Rocket Exhaust Research Articles

Noise suppression of infrared thermal imaging of rocket exhaust plume using SPOD

The environmental noise have a negative influence on the quality of infrared thermal imaging of the rocket exhaust plume. In this study, the noise data of the unsteady rocket exhaust plume flow field was simulated using Gaussian white noise, and the infrared thermal image of the plume was numerically calculated using the narrow band method (SNB) and the line of sight (LOS) method. The denoising of infrared thermal imaging was achieved through the spectral proper orthogonal decomposition (SPOD) inversion method. Results indicate that Gaussian white noise leads to larger infrared thermal image residuals in the intrinsic core of the plume compared to the mixed regions. The infrared thermal image in the 2.7 μm band is greatly affected by the noise with an average error of 21.1%, and the average error in the 4.3 μm band is 17.6%. After SPOD denoising, the error of the plume infrared thermal image is reduced by more than 50%.

Thermal life assessment of laser powder-directed energy deposited NiCrAlY/CuCrZr bimetallic clad for rocket nozzle applications

To enhance the thermal life of rocket exhaust nozzles, the hot side of copper liners is coated with thermal barrier coatings (TBCs) to provide thermal insulation and oxidation resistance. However, interface failures often occur between M-CrAlY bond coats and nozzle liners due to significant differences in their thermal expansion coefficients (CTE). This study explores the use of Laser Powder-Directed Energy Deposition (LP-DED) to clad NiCrAlY onto a CuCrZr substrate, as the process offers localized heating which can offer better bond strength. Optimization trials were conducted using single and multi-track studies to identify optimal parameters. Due to the low energy absorption of the CuCrZr substrate to 1070 nm laser sources, cladding was performed at a high energy density of 135 J/mm2 with a 1.2 g/min feed rate to achieve defect-free clads with sufficient diffusion. The bulk of the NiCrAlY clads showed γ′-Ni3Al, β-NiAl, and γ-Ni phases, while Y4Al2O9 and Y2O3 oxides formed on the top surface due to aluminum and yttrium depletion at high temperatures. The clads exhibited cellular dendritic microstructures at the bulk region, and planar microstructures were observed at the dilution zone. EBSD-KAM maps showed higher dislocation density near the interface due to CTE mismatch across substrate and clad. Scratch tests confirmed strong adhesion with no interface cracks, though crack propagation was observed from the edges after 50 isothermal cycles, driven by copper erosion. With Cu diffusion, interface region exhibited a graded microstructure which could enhance CTE, improving compatibility compared to standard NiCrAlY alloys.

Research on the modeling technique of infrared radiation scaling law for rocket engine exhaust plumes

The complexity of the rocket motor exhaust plume radiation phenomenon, along with the influence of numerous parameters and implicit propagation, makes the theoretical simulation calculations expensive. This hinders design optimization, parameter identification, and other studies. Hence, it is crucial to establish the direct correlation between thrust, velocity, altitude, detection angle, and plume radiation in engineering.The paper proposes a scaling law modeling approach for predicting plume infrared radiation from liquid oxygen (LOX)/kerosene rocket engines to solve the difficulties of quick prediction and lack of explicit models. An explicit scaling law model with four parameters is developed by utilizing the e-exponential scaling law to relate plume radiance to flight velocity, the power scaling law to relate plume radiance to vacuum thrust, and the sinusoidal scaling law to relate plume radiance to detection angle. This model employs a decoupled modeling approach for plume radiance, flight altitude, flight velocity, vacuum thrust, and detection angle. A fast prediction method for the LOX/kerosene rocket engine plume in the altitude range of 11∼61 km is achieved. The theoretical simulation test shows the relative error of the velocity-scaling law in predicting the radiance of the RD-180 template engine during the Atlas III flight trajectory is less than 20 %. The radiance of the RD-170 and RD-191 rocket engines is predicted using the thrust-scaling law based on the radiance of the RD-180 template engine exhaust plume. The radiance of the exhaust plume from the RD-170 and RD-191 rocket motors is estimated using the thrust-scaling law with an accuracy of under 15 %. Subsequently, the radiance of these two motors at detection angles between 10° and 170° is predicted using the angle-scaling with a relative error of less than 35 %.

A numerical study on the influence of multiple nozzles on the infrared radiation signatures of liquid rocket exhaust plumes

A numerical study was conducted to examine the relationship between the infrared radiation characteristics of liquid rocket engine exhaust plumes under the same thrust conditions and the number and expansion degree of nozzles. This study used the Reynolds-averaged computational fluid dynamics method to simulate the flow field of multiple nozzles. A single-line group combined with the G-C approximation was used to calculate the molecular spectral properties, and a three-dimensional radiation transfer calculation model was established based on the line-of-sight method. The infrared radiation characteristics of the exhaust plume were calculated for single, double, three, and four nozzles. The radiation intensity of two typical bands, 2.7 μm and 4.3 μm, was studied, including spectrum, intensity, and radiation intensity distribution. The results show that the integrated infrared radiation intensity of the double-nozzle plume is 51.36 % lower than the average of the other three nozzle configurations under the over-highly under-expansion state. The radiation difference of the double nozzle reaches up to 38.3 % at different detection angles, whereas the differences between the three and four nozzles are below 10.5 % and 6.1 %, respectively, indicating that the radiation difference at each detection angle gradually decreases with the increase in the number of nozzles.

Erosion rate of lunar soil under a landing rocket, part 1: Identifying the rate-limiting physics

Multiple nations are planning activity on the Moon's surface, and to deconflict lunar operations we must understand the sandblasting damage from rocket exhaust blowing soil. Prior research disagreed over the scaling of the erosion rate, which determines the magnitude of the damage. Reduced gravity experiments and two other lines of evidence now indicate that the erosion rate scales with the kinetic energy flux at the bottom of the laminar sublayer of the gas. Because the rocket exhaust is so fast, eroded particles lifted higher in the boundary layer do not impact the surface for kilometers (if at all; some leave the Moon entirely), so there is no saltation in the vicinity of the gas. As a result, there is little transport of gas kinetic energy from higher in the boundary layer down to the surface, so the emission of soil into the gas is a surprisingly low energy process. In low lunar gravity, a dominant source of resistance to this small energy flux turns out to be the cohesive energy density of the lunar soil, which arises primarily from particles in the 0.3 to 3 μm size range. These particles constitute only a tiny fraction of the mass of lunar soil and have been largely ignored in most studies, so they are poorly characterized.

Environmental life cycle assessment of reusable launch vehicle fleets: Large climate impact driven by rocket exhaust emissions

After the success of the reusable Falcon 9 rocket, space actors are pursuing competitive space access by developing Reusable Launch Vehicles (RLVs). While this initiative may enhance recycling rates, it may also trigger the Jevons’ paradox as it amplifies the overall environmental footprint due to increased launch frequencies. It is therefore essential to quantify RLVs’ impacts and identify key design drivers to enable efficient design choices while mitigating undesirable environmental effects.Consequently, this article uses a space specific Life Cycle assessment (LCA) approach to evaluate the environmental footprint, in terms of climate impact, water depletion and land use, of different RLV fleets designed to serve a forecasted European space market. The results show that the LH2 fleet options have 2–8 times lower carbon footprint when compared to the LCH4 fleet as a result of lower propellant consumption and lack of black carbon emissions, suggesting that the environmental burdens are mostly driven by propellant choice. Moreover, the analysis reveals a potential underestimation of climate impacts in previous LCA’s by 2–3 orders of magnitude due to the absence of high altitude characterisation of rocket exhaust emissions and demised aluminium oxides. This increased forcing could lead to fleet choices surpassing the Earth’s carrying capacity given by its planetary boundaries.The methodology and results within this study can support further integration of launch and reentry emissions within LCA by refining modelling techniques, improving impact characterisation and quantifying uncertainties. These advancements can ultimately enable robust eco-design strategies for launch vehicles.

Effect of Different Flame Deflectors Plate Geometry on Supersonic Jet Impingement Flow and Heat Characteristics from Rocket Exhaust

The numerical investigation has been carried out on an under-expanded supersonic jet impingement on the perpendicular and inclined deflector plates of various shapes, such as flat, sinusoidal, and wedge types. The numerical study considers flat plates at Mach number of 2.2, and pressure ratio of 1.2, and results of flow and heat transfers are obtained using the Reynolds averaged Naiver Stokes (RANS) two-equation turbulence models. Various two-equation models (especially Standard k-ε, Reynolds Normalization Group k-ε, Realizable k-ε, Shear Stress Transport k-ω, and Baseline k-ω) were employed to predict the deflector’s surface temperature and pressure ratios. The pressure ratios are plotted with various two-equation turbulence models, and it is observed that among all turbulence models, the Shear Stress Transport k-ω shows close agreement with experimental results. Moreover, among the three deflectors, the wedge-shaped deflector shows improved performance in deflecting the rocket exhaust plume away from the surface. However, the maximum temperature on the deflector surface for a wedge-shaped deflector is 8.19% and 26% higher than the flat and sinusoidal-shaped deflectors, respectively.

Spectral feature extraction of rocket exhaust plume using spectral proper orthogonal decomposition

The spectral characterization of flow-field parameters provides a new perspective for understanding the spatiotemporal evolution of unsteady supersonic exhaust plumes and for extracting typical structures. In this study, a large-eddy simulation is performed to calculate the three-dimensional unsteady supersonic plume flow field of rocket engines, and a spectral proper orthogonal decomposition (SPOD) method with a spatiotemporal separation is established. This approach is used to extract the coherent structural features of the unsteady exhaust plume flow field and analyze the mode space structure at different frequencies. The three-dimensional reconstruction and denoising of the exhaust plume flow-field parameters can be achieved via the frequency- and time-domain reconstructions of the SPOD algorithm and oblique projection method, respectively. The ground rocket exhaust plume of ballistic evaluation motor-II is analyzed. The results indicate that the SPOD method can effectively extract the single-frequency mode structure of the reactive supersonic flow field, and that low-order behavior appears in the m = 0 and m = 1 azimuth modes. The potential core exhibits a high-frequency wave-packet structure that is affected by shock waves and shear layers. Time-domain reconstruction based on the oblique projection method facilitates the capture of the dynamic characteristics of the flow field. For the first-order SPOD mode, the frequency- and time-domain reconstruction errors are 3.3% and 1.5%, respectively. The frequency-domain reconstruction method exhibits a 4% improvement in denoising ability compared to low-pass filters. This study provides a novel method for the spectral characterization and spatiotemporal feature extraction of supersonic exhaust plume flow fields.

Drift and afterburning effect on rocket exhaust plume impingement and optimization of deflector

This paper studies the one-nozzle liquid rocket exhaust plume impinging on the flame deflector. The effect of drift and afterburning on exhaust plume impinging on the deflector during the takeoff phase is studied by numerical simulation. An impact model is established based on three-dimensional compressible Navier-Stokes equations, a two-equation realizable k-epsilon turbulence model, an afterburning model, and second-order total variation diminishing upwind scheme. The study shows that the smoothness of the deflector discharge is affected by the impingement position of the exhaust plume impinging on the deflector due to the drift of such a position during the take-off phase of the rocket. The temperature of the rocket exhaust plume flow field after considering the chemical reactions is higher than that of the frozen flow, and the temperature of the deflector surface is also higher than that of the frozen case. With the increase of the impingement angleat inlet areaof the rocket exhaust plume impinging on the deflector, the deflector discharge performance gradually decreases. At the same impingement angleat inlet area, as the impingement angleat exit areadecreases, the deflector slot discharges more smoothly.

First magnetic and spectroscopic constraints on attenuated space weathering at the Chang'e-5 landing site

Space weathering is a primary process altering the lunar surface, producing submicroscopic iron particles embedded in soil grains, which have diagnostic magnetic and spectroscopic effects. However, the space weathering of lunar soils from young geological units remains mysterious without solid constraints from returned samples. Chang'e-5 mission is the first lunar sample-return mission in decades. The Chang'e-5 landing and sampling unit (∼2.0 Ga) is much younger than all Apollo and Luna units (>3.0 Ga). It has provided an unprecedented opportunity to reveal the recent space weathering characteristics with both returned soil samples and in situ investigation data. In this study, magnetic approaches (Vibrating Sample Magnetometer, Electron Paramagnetic Resonance Spectrometer, Mössbauer Spectrometer) were used to characterize the Chang'e-5 lunar soil for the first time after the return of the samples, strongly supporting it is among the most immature soils on the Moon with a metallic iron abundance of 0.42% and an Is/FeO value of 14−10+6. Additional Landing Camera and Lunar Mineralogical Spectrometer (480–3200 nm) data obtained surrounding the sampling area were used to study the lateral effects of rocket exhaust. All data from soil samples and in situ observations have offered magnetic and spectroscopic constraints on the space weathering state of the Chang'e-5 site. The combined effects of rocket exhaust by removing the top highly weathered soil and the young age of the landing site with lesser degree of impact reworking may attenuate the maturity of Chang'e-5 lunar soils.

Effect of water injection on the cooling performance of flame deflector in rocket engine testing platform

Ground test is an essential step to ensure the reliability of the design and assess the quantity and performance of rocket engine. However, supersonic and high-temperature rocket exhaust gas directly impinging on the flame deflector of test platform cause severe heat damage which brings about a low repeated utilization and construction cost of the deflector. Generally, the water is employed as the coolant for the spray cooling system to protect the flame defector of test platform. In this paper, Euler-Euler approach is adopted to investigate the multi-phase flow in a flame deflection of rocket engine test platform. An evaporation model based on concentration difference is chosen to simulate the mass and heat transfer between water injection and exhaust plume. Model validation shows that the predicted results are in good agreement with the experimental data. The effect of water injection, including spray location and angle, and mass flow rate, on the afterburning and cooling performance of flame deflector is numerically investigated. The results show that the afterburning of exhaust plume could be inhibited by the water injection. In terms of maximum and area-averaged temperature on the bottom and inclined surfaces of flame deflector, the configuration with spray nozzle angle of 60° has the lowest area-averaged temperature. For the simulated cases with different spray nozzle diameters, the difference of maximum and area-averaged temperatures is below 11.0 K and 6.1 K, respectively. In addition, the maximum and area-averaged temperatures increase first and then decrease with the increase of spray nozzle diameter.

Accurate evaluation on infrared spectral radiation for booster phase rocket exhaust plumes

Accurate modeling of rocket exhaust plumes’ infrared spectral radiation is of great significance in the field of aerospace. The existing classical statistical-narrow-band (SNB) model cannot adapt well to the highly heterogeneous plume temperature and concentration fields, with large radiation modeling errors at the spectral level. To overcome this limitation, this study presents an improved SNB model (I-SNB) for rocket plume infrared spectral radiation modeling, which categorizes the spectral lines within narrow band Δη into different groups according to their inconsistency in sensitivity to changes in both temperature and concentration along the plume integration path, ensuring that the spectral lines within the same group have similar behaviors with simultaneous changes in both temperature and concentration. The modeling accuracy of the present I-SNB model has been evaluated using FTIR spectral measurements and LBL benchmark solutions, demonstrating significantly better performance than the classical SNB model. Then, the I-SNB model has been used to investigate the rocket plume spectral radiation during the booster phase, the results of which have engagement guiding significance for the development of booster phase interception (BPI) of the space-based infrared early warning system.

Thermal Modeling of Umbilical Tower During the Lift-off Of a Launch Vehicle

Umbilical Tower (UT) forms a critical structure along with mobile service tower, launch pedestal and jet deflectors of a rocket launch pad. UT will be exposed to severe thermal environments during lift off of the launch vehicle owing to its close proximity. UT experiences thermal loads until the launch vehicle lifts off and clears it. Hence thermal environments need to be quantified to assess the survivability of UT for subsequent missions. This paper brings out the methodology developed for estimating different thermal environments experienced by UT structure during a typical rocket launch. The engineering method result in a quick estimate of flow properties of the plume from the rocket exhaust compared to CFD, which will be computationally intensive and time consuming to arrive a converged accurate solution for heat flux estimation. ISRO’s launch vehicles have first stage with Solid Rocket Motor (SRM) as core and solid motors/liquid engines as multiple strap-ons. Thermal loads to launch pad structures include radiative and convective heating from exhaust plumes. The aluminum particle laden exhaust plumes of solid rockets are dominating radiative heat source. Convective heating occurs from hot jets and direct plume impingement. The relative movement of launch vehicle with respect to UT structure is an important parameter affecting thermal environments. It varies for each launch depending on the vehicle variant and its lift-off trajectory. This calls for a detailed assessment of thermal loads for thermal adequacy studies. A finite element based thermal model for umbilical tower is also generated for thermal response studies and Thermal Protection System (TPS) design. The computed temperature levels at various location of UT structure are compared with measured values during typical lift-off. Results indicate that temperature levels of UT structure are within allowable limits and the adopted methodology can predict temperature within a maximum difference of ±5%.

Exploring the effects of the rocket exhaust of the Chang'E-5 lander on the lunar regolith using LROC NAC and landing camera images

During the Chang'E-5 (CE-5) lander descent process, the rocket exhaust disturbed the lunar surface of the landing area, which may have affected the properties of the collected samples. Therefore, it is necessary to explore the effects of lander exhaust on the landing area. In this paper, the images obtained by the Narrow Angle Camera (NAC) onboard the Lunar Reconnaissance Orbiter (LRO) were used to determine the exhaust-disturbed area. It is found that this area was featured by an increased reflectance of ∼10 ± 1% compared with the background (from I/F = 0.048 to 0.053). According to the magnitude of the change in the reflectance of the landing region, the disturbed area was divided into the focus disturbed zone (FDZ) and the diffuse disturbed zone (DDZ). The DDZ extends from the edge of the FDZ and has a lower reflectance anomaly. The whole exhaust-disturbed area was found to extend mainly via the north-south direction and to have the shape of an irregular ellipse. This spatial extent is highly related to the topographic factors rather than the obstacle avoidance process according to the topography and landing process. The phase function of the lunar surface was used to analyze the photometric performance of the disturbed area. The lower slope of the phase function in FDZ demonstrates less backscattering and a smoothing effect of the disturbed lunar surface. The CE-5 high-resolution LCAM images captured the smoothing and brightening effects of the lunar surface, which confirmed the variation in regolith properties. These effects are consistent with the inference from the NAC observations. The damage to the “fairy-castle” structure of the lunar surface could also contribute to surface smoothing. This study can provide a reference for the subsequent analyses of lunar samples and can help to enhance the understanding of the effect of lander exhaust on the lunar regolith.

A Simulation of Artificial Neutral Gas‐Dust Plasma in the Ionosphere

AbstractDusty plasma is generally defined as a plasma system containing a large number of nano‐ or micron‐sized dust particles. Its plasma properties and propagation properties of wave travel through it is of great scientific significance to the understanding of many fundamental problems in space physics and astrophysics. Based on previous studies on rocket exhaust and Polar Mesosphere Summer Echoes, fluid mechanics and dust microphysics are combined to simulate the ionospheric plasma structure generated by neutral gases and charged aerosols using fluid theory and kinetic theory. The Total Electron Content of the in‐situ detection beacon is simulated and compared with the experimental phenomenon given in Charged Aerosol Release Experiment by Bernhardt et al. The ray tracing algorithm is used to simulate the path deflection of the radio wave when it passes through the artificial disturbance area generated by the gas‐solid mixture, and the simulation results of ionosonde and analysis are given. Since the neutral gas expands substantially isotropically after being released, the resulting depleted structure is an approximately isotropic sphere, or an acorn‐like structure. The solid particles are constrained by the nozzle, and particles will distribute within a spray angle, and the depletion structure will appear as a “broom‐like” or “cone‐like” structure. If the size of dust particles is small enough, the particle movement will be restricted by the geomagnetic field, resulting in a “squid‐like” ionospheric structure. Because of the fishtail‐like sharp boundary, radio waves passing through it are deflected in a specific path, which is markedly different from the effect of pure neutral gas injected into the ionosphere.

Leto Mission Concept for Green Reconnaissance of the Marius Hills Lunar Pit

Lava tubes are potentially important sites for the long-term human presence on the Moon because they provide shelter from surface hazards, including micrometeorites, radiation, extreme temperatures, and dust. The discovery of a lava tube opening or pit at Marius Hills in Oceanus Procellarum is compelling motivation for robotic and eventually human exploration missions to these sites for in situ investigations and site assessments to determine viability for habitation and utilization of lunar resources. We make the case for Marius Hills to be a high-priority landing site and present elements of lunar data analysis, instrument/payload concepts, science justification for robotic missions, and thematic geologic reconnaissance and remote sensing that should be conducted prior to any construction or emplacement of infrastructure. This is described as a “green reconnaissance” approach to lunar exploration and exploitation, which seeks to address such contamination factors as sprayed rocket exhaust and sublimating water in order to preserve science fidelity. We are developing a concept of operations called the Leto mission for a green reconnaissance approach to robotically access the Marius Hills sublunarean void.

Lunar plume-surface interactions using rarefiedMultiphaseFoam

Understanding plume-surface interactions is essential to the design of lander modules and potential bases on bodies such as the Moon, as it is important to predict erosion patterns on the surface and the transport of the displaced regolith material. Experimentally, it is difficult to replicate the extra-terrestrial conditions (e.g. the effects of reduced gravity). Existing numerical tools have limited accessibility and different levels of sophistication in the modelling of regolith entrainment and subsequent transport. In this work, a fully transient open source code for solving rarefied multiphase flows, rarefiedMultiphaseFoam, is updated with models to account for solid-solid interactions and applied to rocket exhaust plume-lunar regolith interactions. Two different models to account for the solid-solid collisions are considered; at relatively low volume fractions, a stochastic collision model, and at higher volume fractions the higher fidelity multiphase particle-in-cell (MPPIC) method. Both methods are applied to a scaled down version of the Apollo era lunar module descent engine and comparisons are drawn between the transient simulation results. It is found that the transient effects are important for the gas phase, with the shock structure and stand-off height changing as the regolith is eroded by the plume. Both models predict cratering at early times and similar dispersion characteristics as the viscous erosion becomes dominant. In general, the erosion processes are slower with the multiphase particle-in-cell method because it accounts for more physical effects, such as enduring contacts and a maximum packing limit. It is found that even if the initial volume fraction is low, the stochastic collision method can become unreliable as the plume impinges on the surface and compresses the regolith particles, invalidating the method’s assumption of only binary collisions. Additionally, it is shown that the breakdown of the locally free-molecular flow assumption that is used to calculate the drag and heat transfer on the solid particles has a strong influence on the temperatures that the solid particles obtain.

Experimental studies on combustion performance of beeswax-paraffin blended solid fuels in a hybrid rocket

The study intends to investigate the physical, chemical and thermal characteristics of paraffin blended fuels to determine their suitability as a solid fuel in a hybrid rocket. Wax fuels are a viable and efficient alternative to conventional rocket fuels, having excellent structural strength and thermal and mechanical properties. By utilizing both axial and swirl injection technique, the combustion performance of paraffin – beeswax blended fuels have been tested with a fabricated cylindrical grain in a laboratory-scale rocket setting along with oxygen. The test outcomes revealed solid fuel compositions of more beeswax content in paraffin wax on an oxygenated gaseous environment with a swirl-flow injection method has the highest average regression rate of 1.649 mm/sec at 181 kg/m2s mass flux. Axially injected oxygen with pure paraffin wax has the lowest value of 0.85 mm/sec at 96 kg/m2s. The regression rate comparisons revealed that oxygen injection by a swirl injector increased the regression rates by 40% for mass fluxes greater than 80 kg/m2s. Compared to other studies, the combustion efficiencies have been obtained in this study are good. Blended fuels can manage and increase combustion efficiencies for axial and swirl flow conditions. Swirl injectors outperform axial injectors for oxygen injection and allow for a higher proportion of Beeswax combined with paraffin. This study exclusively designed and manufactured an axial injector and swirl injector, according to the required dimensions of a lab-scale hybrid rocket's combustion chamber, injector, and exhaust nozzle, and their performances have been evaluated.

Phenomenology of plume–surface interactions and preliminary results from the Tianwen-1 landing crater on Mars

The plume–surface interaction (PSI) is a common phenomenon that describes the environment surrounding the landers resulting from the impingement of hot rocket exhaust on the regolith of planetary bodies. The PSI will cause obscuration, erosion of the planetary surface, and high-speed spreading of dust or high-energy ejecta streams, which will induce risks to a safe landing and cause damage to payloads on the landers or to nearby assets. Safe landings and the subsequent scientific goals of deep-space exploration in China call for a comprehensive understanding of the PSI process, including the plume flow mechanics, erosion mechanism, and ejecta dynamics. In addition, the landing crater caused by the plume provides a unique and insightful perspective on the understanding of PSI. In particular, the PSI can be used directly to constrain the composition, structure, and mechanical properties of the surface and subsurface soil. In this study, we conducted a systematic review of the phenomenology and terrestrial tests of PSI: we analyzed the critical factors in the PSI process and compared the differences in PSI phenomena between lunar and Martian conditions; we also reviewed the main erosion mechanisms and the evolution and development of terrestrial tests on PSI. We discuss the problems with PSI, challenges of terrestrial tests, and prospects of PSI, and we show the preliminary results obtained from the landing crater caused by the PSI of Tianwen-1. From analysis of the camera images and digital elevation model reconstructions, we concluded that the landing of Tianwen-1 caused the deepest crater (depth > 40 cm) on a planetary surface reported to date and revealed stratigraphic layers in the subsurface of Martian soil. We further constrained the lower bounds of the mechanical properties of Martian soil by a slope stability analysis of the Tianwen-1 landing crater. The PSI may offer promising opportunities to obtain greater insights into planetary science, including the subsurface structure, mineral composition, and properties of soil.

Sidney N. Stone (1922–2011)

Stone was a scientist at Los Alamos National Laboratory who conducted defense related research such as the ionospheric effects of rocket exhaust and optical diagnostic systems for laser-produced shock waves.