Research on Key Issues for Mars Landing Missions

DONG Jie,SUN Zezhou,RAO Wei,DONG Tianshu

Beijing Institute of Spacecraft System Engineering,Beijing 100094

Abstract:The environmental conditions of Mars landing missions are much different from that of Earth reentry missions,including the distance between the spacecraft and the Earth,the atmosphere,landform,aerodynamic forces,thermal effects.Therefore there are more autonomy requirements,more flight phases,more uncertainties in modeling and analysis,and more difficulties in validation of the flight sequence.Firstly the challenges of Mars landing missions are summarized in this paper.Then the key issues for Mars landing missions are analyzed according to the phases of the landing process.Finally suggestions and proposals for further development for Mars landing technologies are given.

Key words:Mars,EDL,MSL,parachute,simulation

1 INTRODUCTION

Although the landing on Mars is similar to the Earth reentry,the entire landing process for a landing on Mars requires better deceleration performance with more complex phases and more tense timing due to the different atmospheric composition and physical properties and hence larger uncertainty.

The entry,descent,and landing (EDL) process of a successfully implemented Mars landing mission is mainly divided into four phases (aerodynamic deceleration phase,parachute phase,power descent phase and the landing buffer phase).The uncertainty of various parameters and design constraints will affect the accomplishment of each phase and the switching of key links.So far,the worldwide success rate of Mars surface landing missions has been only about 50%.The mission has extreme high risk and hence strict requirements for spacecraft design.The challenges for Mars landing exploration missions include:

1) The relatively long distance between Earth and Mars

The communications delay between the probe and Earth is large,so the ground station does not have the ability to intervene in the EDL process,hence it must be executed autonomously.

2) The thin atmosphere of Mars

Compared with a deceleration landing on Earth,the same payload landing on Mars requires a larger diameter structure.In order to achieve a safe landing,a retrorocket and landing gear are needed since a steady descent speed has not been reached after the Mars entry probe decelerated due to its aerodynamic configuration and a parachute.In addition,the terrain conditions should meet the elevation requirements for sufficient deceleration.

3) Uncertainty of environmental parameters

At present,there is no coherent measurement data for the Mars atmosphere model.All the data sources are spread over several successful exploration missions conducted by the United States,which have been rare.Therefore,the uncertainty is larger than that of the Earth atmosphere,which affects the theoretical analysis and engineering modeling.

4) Complex aerodynamic forces/thermal conditions

Due to the interaction with the flow field,the force/thermal effects during the flight of entry probe in the Mars atmosphere,separation and engine operation need to be specifically designed.

5) Difficulty in ground verification

It is difficult to fully simulate the Mars environment and predict the actual conditions the entry probe will encounter in the ground environment.Only a full-system simulation analysis and semi-physical verification can be carried out through a mathematical model which is based on the comprehensive data of various individual tests.

In view of the challenges and technical status of the Mars landing mission,along with design experience achieved in other countries’ mission development,this paper summarizes the key issues that need to be carefully considered for each phase of the EDL process,plus provides suggestions for future development.

2 AERODYNAMIC DECELERATION PHASE

2.1 Entry Approaches for Mars

In principle there are two main ways of entering Mars:ballistic and ballistic-liftbased on the technical approach.The different entry methods make different design requirements such as system function,configuration,overload,thermal protection,parachute deployment height,and the landing accuracy of the entire entry system.Strictly speaking,among the Mars landing missions that have been successfully implemented,the Mars Science Laboratory (MSL) and Perseverance landed on Mars using the ballistic-lift method,while the Mars Pathfinder,Mars Exploration Rover (MER) and Phoenix landed on Mars using the ballistic method.The biased center of mass approach was designed for the Viking mission to enable a certain lift-to-drag ratio,but the landing position was not controlled by guidance so the landing distribution was large.The method of ballistic-lift is better than ballistic,as it can better adapt to limited off-nominal conditions encountered in the environment and ensure suitable parachute deployment conditions such as height.

As to the ballistic-lift method,in order to further improve the accuracy of the landing point and control any overload and heat flow conditions during the entry process,it is necessary to adopt a guidance method,including the standard trajectory method and the predictive guidance method.For MSL and Perseverance the standard trajectory method was applied,which was a control method improved from the Apollo’s final reentry guidance algorithm.For this method,the range deviation is controlled by adjusting the bank angle,while the adjustment amount of the bank angle is obtained by feedback of the deviation of the range,altitude rate change and drag acceleration relative to the nominal trajectory.The corresponding attitude control strategy is to control the bank angle through attitude control in the roll direction according to the guidance,at the same time,conduct angular velocity damping in the pitch and yaw directions to ensure that the attitude angle is near the trim angle of attack.

2.2 Aerodynamic Configuration

The requirements for aerodynamic performance for a Mars entry probe are significantly different from those for an equivalent Earth reentry.This is due to the different atmospheric composition at each stage of the entry process,different gas vibration characteristic temperatures,which changes the size and distribution of hypersonic aerodynamic forces.However,no matter which entry method is applied,the basic configuration of the heat shield is a spherical conical large blunt body(non-spherical crown configuration) with a half cone angle of 70°.The main characteristic of this configuration is the dynamic instability in the transonic phase.It is also necessary to fully evaluate the impact on the parachute opening and associated steps.For the entry probes with an attitude control ability,angular velocity damping control is usually required to reduce the aerodynamic interference torque.The rear body adopts an inverted cone,which include a single cone,double cone,triple cone,configurations,selected based on the layout of the equipment in the module.The design of heat shield adopted from the Viking mission was inherited by subsequent successful NASA Mars landing missions.

2.3 Aero thermal Condition

The Mars landing probe enters a special aero thermal environment dominated by carbon dioxide dissociation products during the hypersonic phase.The thermal flow conditions of the entire entry process are first related to the entry method.The total heat generated during the ballistic-lift entry is relatively large,but the peak heat flow during the ballistic entry is more severe.

In addition,the aero thermal conditions are also closely related to the existence of transitions.The Mars atmosphere is relatively thin,and the inlet flow Reynolds number is small.The early Viking and MER could be designed according to the laminar flow due to the low entry speed.Although the Mars Pathfinderand Phoenixmissions had turbulent flow,their sizes and initial angles of attack were small,so the effect of heat flow was limited.When the entry mass and configuration are large,the entry velocity is high,and the angle of attack is large,the effect of turbulence is more significant.For example,in the development of a typical MSL mission,numerical calculations and ground wind tunnel tests have verified that at the time of peak aerodynamic heat generation,the inlet flow will transit to turbulence on the surface of heat shield,which leads to a high heat flow and large shear force.Thus,the following design was adopted:

A total of 113 Phenolic Impregnated Carbon Ablator (PICA)heat-proof blocks with 7 different features were adopted,and the heat shield was divided into four areas including the shoulder,transition area,tapered surface and top,which are seamed according to the airflow direction.The difficulty of PICA is that it can only be combined and spliced by sheet materials,the weight of the heat shield is relatively large considering the gaps and interference between the sheets.

In addition,it is necessary to comprehensively analyze constraints such as the initial atmospheric entry angle,bank angle,and the aerodynamic deceleration time in the design of GNC system,in order to control the total heat generated,the peak heat flow during the aerodynamic deceleration process,and to reduce the difficulty of designing the thermal protection system.

2.4 Special Aerodynamic and Aero thermal Problems

2.4.1 Jet flow efficiency

The entry probe is in a dynamic instability state and the total angle of attack fluctuates obviously near the parachute deployment.It’s hard to ensure the total angle of attack required for the parachute deployment without attitude control.In addition,the bank angle needs to be adjusted for the ballistic-lift entry method.The entry probes for Phoenix and MSL missions were both equipped with propulsion systems for attitude control.

However,CFD simulations of the above missions all show that when some thrusters are working,they will interact with the surrounding flow field,especially when the thruster jet direction is in the wake region of the rear body.The jet flow efficiency is significantly reduced and the control torque is reversed,which will seriously affect the control effect.

The problem of the Phoenix mission was that the attitude control thrust was small,while that of the MSL mission was the adoption of a different thruster orientation.Therefore,the Phoenix controlled the dynamic threshold through relaxing the attitude without using the thruster,while the MSL changed the piping layout and corresponding structure of the propulsion system to adapt a different thruster orientation.

2.4.2 Fluctuating pulsation conditions

The advantage of adopting a large blunt configuration is that it has high resistance during the Mars hypersonic entry process,which is conducive to the completion of aerodynamic deceleration.Due to the blunt body flow around the shoulder,the fluid will be separated after bypassing the shoulder,forming a separation zone on the windward side such as the tail end of the cabin body.In the separation zone,the eddy current pulsates violently,resulting in a high pulsation pressure when transiting the wall.Its frequency range includes the resonance frequency of the common structural plate.There are two main effects:

1) The pulsation pressure will cause a large local load on the wall of the probe,which can easily create a chattering response of the structure;

2) As a random excitation,pulsation pressure is transformed into cabin noise through noise transmission and structural resonance,which is reflected on the installation surface of the cabin equipment and affects the working performance to a certain extent.It will affect the measurement output of key navigation sensors such as Inertial Measurement Unit (IMU),thus effecting the navigation results.It is necessary to evaluate the impact,optimize the installation configuration,and carry out vibration reduction or vibration isolation for the IMU.

2.4.3 Local thermal environment of the structure during thruster ignition

The thruster interacts with the surrounding gas flow field during ignition,which will cause heating of the local structure.It was found that the thruster ignition will bring unbearable aero thermal conditions to the surrounding baffle in the MSL mission.Therefore the attitude control strategy of GNC system has to be optimized to reduce the thruster ignition operation time,thus controlling the temperature.

2.5 Parachute Deployment

2.5.1 Deployment constraints

The constraints need to be considered in the design of the parachute deployment are shown in Table 1.

Table 1 Constraint conditions of parachute deployment

Table 2 Angular velocity control threshold of MSL

2.5.2 Design of parachute deployment control method

Current parachute deployment methods mainly include the time control method,overload-time control method,static pressure height control method,radar height control method,and an adaptive control method.

Control methods based on the Earth environment such as time control,overload-time control and static pressure height control methods cannot ensure that the height and dynamic pressure of the parachute are controlled within the required range due to the large uncertainty of parameters such as in the Mars atmospheric model.

Applying the radar height control method requires the installation of a radar on the heat shield,which will have high requirements for a thermal protection structure.Only the Viking lander applied this method at present.

The adaptive control method takes parameters such as dynamic pressure,Mach number,height,and speed as the control criteria (the premise is that the probe enters a certain range of Mach number and dynamic pressure),so dynamically determining the deployment time by identifying the entry trajectory characteristics through a preset control law from the acceleration measured by an accelerometer.This approach has been adopted in the successful landings during the Viking missions.As an alternative,the MSL uses the ballistic-lift method,hence the flight attitude is higher while the dynamic pressure is lower than those for Viking with the same Mach number.Therefore the Mach trigger is selected.

In the aerodynamic descent phase,the probe usually only relies on the IMU combined with gravity calculation to perform navigation calculation in the inertial coordinate system.The navigation error will accumulate with time due to the influence of the initial navigation error (given by the ground station measurement),gyro,accelerometer drift errors,etc.,increasing the landing point dispersion error.In addition,the probe can only obtain the speed data based on the IMU,and cannot measure environmental parameters such as wind speed,thus,there will be error in the calculation of the Mach number.Usually,the total deviation of the Mach number calculation is within ±0.3 Ma.

2.5.3 Angle of attack before parachute deployment

For ballistic-lift entry,since the probe is always attached to the standard trim angle of attack,the angle of attack needs to return to zero before deployment to ensure correct parachute deployment conditions.There are currently two implementation methods:

1) The internal mass of the ejection adjusts into the center of mass of the capsule

In the MSL mission,2 mass blocks on one side were ejected 9 minutes before the Mars entry to offset the center of mass to establish a balanced angle of attack.6 mass blocks on the otherside were ejected before the parachute deployment to make the center of mass return to zero,then the angle of attack returns to zero.The disadvantages of mass blocks are that their mass accounts for about 30% of the effective landing mass,and the safety of the mass blocks separation needs to be considered to prevent re-collision.

2) Configure trim wings to change flow field before deployment

NASA conducted research on trim wings after the MSL mission.The entry probe is an asymmetric body after the trim tab is unfolded,it realizes the trim angle for attack of the aerodynamic deceleration segment through aerodynamic action.Before deploying the parachute,the trim wings are retracted and restored to a symmetrical body to make the trim angle of attack to zero.The advantage is that the effective landing mass can be increased.The trim tab has only been applied for the Tianwen 1 mission.

3 PARACHUTE DESCENT PHASE

3.1 Dynamic Environment for Parachute Deployment

According to NASA’s Mars parachute dynamics model,the parachute deployment force is not only related to atmospheric density,speed,and aerodynamic parameters,but also related to the dynamic environment of parachute deployment influenced by several factors such as surge.The parachute deployment force will fluctuate significantly due to over-inflation of the parachute and area oscillation especially when the Mach number is above 1.4.

Without the above consideration,the upper threshold of angular velocity was not set to the maximum value in the design of ExoMars 2016 lander,the yaw angular velocity continued to exceed the design range (-187°/s) for 1s,resulting in gyro saturation.The calculated navigation altitude on the lander was too low,and the deceleration engine was turned off prematurely,causing the lander crash.In addition,due to the lander’s lack of autonomous fault handling,it didn’t formulate system reconfiguration measures to ensure a safe landing on the basis of various measurement data such as radar altimeter measurement.

At present,research on an attitude determination scheme based on microwave ranging data has been carried out abroad,and it has been verified by helicopter flight.This scheme can obtain the attitude relative to the local horizontal direction.The local terrain fluctuation will create certain errors in the measurement results,but the accuracy of the attitude determination can meet the requirements.The disadvantage of this scheme is that it cannot obtain the horizontal attitude and pointing.

3.2 Heat Shield Separation

At present,there are two separation triggering methods,based on the Mach number or time delay.The United State’s successful Mars landing missions before MSL were all triggered by a fixed time delay based on the time the parachute deployed,while the MSL was triggered based on the Mach number.After the heat shield separated from the Mars entry vehicle and entered a safe distance,the entry vehicle deployed the landing buffer mechanism,then the microwave radar started to measure.

The heat shield separation needs to meet the following conditions:

1) Separation safety

The separation mechanism needs to create a certain initial separation speed to avoid the near-field flow field characteristics that make the heat shield return and collide with the entry vehicle.There is a sufficient ballistic coefficient difference between the entry/parachute combination and the heat shield to ensure a far-field positive separation state,because the resistance coefficient of the two significantly changes when the Mach number is below 0.8.

2) Effectiveness of microwave radar measurements

After the heat shield separation and before the back shell separation,it is necessary to ensure that the microwave radar has enough time to measure the Mars surface considering the calculation convergence process where the entry vehicle modifies the IMU using microwave measurement data.At the same time,the time of radar measurement should not be too early,since the radar beam will damage the heat shield or return wrong ranging data if the heat shield is close to the entry vehicle,affecting the navigation.

3.3 Back Shell Separation and Avoidance

3.3.1 Back shell separation

The entry vehicle calculated the appropriate height to trigger the separation of the back shell and started the power descent according to the estimated vertical speed.The nominal separation speed of the MSL mission was 80 m/s and the height was 1.6 km.After the entry vehicle entered a specific height and speed range,the relative nominal value of the separation height was appropriately increased when the predicted relative nominal value of vertical speed was too large,avoiding insufficient height allowance in the deceleration process;the relative nominal value of the separation height was appropriately reduced when the predicted relative nominal value of vertical speed was too small,avoiding additional propellant consumption.

In order to ensure the safe separation,the following engine control strategies were implemented after the MSL back shell separation:

Free fall within 1 s to prevent short-term collision;

At 1 s,the thrust of the orbit control engine increased to 20%,and the engine warmed up;

At 1.2 s,the attitude control executed rate damping,and the attitude maneuver control executed back shell avoidance at the same time.

At 3.4 s,it conducted powered descent guidance and control.

3.3.2 Back shell avoidance

After the back shell separation,back shell avoidance was required to prevent the parachute and back shell combination from contacting the lander again.At present,there are two methods:avoidance in orbital plane and perpendicular to the orbital plane.

Both the Phoenix missionand ExoMars 2016 lander missionadopted orbital plane avoidance,that is,when the horizontal velocity was less than a certain threshold,an attitude adjustment of the whole vehicle was performed to generate acceleration in the opposite direction of the horizontal velocity to achieve avoidance.

The MSL mission adopted the avoidance perpendicular to the orbital plane,that is,to prevent collision without carrying out speed judgment,maneuvering 300 m outside the orbital motion plane.Compared with orbital plane avoidance,the propellant consumption of the avoidance perpendicular to the orbital plane was slightly increased,with the advantage of simple logic and high success rate of avoidance.

3.4 Attitude Control

The separation safety of the heat shield/back shell and the microwave radar beam pointing had certain requirements on the attitude of the entry vehicle.Due to the influence of the dynamic environment after deployment,the angular velocity and swing angle of the entry vehicle were large in the initial stage of deployment,so angular velocity damping control was needed.At the same time,as part of the MSL mission analysis it was found that in order to prevent the possible ablative damage to the parachute caused by the attitude control plume,the attitude control jet frequency should be reduced as much as possible.

The MSL satisfied the above constraints by designing attitude and start control strategies and angular velocity control thresholds.

Parachute deployment:attitude control was prohibited when deploying,and it started 10 s after the parachute deployment.

Heat shield separation:attitude control was prohibited during the heat shield separation,and it started 3 s after the separation.

Back shell separation:attitude control was prohibited during the separation,and it started 1 s after the separation.

4 POWER DESCENT PHASE

4.1 Engine Shutdown

The strategies for engine shutdown before landing mainly include shutting down before/after touching the Mars surface.Currently,only the ExoMars 2016 lander is designed to shut down before touching the Mars surface,the Viking,Phoenix (including the failed Mars Polar Lander mission) and the MSL mission all shut down the engine after touching the surface.

The ExoMars 2016 mission selected to shut down the engine after the center of mass of the lander was 2 m away from the Mars surface,and the vertical speed was not greater than 4 m/s when it touched the surface.This method has high requirements on the accuracy of GNC navigation altitude at the end of landing.

The Viking and Phoenix missions shut down the engines based on the first landing footpad signal.This method can guarantee the touchdown signal through the redundant backup of multi-foot pad signals.

After the MSL rover started to touch the Mars surface,the engine thrust gradually decreasd under the control law.It was considered that the rover landed on the Mars when the engine thrust had been reduced to about half.The aerial crane cut off the cable of the crane,conducted avoidance maneuvering up to a distance of 600 m to prevent collision with the rover.This method is customized for an aerial crane,and depends on the accuracy of IMU measurement.

4.2 Landing Buffer

There are mainly four techniques used as a landing buffer,including using landing legs,airbag,collapsible structure and aerial crane.

Airbags were applied in the Pathfinder,Spirit and Opportunity missions,which had strong adaptability to terrain and initial touching speed,but had great limitations in landing quality,they also required many related links and were a complex design.

The collapsible structure alternative was simple in form and used on the ExoMars 2016 lander,absorbing energy through material deformation.It is also mainly applicable to the case of a small landing mass.

Landing legs were used on the Viking,Phoenixand InSight missions.The load-bearing capacity is higher than that of the airbag,but it has high requirements for landing terrain and stability,and it also needs to take into account the envelope of the capsule.

The aerial crane method applied to the MSL mission,precisely controls the soft landing of the suspension and makes up for the deficiencies of the airbag and landing leg in terms of landing mass and landing safety,but the whole suspension mechanism system is relatively complex.

5 LANDING SITES SELECTION

In addition to probe design,landing site selection also has a significant impact on landing safety.

5.1 EDL Constraints

Atmospheric density,elevation,slope and rock distribution are the main factors affecting landing safety.

The atmospheric density is related to altitude and directly affects peak overload,parachute deployment,and dynamic descent speed conditions.

The landing site elevation needs to be low enough to meet the deceleration descent height requirement.According to the analysis,regions in the northern hemisphere (mainly low-latitude plains) below the MOLA datumof -3 km would meet the requirements.

The distribution of slopes and rocks mainly affects the safety of radar measurements and touchdown on Mars at the end of the landing,and is usually used as the main constraint when selecting a small-scale candidate landing area for analysis.

5.2 Surface Operation Constraints

In addition to the EDL constraints,the latitude range of Mars should be selected to meet the solar array power supply and thermal control requirements of the landing probe on the Mars surface.The plain area of 5°N to 30°N is suitable according to light and temperature conditions.The area with significant dust coverage according to the requirements of thermal inertia (＜100 JmsK) and reflectivity should be avoided (＞0.25).

5.3 Candidate Landing Site

At first,two large-scale landing areas on the Mars surface should be selectedaccording to the above constraints.Then,further determination of the slope and rock distribution constraints (typical index of slope and rock distribution is shown in Table 3),determination of the size and shape of the landing zone (usually ellipse) according to the dispersal prediction of EDL landing points.Finally,selection of traverse for a large range of landing areas based on the landing ellipse as the basic research unit,and then select several candidate landing areas.

Table 3 Typical index of slope and rock distribution

6 DEVELOPMENT SUGGESTIONS

Mars landing missions incur large environmental uncertainties,multiple state transitions and high requirements for autonomy.According to the previous analysis,it is difficult to carry out whole-system tests and verification on the ground because multiple disciplines such as GNC,pneumatic,thermal protection,propulsion,parachute,structure,remote sensing mapping,and meteorology are involved.Therefore,it is necessary to form a simulation analysis model (including fault model) that is close to the expected real performance based on a single-machine product design and experimental verification results,which supports full digital and semi-physical simulation analysis covering the whole process with limit conditions,and obtain a data envelope through limit/fault testing,verify the fault diagnosis and system reconstruction capabilities of hardware and software,and find the weak aspects in the design.

The failure of the ExoMars 2016 landing mission is a typical case of the imperfection of the parachute-aeroshell assembly dynamic simulation analysis model,resulting in inadequate fault diagnosis,system reconstruction design considerations and ground verification.

7 CONCLUSION

Based on the research progress of Mars landing missions,this paper analyzes the characteristics and key issues of Mars landing missions,and gives some suggestions for technical development that need to be focused on,which can provide references for the design of subsequent celestial entry and landing missions such as for Mars landing exploration.