continuing the 111 resonance, in the last
paragraph, back to Mercury perihelion. Here
we can use the further sequence: 1019, 8/7,
511, 4 1 3 , 715, 10/7, and 3/2. Again, further Mercury swingbys can change the final
inclination f 3 degrees relative to the
Mercury ecliptic. Launch C 3 is the same,
20 ~ m ~ / s e c 2Total
.
Maneuvering fuel is .6
km/sec. Total flight time to the 312 orbit
is 16.7 years, which can be cut in half
for 1 kmlsec more maneuvering fuel. The
final orbit is .2825 AU by .3075 AU, with
the spacecraft returning to Mercury perih.elion every other Mercury year. Figure 1 2
shows this flightpath, although just a few
of the intermediate orbits are illustrated.
Figure 12
.
V ~ H ~ to
~ G 3/2
A Harmonic NSP
In the event that solar-ion propulsion
systems are allowed to be used, this
option becomes quite attractive. Reference
3 shows that s o l a p i o n propulsion systems
have a synergetic effect with Mercury
Gravity assists. Any contemporary Solarion system available off-the-shelf, could
reach .2 AU to .1 AU, with total flight
times of 4 to 7 years. Payloads will be
about 50% of the mass launched past Venus.
Solar-ion systems will increase the excess
energy, so final Mercury swingbys can
change the inclination by* 5' to f7O. One
of the major advantages of this option is
the greatly reduced flighttimes. The
solar-ion scenario requires the speed relative to Mercury to be reduced at first,
via inverse Mercury Gravity Assists. Then
the orbital pumpdown is increased with the
solar-ion propulsion system. Finally, the
solar-ion system increases the velocity
relative to Mercury, producing orbits
closer to the sun. Thus, if solar-ion can
be used with VHGA, then the circingle between .1 -.2 AU can be explored.
SAMPLE AV EJEGA
----
NSP
The use of a Jupiter Swingby is essential for achieving distances less than
.1 AU from the sun, by ballistic means.
Two options will be discussed. First,
after the Jupiter swingby, multiple Earth
swingbys can be used to lower the spacecraft orbit by orbit. Second, another
option is to perform an inverse AV burn
at closest solar approach, using subsequent Earth swingbys for inclination
changes. Each Earth swingby could change
the inclination about 2O, and several
swingbys may make inclination changes
totalling 6-go!
Figure 7 showed how the 2 year E-E
orbit was nearly a s good a s the 3 year E-E
precurser orbit. The use of one or two
Earth swingbys, after Jupiter, closes the
performance gap, even very close to the
sun, because Earth swingbys, here, lower
the spacecraft perihelion, a s well a s the
aphelion. Thus the 2 year precursor orbit
i s recommended, since the total time is a
year less. Figure 13 shows the effect of
A V burns a t .025 AU on the final orbit.
Figure 13. Final Orbit vs
6"
5
Period
(years)
Capture
"
4.
3 .,
Jup.
t
AV
a t .025 AU
V:~,(AU/Y~)
Escape
/
,[
8
Downloaded by UNIVERSITY OF NEW SOUTH WALES (UNSW) on October 27, 2017 | http://arc.aiaa.org | DOI: 10.2514/6.1990-2894
Notice in Figure 13, that if the perihelim
burn is negative, or less than .39 Milsec,
than the spacecraft remains in orbit about
the sun. If the burn is -.35 miles per sec.
then the period is 2 years. At -.65 Mils.,
the period equals 1 year. Actually, one
Earth swingby can contribute to reducing
the period, hence the necessary velocities
are .25 and .5 Milsecond, respectively.The
hyperbolic escape velocities are listed at
the right, for solar escape missions, in
AU per year at infinity.
If the option of doing one Earth
swingby plus a retarding AV burn is selecte d , then it may be preferable to use the
initial 3fyear Earth-Earth orbit, because
it is more efficient, close to the sun.The
total mission time to first solar perihelion is then 64 years. The one year orbit
should be selected over the 2 year final
orbit, because it makes twice a s many returns to the sun and Earth, for just an
additional .25 miles/second.Thus the sun
will be visited at 64,74,83,93,etc. years.
Each Earth return may vary the inclination
by 2'.
and large inclinations of 10' or
more can be built up, after several years.
This option appears quite promising.
There'sanother option using the 3+ yr
4 V EJEGA method. If a launch velocity just
.2 miles per second greater is used, the
spacecraft will go into a retrograde orbit
close to the sun. A proper selection of
trajectory will allow the Earth to be visited on one, or preferably both legs, less
than a month apart! The use of two Earth
swingbys on each orbit can lower the aphelion faster. Spacecraft can be lowered directly into 3 year, 2year, and 1 year orbits. Thus, (using a 3+year launch) the s m
will be visited at 6 3 , 9 + , 1 1 $ , 1 2 ~ , 1 3 year
~
times. Again, the option to use a r e t a r d i e
AV burn a t solar perihelion, may be preferable, a s the spacecraft will visit the sun
every year after 6 3 years, and the penalty
is only .385 miles per second, if two
Earth swingbys can be used to augment the
slowdown, into a one year orbit. If the
retrograde option i s used in conjunction
with a prograde NSP, the relative velocity
at perihelion, between the two, will be 3%
miles/second!
Thus the &V EJEGA options are very
powerful. However, it is recommended that
preliminary probes explore the region between .1AU t o . 2 AU, first, to determine
how close the vehicle can safely approach
the sun.
may not be available for some time.
Current Solar sail technology6can produce accelerations of 1 mm/sec2 at 1 AU.At
.2 AU the same sail generates 25 mm/s2!
Actually, a high temperature Solar sail may
be 1 0 times heavier, for an acceleration of
2.5 mm/s2 at .2 AU. This is equivalent to
.215 Km/sec per day, which is enough to
change the circular velocity by 5% every
day at .2 AU. With this much acceleration,
it i s possible to spiral into a .1 AU circular orbit in under 4 0 days, or to change
the orbital inclination by 90' in the same
period. It will be possible to lower the
NSP orbit a s close to the sun a s the temperature will allow.
Since solar sails can produce nearly
circular orbits, close to the sun, unreachable in any other w a y , they will,
probably be launched after the turn of
the century. An interesting goal would be
to reach Heliostationary orbit. That i s , m
orbit which circles the sun every 26 days,
always remaining over one location on the
Sun's equator. This is an ideal orbit for
helioseismology work. The corresponding
distance from the sun's center, for this
orbit i s .I71762 AU. Actually, a radius
slightly greater, or less, may be used, a s
the sun rotates slower at higher latitudes.
Also, three heliostationary satellites are
required to give complete 360' coverage of
the sun's surface. Only solar sails can
efficiently inject satellites into such
orbits, using either VGA or VHGA to reach
.2 AU, or thereabouts.
Design of high temperature solar
sails is essential to a Heliostationary
mission. Equilibrium temperatures at .I72
AU are near 450°C. Solar sails using silver
aluminium or tin coabings, will evaporate
in a few days at this temperature, and
parts of the reflector may get even hotter.
Higher temperature reflectors for this
mission might survive with Chromium,gold
or nickle coatings.
Solar sails create a design challenge
for trajectory dynamicists, also.Solar
sails possess a large radial component to
their acceleration vector. The thrust direction must be optimized for each moment of
flight. Since this mission is not likely to
be performed for another 15 years, it will
not be analyzed further herein.
Figure 14.
_ - - --
-
/
H
,,
/
SOLAR SAIL NSP
--Preliminary calculations suggest that
the use of high temperature solar sails
can produce a variety of orbits close to
the sun. Most of these orbits cannot be
achieved in any other way. High temperature
solar sails are essential, for operating in
the region under .2 AU from the sun.
Temperatures from 400°C to 5000°C occur
between .2 AU and the sun's surface.Since
many solar sail designs are based on easily evaporated metal reflectors, hightemperature materials must be studied, and
-.
Heliostationary Orbit
-r
Merc_u_ry-Crbit
/
'
#
,' Sun
I /
I
I
I
\
\
\
\
\
I
I
\,
.- - _ -
,- ..1673-. 189
30,'days,
/
/
...- - - -___ --- -\
\
\
I
\
\
\
!
\
\
'-25 days(3,L'to
\
\
\,~
1
\
I
\
\
-.
1
1
\
\
.
/
/
'
'
1
,
/
/
',Earth
Orbit
V e u s ..Orb_i.
t
AU_ R
4
/
/
,/
-
r
'
/
,
I
NAMES
Each NSP spacecraft will need a different,distinctive name. Helios and Gallileo have already been used. Currently,
spacecraft are named after scientists who
have researched the spacecraft objectives.
Thus Maunder (Maunder Minium) is appropriate. A Solar Sail mission, that has its
sails slowly evaporating, could be named
Icarus ( a s already suggested by the POGO
cartoonists). Other solar scientists names
can also be used.
Downloaded by UNIVERSITY OF NEW SOUTH WALES (UNSW) on October 27, 2017 | http://arc.aiaa.org | DOI: 10.2514/6.1990-2894
COMMENTS
There exist two other techniques to
accelerate the rate of orbital pumpdown
by the terrestrial planets. These are nonharmonic orbits and additional Mercury or
Venus swingbys.
All of the trajectories in this paper
used harmonic orbits, which return to
a
In many
planet using multiples of 360'.
cases, the spacecraft orbit crosses the
swingby planet orbit at two locations, the
second not being an intergral multiple of
360". I n some instances, the pumpdown can
be accelerated by making a larger swingby
angle before flipping the swingby location
to the other orbital crossing point, using
a nonharmonic orbit. The calculation
of
these nonharmonic orbits requires about a
dozen times the number of calculations a s
for the harmonic orbit case. Consequently,
it wasn't performed for these cases. Mission times can be reduced about 30%, for
many cases, where nonharmonic orbits can
be incorporated into the trajectory. Note
that this will affect all subsequent
spacecraft orbits, a s well. Use of nonharmonic orbits will not affect the final
orbit of the spacecraft, only the time
required to reach this goal. See Figure 15.
Figure 15 Nonharmonic Orbit Pumpdown
Orbit of Mercury,Venus,Earth or Mars
--
/
'
and
.Start
Harmonic
'.
\
Return
Aharmonic
Noninterger
-
& Nonperiodic
(both in plane of Planet
ecliptic, for Aharmonic)
Any inclination can be used with Harmonic.
The second technique for improving
these trajectories involves additional
swingbys of Mercury or Venus. During the
pumpdown of a VnGA type orbit, the spacecraft will cross the orbit of Mercury many
times. Frequently, the spacecraft may come
close to Mercury. Occasionally these close
Mercury flybys can be converted into extra
Mercury swingbys. The gravity of Mercury
can then be used in either of two ways. If
Mercury boosts the spacecraft orbit, then
return to Venus (or Earth) will be faster.
This results in a longer mission, which
will get about .O1 AU closer to the sun
for each additional Mercury boost. On the
other hand, if Mercury decellerates the
spacecraft, then the pumpdown may be made
in less time, but will not approach the
sun a s closely. Similarly, an inner planet
may be used during EnGA or MnGA sequences.
Figure 16 illustrates the two planet,orbit
crossing geometry, although the orbital
perturbation due to Mercury is too small
to be perceived on this scale. Mercury may
be flownby a t either of the two crossing
points, on any of the several harmonic
orbits. ( A or B).
Figure 16
-.
Potential HGA during VnGA
- .- ',
Venus Orbit
_-
,
/
/
/
I
I
\
\
\
\
Mercury
'
Orbit
\
craft
Furthermore, both of these techniques
may be used together on a NSP mission.
Occasionally a NSP trajectory may
pass close to a comet, such a s Encke's
comet. It might be desireable to select
such trajectories, a s a comet flythrough
could return valuable scientific information.
Downloaded by UNIVERSITY OF NEW SOUTH WALES (UNSW) on October 27, 2017 | http://arc.aiaa.org | DOI: 10.2514/6.1990-2894
RECOMMENDATIONS
Several NSP's should be launched in
the coming decades. First, a NSP that uses
several VGA to reach a 211 harmonic orbit
with Venus, allowing for many different
solar inclinations up to 15O to 20° is an
essential precurser to closer missions. It
may be launched in the 1993-1995 timeframe, and has a relatively low mission
velocity and cost.
Next, a VHGA to an orbit that is about
.31 AU in radius and circular, is optional
for Helioseismology studies.
After that, a NSP using a A V EJEGA is
capable of getting really close to the sun
for short periods, and can return to Earth
periodically. This should be launched near
2000 AD.
For the far future, Solar sails are
needed to produce roughly circular orbits
close to the sun, such a s Heliostationary
orbits at .17 AU, where 3 spacecraft are
recommended. This may be launched around
2005 AD. Even later, Solar sails may put
spacecraft in circular orbits close to the
sun's surface, for neutrino and other experiments. These missions will all require
high temperature solar sails, and preliminary flybys of Venus.
CONCLUSIONS
A variety of trajectory types were
examined to determine the parameters and
characteristics of a Near Sun Probe (NSP)
mission. Several Candidate trajectories
emerged. For the immediate future,
a
can
multiple Venus swingby trajectory
reach a 211 Venus Harmonic orbit in three
years, and the final orbit can be varied
+20° in inclination, with further
Venus
swingbys, for a few more years.
Toward
the end of the Century spacecraft may use
Venus-Mercury Gravity Assists (VHGA) and
Delta-V Earth-Jupiter-Earth trajectories
to efficiently get close to the sun.
Again, multiple swingbys of the final
planet can produce inclination changes in
the orbit. Finally, solar sails will
be
used next
century to reach orbits, such
a s Heliostationary orbit, which cannot be
accomplished with other techniques.
ACKNOWLEDGEMENTS
I would like to thank Dr. Chen-wan L.
Yen, and the astronomy Department at UC
Berkeley for inspiration and comments on
this paper. All calculations were performed on an HP-45 and Tandy 1000 TX. Data from
references 4 and 5 were used in the calculations.
REFERENCES
1 ) "Space: 20 Years and Countingw, W.Vaning
PTI, 1977
2) "Ballistic Mercury Orbiter Mission via
Venus and Mercury Gravity AssistsW,Dr.
Chen-wan L.Yen, AAS 85-346, JPL 1985
3) "Unique Gravity Assist Missions for the
1990'sM, W.Vaning AAS 89-468,1989
4) "Ephemerides of Minor Planets for 1987",
Marsden,B. & Cincinnatti Minor Planet
Center.
5) "Planetary Flight Handbook, Vo1.3, Part
9-Direct and Venus Swingby Trajectories
to Mercury", NASA SP-35, 1970, General
Dynamics
6 "Some Solar-Sailing Strategies for
Scientific Missions in the Ecliptic
Plane" Willem Stuiver, AAS 89-466
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