Conference PaperPDF Available

Venus Sample Return Mission Concept Development

Authors:

Figures

Content may be subject to copyright.
1
American Institute of Aeronautics and Astronautics
Venus Sample Return Mission
Concept Development
Kevin Carpenter, Samuel Basinait, James Bilello, Patrick Carroll
and
Javid Bayandor
CRrashworthiness for Aerospace Structures and Hybrids (CRASH) Lab
Department of Mechanical and Aerospace Engineering
University at Buffalo – The State University of New York, Buffalo, NY
NASA’s Jet Propulsion Laboratory (JPL) has supported an effort to investigate concepts
for a Venus Sample Return Mission through an undergraduate team of students. An in-depth
study was conducted to determine the overall architecture of the mission and identify feasible
technologies that could be used to accomplish the objectives. Any proposed system was
required to collect and return specimens of regolith and atmospheric samples to Earth or near
Earth orbit. Regolith need to be obtained over a sufficient area to minimize measurement
biases due to their limited quantity. The lander also had to accommodate a minimum payload
of 50 kg while enduring the harsh Venusian surface environment for at least one Earth day.
This paper will detail the methodology proposed for the development of a feasible design
capable of meeting the prescribed mission requirements.
I. Nomenclature
EDL = Entry, Descent, and Landing CD = coefficient of drag
MLI = Multi-layer insulation ag = acceleration due to gravity [m/s2]
TPS = Thermal Protection System Q” = heat flux [W/m2]
V = velocity [m/s] RaL = Rayleigh number
m = mass [kg] Nu = Nusselt number
II. Introduction
The exploration of Venus is a mission that has been sought after for over half of a century. Being the closest planet
to the Earth in distance and size, there is substantial curiosity to what lies beneath its dense clouds. Starting in the
early 1960s, a number of former Soviet Union missions known as the Venera series were sent to answer these
questions. Of these probes, the longest that had ever survived on the surface lasted about two hours and managed to
capture only one image. Further missions were launched, which include Vega 1 and 2, Pioneer 12 and 13, Venus
Express, and Akatsuki. The data collected throughout these campaigns helped gain insight into the terrain, atmosphere,
and surface conditions [1]. While this information has been useful to the scientific community, no new images from
the surface were taken and none of the probes were able to last more than a couple of hours.
The harsh conditions of Venus provide a highly challenging environment for equipment to operate. To successfully
design a lander with sample collection capabilities, all of the following factors must be considered. From the
measurements collected in previous missions, it has been determined that the atmosphere of Venus consists of
approximately 97% CO2. This causes what scientists call a “Runaway Greenhouse Effect” [1], which thermally
insulates the planet so that the atmosphere prevents a significant amount of heat from leaving. Due to this condition,
Venus is the hottest planet in the solar system. These temperatures average around 467 (864 ) on the surface, but
decrease to about 60 at an altitude of 55 km. In addition to temperature, the atmosphere is extremely dense compared
to that of the Earth. The average density on the surface of Venus was measured to be 67 kg/m3 with a considerable air
pressure of 91 atm. To put this in perspective, the pressure of the air on the surface of Venus is equivalent to the
2
American Institute of Aeronautics and Astronautics
pressure felt a half mile beneath the ocean on Earth. When considering higher altitudes, the air pressure and density
reduces to near Earth conditions, but at the cost of high wind speeds. These winds reach velocities more than 200 m/s
in the upper atmosphere to produce what is known as a “super rotation” that only takes about four Earth days to fully
circle the planet. This is a large contrast to the slow body rotation of 243 Earth’s days. As altitude decreases the wind
speed drops to virtually nothing at the surface. A sulfuric acid layer in the atmosphere adds another extreme planetary
condition to consider. Taking up altitudes between 30 and 68 km, sulfuric acid is known for its highly corrosive nature
when in contact with many materials including metals and could be catastrophic without the proper precautions. From
prior missions it was also determined that the crust of Venus was mostly basalt based on what was found and analyzed
at the probe sites. All of these conditions combined make the cumulative environment that will be found on Venus.
For a successful mission, each subsystem element must be capable to withstand these challenges.
Quantitative performance objectives were developed so that a defined goal can be measured: (1) Return to Earth
at least 1 kg of Venus regolith from the top 10 cm of the surface. (2) Return to Earth at least one cubic liter of Venusian
atmosphere from the surface, and at least one cubic liter of atmosphere from an altitude of 40 km or greater. (3)
Accommodate at least 50 kg of scientific payload that can return data for at least one Earth day. (4) Take Venus
regolith samples over an area of at least 10 m2. (5) Collect Venus regolith samples vertically from the top 10 cm of the
surface.
The success of this mission would provide vast insight into the formulation of Venus and potentially other
terrestrial bodies in Solar System, with subsequent studies highlighting how the planetary evolution on Venus has
resulted in the extreme condition fostered at the surface.
III. Mission Architecture Design
A. Orbital Trajectory
The mission is designed to launch on May 14th, 2031, with a direct transit to Venus. This launch window was
selected not only to provide an optimized trajectory with minimum V between Earth and Venus, but to allow time
for maturation of projected technologies required for the development of an effective mission architecture. After 146
days in transit, the craft will arrive traveling at roughly 10.6 km/s to Low Venus Orbit(LVO). From there, a burn will
reduce the velocity of the craft, enabling entry into the Venusian’s gravitational field, transitioning away from orbiting
the Sun. At this point, the craft will enter into a highly elliptical orbit around Venus, where its full multiple orbital
periods would be aligned with the total time for the lander’s EDL, sample collection, and final ascension and re-
docking. This will reduce the effects of aerodynamic drag on the craft imposed by the Venusian atmosphere at
pericythe, allowing enough specific orbital energy to be maintained until the next launch window for return to Earth.
This should occur approximately 470 days after final transfer from LVO to the adjusted elliptical orbit. At the point
of return, the craft will exert another ΔV to transfer from its Venusian orbit to a heliocentric orbit on transit back to
Earth. The arrival date to Low Earth Orbit (LEO) will be on June 12th, 2033, where a final burn will allow the module
carrying Venusian samples to rendezvous with a LEO spacecraft or station for retrieval.
The specifics of the overall orbital trajectory discussed are listed below. The following assumptions were made
based upon the nature of the mission. Given the parameters and assumptions considered, the following ΔVs, shown in
Table 1, need be applied to achieve the required trajectory.
Assumptions:
1) Circular orbits around Earth
2) No solar radiation pressure
3) The orbital trajectories were Hohmann transfers with
instantaneous impulses
4) Initial orbital radius around Earth before transfer
ejection: 100 km
5) Final orbit radius around Earth after return: 400 km
6) 4-hour period of orbit with pericythe at 100 km
7) The pericythe was raised from 100 km to 500 km after
rendezvous
Table 1: ΔV requirements
Required Burn Delta V (km/s)
Earth To Heliocentric 3.52
Heliocentric to Venus -2.48
Raise Venus Perigee .266
Venus to Heliocentric 2.13
Heliocentric to Earth -3.46
3
American Institute of Aeronautics and Astronautics
B. Entry, Descent, and Landing (EDL)
For the initial entry and descent of the lander, a 45˚ sphere coned aeroshell measuring 4 m in diameter will be used
for the initial deceleration stage. It will have an entry path angle of approximately 20º and an entry velocity of 12 km/s.
The geometry selected was modeled after the Pioneer probe aeroshell because it proved to be both safe and reliable
during its mission when subjected to the extreme Venetian atmosphere. A developing three-dimensional woven TPS
design known as HEEET (Heat-Shield for Extreme Entry Environment) [5] will be used to reduce mass while
providing the required peak heat flux.
At 65 km a parachute will be deployed and the aeroshell will be detached away from the lander. The back plate
will remain as the connector between the lander and parachute as it further decelerates for another 15 minutes. Due to
the thick Sulfuric acid clouds stretched between 30 and 68 km altitude over Venus, the parachute must be composed
of materials that can withstand this corrosive environment. Testing conducted at JPL has shown that materials such as
Mylar covered in various fluoropolymers or Polybenzoxazole (PBO) [7, 8] have great promise for sulfuric acid
protection, and are currently in the process of being further developed. The parachute will also need to be large enough
such that the lander decelerates faster than the aeroshell, so that the aeroshell can detach and drop clear from the
lander.
𝑉 = 200 𝑚
𝑠 , 𝐶, = 1.0 , 𝑚 = 500 𝑘𝑔, 𝐷 = 4 𝑚
𝐶, = 1.5, 𝑚 = 1400 𝑘𝑔, 𝜌 = 0.5 𝑘𝑔
𝑚
𝐹 =1
2𝐶𝜌𝑉𝐴
𝐹
= 𝑚𝑎
In order to determine the effective size of the parachute, a force balance between the gravitational force and the
drag force must be computed on both the aeroshell and drag plate. The assumed atmospheric density and velocity at
this height is ρ = 0.50 kg/m3 and V = 200 m/s. From these values, the net deceleration of the aeroshell after separation
is a = 141.9 m/s2. Thus, the parachute must provide a deceleration greater than this to safely separate the aeroshell
from the lander. Assuming a deceleration of 150 m/s2, and a lander and parachute mass of 1400 kg, the effective
diameter of the parachute is calculated to be 8.4 m with an effective area of 55.6 m2.
Fig. 1: Entry and descent overview Fig. 2: Landing leg deployment
(Assumptions)
(1)
(2)
4
American Institute of Aeronautics and Astronautics
After 15 minutes of decelerating with the parachute, the lander will then detach. The remaining drag plate
continuing to slow down and steer the vehicle until landing. The reason for this is to annul the risk of the parachute
draping over the lander upon landing, making the lander inoperable. The drag plate is designed to orient the lander in
descent and slow it to a terminal velocity of 5.5 m/s. The drag plate design determination is discussed further at length
in Section IV.A. At 20 m above the surface the landing legs will be deployed along with the detachment for the drill
module covers. A crushable attenuator [3, 12] will be centered on the bottom of the craft to absorb initial impact upon
landing with the rest of the impact being distributed among the legs [14].
C. Sample Collection
Once successfully landed, the samples must be gathered from a distributed area of at least 10 m2. The lander will
deploy four drilling modules to increase the reach of the drills. Each module will collect three samples at variable
distances from the pressure vessel. Motors will drive lead screws that will provide mobility to the drill as it travels
along the distance of the module. Power will be sourced from primary batteries in the temperature controlled pressure
vessel, as discussed in detail in Section V.A.
The drilling device to be used is what is known as an Ultrasonic Drill [9,10]. It utilizes a piezoelectric wafer to
pulverize a ring into the regolith, taking advantage of high frequency electronic pulses. The device uses a coring
attachment that does not require sharpening which is optimal for sample collection. Studies have shown effective
operation under Venus-like conditions. For this reason, the drills will not need to be contained in a temperature
controlled environment. The drills also consume little power (less than 5 W each) and are extremely small in size and
mass. These qualities make it optimal for the proposed lander to operate multiple drills.
There are two main reasons for selecting four drilling modules for sample collection: one is the redundancy of the
drilling system. If two or even three drills were to become non-operational, there would still be a chance for a
successful mission, as the backup drill can still operate. The second and more important reason is time: to limit the
amount of risk associated with sample collection, it is necessary to collect samples in the shortest amount of time
possible. Since four drilling modules are taking samples simultaneously instead of one drill collecting all, the total
collection time will be drastically reduced to less than two hours, instead of the potential eight hours that it would take
for only one drill to carry out the task.
D. Ascent
The ascent will be a two phase portion of the mission. The initial stage will be done via a balloon in order to raise
the ascent vehicle above the dense lower atmosphere. On the Venusian surface, rocket propellants will be highly
volatile due to the extreme temperatures. Additionally, the amount of thrust required in the dense lower atmosphere
would be incredibly high due to aerodynamic drag. A balloon would provide a most efficient method of ascension in
the 67 kg/m3 dense surface atmosphere [7, 8]. The gas used for inflation will be helium, as it provides the most rise
due to its low molecular density while being an inert gas. The helium will be contained by tanks that would be left on
the surface in order to reduce mass for the return phase.
Fig. 4: Drilling module architecture
Fig. 3: Samples collection
5
American Institute of Aeronautics and Astronautics
A material with specific properties for the balloon must be chosen in order to ensure mission success. This material
must at least satisfy the following conditions:
1) Maintain mechanical properties in temperatures up to 500
2) Tear resistant and high tensile strength
3) Resistance to sulfuric acid
4) Little to no gas permeability
5) Flexible enough to resist high winds and dynamic changes
Finding a material that is able to satisfy all of the criteria above has proven to be a challenge. While many materials
may fit some of the requirements, being able to fulfill all criteria is a must. For example, Teflon was considered due
to the relatively high resistance to sulfuric acid, but has a melting point of less than 350 . The ideal material that
was found to be is PBO [8]. It has a high tensile strength, is durable and flexible enough, and can reach an acceptable
temperature before it decomposes. PBO has no melting point, instead it rapidly decomposes based upon its
temperature. At 500 , it has been shown to maintain its mechanical properties for over six hours at a time, long
enough to meet the success criteria for this phase of the mission. However, one downside is that it has a relatively
poor resistance to sulfuric acid. When exposed to sulfuric acid, samples of PBO have shown to plasticize and lose up
to 75% of their strength relatively quickly. This can be remedied with a coating of a fluoropolymer film or a metallic
film. Fluoropolymers have a relatively high corrosive resistance, but poor thermal resistance. The recommended
coating would be a thin layer of gold, as it has a high melting point (>1000 °C) and is very resistant to corrosion.
Lastly, in order to solve the issue of gas permeability, multiple layers of PBO are recommended with an adhesive in
order to prevent leaks of the internal gas.
The second phase of the ascent will begin where the balloon has reached its maximum altitude. Due to the change
in temperature, density, and pressure of the atmosphere, the balloon will either burst or no longer provide enough
buoyancy for lift. This occurs at an altitude of around 50 km, where the density of Venus’ atmosphere is approximately
the same of that of the Earth. The proposed method for phase 2 is a three stage radial rocket designed to propel the
samples through the remaining trajectory of the ascent to rendezvous with the orbiter. This design was originally
developed by a company known as Bloostar [6] which specializes in balloon assisted rockets. As shown in Fig. 6,
each stage will be embedded within the other as to provide a compact configuration for storing fuel and proper control
of the larger stage to get through the high winds. Once united with the orbiting craft, the samples will remain in orbit
until the next launch window, before beginning their journey back to Earth. The complete mission profile is depicted
in Fig 12.
Fig. 5: Transition to ascent vehicle Fig. 6: Bloostar radial rocket design
6
American Institute of Aeronautics and Astronautics
IV. Computational Analysis
A. Drag Plate Terminal Velocity
A crucial aspect of the drag plate design is to determine its terminal velocity such that it can slow down the lander
enough for a successful impact landing without catastrophic damage to any components. In order to calculate this
terminal velocity, Computational Fluid Dynamics (CFD) was used on the lander and drag plate. Assuming a mass of
1400 kg, the gravitational force of 12,418 N was calculated. ANSYS-Fluent was therefore used in a trial error fashion
to determine the flow velocity corresponding to the drag force of around 12,418 N. When the drag force equals the
gravitational force, the velocity used in the calculations will be the terminal velocity.
The assumed properties of the Venusian atmosphere (fluid flow) near the surface are ρ = 67 kg/m3, a temperature
of T = 735 K, and a viscosity of µ = 5.2e-06 m2/s. The simulation ran 60,000 iterations, ensuring that the force
calculations converged to a steady value. Figure 7 depicts the fluid flow around the lander and drag plate, with Fig. 8
displaying the static pressure contours on the lander. From the results, it can be seen that the seemingly random fluid
path lines around the body of the lander, seen in Fig. 7, are due to highly turbulent flow caused by the sharp edges of
both the lander base and octagonal body, resulting in abrupt directional change within the flow. The low-pressure
region within the top of the drag shield shown in Fig. 8, depicted by the blueish green region, accounts for the vortex
generation effect of the flow on top of the drag plate. The flow begins to converge back to the center of the lander
where the effects of the body are no longer present, and additionally the four cut-outs in the drag plate, which allow
regolith sample storage post-collection, also allow for air to flow through them, which accounts for the streamlines
within the drag plate.
From the analysis, and as it can be seen in Fig. 8, the total drag force on the plate is slightly greater than the
gravitational force, with FDrag = 12,911.169 N. From this, it can be concluded that the terminal velocity will be slightly
slower than 5.5 m/s, which will be safe for landing. The design utilizes a 3 m diameter drag plate with a 30° inclination.
Depending on impact needs, the drag plate size could be increased or decreased in order to be optimized. If a higher
impact velocity is feasible, then the plate diameter could be reduced, allowing for increased payload mass and size
savings. The same argument is valid for a slower impact velocity; however, a greater drag plate diameter would
decrease packing efficiency and increase the ΔV budget. Based on the information available from past Venus missions,
the drag plate dimensions are sufficient for a successful landing.
Fig. 7: Fluid flow during descent Fig. 8: Static pressure contour on lander
7
American Institute of Aeronautics and Astronautics
B. Atmospheric Sample Pressure Vessel
In order to properly house samples at a required atmospheric condition, the pressure vessel must be able to maintain
gaseous samples, taken from roughly 90 atm, or 9.116e+06 Pa, collected by a sealed capsule within the module with
internal vacuum conditions. To find a capsule material that meets these requirements, Finite Element Analysis (FEA)
in ANSYS was used. The design is a cylindrical pressure vessel, with an interior volume of 1 L, as defined by the
requirements. The exterior is defined to have a 2 cm thick titanium alloy shell. A pressure of 9.119+06 Pa is defined
on the interior, with no pressure on the outside, to simulate the worst-case scenario. Figure 9 shows the results from
the simulation.
From this analysis, it can be seen that the maximum stress
experience by the pressure vessel is 0.122 GPa. The titanium
alloy used for this is Ti-6Al-2Sn-4Zr-6Mo, which has a yield
strength of 1.10 GPa. With a factor of safety of N = 9.02, the
atmospheric sampling vessel will withstand the pressure
gradient with no plastic deformation, making it a feasible
material for the atmospheric sample pressure vessel. The
interior will likely be coated in Teflon, which holds up
exceptionally well to Sulfuric Acid concentrations. It should
be noted that a factor of safety of 9.02 is too high for this
design, meaning excess material is being brought. NASA
states that for metallic pressure vessels, a yield strength factor
of safety of 1.4 is necessary [13]. Based on this, the wall
thickness of the pressure vessel could be decreased, allowing
for a higher stress and lower factor of safety. A recommended
thickness to consider would be much closer to 1 cm of shell
thickness instead of 2 cm.
C. Lander Pressure Vessel Thermal Analysis
In order to fulfill the mission requirements, the lander must be able to operate in the extreme conditions on the
surface of Venus. With temperatures of roughly 460 °C and an atmospheric density of 67 kg/m3, the lander must be
designed in such a way that the interior stays cool enough for all components to operate. The interior temperature of
the pressure vessel was chosen to be 30 °C as it will house all temperature sensitive components. At this temperature,
the primary Lithium Thionyl Chloride (Li-SOCl2) batteries and all scientific payload will operate effectively. In order
to keep the payload pressure vessel cool, both MLI insulation
and a Stirling cooler will be used. In order to determine the
MLI thickness and Stirling cooler wattage, ANSYS steady-
state thermal analysis was used.
The pressure vessel modeled will contain 10 cm of MLI
insulation surrounding the vessel. From experimental results,
the thermal conductivity of MLI at 700 K was estimated to
be k = 0.048 W/mK. Thus, the body was modeled with an
assigned thermal conductivity of 0.048. Inside the pressure
vessel, the walls were assigned a temperature of 30 °C.
Convective heat transfer was assigned to the exterior of the
body, with an ambient temperature of 460 °C. The
convection heat transfer coefficient was found assuming free
convection of CO2 at 735 K. Equations (3) and (4) show the
steps taken to find the convection heat transfer coefficient.
Fig. 9: FEA on atmospheric sample
storage pressure vessel
Fig. 10: Pressure vessel thermal analysis
8
American Institute of Aeronautics and Astronautics
𝑇
= 735.15 °𝐾 , 𝑇
= 373.15°𝐾 , 𝑡 = 0.1 𝑚 , 𝐾 = 0.048 𝑤
𝑚𝐾
𝑔 = 8.87 𝑚
𝑠, 𝛽 = 1
𝑇
= .0014 𝐾 , 𝐿 = 1𝑚, 𝑣 = 40.1𝐸 − 6 𝑚
𝑠
𝛼 = 56.3𝐸 − 6
, 𝐶 = .125, 𝑛 = .333
𝑁𝑢=
𝐿
𝐾= 𝐶 𝑅𝑎
=𝐾
𝐿𝐶𝑅𝑎
𝑅𝑎= 𝑔 𝛽(𝑇
− 𝑇
)𝐿
𝑣 𝛼
𝑅𝑎= 1.94𝐸9
= 7.44 𝑊
𝑚𝐾
Using the convective heat transfer coefficient of h = 7.44 W/m2K, the analysis was run. The steady state thermal
simulation utilized 20 iterations to ensure the flux values converged. Figure 10 displays the results of this analysis,
and the average estimated heat flux of the pressure vessel was determined to be Q” = 188.04 W/m2.
The payload vessel has a total surface area of 12.067 m2. Utilizing this surface area, the overall heat transfer is
2,269.07 W. From this value, the Stirling cooler can be designed such that the heat removed by the cooler is equal to
the heat going into the vessel. It should be noted that heat generated from the electronics and batteries housed within
the pressure vessel (Qgen) were neglected. In the future, approximated Qgen values will be needed in order to better
represent the thermal analysis on the lander. Beyond the current analysis performed with atmospheric properties at the
surface, a transient thermal analysis will also need to be performed in order to simulate the heat transfer and interior
temperature during the decent phase. For this, accurate data on temperature versus altitude would be needed as well
as time spent at those altitudes during decent.
V. Lander Subsystems
A. Power
For the lander, two electrical power options were weighed to determine the most feasible option: Li-SOCl2 primary
batteries, and General Purpose Heat Source (GPHS) [2] units in conjunction with a thermal electric generator.
Li-SOCl2 have a high specific energy (~400 Wh/kg) and a proven track record (Mars Sojourner Rover, and New
Millennium Deep Space-2), making it an excellent candidate for primary batteries. The other option considered, GPHS
units with a thermoelectric generator, was chosen due to the nature of the lander’s thermal system. The active thermal
control system on the lander could consist of GPHS units powering a Stirling cooler, and additional GPHS units could
be used to provide electrical power as well. Based on previous studies, the thermal to electrical efficiency of the
thermoelectric generator was assumed to be 38%, and each GPHS houses Plutonium-238, producing 250 W of thermal
power and weighing 1.5 kg.
Due to the high cost and lifespan of Plutonium-238 GPHS modules to power the lander, Li-SOCl2 batteries were
chosen instead, even though their associated mass is greater. Based on the power requirements, the lander will contain
30 kg of Li-SOCl2 primary batteries for EDL, ground, and ascent operations. The batteries will be stored in an insulated
pressure vessel, which also contains all scientific payload and electrical systems, that is thermally controlled by the
Stirling cooler at a constant temperature of 30 °C.
(3)
(Assumptions)
(4)
9
American Institute of Aeronautics and Astronautics
B. Propulsion
Propulsion will be important for the success of this mission and will be used during multiple phases of the mission.
The main use of propulsion will be for the orbital transfers that must be accomplished in order to both go to and return
from Venus. Once the rocket is in orbit around Earth, a second stage will be used to transfer to a heliocentric orbit
until the rocket arrives at Venus, where it will then decelerate and enter into an orbit around the target planet by
aerobraking through the Venusian atmosphere. On return, following the mission success and rendezvous with the
ascending vehicle, the orbiting craft will undergo another burn to reverse this process. It will subsequently transfer
from orbiting Venus to a heliocentric orbit until reaching Earth. Besides the previously noted orbital maneuvers,
propellant will be used to ascend after the balloon has reached maximum altitude. Due to the nature of the mission,
this will be vital in order to escape the Venusian atmosphere and rendezvous with the orbiter.
To determine which propellant can be used, multiple aspects need to be taken into consideration. The specific
impulse of the rocket is the most important design factor, due to the requirement for efficiency on this mission.
However, other key aspects also need to be taken into account. For example, fuels that need to be super cooled cannot
be considered for the ascent portion of the mission due to the challenge of the ambient temperatures on Venus being
over 460 . The propellant used for orbital trajectory will be dictated by the manufacturer of the main rocket, but for
efficiency purposes a bipropellant of either liquid oxygen (LOx) and Hydrogen, or LOx and methane would be
recommended. What is of more concern to this mission is the propellant that will be used in the ascent vehicle. The
propellant has to be efficient, controllable, and able to survive when contained on the surface of Venus.
The two recommended propellants for the ascent vehicle would be hydrazine and LMP-103S [11], both
monopropellants. Monopropellants were chosen over solid propellants and bipropellants for a few reasons. Solid
propellants were not considered due to the fact that they cannot be controlled as needed. Venus’s high winds in the
upper atmosphere reaching speeds of up to 100 m/s, therefore maintaining proper attitude during the ascent phase will
be critical. This requires that there will be a gimballed thrust mechanism as well as a controlled propulsion system;
solid propellant rockets do not offer the latter. Bipropellants, while controllable, require a temperature controlled
environment, which may be infeasible on the surface of Venus. Most bipropellant oxidizers require extensive cooling
(use of Lox or dinitrogen tetroxide), while others are classified as unstable to highly unstable (liquid fluorine). Due to
these reasons, a monopropellant is proposed, which produces a relatively high specific impulse while also being
hypergolic.
Hydrazine has been the industry standard for monopropellants over the past few decades. However, LMP-103S
which is relatively new monopropellant looks to be promising, having about a 6% higher specific impulse and about
24% higher density. In addition, LMP-103S is significantly more stable and less toxic. LMP-103S is yet to be ready
to fly, and will ideally be tested and proven worthy by the time the Venus Sample Return Mission is finalized. LMP-
103S requires a catalyst, for which the materials are still being tested. Further, it is generally required to be heated up
to 350 °C, which should not be a problem due to the high temperatures at the surface of Venus. By exposing the
catalyst to high environmental temperatures, the catalyst should be able to reach the desired temperature without
spending energy from the batteries to be heated. One downside of LMP-103S is the significantly higher burning
temperature, which will require heavier and more expensive materials for the nozzle. Due to the increased efficiency
though, LMP-103S is preferred over hydrazine.
10
American Institute of Aeronautics and Astronautics
C. Active Cooling System
Based on the required heat transfer of 2,269.07 W to maintain an interior temperature of 30 °C, the Stirling cooler
system [1,3] must be designed to remove the same amount of heat. The lander will utilize a two stage Stirling cooler.
A linear, piston style Stirling cooler works by moving a working fluid between hot and cold end spaces. The working
fluid is compressed in the hot end space, which causes it to release heat to the environment due to the temperature
increase of the fluid from compression. The piston continues to compress the working fluid until it is pushed into the
cold end space. The fluid is then expanded in the cold end space, causing it to absorb energy from the cold space. The
piston then pushes the fluid back into the hot end space where the heat is then emitted, and the process begins again.
For the lander design, the working fluid will be helium, which is generally the working fluid used in piston style
Stirling coolers.
Fig. 11: Two Stage Stirling Cooler [1]
In the arrangement displayed in Fig. 11, the purpose of having two embedded stages is illustrated. The first
stage cooler (inner vessel) will push out heat generated by electronics and incoming heat leak to the outer pressure
vessel. The inner vessel will maintain a temperature of 30 ºC. The second stage cooler (outer vessel) will push the
heat exchanged from the first stage cooler and incoming heat leak from the ambient Venusian environment to the
atmosphere outside of the lander. With the Stirling Cooler and MLI insulation, the outer vessel will maintain a
temperature of 250 ºC. By cooling the pressure vessel with stages in this manner, the required operations temperature
for the scientific payload can be maintained within the power limitations of primary batteries. This technology is
promising, however, future research and testing will be required before it is flight ready. [1,3]
Stage 2
Cold Sink Temperature 250 ºC
Hot Sink Temperature 500 ºC
Total
Heat Removed 2,270 W
Heat Rejected 6,721 W
Stage 1
Cold Sink Temperature 30 ºC
Hot Sink Temperature 250 ºC
Table 2: Stirling Cooler stages breakdown
11
American Institute of Aeronautics and Astronautics
Conclusion
The mission to return physical samples from Venus has been desired by scientists for decades. Being named
one of the top priorities on NASA’s Decadal Survey 2013-2022, there is hope to find valuable insight if a mission
were pursued. The largest obstacle for this mission however is the harsh environment on the Venusian surface.
Averaging temperatures of 467 °C and an atmospheric pressures 90 times that of the Earth, making it extremely
difficult for equipment to effectively operate. To overcome this, cutting-edge technologies must be pursued and
utilized. Considering the research presented herein as the basis, much of the selected technology to enable the mission
is still in development.
Two key technologies features in the proposed mission architecture are the Stirling cooler and LiSOCl2
primary batteries. A series of Stirling coolers will be used to provide the needed temperature reduction in which the
scientific payload will operate, and an MLI shell to prevent thermal leakage. Using LiSOCl2 primary batteries as the
main power source, both the Stirling coolers and the scientific payload will have sufficient operating power to last for
the duration of the mission. Alternatives include using combustion with Lithium to produce thermal energy that can
power the Stirling coolers; however, further research in this area will be required.
The ultrasonic drilling system selected for sample collection is an engineering marvel that has received high
praises from the research community. Its light weight-low power design is capable of operating in extremely high
temperatures and can drill through hard terrain such as basalt based rocks. As a result, this drilling system constitutes
an optimal choice for the Venus Sample Return Mission and allows for more engineering freedom in drill mobility
due to its manageable size. The sample collection architecture utilizes four drilling modules that each move an
ultrasonic drill along the length of the module. A single drilling module is capable of taking three samples in a
calculated timeframe of under two hours until full execution. Using four modules minimizes both time and risk due
to their operational redundancy. Once collected, the samples will be elevated up to the ascent vehicle. The ascent
vehicle will then be lifted off the Venusian surface through utilizing an inflated balloon.
The balloon will use helium that will provide the required buoyancy for the initial ascent phase to 50 km.
Once the altitude of 50 km is reached, to escape the Venusian gravity, the balloon will be released and the ascent
vehicle will transition to a three stage radial rocket for the final ascent phase. The selected balloon ascent structure
will use PBO with a gold foil outer layer as the balloon material of choice due to its high tensile strength at high
temperature and acid resistivity. Bloostar’s radial rocket configuration is recommended once the balloon is released
due to its compact propellant storage structure. LMP-103S is the propellant of choice as it is a monopropellant with
high specific energy.
In addition to the selected technologies, the orbital trajectory and architecture for the mission needs to be
developed further. A major complication to be addressed in this mission is the amount of fuel required to get to Venus
and provide a return journey to Earth for samples. Key developments will include fuel efficient series of transfers, or
a larger rocket for expanded fuel storage. Alternatives, to help with propellant requirement, include reducing the mass
of the aeroshield with HEEET, which is currently in the testing phase at NASA. A lighter overall design will ultimately
lead to reduced amount of propellant needed for the mission.
Overall, the Venus Sample Return Mission design proposed provides a feasible strategy, using the available
or near term technology, to accomplish all of the objectives set for such a mission.
12
American Institute of Aeronautics and Astronautics
Acknowledgments
Authors 1 through 4 would like to thank Drs. Javid Bayandor and Steven Matousek for providing this unique
opportunity to work on future NASA exploration goals. Their motivational and technical input helped us aim for high
research standards. Dr. Bayandor has a long history of inspiring greatness in researchers and we are thankful to have
had him guiding the team through the obstacles that this project presented. In addition, Dr. Matousek from NASA’s
Jet Propulsion Laboratory has provided unparalleled guidance and insight in maintaining progress. Without the help
from these mentors we could have not met the objectives of this priority exploration mission, and gained an invaluable
experience from this highly challenging, yet exciting endeavor.
Fig. 12: Venus Sample Return Mission Profile
13
American Institute of Aeronautics and Astronautics
References
[1] Hall, J. L., Senske, D. A., Grammier, R., Bullock, M., Cutts, J. A., “Venus Flagship Mission Study,” JPL NMO710851, Jet
Propulsion Laboratory, Pasadena, California, April 2009.
[2] Kolawa, E., Balint, T., Birur, G., Bolotin, G., Castillo, L. D., Garret, H., Hall, J. L., Johnson, M., Jones, J., Jun, I., Manvi, R.,
Mojarradi, M., Moussessian, A., Patel, J., Pauken, M., Peterson, C., Surampudi, R., Schone, H., Whitacre, J., Matinez, E., Laub,
B., Venkapathy, R., Neudeck, P., “Extreme Environment Technologies for Future Space Science Missions,” JPL D-32832, Jet
Propulsion Laboratory, Pasadena, California. Sept. 2007.
[3] Oleson, S. R., McGuire, M. L., Sandifer II, C. E., Balkanyi, L., Michael, B., Burke, L., Colozza, T., Dankanich, J., Drexler, J.,
Fittje, J., Gyekenyesi, J., Landis, G., Martini, M., Martini, M., Packard, T., Rodriguez, C., Schmitz, P., Sheehe, C., Tenteris,
A., Warner, J., Williams, G., “Advanced Long-life Lander Investigating the Venus Environment (ALIVE),” NASA TM-2018-
219417, NASA Glenn Research Center, Cleveland, Ohio, Jan. 2018.
[4] Bate, R., Mueller, D. D., and White, J. E., Fundamentals of Astrodynamics, Mineola, New York: Dover Publications, June
1971.
[5] Venkatapathy, E., Ellerby, D., Stackpoole, M., Peterson, K., Gage, P., Beerman, A., Gasch, M., Munk, M., Prabhu D., and
Poteet, C., “Heat-shield for Extreme Entry Environment (HEEET); A Game Changing Thermal Protection System for Saturn
and Uranus Probe Missions,” Outer Planets Assessment Group, Jan. 2014.
[6] Schoenmaker, A., Lopez-Urdiales, J. M., and Gonzalez, J. C. C. “Bloostar, The Shortcut to Orbit,” Space Transportation
Solutions and Innovations Symposium, IAC-15-D2.7, Barcelona, Spain., Oct. 2015.
[7] Hall, J. L., “Venus Balloons for High Altitude”, Venus Upper Atmosphere Investigations Science and Technology Interchange
Meeting, Venus Exploration Analysis Group, Cleveland, Ohio, Jan. 2013.
[8] Yavrouian, A., Yen, S. P. S., Plett, G., and Weissman, N., “High Temperature Materials for Venus Balloon Envelopes,” 11th
Lighter-than-Air Systems Technology Conference, AIAA-95-1617-CP, Aug. 2012.
[9] Bar-Cohen, Y., Sherrit, S., Boa, X., Badescu, M., and Chang, Z., “Ultrasonic/Sonic Drilling/Coring (USDC) as a Subsurface
Drill Sampler and Lab-On-A-Drill for Planetary Exploration Applications,” The International Society for Optical Engineering
Smart Structures and Materials Conference, Vol. 5762-22, San Diego, California, March 2005.
[10] Bar-Cohen, Y., Sherrit, S., Dolgin, B., Peterson, T. M., Pal, D., Kroh, J., Krahe, R., “Smart-ultrasonic/sonic driller/corer,” US
Patent No. 6,863,136, US Patent and Trademark Office, Washington D.C., March 2005.
[11] Negri, M., and Grund, L., “Replacement of Hydrazine: Overview and First Results of the H2020 Project Rheform”, 6th
European Conference for Aeronautics and Space Science, Germany, 2015.
[12] Perino, S. V., Bayandor, J., Samareh, J. A., and Armand, S. C., “Contemporary Impact Analysis Methodology for Planetary
Sample Return Missions,” AIAA Journal of Spacecraft and Rockets, Vol. 52, No. 4, 2015, pp. 1217-1227.
[13] “Structural Design and Test Factors of Safety for Spaceflight Hardware,” NASA–STD-5001B, NASA HQ, Washington, D.C.,
Aug. 2014.
[14] Schroeder, K., Bayandor, J., and Samareh, J., “Sizing and synthesis of Venera-class landers,” AIAA Journal of Spacecraft and
Rockets, Vol. 55, No. 3, pp. 561-574, 2018. doi: 10.2514/1.A33912
... Previous Venus sample return concepts focused on surface sample return, with study reports as early as the mid-1980s [143][144][145][146][147]. A more recent review of past activities is in [148]. ...
Preprint
Full-text available
The Venus Life Finder Missions are a series of focused astrobiology mission concepts to search for habitability, signs of life, and life itself in the Venus atmosphere. While people have speculated on life in the Venus clouds for decades, we are now able to act with cost-effective and highly-focused missions. A major motivation are unexplained atmospheric chemical anomalies, including the "mysterious UV-absorber", tens of ppm O$_2$, SO$_2$ and H$_2$O vertical abundance profiles, the possible presence of PH$_3$ and NH$_3$, and the unknown composition of Mode 3 cloud particles. These anomalies, which have lingered for decades, might be tied to habitability and life's activities or be indicative of unknown chemistry itself worth exploring. Our proposed series of VLF missions aim to study Venus' cloud particles and to continue where the pioneering in situ probe missions from nearly four decades ago left off. The world is poised on the brink of a revolution in space science. Our goal is not to supplant any other efforts but to take advantage of an opportunity for high-risk, high-reward science, which stands to possibly answer one of the greatest scientific mysteries of all, and in the process pioneer a new model of private/public partnership in space exploration.
Article
An in-depth investigation of the structural design of the Venera landers was explored. A reverse engineering of the Venera-class lander was performed. The lander was broken down into its fundamental components and analyzed. This provided insight into the structural components of the lander and highlighted the mass drivers of the design. A trade study was performed to find the sensitivity of the lander’s overall mass to several key parameters. A multi-fidelity design tool, used for further investigation of the parameterized lander, was developed. The low-fidelity model was a nonlinear model based on geometric constraints developed to predict the mass of each design rapidly, whereas the medium- and high-fidelity models used an explicit finite element framework to verify the low-fidelity predictions. This methodology allowed for a large variety of designs to be investigated, allowing an optimal configuration to be found for a given payload mass. Key features identified for the design of robust landers will serve as foundations for the development of the next generation of landers for future exploration missions to Venus. Results from this paper represent a benchmark of the current state-of-the-art for the Venus In-Situ Explorers (VISE) mission.
Conference Paper
The goal of the EU Horizon2020 project Rheform is the replacement of hydrazine with liquid propellants based on ammonium dinitramide (ADN) for orbital and launcher propulsion systems. Hydrazine and its derivatives are the standard propellants for spacecraft propulsion system since the 1960s, but they are highly toxic and carcinogenic. New regulations will lead to restriction of their use in the near to mid-term. The first part of this article gives an overview on ADN-based propellants and of the Rheform project. The second part contains the results of thermochemical calculations showing the influence of propellant formulation on performance and combustion chamber temperatures.
Article
Development of an Earth entry vehicle and the methodology created to evaluate the vehicle's impact landing response when returning to Earth is reported. NASA's future Mars Sample Return Mission requires a robust vehicle to return Martian samples back to Earth for analysis. The Earth entry vehicle is a proposed solution to this Mars mission requirement. During Earth reentry, the vehicle slows within the atmosphere and then impacts the ground at its terminal velocity. To protect the Martian samples, a spherical energy absorber called an impact sphere is under development. The impact sphere is composed of hybrid composite and crushable foam elements that endure large plastic deformations during impact and cause a highly nonlinear vehicle response. The developed analysis methodology captures a range of complex structural interactions and much of the failure physics that occurs during impact. Numerical models were created and benchmarked against experimental tests conducted at NASA Langley Research Center. The postimpact structural damage assessment showed close correlation between simulation predictions and experimental results. Acceleration, velocity, displacement, damage modes, and failure mechanisms were all effectively captured. These investigations demonstrate that the Earth entry vehicle has great potential in facilitating future sample return missions.
Venus Flagship Mission Study
  • J L Hall
  • D A Senske
  • R Grammier
  • M Bullock
  • J A Cutts
Hall, J. L., Senske, D. A., Grammier, R., Bullock, M., Cutts, J. A., "Venus Flagship Mission Study," JPL NMO710851, Jet Propulsion Laboratory, Pasadena, California, April 2009.
Extreme Environment Technologies for Future Space Science Missions
  • E Kolawa
  • T Balint
  • G Birur
  • G Bolotin
  • L D Castillo
  • H Garret
  • J L Hall
  • M Johnson
  • J Jones
  • I Jun
  • R Manvi
  • M Mojarradi
  • A Moussessian
  • J Patel
  • M Pauken
  • C Peterson
  • R Surampudi
  • H Schone
  • J Whitacre
  • E Matinez
  • B Laub
  • R Venkapathy
  • P Neudeck
Kolawa, E., Balint, T., Birur, G., Bolotin, G., Castillo, L. D., Garret, H., Hall, J. L., Johnson, M., Jones, J., Jun, I., Manvi, R., Mojarradi, M., Moussessian, A., Patel, J., Pauken, M., Peterson, C., Surampudi, R., Schone, H., Whitacre, J., Matinez, E., Laub, B., Venkapathy, R., Neudeck, P., "Extreme Environment Technologies for Future Space Science Missions," JPL D-32832, Jet Propulsion Laboratory, Pasadena, California. Sept. 2007.
  • S R Oleson
  • M L Mcguire
  • I I Sandifer
  • C E Balkanyi
  • L Michael
  • B Burke
  • L Colozza
  • T Dankanich
  • J Drexler
  • J Fittje
  • J Gyekenyesi
  • J Landis
  • G Martini
  • M Martini
  • M Packard
  • T Rodriguez
  • C Schmitz
  • P Sheehe
  • C Tenteris
  • A Warner
  • J Williams
Oleson, S. R., McGuire, M. L., Sandifer II, C. E., Balkanyi, L., Michael, B., Burke, L., Colozza, T., Dankanich, J., Drexler, J., Fittje, J., Gyekenyesi, J., Landis, G., Martini, M., Martini, M., Packard, T., Rodriguez, C., Schmitz, P., Sheehe, C., Tenteris, A., Warner, J., Williams, G., "Advanced Long-life Lander Investigating the Venus Environment (ALIVE)," NASA TM-2018-219417, NASA Glenn Research Center, Cleveland, Ohio, Jan. 2018.
Bloostar, The Shortcut to Orbit
  • A Schoenmaker
  • J M Lopez-Urdiales
  • J C Gonzalez
Schoenmaker, A., Lopez-Urdiales, J. M., and Gonzalez, J. C. C. "Bloostar, The Shortcut to Orbit," Space Transportation Solutions and Innovations Symposium, IAC-15-D2.7, Barcelona, Spain., Oct. 2015.
Venus Balloons for High Altitude
  • J L Hall
Hall, J. L., "Venus Balloons for High Altitude", Venus Upper Atmosphere Investigations Science and Technology Interchange Meeting, Venus Exploration Analysis Group, Cleveland, Ohio, Jan. 2013.
High Temperature Materials for Venus Balloon Envelopes
  • A Yavrouian
  • S P S Yen
  • G Plett
  • N Weissman
Yavrouian, A., Yen, S. P. S., Plett, G., and Weissman, N., "High Temperature Materials for Venus Balloon Envelopes," 11 th Lighter-than-Air Systems Technology Conference, AIAA-95-1617-CP, Aug. 2012.
Ultrasonic/Sonic Drilling/Coring (USDC) as a Subsurface Drill Sampler and Lab-On-A-Drill for Planetary Exploration Applications
  • Y Bar-Cohen
  • S Sherrit
  • X Boa
  • M Badescu
Bar-Cohen, Y., Sherrit, S., Boa, X., Badescu, M., and Chang, Z., "Ultrasonic/Sonic Drilling/Coring (USDC) as a Subsurface Drill Sampler and Lab-On-A-Drill for Planetary Exploration Applications," The International Society for Optical Engineering Smart Structures and Materials Conference, Vol. 5762-22, San Diego, California, March 2005.