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Abstract and Figures

Small satellites may provide a low-cost platform for targeted science investigations in the Mars system. With current technology, small satellites require ride shares with larger orbiters to capture into orbit, limiting the range of orbits available to small satellite mission designers. Successful development of a small satellite aerocapture capability would allow small satellite mission designers to choose the orbit most appropriate for a science investigation while enabling small satellite ride shares on any mission to Mars. A generic small satellite aerocapture system is assessed for use at Mars across a range of small satellite payloads, approach trajectories, and destinations in the Mars system. The aerocapture system uses drag modulation for trajectory control to ensure successful orbit insertion. Analyses include assessment of the sensitivity of the entry corridor size to the ballistic-coefficient ratio, the effectiveness of real-time aerocapture guidance and control algorithms, aerocapture system-level impacts of different target orbits, and development of requirements and recommendations for the development of a small satellite aerocapture system. Results indicate that a discrete drag-modulation aerocapture system may provide an orbit-insertion capability for small satellites with modest propulsion requirements.
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(Preprint) AAS 18-052
G. Falcone
, J. W. Williams
, and Z. R. Putnam
Small satellites may provide a low-cost platform for targeted science investigations
in the Mars system. With current technology, small satellites require ride shares
with larger orbiters to capture into orbit, limiting the range of orbits available to
small satellite mission designers. Successful development of a small satellite ae-
rocapture capability would allow small satellite mission designers to choose the
orbit most appropriate for a science investigation while enabling small satellite
ride shares on any mission to Mars. A generic small satellite aerocapture system
is assessed for use at Mars across a range of small satellite payloads, approach tra-
jectories, and destinations in the Mars system. The aerocapture system uses drag
modulation for trajectory control to ensure successful orbit insertion. Analyses
include assessment of the sensitivity of the entry corridor size to the ballistic-
coefficient ratio, the effectiveness of real-time aerocapture guidance and control
algorithms, aerocapture system-level impacts of different target orbits, and de-
velopment of requirements and recommendations for the development of a small
satellite aerocapture system. Results indicate that a discrete drag-modulation ae-
rocapture system may provide an orbit-insertion capability for small satellites with
modest propulsion requirements.
βBallistic coefficient
v Change in velocity
γFlight-path angle
ESPA EELV Secondary Payload Adapter
NPC Numeric predictor-corrector
RPM Rotations per minute
Recent interest in Mars has led to an increasing number of large spacecraft en-route to Mars
that may provide rideshares opportunities for small satellites (smallsats). However, current smallsat
high-thrust propulsion technology limits the range of target orbits for rideshare smallsats to within
Ph.D. Student, University of Illinois at Urbana-Champaign (
Graduate Student, University of Illinois at Urbana-Champaign (
Assistant Professor, University of Illinois at Urbana-Champaign (
Table 1. Orbit insertion via rideshare at Mars limits smallsats to near-polar orbits or landing on the
Mission Launch year Inclination Semimajor
Axis Eccentricity
Mars Global Surveyor 1996 93 deg 3708 km 0.88
Mars Pathfinder 1996 Surface
2001 Mars Odyssey 2001 93 deg 3788 km 0.0064
Mars Express 2003 86.7 deg 8593 km 0.57
MER Spirit 2003 Surface
MER Opportunity 2003 Surface
Mars Reconnaissance Orbiter 2005 92.6 deg 3659 km 0.01
Phoenix 2007 Surface
Mars Science Laboratory 2011 Surface
Mars Orbiter Mission 2013 150 deg 39200 km 0.91
MAVEN 2013 74.2 deg 6655 km 0.47
ExoMars Trace Gas Orbiter 2016 7.8 deg 51300 km 0.93
a small deviation of the orbits of their respective host spacecraft. Smallsat mission options are
further limited by the fact that most planetary orbiters insert into near-polar orbits to provide global
coverage of the planet below. For example, with one exception, all spacecraft sent to Mars in the
recent past have either landed on the surface or entered high-inclination orbits at Mars as shown in
Table 1. For smallsats using rideshare opportunities, the large v associated with plane changes
makes inclination change from the host spacecraft’s target orbit difficult: chemical systems require
significant mass and volume, problematic for a smallsat; electric propulsion systems require long-
duration maneuvers to change orbits, on the order of years, limiting a smallsat’s ability to use off-
the-shelf components due to reliability concerns on long, deep-space missions. Development of an
independent smallsat orbit insertion capability would enable smallsat missions to address significant
planetary science objectives at Mars and its moons which may otherwise be ignored by larger, more
risk-adverse missions.
A smallsat orbit insertion capability would enable future smallsat mission designers to determine
(and obtain) the orbit best suited to achieving their mission goals and objectives, enabling low-cost,
independent science-, technology-, and infrastructure-driven missions to be hosted by any planetary
mission departing Earth. Smallsat missions enabled by an independent orbit insertion capability
include, but are not limited to:
missions to Mars’s moons Phobos and Deimos, which require a near-equatorial orbit
highly elliptic orbits to enable in situ study of the upper atmosphere
constellations of multiple smallsats in different orbits
smallsat technology demonstrations and infrastructure accretion, including geostationary com-
munications relays, simultaneous global coverage observation platforms, and high-altitude
maximum coverage remote sensing platforms for weather observation.
The current method for orbital insertion is to perform a large delta-V, high-thrust propulsive
maneuver. This method requires a considerable amount of propellant mass. Moreover, no high-
ΔV at periapsis to
correct for
apoapsis error
ΔV at apoapsis to
raise periapsis
target altitude
interface (AI)
Figure 1. Aerocapture Phases in Planet Reference Frame1
thrust chemical propulsion system with sufficient delta-V capability exists for smallsats. Even if
one did, the propellant mass fraction required for a propulsive insertion would exacerbate packaging
problems for smallsat form factors.
Another method currently used is a propulsive insertion combined with aerobraking: the propul-
sive maneuver places the craft on a highly-elliptic orbit. Then aerobraking is performed through
many passes through the atmosphere which slow down the vehicle, decreasing its semi-major axis
to the desired one. This method reduces the amount of fuel required for orbital insertion compared
to the direct insertion method. However, the insertion maneuver still remains costly in terms of
propellant budget. Furthermore, the vehicle has to perform many passes through the atmosphere
of the planet over a period of months, an operationally intensive and relatively expensive method,
especially for smallsats.
In contrast, an aerocapture maneuver uses a single pass through the atmosphere of the target body
to affect the same v with minimal propellant requirements (Figure 1). Although aerocapture has
not been performed to date, many of the required technologies have been matured through use on
planetary entry systems (e.g., heatshield technology, precision guidance and navigation). Due to
the aforementioned propulsive limitations, any rideshare mission to an interplanetary target which
uses existing technology must have a final desired orbit closely matching that of the host mission.
Through the use of B-plane targeting, aerocapture increases the number of available rideshare op-
portunities for a given target orbit. In the Mars system, aerocapture enables a smallsat mission to
enter orbits including a rendezvous or flyby of Phobos or Deimos, a Mars-synchronous orbit, or a
circular orbit with an inclination markedly different from that of the host vehicle’s.
Figure 2a shows the mass fraction vs. delta-V curves for propulsive and aerocapture insertion
methods, while Figure 2b shows the relative mass and cost savings of using aerocapture over propul-
sive insertion for various interplanetary targets. From this figure, it is apparent that aerocapture has
the potential to provide an increased payload mass fraction relative to propulsive insertion as the
required v increases, as well as large mass and cost savings in general.
Aerocapture requires aerodynamic control to ensure the vehicle exits the atmosphere with the
proper energy. If too much energy is depleted, the vehicle will either fail to exit the atmosphere or
Figure 2. Orbit insertion with aerocapture provides significant cost and mass savings
at many planetary destinations of interest.2,3
undershoot the target apoapsis. Conversely, if too little energy is depleted, the vehicle will either
continue on a hyperbolic trajectory out of the planetary system or overshoot the apoapsis target.
The majority of aerocapture system studies to date have assumed trajectory control during the at-
mospheric pass is achieved through lift-modulation via bank-angle steering.
A simpler approach to aerocapture trajectory control is drag modulation, which provides suffi-
cient control authority for aerocapture by changing only the drag area during flight.1Drag modu-
lation enables flight at zero angle of attack with spin stabilization, eliminating the need for active
attitude control during the atmospheric pass. A drag modulation system may be further simplified
by utilizing only a single control event: A one-time jettison or retraction of a pre-deployed drag
skirt. Despite this simplicity, previous analyses have shown that discrete drag modulation systems
are capable of achieving accuracy competitive with that of lift-modulation systems.1, 4, 5 Addition-
ally, drag-modulation systems typically have low ballistic coefficients, resulting in a more benign
aerothermal environment than that experienced by lifting aerocapture concepts, enabling drag mod-
ulation systems to utilize less expensive and lighter weight thermal protection system materials.
The simplicity and efficiency in cost, mass, and volume of drag modulation aerocapture makes
it an ideal candidate for inserting smallsats into orbit about planetary bodies of interest. Like all
aerocapture systems, drag modulation systems have only modest propulsive v requirements that
can be satisfied with currently available smallsat propulsion systems. Recent developments in de-
ployable aeroshell technology, such as the Adaptable, Deployable Entry and Placement Technology
(ADEPT)6enable relatively large drag skirts to be used for smallsats while maintaining compatibil-
ity with current rideshare opportunities such as EELV Secondary Payload Adapter (ESPA) rings.
Beyond enabling smallsat missions to be driven by their own mission objectives and enabling
new missions, an independent orbit insertion capability reduces the burden of ridesharing on both
the smallsat and the host spacecraft. A deep-space planetary smallsat with an orbit insertion system
may be attached to the interplanetary trajectory injection stage via an ESPA ring or similar,7then
proceed to the planetary destination as a free-flier, eliminating most of the interface between the
host spacecraft and smallsat and freeing the host from having to insert additional mass into orbit
at arrival. Such a system may make free-flier deep space smallsats more attractive to future host
mission program management.
The goal of this paper is to assess smallsat drag-modulation aerocapture system options consistent
with an ESPA-class rideshare opportunity and determine benefits and costs relative to current tech-
nologies. A characterization of the main advantages of drag modulation aerocapture with respect
to orbit insertion using conventional approaches is presented. The influence of several variables on
the drag-modulation aerocapture maneuver is assessed, including the arrival velocity of the vehicle,
the separation time of the drag-skirt with the main bus, and the target orbit. The performance of the
aerocapture maneuver is assessed in presence of expected day-of-flight uncertainties, and a post-
aerocapture maneuver plan is developed to limit the effect of uncertainty. Overall, results indicate
that drag-modulation aerocapture may provide a cost-effective solution for smallsat aerocapture at
Mars, although significant propulsive capability must still be provided to reach many destinations
of interest.
AI state:
high-energy orbit
Atmospheric exit state:
lower-energy orbit
drag skirt
drag area
Low-β decel.
(β = β1)
High-β decel.
(β = β2)
Figure 3. Drag-modulation Aerocapture Phases1
Drag-modulation aerocapture takes advantage of discrete changes in the drag area of a spacecraft
through its flight in an atmosphere to achieve orbit insertion with modest required propellant mass.
The system uses a single control event to control the vehicle energy accurately targeting the desired
apoapsis altitude.
The maneuver is divided into several phases, as shown in Figure 1 and Figure 3. Initially, the
smallsat aerocapture vehicle separates from the host spacecraft and approaches Mars on a hyperbolic
trajectory. The smallsat aerocapture vehicle is composed of the spacecraft and the drag skirt. From
this phase until the atmospheric pass, the vehicle is controlled using spin stabilization at around
2 RPM. The vehicle will then enter the atmosphere of Mars with a defined flight path angle, γ.
Subsequently, the drag skirt is jettisoned; the timing and location of jettison depend on the post-
aerocapture target orbit and onboard navigation. The ballistic coefficient of the vehicle will make
an almost immediate jump to a higher value. The ratio of pre- and post-jettison ballistic coefficient
for this study was fixed at 9, corresponding to a pre-jettison diameter of 1.5 m and a post-jettison
diameter of 0.5 m, consistent with a vehicle that fits in an ESPA-class payload slot Previous studies
have indicated this ballistic coefficient ratio provides sufficient control authority for the aerocapture
The drag-skirt jettison time is determined by the onboard guidance and navigation system. Two
guidance schemes are considered: a numeric predictor-corrector (NPC) and a heuristic velocity trig-
ger. The NPC numerically propagates the current state forward until atmospheric exit with a jettison
time guess. The jettison time is refined with the predicted apoapsis error until the predicted error is
below a preset threshold. The velocity trigger is a simpler and computational lighter algorithm. The
jettison time is determined based on a predetermined velocity.
The aerocapture maneuver will place the vehicle into an elliptical transfer orbit with the required
apoapsis altitude. After the vehicle exits Mars’s atmosphere, it performs a 180-degree slew ma-
neuver to reorient the vehicle to later perform a periapsis raise maneuver at apoapsis. This small
maneuver raises periapsis out of the atmosphere to prevent the vehicle from reentering the atmo-
sphere on subsequent passes.
Mission designers may typically exert a measure of control over the inbound velocity and the
flight-path angle at the top of the atmosphere; however, since the inbound velocity is constrained
by the launch opportunity and rideshare host, the only parameter available to smallsat mission de-
signers is γ. The bounding cases for drag-modulation aerocapture are immediate jettison and no
jettison. The range of entry flight-path angles that will permit a successful aerocapture maneuver is
called the corridor, shown in Figure 4a over a range of arrival velocities. The width of the corridor
is largely determined by the ballistic coefficient ratio. A larger ballistic coefficient ratio (decreasing
β1or increasing β2) results in a wider corridor. Expanding the corridor width is beneficial since
this provides a larger approach navigation target flight control during the aerocapture, decreases
the probability of failure, and reduces control authority requirements for the aerocapture vehicle.
Moreover, increasing the corridor width provides more flexibility in the aerocapture trajectory, en-
abling mission designers to limit peak heat rate and peak acceleration during the atmospheric pass
by placing additional constraints on allowable jettison times.
(a) Aerocapture Corridor with respect to Entry Velocities
(b) Nominal Corridor and Operational Corridor
Figure 4. Drag-Modulation Aerocapture Corridor
The corridor may change according to the atmospheric density on a particular day. A dense
atmosphere may move the corridor to shallower angles because energy is removed more rapidly,
and a less dense atmosphere may require the vehicle to enter at steeper angles. The operational
corridor is defined by the ability to perform aerocapture in the presence of expected day-of-flight
uncertainty in vehicle and environmental properties. Figure 4b shows the corridor for each case,
the operational corridor, and the nominal corridor. While accounting for expected variations in
atmospheric density reduces the width of the corridor, sufficient corridor width remains across all
arrival velocities to accommodate a ballistic coefficient ratio of 9 and current approach navigation
Heat Load, Heat Rate and Acceleration
5000 6000 7000 8000
Entry Velocity [m/s]
Max Acceleration [m/s2]
Corridor Center
Shallow Boundary
Steep Boundary
(a) Max Acceleration vs Entry Veloc-
5000 6000 7000 8000
Entry Velocity [m/s]
Max Heat Rate [W/cm2]
Corridor Center
Shallow Boundary
Steep Boundary
(b) Peak Heat Rate vs Entry Velocity
5000 6000 7000 8000
Entry Velocity [m/s]
Max Heat Load [J/cm 2]
Corridor Center
Shallow Boundary
Steep Boundary
(c) Peak Heat Load vs Entry Velocity
Figure 5. Max acceleration, peak heat rate, and total heat load for boundary and
corridor center cases
The survival of the vehicle relies primarily on its response to the integrated heat load, peak heat
rate and peak acceleration experienced during the aerocapture pass. A high acceleration requires
a more massive aeroshell structure while a high heat rate and heat load requires a more massive
heatshield. These three parameters are function of the vehicle state at the top of the atmosphere,
vehicle properties, and jettison time. Figure 5 shows the impact of arrival velocity on these quantities
of interest; in general, all three increase with increasing velocity.
Simulations for the nominal case of a Phobos rendezvous mission varying the boundary with an
entry velocity of 5.8 km/s have been performed (Figure 6). As the entry flight-path angle approaches
the steep boundary angle, the peak heat rate and peak heat load consistently increase; however, the
case in the corridor center has the highest peak acceleration because the vehicle retains a large drag
area at lower altitudes when the atmosphere is denser. However, the center-of-the-corridor case is
also able to limit peak heat rate to approximately 25 W/cm2.
Using drag-modulation aerocapture, an Areosynchronous orbit (ASO) can be easily achieved.
ASO is an orbit with a semi-major axis of 20,000 km with the same properties as a Geosynchronous
orbit. This parking orbit is extremely appealing for guaranteeing full coverage of the surface of
Mars. The polar mapping orbit baseline is an orbit with 90 degrees of inclination and a semi-major
axis of 4,000 km; the Phobos and Deimos rendezvous orbits are equatorial orbits of 9,378 km and
4000 4500 5000 5500 6000
Velocity [m/s]
Altitude [m]
Corridor Center
Shallow Boundary
Steep Boundary
(a) Altitude Variation
4000 4500 5000 5500 6000
Velocity [m/s]
Acceleration [m/s 2]
Corridor Center
Shallow Boundary
Steep Boundary
(b) Acceleration Variation
4000 4500 5000 5500 6000
Velocity [m/s]
Heat Load [J/cm 2]
Corridor Center
Shallow Boundary
Steep Boundary
(c) Heat Load Variation
4000 4500 5000 5500 6000
Velocity [m/s]
Heat Rate [W/cm2]
Corridor Center
Shallow Boundary
Steep Boundary
(d) Heat Rate Variation
Figure 6. Variation of parameters though an aerocapture maneuver around Mars
with an entry velocity 5.8 km/s
23,459 km, respectively. Phobos and Deimos are respectively inclined 1.08 and 1.79 degrees, thus a
quasi-equatorial orbit must be achieved for their inspection.8The drag-modulation approach takes
into consideration the v budget to perform three consecutive maneuvers: periapsis raise, apoapsis
raise, and circularization of the final orbit. Figure 7 shows a stronger increase in the v budget with
the altitude of the parking final orbit with respect to the impulsive approach.
Polar Mapping Phobos Rendezvous Deimos Rendezvous Areosynchronous Orbit
Delta V [km/s]
Impulsive Approach
Drag-Modulation Aerocapture Approach
Figure 7. v Comparison between Impulsive Approach and Aerocapture Approach
for some baseline missions
The drag-modulation aerocapture phase may be affected by two parameters: The arrival state
at the top of the atmosphere and the time of drag skirt jettison. The former is a function of the
transfer trajectory between Earth and Mars and the latter is determined by the onboard guidance
and navigation system.
Arrival state
Several launch dates for an Earth to Mars transfer between 2020 and 2037 have been taken into
consideration. Figure 8 shows the connection between the Mars entry velocity and the Earth-Mars
transfer time.9The possible entry velocities laid in a range between 5.5 to 13 km/s, while the
transfer time varies from 120 to 270 days. For all the considered years, a range of transfers with
an entry velocity between 5 and 7.5 km/s exist. The reduced range of velocities coincides with the
studied case in Figure 4(a). The transfer time remains spread out in the original range from 120 to
270 days. A launch that guarantees a lower transfer time is preferred.
Figure 8. Entry Velocity Changing with respect to Transfer Time9
As already stated, the majority of rideshares towards Mars target polar orbits. To highlight the
flexibility of the aerocapture maneuver, an investigation of the variation in the B-plane targeting
performed by the spacecraft has been conducted. After the smallsat separates from the host, the
smallsat must perform a small corrective maneuver to modify its B-plane target if the smallasat is
detached after 10 days the departure from Earth. Figure 9 shows that from an initial inclination of
90 deg, coinciding with the host inclination, the spacecraft has to perform a maneuver of 5.5 m/s in
its B-plane targeting to reach an equatorial orbit. This is of high relevance since it allows the use of
rideshares with broader margins with respect to the mission trajectory of the host. Drag-modulation
aerocapture enables a smallsat mission to target every inclination which saves the propellant needed
to perform an inclination change.
Real-time Guidance Trade Study
The onboard guidance algorithm evaluates the time in which the drag-skirt has to be jettisoned
Consequently, if an error in the timing is present, the energy of the vehicle at the exit of the at-
mosphere will be incorrect. This results in an error in the apoapsis of the target maneuver, which
requires additional V to correct
Two guidance algorithms have been tested, a numerical predictor-corrector and the velocity trig-
ger. The velocity trigger is a simple and cheap guidance option that determines the jettison time
based on a predetermined velocity. A 1,000-sample Monte Carlo analysis was performed to eval-
uate the performance of the guidance algorithms with day-of-flight uncertainty. The dispersions
0 20 40 60 80 100
Inclination Final Orbit [deg]
Delta V [m/s]
Figure 9. Incoming Velocity Changing wrt Target Inclination
in entry states and vehicle aerodynamic properties were modeled as uniform distributions. Each
case used a randomly generated atmosphere from Mars-GRAM to determine atmospheric proper-
ties. The simulation was implemented on the basis of a mission to rendezvous with Phobos. All
the cases were able to successfully aerocapture into orbit. The obtained results are shown in Table
2. The NPC presents a lower mean and standard deviation of apoapsis error and v compared to
the velocity trigger at the expense of increased software complexity. The v includes the raise
periapsis, raise apoapsis, and circularization.
To ensure the closure of a smallsat design using drag modulation aerocapture, a systems-level
analysis was performed on a smallsat aerocapture system. intended to deliver a 12-U CubeSat to an
Areocentric, Phobos-trailing orbit. The analysis showed that drag-modulation aerocapture can be
used to deliver a CubeSat to a broad variety of target orbits while fitting in the packaging constraints
of an ESPA ring. Figure 10 shows a concept of operations for the analyzed mission.
Figure 10. Concept of operations for a smallsat aerocapture system with CubeSat payload
Table 2. Monte Carlo results for Phobos rendezvous mission
Max Acceleration
Peak Heat Rate
Heat Load
Apoapsis Error
Mean 26.3 26.2 24 24.1 4.9 4.9 389 767 586 612
σ1 1.3 1.5 0.9 0.2 0.2 428 1184 10.3 31.8
Mean +3σ29.4 30.2 28.5 26.7 5.5 5.6 1673 4318 617 707
Regardless of the choice of the guidance algorithm, the aerocapture maneuver saves fuel while
maintaining good targeting performance. Moreover, the error in the apoapsis may be used in favor
of the rest of the mission. In the baseline missions, both for Phobos/Deimos rendezvous or flyby and
for the ASO mission a phasing maneuver and orbit raising maneuver must be performed to achieve
the required orbit. The polar orbit baseline may require a phasing maneuver as well according to
its mission profile. Furthermore, the phasing time increases with the altitude of the target, since the
target period increase with the cube of the mean distance between the planet and the satellite. The
main idea consists of using this uncertainty to reduce the phasing time with a negligible increase in
the overall propellant budget.
-1 -0.5 0 0.5 1
!'(.*+#.#-/*. #'-
Figure 11. Variable apoapsis maneuver provides multiple options to reduce total
phasing time without a significant increase in V.
To achieve this result, two variable-magnitude impulsive maneuvers burns are executed to reach
the correct apoapsis altitude with appropriate phasing for rendezvous with Phobos or Deimos, as
applicable. Reaching the apoapsis after the periapsis raise maneuver, the first high-thrust impulse
may be performed in that location, moving the periapsis of the orbit to a convenient position, further
outside of the atmosphere. The new periapsis is bounded between the initial orbit and the target
semi-major axis. The second impulse will be performed on the temporary periapsis and will locate
the new apoapsis on the goal orbit. After this, a circularization maneuver may be performed at any
subsequent apoapsis passage. Figure 11 shows the proposed procedure, taking in consideration a
mission to rendezvous with Phobos.
The advantage of this combined maneuver lays in the phasing time, which decrease from 5 days
without its application to 1.1 days. In the case of Phobos, the decrease in phasing time is marginal
because of the low altitude of the moon with respect to Mars. However, in the case of the farther
moon Deimos, the maneuver helps to drastically reduce the phasing time. The periapsis altitude
is strictly combined with the angular position of Phobos. Once the true anomaly of Phobos is
determined, the value of the periapsis altitude can be estimated matching the time necessary to
Phobos to reach the next apoapsis to the time necessary to complete the variable apoapsis maneuver.
If no solution can be found, the spacecraft will orbit of a number of periods necessary to meet a
feasible result, as displayed in Figure 12(a).
0 100 200 300
True Anomaly of Phobos [deg]
Phasing Time [days]
(a) Phasing time vs true anomaly of Pho-
0 0.5 1
Phasing Time [days]
Delta-v [km/s]
Delta-v for Phasing
Delta-v Phasing+Circularize
Minimum delta-v to Circularize
(b) v vs Phasing Time of the Variable
Apoapsis Maneuver
Figure 12. Phasing Time in function of True Anomaly and velocity budget
Finally, the v magnitude to complete this maneuver is not negligible and depends on the peri-
apsis altitude. The value varies from 10 m/s to 530 m/s, as shown in Figure 12(b). However, since
the target orbit is circular, another impulsive maneuver must be performed to increase the apoapsis
altitude to the correct, circular value. This leads to a a decrease in the v magnitude of the succes-
sive circularization maneuver. Figure 12b also shows that the v magnitude of the circularization
maneuver is of 530 m/s, instead, the maximum v magnitude of the circularization and phasing
maneuver is only 0.75% more, at 534 m/s.
Figure 13 shows the v required for each of the four baseline missions. As the semi-major axis
of the final orbit increases, the total required v also increases, but the periapsis-raise maneuver
decreases in v. Figure 7 displays the comparison for several baseline missions between an all-
propulsive approach and the proposed approach. All the considered parking orbits are circular.
Mass budgets were developed for the smallsat aerocapture system to ensure the design could
successfully perform its mission. In order to be integrated with an ESPA ring rideshare opportunity,
the smallsat aerocapture system must have a mass of less than 181 kg. A subsystem-based mass
budget was developed for four separate mission cases to ensure this requirement was met, shown in
Table 3.
Polar Mapping Phobos Rendezvous Deimos Rendezvous Areosynchronous Orbit
Delta V [km/s]
Raise Periapsis
Raise Apoapsis
Figure 13. v maneuvers post-aerocapture for the four baseline missions
Table 3. Mass budget summaries for the four smallsat aerocapture cases analyzed. Payload is a 24 kg
CubeSat and mass constraint is 181.0 kg.
mass (kg)
mass (kg)
Wet mass
Mass (kg)
Mass Mar-
gin (%)
Mapping Orbit 68.0 10.2 102.2 113.5 37.3
rendezvous 69.5 33.1 126.7 140.7 22.2
rendezvous 69.9 38.2 132.1 146.8 18.9
69.7 35.4 129 143.4 20.8
The increasing ability of smallsats to perform science missions at relatively low cost has led to
interest in using them as rideshare on interplanetary missions. Packaging and mass constraints limit
the application of onboard propulsion systems on rideshare missions, leading to a limited range
of orbits available to these interplanetary smallsat missions. The use of aerocapture drastically in-
creases the range of orbits available to smallsat missions by decreasing their required propellant
mass fraction. Four potential smallsat missions are assessed, highlighting the savings in the ve-
locity budget. A relatively simple method of aerocapture is to use drag-modulation to control the
energy of the atmospheric exit state. Furthermore, although drag-modulation aerocapture maneu-
ver is slightly sensitive to the uncertainties that can lead to an small error in the apoapsis altitude, a
post-maneuver plan to relieve this effect is proposed. A configuration of a smallsat drag-modulation
aerocapture orbit insertion system is proposed, and a first-order mass budget for the four baseline
cases is provided.
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Conference Paper
Despite growing international interest in small satellites, high dedicated expendable launch vehicle costs and the lack of secondary launch opportunities continue to hinder the full exploitation of small satellite technology. In the United States, the Department of Defense (DoD), NASA, other government agencies, commercial companies, and many universities use small satellites to perform space experiments, demonstrate new technology, and test operational prototype hardware. In addition, the DoD continues to study the role of small satellites in fulfilling operational mission requirements. However, the US lacks sufficient small satellite launch capacity. Furthermore, US government agencies are restricted to the use of US launch vehicles, which eliminates many affordable launch opportunities. In an effort to increase the number of space experiments that can be flown with a small, fixed budget, the DoD Space Test Program (STP) has teamed with the Air Force Research Laboratory Space Vehicles Directorate (AFRL/VS) to develop a low-cost solution for the small satellite launch program. Our solution, which can be implemented on both Boeing and Lockheed-Martin Evolved Expendable Launch Vehicle-Medium (EELV-M) boosters, is called the EELV Secondary Payload Adaptor (ESPA). ESPA will increase the number of launch opportunities for 180kg-class (or smaller) satellites at prices highly competitive with other secondary launch services worldwide.
Drag-modulation flight control may provide a simple method for controlling energy during aerocapture. Several drag-modulation flight-control system options are discussed and evaluated, including single-stage jettison, two-stage jettison, and continuously variable drag-modulation systems. Performance is assessed using numeric simulation with real-time guidance and targeting algorithms. Monte Carlo simulation is used to evaluate system robustness to expected day-of-flight uncertainties. Results indicate that drag-modulation flight control is an attractive option for aerocapture systems at Mars, where low peak heat rates enable the use of lightweight inflatable drag areas. Aerocapture using drag modulation at Titan is found to require large drag areas to limit peak heat rates to nonablative thermal-protection system limits or advanced lightweight ablators. The large gravity well and high peak heat rates experienced during aerocapture at Venus make drag-modulation flight control unattractive when combined with a nonablative thermal-protection system. Significantly larger drag areas or advances in fabric-based material thermal properties are required to improve feasibility at Venus.
An entry, descent, and landing architecture capable of achieving Mars Science Laboratory-class landed accuracy (within 10 km of target) while delivering a Mars Exploration Rover-class payload to the surface of Mars is presented. The architecture consists of a Mars Exploration Rover-class aeroshell with a rigid, annular drag skirt. Maximum vehicle diameter, including drag skirt, is limited to be compatible with current launch-vehicle fairings. A single drag-skirt jettison event is used to control range during entry. Three-degree-of-freedom trajectory simulation is used in conjunction with Monte Carlo techniques to assess the flight performance of the proposed architecture. Results indicate that landed accuracy is competitive with preflight Mars Science Laboratory estimates, and peak heat rate and integrated heat load are significantly reduced relative to the Mars Exploration Rover entry system. Modeling parachute descent within the onboard guidance algorithm is found to remove range error bias present at touchdown; the addition of a range-based parachute deploy trigger is found to significantly improve landed accuracy.
Trajectory options for conjunction-class human Mars missions are examined, including crewed Earth-Mars trajectories with the option for abort to Earth, with the intent of serving as a resource for mission designers. An analysis of the impact of Earth and Mars entry velocities on aeroassist systems is included, and constraints are suggested for interplanetary trajectories based upon aeroassist system capabilities.
Calculations have been performed to quantify the cost and delivered mass advantages of aerocapture at all destinations in the Solar System with significant atmospheres. A total of eleven representative missions were defined for the eight possible destinations and complete launch-to-orbit insertion architectures constructed. Direct comparisons were made between aerocapture and competing orbit insertion techniques based on state-of-the-art and advanced chemical propulsion, solar electric propulsion, and aerobraking. The results show that three of the missions cannot be done without aerocapture: Neptune elliptical orbits, Saturn circular orbits, and Jupiter circular orbits. Aerocapture was found to substantially reduce the cost per unit mass delivered into orbit for five other missions based on a heavy launch vehicle: Venus circular orbits (55% reduction in $/kg costs), Venus elliptical orbits (43% reduction); Mars circular orbits (13% reduction), Titan circular orbits (75% reduction), and Uranus circular orbits (69% reduction). These results were found to be relatively insensitive to 30% increases in both the estimated aerocapture system mass
Proceedings of the 3rd International Planetary Probe Workshop June 2005, Athens, Greece. Ballute aerodynamic decelerators have been studied since early in the space age (1960’s), being proposed for aerocapture in the early 1980’s. Significant technology advances in fabric and polymer materials as well as analysis capabilities lend credibility to the potential of ballute aerocapture. The concept of the thin-film ballute for aerocapture shows the potential for large mass savings over propulsive orbit insertion or rigid aeroshell aerocapture. The mass savings of this concept enables a number of high value science missions. Current studies of ballute aerocapture at Titan and Earth may lead to flight test of one or more ballute concepts within the next five years. This paper provides a survey of the literature with application to ballute aerocapture. Special attention is paid to advances in trajectory analysis, hypersonic aerothermodynamics, structural analysis, coupled analysis, and flight tests. Advances anticipated over the next 5 years are summarized.