Conference PaperPDF Available

Microgravity restart of liquid rocket engine with low propellant residuals



The active deorbiting and passivation of launch vehicles has become key for the implementation of modern space debris mitigation guidelines. Appropriate engine restart conditions must be provided as part of this process. Ullage motors have been traditionally employed to induce active settling and ensure a gas-free propellant supply to the engines. Although robust and reliable, ullage rockets are also heavy, which motivates the study of alternative approaches to the problem. This paper explores for the first time several high-risk-high-return propellant settling strategies that may result in significant benefits for future space systems. In particular, three distinct Magnetic Positive Positioning concepts, a hydrogen-peroxide-based Propellant Gasification System, and a hybrid device that combines both approaches are introduced. The preliminary feasibility analysis indicates that the successful development of these technologies may lead to mass savings of hundreds of kilograms and economic gains of several hundred thousand dollars per launch. However, the robustness of some of these methods may be compromised by complex fluid-structure interactions that require a careful numerical and/or experimental analysis.
AAS 091
Alvaro Romero-Calvo
, Vladislav Urbansky
, Vadim Yudintsev
, Hanspeter
, and Valeriy Trushlyakov
The active deorbiting and passivation of launch vehicles has become key for the implemen-
tation of modern space debris mitigation guidelines. Appropriate engine restart conditions
must be provided as part of this process. Ullage motors have been traditionally employed
to induce active settling and ensure a gas-free propellant supply to the engines. Although
robust and reliable, ullage rockets are also heavy, which motivates the study of alternative
approaches to the problem. This paper explores for the first time several high-risk-high-
return propellant settling strategies that may result in significant benefits for future space
systems. In particular, three distinct Magnetic Positive Positioning concepts, a hydrogen-
peroxide-based Propellant Gasification System, and a hybrid device that combines both ap-
proaches are introduced. The preliminary feasibility analysis indicates that the successful
development of these technologies may lead to mass savings of hundreds of kilograms and
economic gains of several hundred thousand dollars per launch. However, the robustness
of some of these methods may be compromised by complex fluid-structure interactions that
require a careful numerical and/or experimental analysis.
The exploration and commercialization of space has led to the increasing contamination of the Low Earth
Orbit (LEO) by non-functional man-made objects. Space debris represents a serious safety hazard for current
and future satellites due to the risk of in-orbit collisions, and a concern for the general population during
uncontrolled re-entry events. The minimization of debris release during normal operations has consequently
become a major goal for the international space community.1
Launch vehicles (LVs) represent more than 42% of the abandoned intact objects in orbit and account for
57% of the abandoned mass.2Recent studies have identified the most dangerous elements in an attempt
to guide future active debris removal efforts,2–4 resulting in a list that includes 290 second stages of the
Soviet/Russian “Cosmos-3M” LV, 7 of which have been considered among the 50 most concerning space
debris objects,4110 third stages of the Soviet/Russian “Cyclone-3” LV, 54 units of the American first and
second stages of the “Delta” LV, as well as 38 third stages of the Chinese “CZ-4” and “CZ-2D” LVs. Further
concerns arise due the presence of propellant residuals in the tanks, which account for up to 3% of the initial
fuel mass.5During the long stay of a stage in orbit, the remaining fuel evaporates under the influence of solar
radiation, which leads to an increased risk of explosion in orbit and, therefore, to the generation of space
debris.1The uncontrolled descent of first LV stages in sensitive drop areas can also lead to environmental
Graduate Research Assistant, Department of Aerospace Engineering Sciences, University of Colorado, 431 UCB, Colorado Center for
Astrodynamics Research, Boulder, CO 80309-0431. Contact:
Graduate Research Assistant,Aircraft and Rocket Building Department, Omsk State Technical University, pr. Mira 11, 644050 Omsk,
Russian Federation. Contact:
PhD, Aircraft and Rocket Building Department, Omsk State Technical University, pr. Mira 11, 644050 Omsk, Russian Federation.
§Professor, Glenn L. Murphy Chair, Department of Aerospace Engineering Sciences, University of Colorado, 431 UCB, Colorado Center
for Astrodynamics Research, Boulder, CO 80309-0431. AAS Fellow. Contact:
Professor, Aircraft and Rocket Building Department, Omsk State Technical University, pr. Mira 11, 644050 Omsk, Russian Federation.
pollution caused by the depressurization of toxic fuels, fires in the drop sites, and the contamination of water
bodies. This problem is more relevant for Russian LVs like “Soyuz”, “Proton”, and “Angara”, where most of
the drop zones are located on land.6, 7
Modern launch vehicle operations are subjected to strict space debris mitigation policies.1When graveyard
orbits are not attainable, the orbital lifetime is limited and systems are passivated by removing all energy
sources. Active deorbiting represents a highly attractive alternative to those strategies, but it is not exempt
from risks and technical challenges.8Among them, proper engine restarting conditions must be provided
once the stage is separated from the rest of the vehicle in order to ensure a safe reorbiting or reentry. This
decoupling induces strong disturbances on the propellant residuals and leads to highly non-linear sloshing
dynamics, compromising the operation of the engine feed system.9
Ullage engines have been traditionally employed to settle the propellant residuals during insertion, orbital
coast, or on-orbit operations in an approach known as active settling.10 These independent rockets induce
accelerations that can be as weak as 104to 103m/s2and involve solid, mono-propellant, bi-propellant, or
cold gas technologies, sometimes fed by vaporized propellant vented from the main tanks.11 Some examples
include the Saturn IV-B’s hypergolic liquid bi-propellant Auxiliary Propulsion System (APS),12,13 SpaceX’s
Falcon 9 nitrogen cold gas thrusters for coast attitude control,14 or the two Sistema Obespecheniya Zapuska
(SOZ) ullage engines of the Blok DM-2 Proton upper stage. This last unit has raised concerns in the space
debris community after being responsible for up to 50 on-orbit explosions between 1984 and 2019.15
The technical specifications of ullage engines are not usually accessible to the general public, which ham-
pers any effort to perform an “external” evaluation of these systems. However, numerous reports from the
Apollo era can still be consulted. The two Saturn IV-B APSs were usually fired in three consecutive ullaging
burns for a total of 245 s, consuming 13.5 kg of propellant (23.5% of the total propellant mass of each
APS).16 The dry mass of the APS is unknown to the authors but seems of the order of several hundred kg
judging by the volume of the system. The dry mass of Saturn IV-B was about 13.5 t. On the other hand, each
one of the two SOZ units of the Block DM-2 upper stage had a dry mass of 106 kg and a total propellant
mass of up to 114 kg, while the stage itself weighted 2.1 t. Although determined by the characteristics of the
vehicle and its mission profile, the total mass of the ullage rocket system should be around 200 kg.17 With
a Falcon 9’s launch cost to LEO of 2700 $/kg,18 an economic penalty per launch and stage of 500.000
USD may be estimated. This value is doubled for GEO orbits, and multiplied by an even larger factor in a
Mars mission.
Ullage engines are a robust and well-established solution to deal with the restart of rocket engines in
microgravity conditions. However, that should not preclude the study of alternative approaches to the problem
with lower mass budgets and/or enhanced reliability. In this paper, the feasibility and performance of a
Magnetic Positive Positioning (MP2),19 an on-board Propellant Gasification System (PGS),20 and a hybrid
device are explored for the first time. The historical background of each system is presented together with a
preliminary technical analysis. The ultimate goal is to initiate an open discussion on these technologies and
inform the design of new-generation propellant settling systems.
Although applicable to multiple low-gravity propellant settling systems, the discussion that follows focuses
on the first and second launch stages of a Falcon-9-like LV. The basic parameters of the vehicle are reported in
Table 1, with the geometrical definitions being depicted in Fig. 1. Some of these values are found in SpaceX’s
Falcon User’s Guide,14 while others can only be estimated from unofficial sources.
Propellant behavior during stage separation
The acceleration profile experienced by each stage during separation is key to understand the dynamic
behavior of the propellant. A simple mechanical model is introduced in the Appendix A and employed to
See Consulted on 13/01/2022.
Table 1: Geometrical and inertial parameters of the launch vehicle.
Parameter First Stage Second Stage
Propellant type LOX + RP1 LOX + RP1
Empty mass [t] 22 4.5
Propellant mass [t] 411 111.5
Oxygen tank capacity [t] 287.4 78
Kerosene tank capacity [t] 123.5 33.5
Total mass [t] 433 116
Propellant mass after stages separation [t] 13 3.5
Residual propellant mass after landing [t] 1 0.3
Thrust (stage total) [kN] 7686 981
Throttle capability [kN] 4381 to 7686 626 to 981
Number of engines 9 1
Diameter [m] 3.66 3.66
LOX tank height L0[m] 22.5 8.7
Length of 1 stage h0[m] 29 11.2
Length of fuel tank hg[m] 19.2 7.4
First stage mass center xc[m] 14.9 5.7
Moment of inertia [kg·m2]2.68 ·1063·104
Length to PGS nozzles hn[m] 39 15
Figure 1: Geometrical parameters of a launch vehicle stage.
obtain the acceleration curves reported in Fig. 2 using representative values. Peak accelerations of 1m/s2
are applied to the system and sustained for less than 1 s. Additional effects that may impact the propellant
behavior include the release of strain energy from the walls of the tank, the flow movement induced by engine
suction, or thermal convection.11
Modeling this problem is far from trivial, and experimental data is not easily available because of its
consideration as Export-Controlled material. However, a partial recording of Falcon 9’s first and second
stage liquid oxygen (LOX) tanks during the CRS 5 mission is publicly available, allowing for a qualitative
analysis of the problem. Figure 3 shows the sequence of video frames for (a) the instant before second-stage
engine cut off (SECO), (b) the lateral sloshing wave caused by the structural relaxation after SECO, (c) the
cloud of LOX bubbles generated after separation, and (d) the state of the cloud 45 s after separation. It can be
readily concluded that (i) the SECO induces a mild lateral sloshing wave, but does not significantly disturb
the liquid, (ii) the stage separation atomizes the residual LOX in a myriad of droplets that reach the top of
the tank in less than 40 s (i.e. the droplets move at least at 20 cm/s), and (iii) the droplets keep moving for
at least 6 minutes while coalescing with each other. This behavior is also (briefly) observed in the first stage,
where the droplets seem to move at about 0.5 m/s. This value has been employed in the derivation of the
The interested reader is referred to Consulted on 13/01/2022.
2nd Stage
1st Stage
a [m/s2]
t [s]
0 0.2 0.4 0.6 0.8
Figure 2: Estimated acceleration profile of the 1st and 2nd stages after separation.
(a) Pre-SECO (b) Post-SECO (c) Post-stage-separation (d) 45 s after stage separation
Figure 3: Top view of the second-stage LOX tank of SpaceX’s Falcon 9 during the CRS-5 mission. Source:
acceleration profiles shown in Fig. 2.
Engine restart conditions
The propellant must be settled over the fuel outlet to ensure a safe engine restart. Although this usually
implies bringing all the liquid back to the bottom of the tank before ignition, the requirement may be instead
reformulated by noting that the engines will also accelerate the stage. The goal is thus to have enough
propellant at the outlet so that, when the thrust-induced settling brings all the liquid to the bottom of the tank,
no gas bubbles have made their way into the engines.
The first factor to consider is mass flow rate: higher thrust will settle the propellant droplets faster, but will
also require a larger initial mass. The relation between thrust Tand mass flow rate ˙mpis given by
where Isp is the specific impulse and g0= 9.81 m/s2is the standard gravity acceleration. For a LOX-RP1
chemical engine with a mass ratio of 2.3 the specific impulse should be around 285 s at sea level.21 In fact,
the old SpaceX website attributed to Falcon 9’s Merlin engines the values of 282 s at sea level and 311 s in
vacuum. The second factor is the propellant settling time, that can be divided into two phases. On the first,
the propellant residuals return to the intake device, while on the second, gas bubbles are removed from the
liquid. The total settling time can thus be expressed as
ts=tI+tII .(2)
For a tank of length L, the duration of the first phase is bounded by the kinematic result
See Consulted on:
with m0being the dry mass of the stage. The duration of the second phase, assuming a constant terminal
velocity of the bubbles in the liquid, is22
tII =l8
where rbis the bubble radius, Cdis the drag coefficient, ρgis the gas density, ρlis the liquid density, mtis
the total mass of the stage after settling, and lis the height of the longest liquid column. Consequently, the
initial mass of propellant required to complete the maneuver for a giving thrust level is
mp= ˙mpts=T
which increases with T, showing that small thrust values are convenient to minimize the mass of propellant
required to restart the engines. Table 2 reports the stage acceleration, settling time, and initial oxidizer and
fuel masses for different thrust configurations of Falcon 9’s first and second stages. The values Cd= 0.47
(sphere), rb= 5 mm, ρg= 0.1785 kg/m3(He), ρl= 1141 kg/m3(LOX), and l=mr/(πR2ρl), with Rbeing
the tank radius and mrthe residual LOX mass, are employed in combination with those presented in Table 1
using the larger LOX tank as a reference. The masses reported in Table 2 are the minimum absolute values
required near the fuel outlet to initiate the restart maneuver. Unofficial sources claim that Falcon 9’s first stage
restart acceleration is less than 50 m/s2by employing reverse engineered telemetry data§, which indicates that
either the central engines (maximum thrust) or three outer engines (minimum thrust) are actually employed
in this process. However, the authors were not able to verify this information.
First Stage Second Stage
Maximum Thrust 350 0.66 1287 401 218 0.39 88 38
Minimum Thrust 200 0.88 971 303 140 0.49 70 30
Single Engine 22 2.63 324 101 140 0.49 70 30
Table 2: Stage acceleration, settling time, and minimum initial oxidizer and fuel masses for different restart
configurations of Falcon 9’s first and second stages.
Concept and overview
The ability of controlling the position of susceptible liquids by means of magnetic fields in microgravity
leads to several potential space applications. Those include, but are not limited to, mass transfer,23–25 thermo-
magnetic convection,26,27 or micropropulsion.28, 29 The volume force density that enables these technologies
is induced by inhomogeneous magnetic fields on susceptible liquids, and adopts the form
fm=µ0MH, (6)
with µ0being the permeability of free space, and Mand Hdenoting the magnetization and magnetic fields,
respectively. In addition, the magnetic normal traction
should be considered at the liquid interface, where Mnis the normal magnetization component.30 This
pressure-like term is usually neglected for natural liquids, such as LOX, but becomes relevant for highly
§See for a remarkable example of reverse engineering. Consulted on
Figure 4: Magnetic Positive Positioning
susceptible materials like high-density ferrofluids.31 Since both Hand its gradient decay with the distance to
the source, the magnetic force vanishes relatively quickly. Therefore, powerful magnets or coils are needed
for most applications.
The Magnetic Positive Positioning approach, sketched in Fig. 4 for the system under study, seeks to induce
a magnetic acceleration that holds, collects, and/or traps the liquid near the fuel outlets. The concept was
first proposed in 1963 by Steve Papell in the same patent where he invented ferrofluids.32 The idea was
abandoned until 2001 when, motivated by the advent of stronger permanent magnets and high-temperature
superconductors, the NASA Magnetically Actuated Propellant Orientation (MAPO) experiment explored
the positioning of ferrofluid solutions in a series of parabolic flights.33 Such ferrofluids were selected to
approximate the magnetization curve of LOX for different magnetic field intensities. It should be noted
that LOX is the most susceptible natural paramagnetic liquid,34 making it particularly appropriate for this
application. Subsequent publications by Marchetta and coworkers presented refined numerical models and
results of technical relevance for the development of liquid oxygen magnetic positioning devices.35–43 Recent
works have also explored the free surface oscillations of ferrofluids in microgravity, which may be relevant for
slosh control and the development of novel propellant management devices (PMDs).44–49 A comprehensive
review of the field can be found in Ref. 19.
Significant advances have been made in the modeling and fundamental understanding of magnetic positive
positioning devices during the last two decades. However, none of the aforementioned works explored the
feasibility of this approach as part of the operation of LVs. Although limited by the lack of reliable technical
information, this study aims at covering this knowledge gap by exploring the application of magnetic positive
positioning to the restart of Falcon 9’s first and second stages.
Passive retention strategy
The first and most intuitive approach to magnetic positive positioning is the liquid retention strategy, where
a magnet or coil is used to hold the paramagnetic liquid in the presence of adverse accelerations that tend to
destabilize the free liquid surface. In the classical literature, the critical Bond number
is employed to compute the critical acceleration load gfor which surface tension cannot longer stabilize
the meniscus. Myshkis and coworkers provide a best-case Bo=3.32 for cylindrical tanks at a contact
angle of 90,50 which results in g= 2.9·106m/s2for the LOX tank considered in this work. In other
words, surface tension does little or nothing to prevent the atomization of the residual LOX volume observed
in Fig. 3 under the action of the acceleration loads estimated in Fig. 2. An obvious questions is whether the
magnetic force can hold the liquid against adverse accelerations of 1m/s2.
Marchetta and coworkers explore the problem of magnetic LOX retention in a 12 cm diameter 24 cm
height cylindrical tank under the influence of a point dipole of 1.4 cm diameter and an inertial acceleration
along its major axis.42 The dipole strength required to hold 30 ml of liquid is shown to be about 10 Am2
for g= 1m/s2. Similarly, in Ref. 19 an analytical model is developed to study the stability of magnetic
interfaces and applied to a 10 cm diameter cylindrical tank, showing that a 60 g magnet can increase the
critical load by 31.5%. None of these low-gravity studies can be easily extended to Falcon 9’s 3.66 m
diameter LOX tank, where the liquid outlet has a diameter of about 90 cm and the maximum estimated
Figure 5: Magnetic acceleration contours induced on LOX by a 35 cm diameter coil operating at 1 At.
acceleration is about 1 m/s2. The problem, however, can be easily addressed by plotting the axisymmetric
magnetic acceleration contours induced by a 1 A cylindrical coil on the LOX tank volume as done in Fig. 5.
The figure depicts the magnetic acceleration levels in a logarithmic scale and its direction using black arrows.
A coil mean diameter of 35 cm is chosen to ensure that the liquid gets attracted toward the PMD located at the
the tank outlet. The minimum mass to be retained is 70 kg (second stage), which translates to a LOX sphere
of 25 cm radius. The magnetic acceleration at this distance is about 1011 m/s2. Since the acceleration scales
with the square of the current intensity,19 values of 106At (i.e. coils current intensity times number of coil
turns) would be required to retain the oxidizer against accelerations of 1 to 10 m/s2. Further computations are
not required to conclude that the mass and/or power requirements of this approach are well beyond reason
with existing technologies, particularly for the first stage.
Recovery strategy
The passive magnetic retention strategy sets an upper limit for the magnetic field strength. Since this limit
is hard to reach with existing technologies, alternative strategies must be explored. The first of them is here
introduced and seeks to collect the LOX droplets after they are atomized rather than holding part of the liquid
at the bottom of the tank. In order to evaluate this idea, the time required to settle a LOX droplet is first
derived with a simplified framework of analysis.
Magnetic settling time: If the settling of the propellant is induced by the magnetic interaction and not by a
uniform acceleration g, the derivation of tsis complicated by the presence of an inhomogeneous acceleration
field. A strict approach to the problem would require solving the Navier-Stokes equations with a magnetic
force source term. Although less computationally expensive than the fully coupled fluid-magnetic simulations
that are necessary for highly susceptible ferrofluids,19 this approach is still prohibitive for a preliminary
study. Instead, the movement of a perfectly spherical droplet along the symmetry axis of an axisymmetric
coil or magnet is analyzed. The droplet is small in comparison with the variation of the magnetic field
and exhibits linear magnetization with susceptibility χ1. It is further assumed that external, internal,
and magnetization fields are collinear, that residual liquid volumes do not contribute to the magnetic field,
Following the popular saying, “if you can’t beat them, join them”.
and that magnetic surface force terms are negligible. In this simplified framework, the total magnetic force
induced on the liquid droplet by a circular coil with nturns, radius R, and current intensity Iat a distance z
along the symmetry axis ezisk
Fm≈ −3µ0χ(nI)2R4
This expression can also be applied to axially magnetized cylindrical magnets with magnetization Mm, radius
Rand height lmby considering an equivalent circular coil with the same radius and current intensity nI =
Mmlm. Its main advantage is that it allows the derivation of a quasi-analytical expressions for tI. After
considering Newton’s second law and solving the resulting second-order differential equation with initial
position z(0) = Land initial velocity ˙z(0) = 0, the duration of the first phase becomes
tI(L) = s4πρl
dz, (10)
where it should be noted that tIis inversely proportional to nI (or, if a magnet is employed, to Mmlm) and
R2. Of these, only the current intensity can be considered a design parameter, because Ris bounded by the
fuel intake radius (see previous subsection).
The time required to debubble the multiphase mixture near the fuel outlet using the paramagnetic force
can be derived in a similar way, as done in Ref. 52 after adopting the Stokes’ law for viscous drag (which is
valid for Re 1) instead of the drag coefficient. However, the magnetic debubbling process is much faster
than the first phase because the liquid is closer to the magnetic source, and hence it is further assumed that
Even though the assumptions employed in the derivation of Eq. 9 are not appropriate for highly susceptible
ferrofluids, the volume magnetic susceptibility of the liquids employed in this work is bounded by that of
LOX (χLOX = 0.0034 at 90 K and 1 atm34). It would not make sense to use high density ferrofluids in the
fuel tank when the most demanding requirements are associated with the LOX tank.
Performance analysis: The time of flight of the droplets for a coil diameter of 35 cm is represented in
Fig. 6 as a function of the initial droplet distance to the coil and the applied current intensity. A LOX settling
time of tmax 6minutes, estimated in the analysis of the second stage in Fig. 3, is superposed and treated as
kEquation 9 can also be found in the literature divided by the term (1+χ)2or (1 + )2, with D= 1/3being the demagnetization
factor of a sphere.19,31, 51 Each choice denotes a different modeling of the internal magnetic field inside the droplet. Since in the problem
here considered χ1, the simplest approach, that assumes H0H, is adopted.
Max. T.
Min. T.
Max. T.
Min. T.
1st Stage
2nd Stage
102 A·turn
103 A·turn
104 A·turn
105 A·turn
106 A·turn
Time of flight [s]
Initial droplet distance to fuel outlet [m]
0 1 2 3 4 5
Figure 6: Time required by a LOX droplet to reach the bottom of the tank as a function of its distance to the
coil and the applied current intensity. The minimum tank settling length required by the single engine (S.E.),
minimum thrust (Min. T.) and maximum thrust (Max. T.) configurations is superposed.
a deadline for the collection process. However, unofficial telemetry data shows that the first stage restarts 2
to 3 minutes after stage separation (see footnote at the end of page 5). As it will be seen, this does not change
the qualitative results of the analysis. It is arbitrarily assumed that, after atomization, the LOX droplets are
uniformly distributed in the tank volume. In this framework, the vertical lines represent the tank length that
needs to be settled for each one of the configurations detailed in Table 2 before the LOX droplets stop moving
(i.e. get attached to the walls of the tank).
A qualitative difference is first observed between first and second stages. The LOX mass required to restart
the engines, listed in Table 2, drops by an order of magnitude in the second stage, and thus the length of the
tank that needs to be settled is much smaller under the uniform droplet distribution assumption. The second
factor that should be considered is the density of residual LOX per unit tank length. Although the lower
LOX tank has a length of 22.5 m and the upper of just 8.7 m, the density of residual LOX per unit length
is relatively similar, decreasing from 383.5 kg/m to 267 kg/m, respectively. Figure 6 shows that a coil with
a configuration of 103A·turn can satisfy the requirements of the second stage. However, 104to 105A·turn
are needed to settle the first stage using the single engine restart configuration, with 106to 107A·turn being
required for the rest. In other words, the liquid recovery strategy can potentially reduce the coil strength
requirements by one order of magnitude in the first stage and three orders of magnitude in the second. These
conclusions do not change if tmax drops to 120-180 s for the first stage, as indicated by unofficial telemetry
These results should be taken with care due to the number of assumptions employed in the derivation of
Eq. 10. In particular, fluid-structure interactions have been completely neglected, but Fig. 3 shows that, after
a few minutes, the liquid droplets tend to get stuck to the walls of the tank. This is a natural consequence
of the presence of corner geometries in the interface between PMDs and the walls. The robustness of the
liquid recovery strategy may thus be compromised by this effect, which should be evaluated with flight data
that is not available to the authors. Possible mitigation strategies include the elimination of gaps and corner
geometries or the application of a LOX-phobic treatment to the internal surfaces.
Active retention strategy
The third and final approach seeks to soften the requirements imposed on the magnetic system by storing
the restart liquid in a smaller auxiliary tank during the stage separation process. The propellant is later re-
leased near the fuel outlet, where it is held against disturbing accelerations by means of a permanent magnet
or coil. The restart liquid remains thus unaffected by large stage separation accelerations. The main draw-
backs of this method are the addition of ancillary components and the limitation of the number of restart
A representative magnetic configuration is displayed in Fig. 7, which depicts the magnetic acceleration
field on LOX of a 5 cm radius 10 cm height cylindrical magnet magnetized at 1300 kA/m. The magnet
geometry differs from that of Fig. 5 in the adoption of a smaller magnet radius, which increases the magnetic
force close to the source.19 The magnet volume is chosen to impose an acceleration of 104m/s2(one
order of magnitude larger than microgravity disturbances11) at 40 cm from the magnet. This leads to
the approximate LOX volume that needs to be retained in the single-engine first stage scenario reported in
Table 2. Although the mass of the magnet is just 5.5 kg, ancillary components like the tanks or PMDs must
also be taken into account.
Four 26 cm radius spherical containers are first considered to store the LOX volume in the first stage. The
four tanks are located around the fuel outlet and outside the main tank space, as shown in Fig. 8a. Although
this choice leaves the internal configuration virtually unaffected, an alternative internal positioning would
minimize the pressure difference supported by the walls, and thus the mass requirements. The total mass of
this system is hard to estimate without access to further technical details, but it would not be surprising to
discover that ancillary components account for most of it. For instance, emptying the tanks in the microgravity
environment would require a bladder12 or outlet magnet45 to ensure a gas-free expulsion and a small pressure
difference applied, for instance, by means of the same helium tank employed to pressurize the LOX vessel. An
alternative system that minimizes the number of ancillary components at the expenses of providing a single
Figure 7: Magnetic acceleration field induced by a 5 cm radius 10 cm height praseodymium magnet mag-
netized at 1300 kA/m.
engine restart opportunity is shown in Fig. 8b. This configuration employs one 40 cm radius internal LOX
tank with a top gas valve and an open-ended bottom. After the stage separation, the gas valve is opened and
the LOX flows under the influence of the magnet to the minimum-energy configuration given by the constant
acceleration lines in Fig. 7. The total mass of this system would be close to 20 kg after considering a 1
mm thick Aluminum LOX tank. The major risk of this approach is the potential onset of Rayleigh-Taylor
(a) External (b) Internal
Figure 8: Conceptual active liquid retention configurations for Falcon 9’s first stage LOX tank.
instabilities (see Ref. 53) at the open bottom end if the LOX interface is accidentally exposed to the gas. A
careful fluid dynamic analysis of the problem is thus needed to ensure a robust operation.19,40, 41
The same system can be applied to the second stage and would require four 16 cm radius tanks or a single
25 cm radius vessel with a magnet of just 0.5 kg. The mass of the single tank aluminum walls would account
for 2.2 kg. That is, the active liquid retention strategy may potentially achieve mass savings of one to two
orders of magnitude with respect to existing ullage engines.
Magnetic field generation
With the nI parameter already sized for the different engine restart configurations, the next logical step
is to determine how to produce the required magnetic fields. Three technologies are subsequently studied:
copper/aluminum coils, rare earth permanent magnets, and superconducting coils.
The magnetic field generated by a coil is linearly dependent on the nI parameter, that can be increased by
adding more wire turns or employing higher currents. Total coil mass and power dissipation are the driving
factors of the design. The mass of the coil can be estimated as
m= 2πRnSρw,(11)
where Sis the cross-section of the wire and ρwits density. The heat dissipated by the coil can be derived
from Ohm’s law, resulting in
P= 2πRIe
with ρebeing the resistivity of the material. In a worst-case scenario this heat is stored in the coil instead of
being dissipated, causing a temperature increase of
T=P tmax
where Cpis the heat capacity of the wire. In order to constrain the design, the heat dissipated by the coil is
limited by considering two worst-case scenarios: in the first, the heat is fully transferred to the residual LOX.
The maximum power is arbitrarily set to the one that vaporizes 1 kg of LOX during the 6 minutes operation of
the coils. The latent heat of vaporization of LOX is 6.82 kJ/mol (or 213.13 kJ/kg) at atmospheric pressure,34
which results in a maximum coil power of 592 W. Because nI is fixed, the ratio I/S is determined by Eq. 12.
In the second scenario, the heat is stored in the coil, causing a temperature increase that is limited by choice
to 10 K. Therefore, from Eq. 13 the I/S parameter is extracted. The most restrictive constraint is chosen for
each design so that the thermal runaway of the material and the vaporization of the residual LOX volume are
avoided. Then, the mass is computed for the I /S value from Eq. 11. The second requirement concerns the
voltage of the coil, set to 24 V to ease integration with Falcon 9’s power subsystem. After inserting the I/S
value in Eq. 12 and employing Ohm’s law, the current intensity of the coil and its resistance are computed.
At the boiling temperature of LOX (90 K) the resisitivities of copper and aluminum are 3.5·109m and
4.5·109m, respectively.54 Although copper is slightly more conductive than aluminum, its density and
heat capacity are 8960 kg/m3and 0.385 kJ/kgK, while aluminum has a density of 2700 kg/m3and a heat
capacity of 0.89 kJ/kgK. Therefore, aluminum is chosen to minimize the mass of the design.
The second approach focuses on employing rare earth permanent magnets to generate a constant, unpow-
ered magnetic field. Neodymium (Nd2Fe14B) is the most popular rare earth material, has a density of 7008
kg/m3,34 and exhibits a remanent magnetization of Mm1200 kA/m. It is classified as a “hard material”,
implying that it can be used to manufacture magnets of any shape.55 As previously noted, the sizing parame-
ter nI of a cylindrical coil can be translated to the length lmof an equivalent cylindrical magnet with the same
diameter by means of the expression In =Mmlm. Magnet tessellation strategies such as Halbach arrays can
be employed to boost the paramagnetic force on one side, while partially canceling it on the other.56 Halbach
arrays have already been proposed for space applications52 and would be particularly well suited to the LOX
settling problem for two reasons: the reach of the magnet is increased, and the interaction between the LOX
magnet and the droplets generated in the fuel tank is reduced.
Neodymium magnets experience a slight increase of their magnetic flux as temperature decreases. At
around 135 K, a transition point is reached and the magnet undergoes spin reorientation (i.e. a change in the
preferred direction of the magnetization vector) that decreases the flux by no more than a 14%. This process
is reverted as soon as the temperature increases.57 If needed, the transition point could be avoided by isolating
the magnet in the LOX tank and actively controlling its temperature. A more elegant solution is, however, to
employ praseodymium magnets to avoid the spin reorientation. Praseodymium magnets do not suffer from
spin reorientation and have been shown to reach a remanent magnetization of 1300 kA/m at 85 K,58,59
which makes them ideal for LOX control applications.
The design points of the aluminum coil and praseodymium magnet are shown in Table 3 as a function of the
nI parameter. In all cases but 106At, the design of the coil is driven by the thermal requirement (maximum
temperature increase of 10 K). Magnets are orders of magnitude lighter for all nI values, incurring in a
– still reasonable– mass penalty of 52 kg at 105At. nI values beyond 105At seem unreachable without
incurring in large mass penalties, and it is in this context where high-temperature superconductors (HTC)
can become a game-changing alternative. A HTC wire exhibits zero resistance in a certain operational range,
resulting in no heat loss and a potential reduction in mass and power requirements. This happens when (i)
it is operated below its critical temperature Tc–greater, by definition, than the boiling point of nitrogen (77
K)–, (ii) it is subjected to a magnetic field below the critical field Bc, and (iii) the critical current Icis not
exceeded. The simultaneous satisfaction of these three requirements is far from trivial; in fact, Icdecreases
continuously with increasing temperature and magnetic field.60 For example, Bi2223 (Bi2Sr2Ca2Cu3O10+δ)
has a critical temperature of 110 K, but its critical current drops to zero when the material is exposed to a
field of less than 1 T. RE-123 ((RE)Ba2Cu3O7, where RE stands for Rare Earth element) superconductors
(also known as REBCO), on the contrary, can resist up to 10 T, but only well below a critical temperature of
around 90 K.61 It is nowadays feasible to generate very strong magnetic fields at the boiling point of Helium
(4.22 K), the best example being the 32 T superconducting magnet62 of the National High Magnetic Field
Laboratory∗∗. Reaching similar values at higher temperatures seems, unfortunately, still beyond our technical
capabilities. In the application here discussed the superconductor would be immersed in LOX, which would
act as a cooling agent only if Tc90 K, and the maximum magnetic field imposed near the coils would be
10 T at In = 106At. The results presented in Ref. 60 for different commercial REBCO conductors seem
to indicate that such operation point cannot be reached with current technologies. However, the Icvalue of
4 mm wide superconductors is shown to be 450–1000 A/mm at 12 T and 77 K, a value that jumps up to 60
kA/mm2at 18 T and 4.2 K. This indicates that cooling mechanisms need to be put in place to reach the 106
At configuration with HTCs, which may open an opportunity for multiple-use of the helium tanks employed
for tank pressurization. This possibility, although attractive, would require a deeper technical analysis that is
beyond the scope of this paper.
∗∗See development/magnet-science- technology/
magnet-projects/32- tesla-scm. Consulted on: 26/12/2021
Table 3: Mass and power budget for different magnetic configurations.
Aluminum coil1Praseodymium magnet2
Current ·Turns
[# turns]
1022.51·1030.06 4 1.62·1012.44 0.08 0.052
1032.51·1020.60 40 5.12·10124.4 0.77 0.52
1042.51·1016.03 405 1.62 244 7.7 5.2
1052.51 60.26 4054 5.12 2438 77 52
10624.67 592 40541 16.2 24814 769 519
1Coil of 35 cm diameter operating at 24 V and 90 K.
2Cylindrical magnet of 35 cm diameter magnetized at 1300 kA/m at 90 K.
Fuel tank
From the magnetic actuation perspective, LOX determines the design envelope of the system. On one
hand, the LOX tank is more than two times larger than the fuel tank, and therefore a given magnetic source
will reach a larger portion of the latter. On the other, LOX is a paramagnetic substance with volume magnetic
susceptibility χLOX = 0.0034, while kerosene is a diamagnetic with χKe ≈ −8·106.63 In order to apply the
magnetic retention and settling approaches to the fuel tank, it must be transformed into a para/ferromagnetic
by adding magnetic nanoparticles and creating a kerosene-based ferrofluid. With this approach, the suscepti-
bility of the solution is bounded by the concentration of magnetic nanoparticles.
Kerosene has been employed as a carrier liquid since the invention of ferrofluids in 196332 and kerosene-
based ferrofluids are synthetized and used in numerous fields.64–67 Commercial solutions like Ferrotec’s
EMG-905†† are now widely available at a relatively low cost. The initial susceptibility of a monodisperse,
colloidal ferrofluid can be estimated as30
χini = 8φλ, (14)
where φis the volume fraction of magnetic solids and λis the coupling coefficient, given by
24kT ,(15)
with Mdbeing the saturation moment of the bulk magnetic solid, Vthe nanoparticle volume, kthe Boltzmann
constant, and Tthe absolute temperature. Assuming an iron oxide nanoparticle radius of 5 nm, an absolute
temperature of 293 K, and a saturation moment of 446 kA/m,68 the approximate volume fraction required to
match the magnetic susceptibility of LOX starting from the value of kerosene would be just φ3.2·104.
Not surprisingly, this value is within the range tested by Martin and Holt in the NASA MAPO experiment.33
If the whole kerosene tank volume is magnetized, the magnetic nanoparticles add 40 kg and 11 kg to
the first and second stages, respectively, in addition to negligible variations in density and specific impulse.19
The very low ferrofluid concentration should prevent damage to the engines. Although simple, this approach
is expensive and inefficient in comparison with the active liquid retention strategy, where only the kerosene
employed to restart the engine would need to be enhanced with ferromagnetic nanoparticles. In this case, the
mass penalty associated with such nanoparticles would be of just 40 to 140 g based on the total RP-1 masses
presented in Table 2. The same approach could be applied to the passive liquid retention and liquid recovery
strategies if a concentrated ferrofluid volume is mixed with the RP-1 residuals shortly before MECO/SECO.
The discussion provided in this section focuses on moving the propellant toward the bottom of the tank. A
logical follow-up question is what to do once that happens. Unless some sort of valve or screen mesh is put
in place, it must be assumed that the propellant stored in the tubes leading to the engines is partially released
to the tank. Thus, once the liquid is attracted back to the tank outlet, a gas bubble is generated in the tubes
leading to the engines. This would represent an unacceptable safety risk.
There are several ways to face this problem, and all of them require a deep knowledge of Falcon 9’s tank
outlet design. A possibility would be to employ a combination of PMDs to hold the liquid at the entrance
of the engine against accelerations of 1 m/s2. This should effectively prevent the development of a Rayleigh
Taylor instability and the entrance of gas bubbles. Active strategies may involve purge valves and membranes
located in strategic positions along the liquid path. Unfortunately, a more detailed analysis requires technical
information that is not available to the authors.
††See emg/oil/emg-905/. Consulted on:
Figure 9: Propellant Gasification System
Concept and overview
The injection of hot gases into the tanks for chemical pressurization through propellant evaporation and
combustion has long been known and used in the “Proton”, “Rokot” and “Dnepr” launch vehicles.69 These
systems do not require heat exchangers, which are used to heat cold helium gas from 90 K to 300 K in current
pressurization systems, and are instead based on feeding a nitric acid-based oxidizer and fuel into the fuel and
oxidizer tanks, respectively. The components ignite and heat is released, causing the fuel to vaporize, which
increases the pressure in the tanks. Referring to Sutton, while this type of pressurization system is small
and lightweight, it has generally not yielded reproducible tank pressures due to the difficulties to stabilize the
combustion reaction. For instance, fuel sloshing caused by vehicle maneuvers results in sudden cooling of the
hot pressurizing gas and leads to erratic tank pressure changes. This problem can be avoided by physically
separating the hot gas from the liquid propellant. If the hot gas is generated from a solid propellant reaction
or from mono-propellant decomposition instead of a high-pressure gas supply, a significant reduction in the
mass of the pressurizing system can be achieved.21
The Propellant Gasification System concept was first proposed in the early 2010s as a method to vaporize
the propellant residual of 2nd launch vehicle stages and provide attitude and orbit control capabilities by
means of dedicated vapor-fed thrusters.70–72 The original idea was to inject the combustion products of two-
component propellants (AA and NDMH) in the tank to move the stage from its initial circular orbit to an
elliptical orbit. Such orbit would ensure a successful deorbiting in the time frame of 25 years. In 2015, the
use of solid fuel instead of a two-component propellant was investigated to simplify the design and improve
the energy performance of the PGS. Further analyses on the Soyuz 2.1v launch vehicle showed that the PGS
could also lead to launch vehicle characteristic speed enhancements of up to 5%.73 The PGS baseline design
has currently evolved to reduce its mass and environmental impact using a green mono-propellant (hydrogen
peroxide) that adds the possibilities of i) controlling the movement of the stage to reach a given drop area, ii)
providing conditions for LRE restart by executing the flip around and propellant settling maneuvers, and iii)
passivating the propellant after a normal or emergency cutoff of the LRE.74 Highly concentrated hydrogen
peroxide (85%) has already been employed as a green mono-propellant in substitution of hydrazine on the
“Soyuz” launch vehicle for the operation of turbo-pump units.75
The PGS considered in this work vaporizes the propellant residuals in the oxidizer tank of the launch
vehicle using the catalytic decomposition of hydrogen peroxide, which is placed in an auxiliary tank and
used as a heat source. The mono-propellant is passed through a catalyst chamber that leads to the formation
of up to 823K hot oxygen and steam. The vapor-gas mixture is then transferred into the oxidizer tank, which
leads to vaporization of the liquid phase and a pressure increase. The gas, consisting of vaporized propellant
and pressurizing agent (helium), is used to feed a set of gas thrusters that are employed for attitude control and
tank settling, as sketched in Fig. 9. This approach can be regarded as the active equivalent of the hydrogen
venting strategy employed in the Apollo era.11
It is important to remark that the risk of combustion or explosion is virtually non-existent in the proposed
PGS approach. Since only hot hydrogen peroxide decomposition byproducts (water vapor and oxygen) are
injected into the oxidizer tank and a reducing agent at auto-ignition temperatures is not present, the combus-
tion reaction cannot be produced. Such a hazard could only occur if the PGS was used in the propellant tank.
However, this approach has already been applied to the aforementioned chemical pressurization systems, and
the reaction was safely controlled by tuning the pressurizer gas flow rate.
The following main subsystems compose the PGS: (i) a hot gas generator that includes a bladder-controlled
hydrogen peroxide tank and a catalyst chamber where the exothermic decomposition of hydrogen peroxide
happens, producing a high-temperature vapor-gas mixture (VGM) with a 34% H2Oand 66% O2composi-
tion, (ii) a system of nozzles installed after the catalyst chamber that injects the VGM into the LOX tank
minimizing tank wall heating, and (iii) a system of gas nozzles used to discharge the VGM from the tank and
produce the required thrust. The PGS provides control over the tank discharge valves, hydrogen peroxide
feeding, and gas nozzles. Cold helium gas, which is stored in balloons at the bottom of the oxidizer tank, can
also be used to reduce the temperature of the VGM (see next section).
Concept of Operations:
Figure 10 shows the ConOps for the PGS approach. Three seconds after MECO, the first stage separates
from the second stage using a pneumatic pusher (or equivalent). The control system of the 1st stage starts
the PGS to increase the pressure in the oxygen tank. Nine seconds after MECO, the PGS opens the attitude
control nozzles with a total thrust of 1 kN to start the flip around maneuver, which involves an acceleration
phase, a constant angular velocity phase, and a deceleration phase. About 55 seconds after MECO, a set of
nozzles provide 5 kN of axial acceleration before engine re-start.
The PGS can be operated in three distinct modes in combination with the helium pressurization system:
1. Standalone PGS operation: In this case, the main task of the PGS is to vaporize the liquid residuals.
The highest gas temperature is employed (823 K).
2. Alternate operation: The residual cold helium pressurizer gas (90 K) is fed into the tank in order to
reduce the VGM temperature. This reduces or stops the evaporation of liquid propellant and ensures
that the proper quantity of liquid is vaporized to operate the gas thrusters.
3. Combined operation: In this case, the VGM does not vaporize the propellant residuals, acting instead
as a heat exchanger that warms up the pressurizer gas up to an accepted operational temperature (300
K), replacing the engine heat exchanger.76
Flip around maneuver:
The flip around maneuver to the required φangle consists of three phases that are evaluated in Fig. 11.
The acceleration phase is produced with constant angular acceleration until the speed ωis reached, and lasts
Flip around maneuver
Propellant settling
+3 s
+95 s
1st burn
Starting PGS
LOX evaporation
0 s
5 kN
1 kN
1 kN
Figure 10: PGS Operation cycle
Ang. Vel. [m/s]
t [s]
10 20 30 40 50
Ang. Acc. [deg/s2]
t [s]
10 20 30 40 50
Figure 11: Angular velocity and angular acceleration profiles of the 1st stage during flip around maneuver
using the attitude control nozzles with a total thrust of 1 kN
where zero initial speed and angle are assumed. The second phase is produced with constant angular velocity
ωand duration tω. Finally, the deceleration phase happens with the same angular deceleration. The total
turnaround time tris given by
The magnitude of angular acceleration is determined by the torque produced by the attitude control nozzles
relative to the center of mass and by the lateral moment of inertia of the stage. For instance, around the z-axis
the acceleration becomes
with Fbeing the total thrust of the nozzles, hnxcthe thrust arm, and xcthe distance from the bottom
section of the stage to the center of mass. The parameters of the flip around maneuver are given in Table 4.
Oxygen vaporization model
Vaporization in a tank can occur according to three mechanisms: evaporation without boiling, film boiling,
and nucleate boiling. The criterion that determines whether vaporization belongs to one of three types is the
temperature of the liquid. The mass rate of vaporization during bubble and film boiling is calculated following
a thermodynamic approach that assumes that all the heat supplied is employed to vaporize the liquid.
The mathematical model of the oxygen vaporization process is based on the first law of thermodynamics,
and leads to74
VHu˙mhc +ihc ˙mhc +iev ˙mev iout ˙mout k
dt ,(19a)
V˙mhc + ˙mev ˙mout ρdV
dt ,(19b)
Table 4: Parameters of the flip around maneuver
Parameter Value
Angular velocity, ω[deg/s] 5
Total attitude nozzle thrust force, F[N] 1000
Acceleration and deceleration phases, t[s] 9.7+9.7
Constant angular velocity phase, tω[s] 26.3
Total flip around time, tr[s] 54.7
rad +qmix-w
con +qw-lox
rad +qw-lox
con qw
rad +qext
con qmix-w
rad +qw-lox
rad +qmix-lox
con qev +Hu˙mhc
rad qmix-lox
con +qw-lox
rad +qw-lox
con qev
The system given by Eqs. 19a-e describes the change of the pressure in the tank p, density of the vapor-
gas mixture ρ, and the temperatures of the vapor-gas mixture Tmix, LOX Tlox, and tank walls Tw. The
temperature of the vapor-gas mixture depends on the radiative and convective heat flux to the wall of the tank
con ,qmix-w
rad ), radiative and convective heat flux to LOX (qw-lox
rad ,qmix-lox
con ), heat of vaporization of LOX qev,
and the heat generated during decomposition reaction of H2O2with mass of mhc,Hu˙mhc. The temperature
of the walls of the tank depends on the external heat flux from the atmospheric heating qext, heat flux from the
vapor-gas mixture in the tank qmix-w
rad and qmix-w
con , heat flux to LOX (qw-lox
rad ,qw-lox
con ), and radiative heat flux qw
The temperature of LOX is determined by the heat flux from the wall, vapor-gas mixture in the tank and heat
of vaporization of LOX.
The key heat fluxes that determine the vaporization process of LOX are now considered. The radiant heat
flux from the vapor-gas mixture to the LOX surface is
rad =σmix Flox T4
mix T4
where the area of the surface of vaporization of LOX, Flox, depends on the mechanical condition in the tank.
If LOX is accumulated at the bottom of the tank, then Flox equals the cross section area of the tank. If the
liquid oxygen is in the drop state, then Flox is the total surface of all the drops. The convective heat flux from
the vapor-gas mixture to the oxygen surface is
con =αmixFlox (Tmix Tlox ),(21)
αmix =λmixN umix
being the heat-exchange coefficient of the vapor-gas mixture to the surface of evaporation of the oxygen in the
tank, and where λmix is the thermal conductivity of the vapor-gas mixture in the tank, and lis the characteristic
dimension that is equal to the diameter of the oxygen tank. αmix depends also on the Nusselt number N umix,
which is determined by the vaporization type of LOX: evaporation from the free surface, nucleate boiling or
film boiling. qcon and qrad , among other contributors to the heat exchange process, are defined similarly to
Eqs. 21 and 22.
If free surface vaporization is the dominant vaporization mechanism, the Nusselt number can be expressed
Nuev =C1Re0.8
ev P r0.43
ev (23)
where Reev is the Reynolds number, P rev is the Prandtl number for the vapor-gas mixture, and C1is equal to
In the case of nucleate boiling, the Nusselt number becomes77
Nunuc =αd0
λ= 75(P e ·F o)0.7P r 0.2.(24)
For the nucleate boiling case, the characteristic dimension is the average diameter of the bubble departure
diameter d0, that can be considered independent on the heat flux q. Consequently, the Peclet number for heat
exchange during nucleate boiling is
P e =qd0
with an average diameter of the bubble departure diameter defined by Fritz’s equation78
d0= 1.2θrσ
where θis the contact angle in rad, σis the surface tension coefficient, gis the inertial acceleration, ρland ρv
are the densities of the liquid and gas oxygen, qis the heat flux density, ris the specific heat of evaporation,
and ais the temperature conductivity coefficient. The Fourier number for the nucleate boiling of oxygen is
F o =a
fpd0= 0.56 ρlρv
being the frequency of formation of the vapor bubbles.
Finally, in a film boiling scenario with a large liquid volume on vertical surfaces as well as on horizontal
cylinders and spheres, it is recommended to use the similarity equation
Nuf= 0.18Ra1/3,(29)
Ra =gl3
is the Rayleigh number that characterizes the behavior of a fluid with a temperature gradient, and where
is the capillarity constant, selected as a characteristic length.
Performance analysis
The alternate PGS mode, that reduces the average tank temperature in comparison with a standalone PGS
operation, is employed in the analysis. A combined operation with the tank pressurization system reduces
oxygen evaporation and preserves it for the operation of the LRE. The ratio of hydrogen peroxide and helium
mass flow is not optimized in the results presented in this work, where one of the possible combinations was
chosen. However, even this sub-optimal set point reflects the efficiency of the alternate operation of the PGS
and helium pressurization systems.
Figure 12 shows the time evolution of different variables of the PGS for the nozzle thrust values reported
in Table 4. After stage separation, the supply of hydrogen peroxide to the LOX tank leads to an operating
pressure of 4 atm in the oxygen tank. Figure 12a shows the time evolution of this variable. During the
first six seconds after separation of the stage and supply of heat to the tank, the pressure increases from 2.7
to 4 atm. As noted above, the given pressure level in the tank is maintained by the joint operation of the
PGS and the supply of helium to the LOX tank, that are set manually and may be optimized in future works.
From 9 to 95 s, the stage performs the flip-around and propellant settling maneuvers as described in Fig. 10.
Figures 12b and 12c show the mass flow rates of hydrogen peroxide and helium during this process. Figures
12f and 12e show thrust plots and mass flow rates of the PGS nozzles. During the entire operation, the gas
temperature in the tank increases up to 275 K, which does not violate the standard thermal stability limits of
300-350 K.
Table 5 shows the results of the flip around and propellant settling maneuvers separately. A total of 327.9
kg of gas are employed by the nozzles during the operation of the PGS, with 354 kg of oxygen, 95 kg
of hydrogen peroxide, and 0.7 kg of helium being consumed. After the PGS operation, the gas residuals
(composed of vaporized oxygen, water vapor, and helium) are needed to maintain the tank pressure of 4 atm
Pressure [atm]
t [s]
0 20 40 60 80 100
(a) Pressure in the oxygen tank
H2O2 mass flow rate [kg/s]
t [s]
0 20 40 60 80 100
(b) Hydrogen peroxide mass flow rate
He mass flow rate [kg/s]
t [s]
0 20 40 60 80 100
(c) Helium mass flow rate
Thrust [N]
t [s]
0 20 40 60 80 100
(d) Nozzle thrust in turning and fuel shrinking modes
Gas mass flow rate [kg/s]
t [s]
0 20 40 60 80 100
(e) Gas mass flow rate through gas jet nozzles
Tank gas temperature [K]
t [s]
0 20 40 60 80 100
(f) Gas temperature in oxidizer tank
Mass evaporation rate [kg/s]
t [s]
0 20 40 60 80 100
(g) Mass oxygen evaporation rate
Figure 12: Performance of the PGS.
Table 5: Mass from different components required for the flip around and settling maneuvers.
System Flip around maneuver Propellant settling
Gas through nozzles [kg] 22.9 305
Vaporized oxygen [kg] 53.9 304
PGS [kg] 40 -
Catalyst [kg] 0.5 -
H2O2balloon [kg] 4.6 5.8
H2O2[kg] 42 53
Helium [kg] 0.069 0.668
as a working body for the gas thrusters, that are in charge of stage orientation and stabilization during the
flight. The PGS mass budget of 40 kg includes the gas generator (5 kg), nozzles (10 kg), valves, pipes and
other hardware components. The total mass of the nozzles depends on the selected configuration: 8 nozzles
are arbitrarily selected in this work, with 2 of them having a higher thrust. The mass values are strongly
dependent on the technical decisions of the designers, who will have their own constrains and engineering
vision of the PGS, including the length of the pipes, number of valves, etc. Active catalysts (e.g. potassium
permanganate grains, silver mesh, copper, or other metals79) are used for hydrogen peroxide decomposition.
The catalyst does not lose its properties when exposed to the low temperature of the cryogenic tank or the high
temperatures of the combustion reaction. The temperature of the hydrogen peroxide decomposition products
can be controlled by tuning the hydrogen peroxide concentration,80 which allows the mass evaporation rate
to be modified when the PGS is used in combination with the pressurizer gas helium.
In order to minimize propellant residuals in the gas phase at the moment of PGS operation termination,
it is necessary to optimize the helium and hydrogen peroxide input cyclogram (e.g. using the Pontryagin
maximum principle or dynamic programming81). An optimum flow rate combination reduces the residual
gas both in the balloons and in the tank after vaporization, and also the final PGS mass.
Computations for the second stage would follow the same process as for the first stage and will be addressed
in future works.
A combination of the MP2and PGS technologies can potentially enhance the robustness and performance
of the overall propellant settling system. This hybrid approach involves a permanent magnet located at the
fuel outlet and a smaller PGS aimed at carrying out the flip-around phase and a shorter propellant settling
maneuver. The initial acceleration induced on the liquid residuals during the settling phase starts a slow
movement toward the bottom of the tank, where the magnetic force is stronger and thus able to efficiently
collect the liquid droplets.
Figure 13 depicts the time required by an LOX droplet to reach the bottom of the tank when subject to the
magnetic field generated by a 104At magnet as a function of its initial velocity and distance to the tank outlet.
Based on Table 3, such magnet would have a mass of 5.2 kg. The curves are computed from an extended
version of Eq. 10, and prove that an initial inertial kick can significantly extend the reach of the MP2system.
Initial droplet velocities of just 5 to 10 mm/s allow the magnet to collect the necessary residual propellant
mass under all engine restart configurations listed in Table 2 for both stages assuming that the droplets are
uniformly distributed in the tank volume. The 5 kN PGS nozzles, that induce accelerations of 0.23 m/s2
in the first stage, would theoretically need to operate for less than 0.05 s to induce these droplet velocities,
reducing the propellant settling window in about 40 s. The operation window of the nozzles would need to
be extended to account for transient effects and fluid-structure interactions, but this would only increase the
performance of the system. Based on Table 5, the associated PGS mass savings would be close to 60 kg,
resulting in a total hybrid system mass of 92 kg.
Max. T.
Min. T.
Max. T.
Min. T.
1st Stage
2nd Stage
0 mm/s
5 mm/s
10 mm/s
15 mm/s
20 mm/s
Time of flight [s]
Initial droplet distance to fuel outlet [m]
0 1 2 3 4 5
Figure 13: Time required by a LOX droplet to reach the bottom of the tank as a function of its initial position
and velocity when subject to a 104At coil located at the tank outlet. The minimum tank settling length
required by the single engine (S.E.), minimum thrust (Min. T.) and maximum thrust (Max. T.) configurations
is superposed.
Five novel high-risk-high-return propellant settling approaches have been explored in this work: passive
magnetic retention, magnetic recovery, active magnetic retention, propellant gasification, and hybrid mag-
netic gasification. The general advantages and disadvantages of each of them have been discussed, and
preliminary mass budgets have been derived. Table 6 summarizes and extends the main results from the
While the passive magnetic retention strategy exceeds any reasonable mass budget, the active magnetic
Table 6: Comparison of different propellant settling approaches and mass budgets for first and second stages.
Advantages Disadvantages Mass, 1st
Mass, 2nd
- Simple
- Robust
- Thoroughly studied
- Beyond current technological capabilities
- Massive unless stage operation is adapted
- Limited control → ∞ >1000
- Lightweight
- Simple
- Sensitive to fluid-structure interactions
- Slow
- Requires tank outlet redesign
- Very low TRL
104 1-10
- Lightweight
- Simple
- Limited reusability
- Requires careful tank design
- Requires tank outlet redesign
- Very low TRL
40 6
- Robust
- Provides settling and attitude control
- Fast to operate
- More traditional design
- Complex
- Sensitive to liquid movement
- Very low TRL
147 -
- Lightweight
- More robust than MP2systems
- Provides settling and attitude control
- Fast to operate
- Boosts magnet performance
- Complex
- Sensitive to liquid movement
- Requires tank outlet redesign
- Very low TRL
92 -
retention approach can potentially reduce the mass of existing propellant settling systems by one to two
orders of magnitude, leading to more than half a million dollar savings per launch and stage. This comes
at the cost of higher complexity and potential robustness and reliability issues. Although less efficient, the
magnetic recovery system is also very competitive with respect to current technologies, particularly for upper
stages. Because this approach depends on the availability of uniformly distributed free-floating propellant
droplets, fluid-structure interactions can easily undermine its performance. The PGS, which represents a
relatively more conventional approximation to the problem, can also lead to moderate mass savings that are
significantly increased when operated in combination with a magnetic retention system. As with any other
low-TRL technology, numerous technical challenges remain that can only be addressed with a more detailed
numerical and experimental analysis.
Ullage engines have been employed since the beginning of the space era and are nowadays regarded as a
robust active settling solution. However, publicly available data indicates that they also involve significant
mass and economic penalties that may be reduced with novel approaches. Such approaches must demon-
strate the same level of reliability and robustness in order to become competitive. Although the magnetic
positive positioning and propellant gasification systems introduced in this work are still in a very early stage
of development, the analysis here presented offers reasons to persevere in their development.
The accelerations experienced by each stage during separation are modeled after assuming that the process
is carried out by spring pushers with total initial force Fb, final force Fe, and stroke h. Thus, the total force
acting on the stage is
F(x12) = Fbcx12 if x12 h, 0otherwise,(32)
with c= (FbFe)/h being the total stiffness of the springs, and where the geometrical parameters of the
problem are defined in Fig. 14. After applying Newton’s second law and solving the resulting differential
equation, the duration of the maneuver becomes
carccos Fe
where m12 =m1m2/(m1+m2). The acceleration of the stage with respect to the orbital frame is
cos rc
tif tth,0otherwise.(34)
The values Fe= 0.3Fb, and h= 0.2m are adopted in this work, with Fb= 12.1kN and Fb= 4.5kN in the
first and second stages, respectively. Empty and total masses are considered for the deployed and remaining
stages with a payload mass of 15 t. The resulting propellant dispersion velocities range between 0.45 and 0.5
m/s, a range that seems to agree with observations from SpaceX’s CRS-5 mission.
Figure 14: Stage separation model
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Liquid sloshing represents a major challenge for the design and operation of space vehicles. In low-gravity environments, a highly non-linear movement can be produced due to the lack of stabilizing forces. This gives rise to significant disturbances that impact on the propulsion and attitude control systems of the spacecraft. The employment of magnetically susceptible fluids may open an interesting avenue to address this problem, but their dynamics in low gravity remain practically unexplored. The UNOOSA DropTES StELIUM project aims at filling this gap by studying the lateral sloshing of a ferrofluid solution subjected to an inhomogeneous magnetic field in microgravity. This paper describes the design process, challenges and preliminary results of the experiment, which was successfully launched at ZARM's drop tower in November 2019. The outcomes will be employed to validate the quasi-analytical models developed by the authors and set the path for the design of magnetic propellant positioning devices in space.
This paper describes a process for identifying the intact objects in orbit that (a) pose the greatest debris-generating potential risk to operational satellites or (b) would reduce the risk the most if they were removed or prevented from colliding with each other (i.e., remediated). To accomplish this, a number of diverse, international space organizations were solicited to contribute their lists of the 50 statistically-most-concerning objects. The results of the multiple algorithms are compared, a composite ranked list is provided, and the significance of the consolidated list is presented including critical assumptions and key factors in determining this “hit list.” It is found that the four primary factors used in these processes are mass, encounter rates, orbital lifetime, and proximity to operational satellites. This cooperative international assessment provides a useful ranking of the most hazardous massive derelicts in low Earth orbit as a prioritized list for remediation to (1) enhance space safety and (2) assure long-term space sustainability. This will hopefully catalyze international action in debris remediation.
The concept of reducing environmental impact from a launch vehicle after the emergency cutoff of the liquid-propellant engine is proposed. A key component of the proposed concept is an onboard propellant gasification system, installed on the rocket blocks of the launch vehicle. The onboard propellant gasification system performs forced gasification of the propellant residues in the tanks of the launch vehicle, using the catalytic decomposition of hydrogen peroxide. The vaporized propellant and pressurizing agent (helium) from the tanks are used in cold-gas thrusters that control the attitude motion of the launch vehicle and the motion of the center of mass. The onboard propellant gasification system drives the failed launch vehicle to the preselected emergency drop area with minimal environmental impact. A methodology for the selection of design parameters of the onboard propellant gasification system based on the first law of thermodynamics is developed. A simulation example is included to demonstrate the use of onboard propellant gasification system after the emergency cutoff of the main engine during the second stage of a launch vehicle.