NASA Technical Reports Server (NTRS) 20010007232: CNES-NASA Studies of the Mars Sample Return Orbiter Aerocapture Phase

11 Pages·0.66 MB·English

by  NASA Technical Reports Server (NTRS)

IAF-00-A.6.05
CNES-NASA STUDIES OF THE MARS SAMPLE RETURN
ORBITER AEOCAPTURE PHASE
H.FRAYSSE*, R.POWELL**, S.ROUSSEAU*, S.STRIEPE**
* Centre National d'Etudes Spatiales, 18, avenue Edouard Belin,
31401 Toulouse Cedex 4, France
** NASA-Langley Research Center, 8 Langley Boulevard, Hampton,
Virginia 23681-0001, USA
51st International Astronautical Congress
2-6 Oct, 20001Rio de Janeiro, Brazil
For permission to copy or republish, contact the International Astronautical Federation
3-5 Rue Mario-Nikis, 75015 Paris, France
CNES-NASA STUDIES OF THE MARS SAMPLE RETURN ORBITER AEOCAPTURE
PHASE
H.FRAYSSE*, R.POWELL**, S.ROUSSEAU*, S.STRIEPE**
*Centre National d'Etudes Spattales, 18, avenue Edouard Be(cid:0)in," 31401 l'oulouse ( 'edex 4, France
** NASA-Langley Research (_enter, 8 Langley Boulevard, Hampton, Virginia 23681-0001, USA
Langley Research Center and NASA-Johnson Space
Center. This working group is tasked with developing
and testing guidance algorithms, as well as cross-
validating CNES and NASA flight simulators for the
Mars atmospheric entry phase of this mission. The
final result will be a recommendation to CNES on the
algorithm to use, and an evaluation of the flight risks
associated with the algorithm. This paper will describe
the aerocapture phase of the MSR mission, the main
principles of the guidance algorithms that are under
development, the atmospheric entry simulators
developed tbr the evaluations, the process for the
evaluations, and preliminary results from the
evaluations.
Acronyms
AFE: Aeroassist Flight Experiment
Abstract CNES: Centre National d'Etudes Spatiales
EMCD: European Martian Climate Database
A Mars Sample Return (MSR) mission has been
FPA: Flight Path Angle
proposed as a joint CNES and NASA effort in the
/_"Altitude rate
ongoing Mars Exploration Program. The MSR
IMU: Inertial Measurement Unit
mission is designed to return the first samples of
Martian soil to Earth. The primary elements of the JPL: Jet Propulsion Laboratory
mission are a lander, rover, ascent vehicle, orbiter, and L/D: Lift-to-drag-Ratio
an Earth entry vehicle. The Orbiter has been allocated MarsGRAM: Mars Global Reference Atmosphere
Model
only 2700kg on the launch phase to perforna its part of
the mission. This mass restriction has led to the MCI: Mars Centered Inertial
MCMF: Mars Centered Mars Fixed
decision to use an aerocapture maneuver at Mars for
the orbiter. Aerocapture replaces the initial propulsive MSR: Mars Sample Return
capture maneuver with a single atmospheric pass. This NASA: National Aeronautics and Space
Administration
atmospheric pass will result in the proper apoapsis, but
a periapsis raise maneuver is required at the first 3DOF: Three Degrees Of Freedom
apoapsis. The use of aerocapture reduces the total 6DOF: Six Degrees Of Freedom
mass requirement by -45% for the same payload This
Introduction
mission will be the first to use the aerocapture
technique. Because the spacecraft is flying through the
The Mars Sample Return (MSR) has been
atmosphere, guidance algorithms must be developed proposed as a French-US mission in the framework of
that will autonomously provide the proper commands the Mars Exploration Program. Its purpose will be to
to reach the desired orbit while not violating any of
bring back Martian samples and to deploy a
the design parameters (e.g. maximum deceleration, geophysical station network (the Netlander network).
maximum heating rate, etc.). The guidance algorithm
The initial timeline had the mission occurring in the
must be robust enough to account for uncertainties in 2005 timeframe. The Mars exploration program is
delivery states, atmospheric conditions, mass
currently under review, and the current timeline for
properties, control system performance, and
the MSR mission is not known. All the comments,
aerodynamics. To study this very critical phase of the
results, scenarios, etc. that follow in this paper deal
mission, a joint CNES-NASA technical working with the originally designed MSR mission.
group has been formed. This group is composed of Two kinds of vehicles were to be used. The
atmospheric trajectory specialists fi'om CNES, NASA- US would launch one or two landers to Mars in order
Copyright (_ 2000by the authors. Published bythe American Inslituteof Aercmauticsand Astronautics,Inc.,withpermission. Released to
IAF/IAA/AIAAtopublishinallforms.
tocollecstampleasndputthemintoaMartianorbit.
addition, aerocapture requires less fuel than
TheAriane5 wouldlauncha Frenchorbiter.Its
aerobraking. Up to now, this insertion method has
missionconsistsof deliveringtheNetlandetrhen
never been used for space missions. For the MSR
collectingthesampleasndbringingthembackto orbiter, aerocapture was chosen because this provided
Earth.DuringtheMarsapproacphhaseN,etlanders the best solution for the launch mass constraint.
wouldbetargetetdotheiratmosphereicntrypoints
Initial feasibility studies were performed by
andreleaseadn,dthentheorbitewr ouldpreparfeor
CNES during the development phase A. These studies
theMarsinsertiopnhase.
indicated that the technologies required for
Any insertionaroundMarsrequiresa
aerocapture were sufficiently mature to baseline for
significandtecreasoefthespacecraafrtrivalenergy.
this mission. To further the study of this very critical
Forthismission,therequiredvelocitychangeis
phase of the mission, ajoint CNES-NASA technical
generallgyreatetrhan1.5km/storeachthedesired
working group (led by CNES) has been formed. This
apoapsis.
Traditionally, the insertion is achieved by a group is composed of atmospheric trajectory
propulsive maneuver at the periapsis of the hyperbolic specialists from CNES, NASA-Langley Research
arrival orbit to put the spacecraft in an elliptic orbit Center and NASA-Johnson Space Center. This
around the planet. Other maneuvers are then working group is tasked to review the aerocapture
pertbrmed to correct the orbital elements to attain the study hypotheses, to develop and test guidance
desired mission orbit. This second part of the insertion algorithms, as well as cross-validate CNES and NASA
has been done two ways. The first, as employed by aerocapture flight simulators for this mission The
Viking, is to use the propulsion system. The second, final result will be a recommendation on the guidance
as used by Mars Global Surveyor, and will be used by algorithm to use and an evaluation of the flight risks
the Mars Surveyor 2001 Orbiter, is aerobraking. For associated with that algorithm.
aerobraking, the spacecraft uses multiple passes In parallel to this effort, another working
through the upper atmosphere to gradually reduce the group involving CNES, ONERA and NASA-Langley
apoapsis to that required for the mission. Using this Research Center aerodynamicists and aerothermal
methodology, a significant amount of propellant is specialists has been formed to study the
saved, however a long (weeks) and complex aerothermodynamic issues related to the atmospheric
operational sequence is needed before reaching the transit of the spacecraft during aerocapture
final orbit. (simulations, wind-tunnel tests...). This paper does
Another solution is to use the aerocapture not address these very important issues.
technique. Aerocapture replaces the initial propulsive
capture maneuver with an atmospheric pass; in Aerocapture description
general, its captured orbit is closer to the desired orbit
than the captured orbit resulting from the use of
Main phases
propulsive capture followed by aerobraking, in
Amval hyperbola
AUnosphenc guided phase ;pbere ent1_r
start ,and stop of the guidance
\
i,
Atmosphere exit
Heat-slueld .iettlsonm8
I
i
I
stable orbtt
/
i
1 Penapsts rease boost
The MSR orbiter will arrive at Mars with an MSR Mars Orbit Insertion maneuvers (aerocapture
excess velocity of about 3.2 krn/s. The objective of the
2
followed by a periapsis raise maneuver) is to reach the vehicle. The principle is the same as for a reentry
following Martian orbit: vehicle such as the Atmospheric Reentry
• Apoapsis: 1400 km + 100 km Demonstrator or the Space Shuttle except the target
• Periapsis: 250 km parameter is not a landing point but orbit parameters
• Inclination: 45 deg + 0.5 deg. (apoapsis and inclination). All of these vehicles use
the lift direction to tune the vertical velocity and then
Until Mars Orbit Insertion minus 12 (MOI-12) days,
the drag deceleration level to control the reduction of
the major activity of the Mars approach phase is the
Netlanders release. From MOI-12 days until insertion, velocity. The angle that defines the lift direction is the
the activity is dedicated solely to the MOI of the bank angle (rotation angle around the aerodynamic
Orbiter. The last maneuver prior to aerocapture is velocity vector). As the bank angle magnitude
increases, the Lift vector will force the vehicle into the
planned at MOI-6 hours in order to achieve good
precision at the atmosphere entry point. This higher density portions of the atmosphere than would
maneuver will nominally be designed with a minus 12 be achieved with no bank angle change. Conversely, a
hours Orbit Determination (OD) solution. This decrease in the bank angle magnitude will direct the
solution will also be used to initialize the on-board vehicle to a lower density portion of the atmosphere.
navigation instruments. (See figure below and note that zero bank angle is Lift
The orbiter will then enter the Mars up.) The effect of bank angle on the vertical
atmosphere after having separated from its cruise acceleration is shown through the following relation
stage. The next steps are the following: (assuming a small flight path angle):
• Orbiter attitude control is commanded to the
V 2 L cos(O)
jj= ---g+
inertial attitude corresponding to the desired
R m
aerodynamic attitude at the beginning of the
guided atmospheric trajectory. where _ is the bank angle.
• Start Angle of Attack (AoA) limits (- At this time, there are two solutions to obtain
4deg.<AoA<+4deg.) and "rate damping mode" the same deceleration level: +(11. The orbit
(maximum pitch and yaw angular velocity inclination control will be performed based on the sign
limitation). Start of the guidance algorithm of the bank angle. Indeed, the lift force generates a
(drag deceleration criteria of about 0.05 m/s2). component out of the orbit plane that affects the
• Guided phase: during this phase, the guidance orbital inclination. It is important to note that there is a
computes at each guidance step the bank angle natural drift of the inclination, even if (I)= 0 or
that leads to the desired apoapsis and inclination
(1400 km and 45 deg. for the MSR mission). _) = 180 due to the out of plane drag force
Navigation data are computed from angular rate component.
and deceleration measurements provided by an
Inertial Measurement Unit (IMU). The
minimum altitude reached during this phase is
about 35 km
• End of guidance (drag deceleration criteria) and Drag
transition to inertial attitude control mode.
y: !
• Heat shield jettisoning.
• Periapsis raise maneuver at the end of the
atmospheric pass. The periapsis altitude must be Lowdensity,
lowdeceleration,
above -50 km (system requirement) after the Relative lowheat rate
aerocapture phase. A first periapsis raise
velocity
maneuver will be performed near the apoapsis
highdeceleration,
to increase the periapsis value to at least +165 high heat rate "_
km. /
• Solar panels deployment, Earth acquisition,
system health check of the orbiter and orbit
determination.
• At the next apoapsis, a second periapsis raise
maneuver will be performed to achieve the
+250 km altitude.
Aerocapture principle
The aerocapture technique is based on the use
of the aerodynamic force to slow and to control the
The aerocapture vehicle • Periapsis altitude: >-50 km
• Inclination: 45 deg +0.5deg
• To have an on-board attitude control system that
uses thrusters to control the spacecraft:
• Attitude "rate damping" mode for pitch and
yaw,
• Pitch angle control in the range [-4deg.,
+4deg.],
• Bank angle control sufficient to follow the
commanded value provided by the guidance
algorithm.
The MSR Orbiter vehicle has been allotted a mass
of 2200 kg for the insertion phase (Netlander and
The aerocapture phase leads to specific design cruise stage jettisonned).
requirements for the orbiter. These requirements
include:
CNES-NASA aerocapture study ioint
# To provide a sufficient Lift capability in the Mars
aroup mission
atmosphere to provide the required vertical
acceleration. This requirement necessitates adding Because the aerocapture phase of the MSR is
a heatshield to the spacecraft, which is shaped to critical and because it would be the first aerocapture
provide Lift. The shape chosen is based on the space mission, it was decided to form a dedicated joint
NASA Aeroassist Flight Experiment (AFE) that CNES-NASA project team.
was extensively designed and studied in the The main technical points to be jointly investigated
nineties by the NASA. are_
The aerocapture phase validation methodology
To withstand the loads due to high velocity into The flight simulation cross-validations (3DOF
the atmosphere, i.e.: and 6 DOF),
• Maximum deceleration: 3.5 g's The flight simulation hypotheses and
• Maximum heat flux: 500 kw/m 2 assumptions review,
• Maximum total heat load: 70 MJ/m 2 The guidance algorithm development,
• Maximum dynamic pressure: 4.8 kPa The atmospheric models comparisons,
The stress cases development and simulation,
This requirement leads to a heatshield design that The guidance algorithms comparison and
protects the orbiter from thermal loads and selection (criteria, recommendation).
dynamic pressure. This heatshield should be These points are described in the following
jettisoned after aerocapture to minimize the paragraphs.
propellant use during on-orbit maneuvers, and
must give a sufficient Lift-to-Drag ratio capability Trajectory simulations
to the orbiter, as well as dissipate the thermal
energy imparted to the spacecraft. The NASA-LaRC 3DOF/6DOF simulation
software is based on simulations developed for the
• The spacecraft center of gravity (COG) must
Mars 2001 aerocapture and entry studies, as well as
provide atrim AoA in the range [-2deg, +2deg].
Mars Pathfinder and Mars Polar Lander. (Note the
Mars 2001 orbiter mission has been modified to use
To have an on-board navigation system that
aerobraking and not aerocapture.) The CNES
provides the guidance algorithm and attitude simulation software was derived from the
control system the current attitude, position and
Atmospheric Reentry Demonstrator trajectory
velocity of the orbiter. This information is
software that was developed from earlier Hermes
computed by the integration of gyroscope and
project studies.
accelerometer measurements.
These simulators include Guidance, on-Board
Navigation and Control functions and allow Monte-
To have an on-board guidance algorithm that
Carlo statistic analyses. The 3DOF simulator principle
leads the orbiter to the desired orbit upon exiting
is presented below:
the atmosphere. The system specification for this
orbit is
Apoapsis altitude: 1400 km +100km
4
¢C' roll dlrectlo_
* A numerical predictor-corrector (NASA-
LaRC).
On the CNES side, based on the ARD experience and
on the American and Russian work on the aerocapture
guidance, two algorithms were developed:
F =rna • An analytical predictor-corrector,
DCM Xr111o'11
Integration "_ o A numerical predictor-corrector.
performeO in|
inert_alfram e]
i
All of these guidance principles use the absolute value
Actual Estimated of the bank angle to control the "in-plane trajectory"
x y
(orbit apoapsis) and the sign of the bank angle to
,¢_ix m
control the "out-of-plane trajectory" (orbit
inclination). A bank reversal occurs whenever the
commanded bank angle changes sign for inclination
As a first task, cross-validation of the NASA and control.
CNES 3DOF trajectory simulation software for the
The CNES analytical predictor-corrector is
MSR aerocapture was done. Test cases including
very close to the NASA-JSC version because both are
guidance, control and navigation functions were
based on the AFE studies. The principle of these
jointly defined and performed independently on each
algorithms is for in-plane guidance (apoapsis control)
simulation. Except for tests including the navigation
to focus on the vertical velocity and the drag
function that are not yet completed, all the cross-tests
deceleration of the vehicle during the atmospheric
give very good results.
trajectory (in-plane guidance to control the apoapsis).
As the atmospheric model for the MSR
The trajectory tails into two pans. The purpose of the
studies, CNES use the European Martian Climate
first part of the trajectory (referred to as capture phase,
Database (EMCD) developed for ESA by the
or inbound trajectory) is to control the vehicle
Laboratoire de Meteorologie Dynamique (Paris) and
deceleration until aerodynamic equilibrium is
the Oxford University. For the Martian studies, NASA
achieved (when aerodynamic lift force balances the
uses the Mars-Gram Model. For these common work
inertial and gravitational forces). The desired
and comparisons, CNES has implemented the Mars-
aerodynamic equilibrium is designed to maximize
Gram model in its simulator and NASA has
control margins. After the vehicle slows to a certain
implemented the EMCD model in its simulator. Then,
velocity, the second phase is initiated. The purpose of
CNES and NASA use both models to perform all the
this phase is to benefit from the equilibrium glide
MSR aerocapture simulations. The currently
condition to target the desired orbit (outbound
implemented versions are the 3.8 for Mars-Gram and
trajectory). The atmosphere exit conditions (target
2.3 for EMCD
orbit) are predicted analytically and corrected
assuming a constant vertical velocity from the current
Guidance algorithms
altitude to the altitude at the exit of the atmosphere.
For the MSR mission CNES has decided to This vertical velocity is used by the guidance to
develop the orbiter aerocapture guidance algorithms determine the magnitude of the bank angle. This
until the end of phase B of the orbiter development. At algorithm does not need any pre-loaded on-board
that time, the principal guidance algorithm will be trajectory and has shown good performance during
specified to the orbiter contractor. That is, CNES will testing. As CNES and NASA principles as well as
choose, develop and validate the guidance algorithm results were very close with this algorithm, it was
during phases A and B of the orbiter development. decided to jointly work on this principle to develop the
About ten years ago NASA extensively studied an best possible version for MSR (cfRef I).
aerocapture demonstration around the Earth: the The principle of the numerical predictor-
Aeroassist Flight Experiment. At this time, multiple corrector is to determine at each guidance step the
guidance algorithms were developed by the NASA constant bank angle value to be applied to reach the
teams. target apoapsis. The prediction of the trajectory with
Three years ago, an important NASA effort this constant bank angle is performed on-board by a
was done to develop aerocapture guidance algorithms numerical integration of the equations of motion until
for the early 2001 orbiter mission. Some of them were the exit of the atmosphere.
derived from the AFE studies. Three algorithms This kind of method allows the coupling of
developed for the early 2001 orbiter were proposed by the "in-plane" and "out of plane" guidance due to the
NASA for the MSR aerocapture algorithm: on-board integration of the trajectory. This
• An analytical predictor-corrector (NASA-JSC), methodology was done in CNES version using
• A terminal control algorithm ("Apollo derived", Russian work on that method (Keldysh Institute of
NASA-JSC and NASA-LaRC), Applied Mathematics). For the NASA numerical
5
predictor-correcatlogrorithmth, ebank-reverstiamle
influence functions provide a means to calculate a new
wasdeterminebdaseodnwhentheinclinatiolnimits
trajectory "on-the-fly" when dispersions relative to the
wereexceedeadn,dthenthereversawlascommanded
nominal trajectory are encountered. This new
whenthattimewasreachedA.lltheotheralgorithms
trajectory will theoretically have the same in-plane
performabank-reversthaelinstanitnclinationlimits
terminal conditions as the nominal trajectory; hence,
areviolated.
this type of guidance scheme is referred to as a
TheTerminal Point guidance algorithm uses terminal point control algorithm. As indicated above,
the calculus of variations optimal control approach out-of-plane (inclination) control is achieved by
with influence coefficients to control a vehicle about a monitoring inclination and commanding a reversal
nominal trajectory. Influence coefficients are solved when limits are exceeded.
from a set of differential equations adjoint to the As a common subroutine interface for all the guidance
linearized perturbations of the equations of motion subroutines, the following parameters are used as
about a nominal trajectory. Conceptually, the INPUT/OUTPUT:
INPUTS
Name Description Unit T_pe
time time since simulation start s real*8 time
bank current estimate of vehicle bank an/fie rad real*8 bank
pos abs current estimated vehicle position in MCI coordinates in real*8 pos abst3 )
vit abs current estimated vehicle inertial velocit_ in MCI coordinates tWs real*8 vit abs(3)
pos,,rel current estimated vehicle position in MCMF coordinates m real*8 pos, rel(3 )
current estimated vehicle relative velocity in MCMF
vit tel m/s real*8 vit_rel(3)
coordinates
ace current sensed acceleration in body coordinates flame m/s 2 real*8 ace(,3)
mci2bodl¢ MCI to bod_,"coordinate system transformation matrix real* 8 mci2bod_'(3,3)
mci2mcmf MCI to MCMF coordinate system transformation matrix real*8 mci2mcmt_3_3)
OUTPUTS
Name Description Unit _e
bank com commanded bank anl_le rad real*8 bank com
rotation option flag that specifies the way of achieving the
commanded bank angle :
Q 0 to roll through the smallest angle to reach commanded
idir Integer idir
angle (shortest path),
[] 1to roll in a positive sense,
[] -I to roll ina nel_ative sense
Aerocapture simulation hypotheses FI)A
For the aerocapture simulations, the insertion Vclncity
conditions of a nominal 2005 mission have been used.
The excess velocity of the arrival hyperbola is about
3.2-km/s; the targeted inclination is 45 deg. The
nominal targeted Flight Path Angle (FPA) has been
defined as the middle of the theoretical entry corridor
(about 1.2 deg width). This corresponds to a targeted
FPA of-10.5 deg. This value may be optimized in the
future, depending, for example, on the exact arrival
day.
6
Corridor definitions: theoretical range). This reduction depends on the
1 - Theoretical corridor: FPA range defined guidance algorithm principles and tuning.
by the steepest and shallowest FPA that reach the 3 - $_trvivability corridor: this FPA range in
desired apoapsis. The steepest value corresponds to a which the mission is not lost (survivability criteria to
full lift-up trajectory. The shallowest value be defined). For example, apoapsis between 250 km
corresponds to a full lift-down trajectory. This and 2000 km, as well as maximum loads up to orbiter
corridor is linked to the Lit_-to-Drag ratio capability of design values with no margin would be considered
the vehicle (that is, the corridor is linked to the acceptable.
orbiter's performance). This corridor depends slightly
on the atmospheric model that is used to compute the The other computation assumptions are:
trajectories. It also depends slightly on the arrival date • Maximum bank rate of 20 deg/s
(due to the excess velocity value). It is computed with • On-board navigation initial position
nominal aerodynamic characteristics and nominal corresponding to the nominal entry point
atmosphere (both scenarios when EMCD is used). determined at 12 hours prior to atmospheric
2 - Corridor with guidance limitation: due to entry
the guidance response to the FPA dispersions, the • Perfect initial attitude
previous corridor may be decreased somewhat (that is,
the guidance may not be able to use the whole
MSR Orbiter Atmospheric Mission Dispersions
Quantity 3 6 Nominal Value Distribution 3-c or
DOF DOF Type min/max
Mission Uncertainty
Initial Bank, deg • • -90 or +90 Gaussian 5.0
Close window Initial inertial FPA, deg ° ° -10.5 Gaussian 0.4
Close window Initial inertial FPA, deg ° • -10.5 Gaussian 0.5
Close window Initial inertial FPA, deg ° • -10.5 Gaussian 0.6
Close window Initial inertial FPA, deg ° ° -10.5 Gaussian 0.8
Open window Initial inertial FPA, deg • • -10.5 Gaussian 0.4
Aerodynamic Uncertainty
Trim Angle of Attack Incr, deg ° Nat. trim" Uniform 2.0
Axial Force Coeff Increment • ° Uniform
nominal 10 %
Normal Force Coeff Increment • ° Correlated
Mass Property Uncertainty
Mass, kg * * 2200.0 Gaussian 20.0
X CG position, m ° ° -1.24432 m Gaussian 5E-03
Y CG position, m ° 0 Gaussian 5E-03
Z CG position, m • ° 0.74335 m Gaussian 5E-03
Atmospheric Uncertainty
Initial Seed Value • ° 0 Uniform 1/29999
Scenario for EMCD ° ° "clear"&"Viking" Small and large scale
Tau value for Mars-Gram • • TBD Uniform TBD
Control System Uncertainty
Bank Acceleration, deg/sec" ° • 5.6 Uniform _O",o/,
IMU Uncertainty
Initial angular misalignment, grad • ° 0 Gaussian 200
gyro bias drift, grad/hr • • 0 Gaussian 500
gyro scale factor, ppm ° • 0 Gaussian 60
gyro nonorthogonality, ppm • ° 0 Uniform 20
gyro random walk (PSD), deg/rt-hr • • 0 Gaussian 0.07
accelerometer bias, milligees • ° 0 Gaussian 180
accelerometer scale factor, ppm • • 0 Uniform 100
accelero scale factor asymmetry, ppm ° • 0 Uniform 50
Accelerometer nonlinearity, gg/g2 . ° 0 Uniform 10
accelerometer random noise ,fps/qh " ° 0 Gaussian TBD
accelero input axis nonorthogo., arc-s • ° 0 Uniform 20
• -4 to +4 deg absolute limit
• delta V cost to correct the insertion orbit (linked
Stress cases
mainly to the periapsis altitude after
To test robustness of the guidance algorithms, aerocapture)
some "stress cases" have been defined and will be • number of source lines of code
simulated: - number of parameters to be stored on-board
t .6 sigma IIVlUgyro initialization error, • Complexity of the source codes (numbers of
2-Extreme-density shear/wave, loops, tests, etc.)
3 •10 %-uncorrelated variation of CA and CN
inaddition to the 10% correlated variation, First results
4• Accelerometer signal degradation,
The first round of Monte-Carlo simulations performed
5• Gyroscope signal degradation,
with a FPA uncertainty up to +0.4 deg 3_ has
6.Single roll-thruster failure,
7-Computer swaps during aeropass. demonstrated that all the proposed algorithms are able
Other stress cases may be added in the near future to reach the specifications of precision orbital delivery
considering the results of a Failure Mode Analysis and delta V cost to correct the periapsis at 3 sigma• As
relative to the aerocapture phase of the MSR mission. typical results, the following figures show Monte-
Carlo results obtained with an analytical predictor-
corrector guidance algorithm (on-board navigation
errors included), and Monte-Carlo results obtained
Algorithms comparison criteria
with a numerical predictor-corrector (on-board
In order to choose the best guidance navigation errors not included).
algorithm for the MSR mission, an initial list of Note that the NASA numerical predictor-corrector
comparison criteria has been defined (other criteria algorithm was developed using the Mars-GRAM
may be added): atmosphere model• Further work to tune the algorithm
• guidance precision and robustness to for both the Mars-GRAM and EMCD will reduce the
uncertainties (Monte-Carlo results) delta V and inclination error.
• robustness to stress cases
Peak Heat Rate DV 1400 250 _ 45
<=,
r,._- IC_, t_C, X:< ./'" ,"
li_clinat.l )n Apoaps is Altitude Per iapsis Altitude
!ii i i'
c, '
lnc Hap Hper Heat Rate Heat load G loads Pdyn DeltaV
deg. km. kin. kw/m z Mj/m z Kpa. (n_s)
Mean 45.04 1416.7 25.9 394.5 50•2 2.30 3.0 59.9
St Dev 0.13 31.74 7.57 16.98 1•84 0.I1 0•24 5.40
3000 runs with EMCD and Mars-Gram using analytic predictor-corrector
8
NASA PC Guidance: 3DOF MSRO NASAJCNES Simulation Results
Closed End of Launch Window: ._.4 °Initial _ Dispersions
180
::;:_ _: _:: o2200;
3
_-2ooo_
i
1,oop
44.5 45.0 45,5 < 0 10 20 30 40 50 6O
Inclination (deg) Periapsis Altitude (km)
Total Number ofCases: 2000 600 • - ,
100% betwee_c: 44.55 _'and 45.51'4
D.: 50.67° and 51.19_ o°soo_
AV: 46.58 and 163.67 m/s
hp: 19.7 and 46.3 km _4oo
L@
99.7 %between stated performance requirements_ 3ooi
100 centile cases are characterized by:
Max QBAR between 2,01 and4.04 kPa "6200
Max Heat Rate between 318 and 448 kW/m 2 z_
Max Tot Heat Load betwe4¢,.75 and 63.11 MJ/m ,- loo
Peak g Load between 1.43 and 2.86 Earth g'_ I
Z
60 80 100 120 140 180
V (m/s)
2000 runs with EMCD for the NASA numerical predictor-correcter
NASA PC Guidance: 3DOF MSRO Simulation Results
Closed End of Launch Window: ._.4 ° Initial fpa Dispersions
110 • ........... o_16oo
o
o t550
_100_
_ go
_E_1450
__ 1400
--I
_70_ _. 1350
<
.__ 1300
>_ 60[ ¢)
_0-1250
5O ,,,.s.... 45.o.... 45.5 _.12o0 (o 2o _o _ ._ so
<
Inclination (d_) Periapsis Altitude (km)
Total Number of Cases: 2000 4oo
100% betwee_c: 44.53 ° and 45.46 '_ _ 35o
AV: 50.06 and 59.74 m/s = 30o
hp: 10.3 and 43.5 km _25o
o
88.2 '/obetween stated performance requirements_
100 centile cases are characterized by: L)_
Max QBAR between 2.69 and 4".89kPa "6 "_
Max Heat Rate between 376 and 491 kW/m 2 $ lo0t
Max Tot Heat Load betwe_.67 and 61.07 MJ/m@ [
Peak g Load between 1.86 and 3.45 Earth g's_ 5oI
z 0i _ -
45 50 55 60 65 70 7"5 8O
V (m/s)
2000 with Mars-GRAM for the NASA numerical predictor-correcter
9