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Abstract This article reviews South Korea's response model for space object impacts, re-entries, and collisions. These are countermeasure models that are prepared for natural or artificial space objects falling in South Korea or colliding in space, and they employ the Ministry of Science and ICT (MSIT) to perform 4 stages that include prevention, preparation, response, and restoration. However, the article suggests that this process is inefficient given that the MSIT specializes in research and development but is expected to perform disaster response tasks in the event of natural or manmade space object impacts and re-entries. The research presents 2 alternative models to the current ones: the first model specifies that the MSIT and the Ministry of the Interior and Safety (MOIS) work together, and the second model designates the MOIS as the only agency responsible for the disaster response. These alternative models would more effectively handle disaster management than would the current models. [ABSTRACT FROM AUTHOR]

Flush air data sensing system (FADS) forms a mission-critical subsystem in re-entry vehicles. It makes use of surface pressure measurements from the nose cap of the vehicle for deriving air data parameters such as angle of attack, angle of sideslip, Mach number, etc. of the vehicle. These parameters are used by the flight control and guidance systems, and also assist in the overall mission management. The overall system engineering of FADS, including selection of pressure transducers, tubing size, port geometry, FADS algorithm and associated processing electronics along with the integration scheme is addressed in this article. Details of the qualification tests carried out in wind tunnel for end-to-end verification of the entire FADS system are covered in brief. Majority of the tests were carried out in a low-speed wind tunnel at a wind speed of 65 m/s (Mach number 0.2). The flight performance of FADS is also discussed in this article. [ABSTRACT FROM AUTHOR]

Spent rocket bodies in geostationary transfer orbit (GTO) pose impact risks to the Earth’s surface when they reenter the Earth’s atmosphere. To mitigate these risks, reentry prediction of GTO rocket bodies is required. In this paper, the reentry prediction of rocket bodies in eccentric orbits based on only Two-Line Element (TLE) data and using only ballistic coefficient (BC) estimation is assessed. The TLEs are preprocessed to filter out outliers and the BC is estimated using only semimajor axis data. The BC estimation and reentry prediction accuracy are analyzed by performing predictions for 101 rocket bodies initially in GTO and comparing with the actual reentry epoch at different times before reentry. Predictions using a single and multiple BC estimates and using state estimation by orbit determination are quantitatively compared with each other for the 101 upper stages. [ABSTRACT FROM AUTHOR]

The Hayabusa 2 mission is targeted to explore the asteroid (162173) 1999 JU 3 and return surface as well as sub-surface samples through a novel impactor. Upon its return, at the end of 2020, the spacecraft will release the capsule for Earth re-entry and drift away from the planet. Based on the current mission profile, the spacecraft is expected to retain 30 kg of xenon propellant for trajectory maneuvers after the capsule is released. This remaining fuel can be used to extend the mission and improve its scientific return by exploring a new target. Work herein outlines an extension plan for Hayabusa 2, detailing the target selection process and its subsequent trajectory design. Due to final Earth escape trajectory, considering the excess velocity and orbital geometry, the only available extension option is an asteroid flyby. One of the most important trajectory characteristic is to maximize the spacecraft's optical detection capabilities. As a result the asteroid 2001 WR 1 is identified as the most promising target candidate. The resulting trajectory uses all the available xenon with 100% duty cycle. Furthermore, the extension lasts for 932 days and offers 1.57 days of optical navigation time for a flyby on June 27, 2023. [ABSTRACT FROM AUTHOR]

In order to mitigate debris in orbit and to avoid dramatic collisions on Earth after atmospheric re-entry at its end-of-life, the spacecraft missions have to take into account the influence of atmospheric re-entry conditions on the spacecraft survivability. In this way, oxidation in air plasma conditions was studied on Inconel 718 alloy in order to obtain degradation kinetic laws at high temperature for short time duration (few minutes). Moreover, measurements of the total emissivity at high temperature up to the melting point were carried out on the same alloy – in high vacuum to avoid oxidation – and on samples oxidized in standard or in air plasma conditions. An increase of the total emissivity for oxidized samples by a factor 3 to 4 is obtained compared to as-received samples. This effect can strongly influence the ablation during a trajectory and consequently the casualty area calculation. Experimental results on oxidation and emissivity are presented together with the microstructural characterization of the samples (surface and cross-section) using several techniques. [ABSTRACT FROM AUTHOR]

An integrated finite-time disturbance observer (FDO) and attitude controller is designed for a reusable launch vehicle (RLV) in this paper. In accordance with the multiple-timescale features, RLV attitude dynamics are divided into an outer-loop subsystem and an inner-loop subsystem. Based on the recently developed sliding mode control (SMC), a novel multivariable supertwisting sliding mode controller driven by a FDO is designed to achieve a fast and accurate reentry attitude tracking. This integrated design can generate a continuous control law which has excellent robustness to uncertainty and disturbances with known bounds while achieving an arbitrarily fast convergence. The finite-time stability of the overall system is proved by using the Lyapunov function technique and the multiple-timescale separation principle. In addition, an optimal control allocation for allocating torque commands into aerodynamic surface deflection commands with constraints is also proposed. Finally, the effectiveness and the robustness of the integrated control scheme are verified by the simulation results of a six-degree-of-freedom (6-DOF) RLV. [ABSTRACT FROM AUTHOR]

The probability of KufaSat collisions with different sizes of orbital debris and with other satellites which operating in the same orbit during orbital lifetime was determined. Apogee/Perigee Altitude History was used to graph apogee and perigee altitudes over KufaSat lifetime. The required change in velocity for maneuvers necessary to reentry atmospheric within 25 years was calculated. The prediction of orbital lifetime of KufaSat using orbital parameters and engineering specifications as inputs to the Debris Assessment Software (DAS) was done, it has been verified that the orbital lifetime will not be more than 25 years after end of mission which is compatible with recommendation of Inter-Agency Space Debris Coordination Committee (IADC). [ABSTRACT FROM AUTHOR]

The capsule SHARK (Sounding Hypersonic Atmospheric Re-entering “Kapsule”) was designed and realized for testing in real flight conditions a ZrB 2 -based Ultra High Temperature Ceramic (UHTC) Nose Tip. The capsule was instrumented with accelerometers, rate and pressure sensors, and thermocouples of which three were inserted inside the UHTC Nose Tip. The capsule was launched by the European Space Agency (ESA) rocket MAXUS-8 performing an atmospheric re-entry after a parabolic trajectory with an apogee at 700 km of altitude. The temperature profile recorded by the thermocouples inserted inside the Nose Tip reported a discontinuity at about 762 s after the release from the rocket, revealing the failure of the ceramic component during the flight. The capsule was recovered and the Nose Tip was found damaged: the fore half was missing while the inner part, fractured in three pieces, was recovered. The fractographic analysis of the three recovered fragments was carried out by optical microscopy techniques which allowed identifying the main fracture marks, such as fracture mirror, hackle lines, arrest lines, and Wallner lines. The fracture dynamic has been outwardly reconstructed and the stress at fracture was evaluated by means of the Orr's equation involving the measurement of the fracture mirror radius. Then applying the Griffith's relationship, the flaw size was estimated. The origin of the main fracture was identified on the edge of the hole manufactured by Electrical Discharge Machining (EDM) to lodge thermocouples inside the nose tip, causing the strength-limiting flaw. Thermal shock was recognized as the probable failure cause as already observed by on-ground tests performed in the Plasma Wind Tunnel (PWT) “Scirocco” by the Italian Aerospace Research Centre (CIRA) on a similar UHTC Nose Tips manufactured with the same machining technology. [ABSTRACT FROM AUTHOR]

The effect of shock interactions on the attitude stability of a reentry vehicle system with a toroidal ballute was investigated. The hypersonic wind tunnel experimental results showed that when the shock interaction occurred near or outside the ballute, an unstable oscillation of the ballute was observed. This was caused by the local high-pressure region on the ballute surface created by the shock interaction between the shock from the reentry capsule and the shock from the ballute. To avoid this unstable oscillation, the radius of the ballute should be designed to be large enough so that the shock from the capsule will be located inside the ballute, which can avoid the local high-pressure region on the ballute surface. [ABSTRACT FROM AUTHOR]

The paper deals with a microsatellite and the related deployable recovery capsule. The aero-brake is folded at launch and deployed in space and is able to perform a de-orbiting controlled re-entry. This kind of capsule, with a flexible, high temperature resistant fabric, thanks to its lightness and modulating capability, can be an alternative to the current “conventional” recovery capsules. The present authors already analyzed the trajectory and the aerodynamic behavior of low ballistic coefficient capsules during Earth re-entry and Mars entry. In previous studies, aerodynamic longitudinal stability analysis and evaluation of thermal and aerodynamic loads for a possible suborbital re-entry demonstrator were carried out in both continuum and rarefied regimes. The present study is aimed at providing preliminary information about thermal and aerodynamic loads and longitudinal stability for a similar deployable capsule, as well as information about the electronic composition of the plasma sheet and its possible influence on radio communications at the altitudes where GPS black-out could occur. Since the computer tests were carried out at high altitudes, therefore in rarefied flow fields, use of Direct Simulation Monte Carlo codes was mandatory. The computations involved both global aerodynamic quantities (drag and longitudinal moment coefficients) and local aerodynamic quantities (heat flux and pressure distributions along the capsule surface). The results verified that the capsule at high altitude (150 km) is self-stabilizing; it is stable around the nominal attitude or at zero angle of attack and unstable around the reverse attitude or at 180° angle of attack. The analysis also pointed out the presence of extra statically stable equilibrium trim points. [ABSTRACT FROM AUTHOR]

The following document presents a trustworthy method for validating dynamic stability results obtained from wind-tunnel test campaigns. The first phase comprises the calibration of the electromagnetic test setup, in order to obtain a relation between the supplied current and the generated magnetic damping. In the validation phase this setup is attached to the dynamic stability forced oscillations test structure, outside of the wind tunnel. Several runs are performed with different oscillating speeds and a wide range of known values of magnetic damping for configurations with and without a spring element resembling the aerodynamic stiffness experienced by the model in the wind-tunnel. The Energy Method is then used to process the momentum and angular position recorded during the test and compute the damping acting on the system. The magnetically applied damping and the computed damping are then compared, in order to access the magnitude of the error. For the system without spring applied, results show that low values of error and uncertainty are obtained for most part of the considered curve. The magnitude of the error usually diminishes for an increase in the oscillating speed, but increases for lower values of damping. [ABSTRACT FROM AUTHOR]

The paper deals with the aerodynamic performance analysis at a phase-A design level of a reusable and unmanned flying laboratory designed to perform a high lift return from low Earth orbit. The flying test bed aims to provide experimental data in the framework of re-entry technologies. The vehicle concept is a wing body with rather sharp leading edges and slender aerodynamic configuration. Several design approaches, ranging from engineering-based to computational fluid dynamics analyses, have been addressed in this work. In particular, vehicle aerodynamic performances for a wide range of freestream flow conditions, from subsonic to hypersonic regime, including reacting and non-reacting flow and different angles of attack have been provided and compared. Highlights on aerodynamic control surfaces' effects, vehicle longitudinal and lateral–directional static stability are provided as well. Computational fluid dynamics results confirm that high temperature real gas effects seem to be fundamental for the assessment of the concept aerodynamics, especially concerning pitching moments evaluation. [ABSTRACT FROM AUTHOR]

The purpose of this work is to formulate a simple method which can be used for the transient thermal load identification. The proposed algorithm can reconstruct the transient temperature distribution in a whole component based on measured temperatures in selected points on the component surface. The presented method may use some commercial heat transfer modelling software. In this work, ANSYS Multiphysics is applied. Two numerical examples of identification of the transient thermal load taken from literature are presented. The determination of the heating process of a simple two-dimensional plate and the recognition of the aerodynamic heating on the atmospheric reentry capsule will demonstrate high accuracy and stability of the proposed algorithm. Future time steps and smoothing filters are used to stabilize the solution of an ill-posed problem. The presented method is easy to use for engineers working in the aerospace or power industry. [ABSTRACT FROM AUTHOR]

Dynamics and control of a tether-assisted return mission of a re-entry capsule are considered. Efficiency of the braking process of the capsule depends on a deflection angle of the tether from a local vertical before separating the capsule from the tether. The aim is of this paper to find a tether length control law that allows to increase the deflection angle of the tether from the local vertical. The control law is based on a principle of a swing. This control law may be applied to the final phase to two possible options to perform a tether-assisted deorbit maneuver: static and dynamic release. An approximate analytical solution for the deflection angle from the local vertical is obtained for the control law. The numerical simulations have shown that the application of the control law allows a reduction in the required tether length of a tether-assisted deorbit maneuver. The proposed control law can be applied to develop new space tethered systems. [ABSTRACT FROM AUTHOR]

An attitude controller is designed for a reusable launch vehicle (RLV) during the re-entry phase with input constraint, model uncertainty and external disturbance. A control-oriented model with matched and unmatched uncertainty is first derived for the control design. To proceed, the control scheme is designed taking advantage of the robust stability property of sliding mode control, the compensation ability of the non-linear disturbance observer (DOB) technique and the systematic design procedure of the backstepping technique. An additional system, the states of which are used for controller design and stability analysis, is constructed to handle the input constraint. The time derivative of the virtual control input is considered as an uncertainty to eliminate the ‘explosion of terms’ problem inherent in backstepping control. Moreover, the proposed method alleviates the chattering problem of the traditional sliding mode backstepping control scheme. Stability analysis of the composite control scheme consisting of the designed controller and the DOB is performed via Lyapunov theory. Finally, simulation results are compared to show that the proposed method is able to achieve better tracking performance and tackles input constraints more effectively than an adaptive filter backstepping control scheme. [ABSTRACT FROM AUTHOR]

Surface roughness, especially if enhanced due to ablative form change, increases skin friction drag and convective heat transfer over reentry vehicles. Although the corresponding heat flux augmentation is usually lower compared to increased friction, careful consideration in the prediction of the resulting heat load levels is required. Within the European Mars mission ExoMars, the potential roughness impact on the thermal protection system of the descent module has been analyzed based on analytical predictions, numerical calculations, and dedicated experimental campaigns. This paper describes the experimental efforts in the compressible flow regime to study the impact of roughness at representative conditions. The data are discussed based on comparisons with prediction methods and results of other investigators. Based on these data, the numerical predictive capabilities within the ExoMars program are characterized and validated. [ABSTRACT FROM AUTHOR]

Thermal protection system design for atmospheric reentry vehicles remains a challenging and complex problem. Recent advances in computational modeling of air-carbon interactions consider competing finite-rate reactions on a limited number of available surface sites. One of the most advanced kinetic models is due to Zhluktov and Abe. However, the Zhluktov and Abe model only describes the oxidation and sublimation of carbon and has no nitridation mechanism. The following study develops several modifications to the Zhluktov and Abe air-carbon model that account for all three reaction mechanisms with the goal of improving cyanogen shock-layer radiation predictions to recent experimental results. First, the study examines two possible paths for carbon nitridation and then assesses the augmented surface reaction model in a representative blunt-body reentry flow. Second, a sensitivity analysis is performed to determine which surface reactions have the most impact on altering cyanogen radiance predictions. It was found that modifications to the oxidation mechanisms significantly improved the comparisons and the nitridation mechanism had a minor effect. Specifically, it was determined that carbon monoxide must be the principal oxidation product at high surface temperatures instead of carbon dioxide, as the Zhluktov and Abe model had originally predicted. More detailed measurements of both the carbon oxidation and nitridation reactions at higher surface temperatures are required to further validate and improve the rate parameters derived in this study. [ABSTRACT FROM AUTHOR]

A gossamer structure for end-of-life disposal of spacecraft to mitigate space debris is considered in comparison with other end-of-life disposal concepts to determine when it would be preferable. A needs analysis, potential use cases, and concept of operations are developed. A survey of disposal strategies is presented for comparison before a downselection of viable competing techniques: solar sailing, high- and low-thrust propulsion, and electrodynamic tethers. A parametric comparison of the downselection competing techniques is presented. Exploiting solar radiation pressure on the structure is of limited value. Atmospheric drag augmentation was found to be of most benefit for end-of-life disposal when an entirely passive means is required, allowing the gossamer device to act as a "fail-safe". This is applicable to only low- and medium-mass spacecraft or spacecraft that are unlikely to survive atmospheric reentry, hence minimizing risk to human life. It does not significantly alter the operating ceiling altitude but does alter the maximum allowable end-of-life mass. Peak mass benefit occurs in the altitude range 550-650 km and is largely independent of deorbit time. [ABSTRACT FROM AUTHOR]

The objective of this study was to introduce an efficient and accurate approach to the quantification of margins and uncertainties for integrated spacecraft systems models. In this study, stochastic expansions, based on nonintrusive polynomial chaos, were used for efficient representation of uncertainty both in design metrics and associated performance limits of a system. Additionally, procedures were outlined for analyzing systems that contain different uncertainty types between the performance metrics and performance limits. These methodologies were demonstrated on two model problems, each possessing mixed (epistemic and aleatory) uncertainty, which was propagated through the models using second-order probability. The first was a complex system model of highly nonlinear analytical functions. The second was a coupled multisystem, physics-based model for spacecraft reentry. The performance metrics consisted of two systems used to determine the maximum g load, the necessary bank angle correction, and maximum convective heat load along a reentry trajectory. Overall, the methodologies and examples of this work have detailed an efficient approach for measuring the reliability of complex spacecraft systems models, as well as the importance of quantifying margins and uncertainties for the design of reliable systems. [ABSTRACT FROM AUTHOR]

The present work investigates plasma effects around two lifting hypersonic vehicles and their implications on the radiofrequency communication blackout performances during the Earth atmospheric reentry phase. In particular, the Intermediate Experimental Vehicle developed by the European Space Agency and a slender, high lift-over-drag space plane, investigated during previous studies, have been taken into account. Chemical nonequilibrium analyses have been carried out by means of a numerical code, updated with a customized chemical model comprised of several reacting species and different reaction mechanisms. Different flight conditions, including the most critical for communications, have been considered for both vehicles. The main outcomes of the study show that, for very similar freestream conditions, the slender space plane exhibits better communication capabilities when compared with the Intermediate Experimental Vehicle. Particularly, the results suggest that the Intermediate Experimental Vehicle would be able to communicate from the lee side for a limited range of the angle of sight, while no communications from the wind side would be possible. On the other hand, for the slender space plane, flying at relatively low angles of attack, communications are possible either from the lee side or the wind side with very wide angles of sight. [ABSTRACT FROM AUTHOR]

The complexity of the GNSS and the several types of satellites in the MEO region turns the creation of a definitive strategy to dispose the satellites of this system into a hard task. Each constellation of the system adopts its own disposal strategy; for example, in the American GPS, the disposal strategy consists in changing the altitude of the nonoperational satellites to 500 km above or below their nominal orbits. In this work, we propose simple but efficient techniques to discard satellites of the GNSS by exploiting Hohmann type maneuvers combined with the use of the 2ω˙+Ω˙≈0 resonance to increase its orbital eccentricity, thus promoting atmospheric reentry. The results are shown in terms of the increment of velocity required to transfer the satellites to the new orbits. Some comparisons with direct disposal maneuvers (Hohmann type) are also presented. We use the exact equations of motion, considering the perturbations of the Sun, the Moon, and the solar radiation pressure. The geopotential model was considered up to order and degree eight. We showed the quantitative influence of the sun and the moon on the orbit of these satellites by using the method of the integral of the forces over the time. [ABSTRACT FROM AUTHOR]

Aerodynamic configuration of the capsule for manned re-entry at the second cosmic speed requires a relatively higher lift-to-drag ratio compared to the configuration for manned re-entry at the first cosmic speed. The present capsule configurations with high lift-to-drag ratio, for instance, the Apollo and the Orion, have secondary statically stable trim points, which is highly undesirable since it would cause disastrous result once the vehicle flies at the secondary trim point during re-entry. In the present study, effective design methodology for improving the monostability characteristics of re-entry capsules without compromising the lift-to-drag ratio is revealed based on flow field characteristics analysis. As a result, aerodynamic configuration design of a capsule with high lift-to-drag ratio and monostable characteristics is proposed. And the aerodynamic characteristics of the capsule ranging from subsonic to hypersonic speed are predicted by numerical simulations. [ABSTRACT FROM AUTHOR]

A transient inverse heat conduction analysis enables the identification of the unknown boundary heat flux from finite number of temperature data obtained, e.g., in the high temperature structural tests or in the actual operation. Since the prediction of thermal load (heat flux) is difficult, the inverse analysis is expected to improve structural design of high temperature components. The present study develops a computational method of transient inverse heat conduction analysis. The developed code is applied to problems of a simple two-dimensional plate and of an atmospheric reentry capsule. Sequential Function Specification (SFS) method and Truncated Singular Value Decomposition (TSVD) are employed to improve the stability of the inverse analysis. Effects of these regularization methods are numerically discussed. [ABSTRACT FROM AUTHOR]

Purpose: Since the aerodynamic coefficients of the hypersonic vehicle vary quickly during the reentry phase, a robust Angle-of-attack (AOA) tracking control system is designed for the Hypersonic Reentry Vehicle (HRV) based on the Linear Parameter Varying (LPV) theory. Design/methodology/approach: Firstly, longitudinal moment trim is done along the desired flight trajectory. The linearized system at each trim point is built and the dynamic characteristics analysis is made. Then the LPV control law with Parameter Dependent Quadratic Lyapunov Function (PDQLF-LPV) is applied to design the AOA tracking autopilot at each trim point. Frequency performance of the autopilot is assessed and the step response simulation is conducted to validate the effectiveness of the control system. Finally, actual AOA command tracking simulations based on the time-varying nonlinear model are carried out to test the correctness and robustness of the PDQLF-LPV autopilot. Findings: Analysis results demonstrates that the PDQLF-LPV control system can track the AOA command perfectly during the whole flight envelop with dynamics parameter variation or disturbances, which indicates that it is effective to integrate the PDQLF-LPV control theory with a parameter-varying reference model for control system design of HRV. Originality/value: A reference model with varying parameters is utilized to guarantee the transient performance of the autopilot, and induced L[sub 2]-norm analysis is introduced to describe and guarantee the robust stability of the autopilot. [ABSTRACT FROM AUTHOR]

Aerothermal heating of hypersonic reentry bodies can generate significant radiation of visible light. This study investigates the conditions for ground-level visibility and what the authors term noticeability of that radiated light to the unaided human eye. This study considers radiated visible light, transmission to ground level as a function of relative observer position, sky background contrast, human contrast perception limits, and psychophysical considerations for attention capture. Considering both pure ballistic trajectories and lifting-body trajectories, this study presents examples of spatial and temporal viewing zones for both detectability and noticeability. In some cases, viewing zones extend a significant distance from the impact location and/or include the impact location for a potentially significant period of time before impact. [ABSTRACT FROM AUTHOR]

An uncertainty qnantification and sensitivity analysis computational technique based on the Monte Carlo method has been used for CubeSat drag studies for exploring issues in satellite decommissioning. These studies included computational simulations for altitudes of 100 and 400 km using the Hypersonic Aerothermodynamic Particle code, a direct simulation Monte Carlo solver with uncertainty quantification and sensitivity analysis capabilities allowing for variation of the solver input quantities, propagation of the input parameter variations through the model, and prediction of output uncertainties, all in an automated manner. Satellite atmospheric drag was identified as the main parameter of interest because uncertainty in drag may alter the satellite orbit and influence its lifespan. This study considers relative contributions of the input and model parameter uncertainties to the resulting satellite drag uncertainty, which enable reduction of the initial parameter space to a few major contributors; these major contributions are then examined in detail. The effect of surface thermal accommodation coefficient is one of the parameters considered, and it is found that its contribution has opposite effects on the drag force in free molecular and collisional flows. The study demonstrates the ability of the numerical techniques to solve the flow cases for a specified accuracy. [ABSTRACT FROM AUTHOR]

Emission signatures of the Hayabusa capsule reentry, acquired by the Australian Ultraviolet Spectrograph instrument on the NASA DC-8 mission over the South Australian Woomera landing site on 13 June 2010, are analyzed to identify thermal characteristics of the shock layer and the ablative heat-shield performance. The spectral range 330 to 470 nm of the Australian Ultraviolet Spectrograph instrument contains signatures of the air plasma species nitrogen and the ablation-related species cyanogen. The time-resolved evolution of averaged species number densities, vibrational and rotational plasma temperatures, and the heat-shield surface temperature are deduced through a multiparameter error-minimization analysis based on the European Space Agency plasma radiation database PARADE. The plasma temperatures indicate noneqnilibrium conditions for cyanogen at altitudes above 70 km and equilibrium conditions after peak heating. Two separate thermal protection system activity periods are identified by tbe analysis. Early pyrolysis activity is indicated by a peak in cyanogen and nitrogen-hydrogen emissions at altitudes around 80 km. Broader peaks in the number of cyanogen molecules and reduced nitrogen-hydrogen emissions were identified around the time of peak beating at an altitude of about 57 km and elevated surface temperatures, above 3000 K. [ABSTRACT FROM AUTHOR]

A methodology to correlate flight aeroheating environments to the arc-jet environment is presented. For a desired hot-wall flight beating rate, the methodology provides the arc-jet bulk enthalpy for the corresponding cold-wall heating rate. A series of analyses were conducted to examine the effects of the test sample model holder geometry to the overall performance of the test sample. The analyses were compared with arc-jet test samples, and challenges and issues are presented. The transient flight environment was calculated for the Hypersonic Inflatable Aerodynamic Decelerator Earth Atmospheric Reentry Test vehicle, which is a planned demonstration vehicle using a large inflatable, flexible thermal protection system to reenter the Earth's atmosphere from the International Space Station. A series of correlations were developed to define the relevant arc-jet test environment to properly approximate the vehicle flight environment. The computed arc-jet environments were compared with the measured arc-jet values to define the uncertainty of the correlated environment. The results show that, for a given flight surface heat flux and a fully catalytic thermal protection system, the flight-relevant arc-jet heat flux increases with the arc-jet bulk enthalpy, while for a noncatalytic thermal protection system, the arc-jet heat flux decreases with the bulk enthalpy. [ABSTRACT FROM AUTHOR]

The article covers the outcome of the method for optimal filtering of measurements developed by the author and aimed at the determination of the time and place of reentry of the Phobos-Grunt spacecraft. So-called two-line elements (TLE) of the orbit of the American Space Surveillance System are used as measurements. [ABSTRACT FROM AUTHOR]

Abstract: Traditional air pollution epidemiology studies being conducted in large cities can be limited by the availability of monitoring. Satellite Aerosol Optical Depth (AOD) measurements offer the possibility of exposure estimates for the entire population. We previously demonstrated that daily calibration substantially increased the predictive power of satellite AOD measurements for fine particles (PM2.5) in New England, allowing estimation of exposure in locations without monitors. Similar results have not been reported for larger particles (PM10), which are often the only measures that can be used in locations worldwide that do not have comprehensive PM2.5 monitoring; this also applies to PM estimation of historical exposures. Here we extend this methodology by applying it to 2000–2009 PM10 data from Lombardy, Northern Italy a region with high altitude differences, frequent temperature inversions and stationary fronts. Specifically, by 1) incorporating a model for missing AOD data to deal with non-randomness in the missing data; and 2) modeling interactions between land use and meteorological parameters to better capture space–time interactions. We calibrated AOD data through mixed-effect models regressing PM10 measurements using day-specific random intercepts, and fixed and random AOD and temperature slopes. We used inverse-probability weighting to account for nonrandom AOD missingness, while reducing the dimensionality of spatial and temporal predictors and avoiding selecting different predictors in different locations, as common in land-use regression. We take advantage of associations of grid-cell AOD values with PM10 monitoring located elsewhere and with AOD values in neighboring grid cells to develop grid-cell predictions when AOD is missing. By using ten-fold cross-validation to test the accuracy of our predictions, we found high out-of-sample R 2 (R 2 = 0.787, year to year variation 0.738–0.818) for days with available AOD data. Even in days with no available AOD, our model performance was comparable (R 2 = 0.787, year to year variation 0.736–0.841). Our results demonstrate that PM10 can be reliably predicted using this AOD-based prediction model, even in a geographical area with complex geographic and weather patterns. [Copyright &y& Elsevier]

The aerothermodynamic environment around an 8.3 m High-Energy Atmospheric Reentry Test vehicle is assessed. Twelve nose-shape configurations are generated, and their responses at the peak heating trajectory point are compared against the baseline nose shape. The possibility of a two-piece thermal protection system design at the nose is also considered, as are the surface catalytic affects of the aeroheating environment of such a configuration. Based on these analyses, an optimum nose shape is proposed to minimize the surface heating. The heat flux sensitivity of the optimum geometry to the angle of attack variations are also studied. A recommendation is also made for a two-piece thermal protection system design, for which the surface catalytic uncertainty associated with the jump in heating at the nose-inflatable aerodynamic decelerator juncture is reduced by a minimum of 93%. In this paper, the aeroshell is assumed to be rigid, and the inflatable fluid interaction effect is left for future investigations. [ABSTRACT FROM AUTHOR]

This paper presents a new concept for atmospheric reentry online optimal guidance and control using a method called MARE G&C that exploits the different time scale featured by reentry dynamics. The new technique reaches a quasi-analytical solution and simplified computations, even considering both lift-to-drag ratio and aerodynamic roll as control variables; in addition, the paper offers a solution for the challenging path constraints issue, getting inspiration from the inverse problem methodology. The final resulting algorithm seems suitable for onboard predictive guidance, a new need for future space missions. [ABSTRACT FROM AUTHOR]

During atmospheric reentry, radiative heating is one of the most important component of the total heat flux. In this paper, we investigate how the thermal radiation coming from the postshock region interacts with the spacecraft structure. A model that takes into account the radiation reflected by the surface is developed and implemented in a solid solver. A partitioned algorithm performs the coupling between the fluid and the solid thermal fields. Numerical simulation of a hollow cone head and a deployed flap region shows the effects of the radiative cooling and the significance of the surface radiation. [ABSTRACT FROM AUTHOR]

Abstract: The presence of random fluctuations of air temperature within the Earth’s atmosphere is a well-documented phenomenon. During the past seventy years there have been numerous experimental efforts to accurately measure air temperature as a function of altitude and, through careful data analysis, provide statistics describing these fluctuations and the associated fluctuations in temperature gradients. In addition, several researchers suggest the presence of atmospheric layers or “sheets” where the statistics describing fluctuations in air temperature can vary significantly from layer to layer. Herein, we propose a model to represent fluctuations of air temperature within a layered atmosphere. The model is a special type of inhomogeneous non-Gaussian differentiable random process and can be calibrated to available data on the marginal statistics and spectral content of the fluctuating temperature field, as well as the associated first derivative of the process representing fluctuations in temperature gradients. Properties of the proposed model are presented, and statistical realizations of the fluctuating temperature field and its gradient are computed and presented for illustration. The random vibration response of a spacecraft falling to Earth through these fluctuating conditions is then considered to demonstrate the usefulness of the proposed model. [Copyright &y& Elsevier]

Abstract: The present paper aims at improving the efficiency of the robustness analyses of flight control laws with respect to conventional techniques, especially when applied to vehicles following time-varying reference trajectories, such as in an atmospheric re-entry. A nonlinear robustness criterion is proposed, stemming from the practical stability framework, which allows dealing effectively with such cases. A novel approach is presented, which exploits the convexity of linear time varying systems, coupled to an approximate description of the original nonlinear system by a certain number of its time-varying linearizations. The suitability of the approximating systems is evaluated in a probabilistic fashion making use of the unscented transformation technique. The effectiveness and potentials of the method are ascertained by application to the robustness analysis of the longitudinal flight control laws of the Italian Aerospace Research Center (CIRA) experimental vehicle USV. [Copyright &y& Elsevier]

Abstract: The results of static and hot tests of ceramic matrix composite fasteners for heat shield attachment to a re-entry vehicle are presented. Herein, the effect of the high temperature cycle occurring during atmospheric re-entry on the reduction of the tightening torque of the CMC fasteners is investigated. From cold testing, it is found that the Young’s modulus of the material decreases with first-time loading and, therefore, previous loading of the fasteners before their application is recommended. In hot testing of a fully integrated fastener connection, the material temperatures within the connection are experimentally simulated. It is shown that with previous straining, a significant improvement of the fastener is achieved. The loosening torque is no longer reduced by thermally induced stretching of the fastener. [Copyright &y& Elsevier]

A multi-physics frame work has been setup for the simulation of surface heat flux for nonablating hypersonic reentry vehicles and presented in this paper. The main goal of this work was to set up a simple approach for the heat flux prediction during the reentry of the vehicle. The vehicle considered in the calculation is an axisymmetric vehicle flying at zero degree angle of attack. Chemical nonequilibrium in the flowfield is simulated by implementing a set of finite rate equations in the laminar finite rate model in FLUENT. The frame work set up was validated with the results available in the literature. Good correlation was observed between the results from the commercial code with the implemented equations and the results from the literature. [ABSTRACT FROM AUTHOR]

Abstract: The studied type of Crew Transfer Vehicle (CTV) is basically a modified ATV and a capsule for the crew. The emergency trajectories of the CTV consist of a propelled exo-atmospheric arc and the re-entry of the capsule only. The starting point is a failure in the Ariane5 EPC (Étage Principal Cryotechnique) after the Launch Escape System (LES) has been jettisoned, i.e. in exo-atmospheric flight. The CTV emergency trajectories must assure the safe re-entry of the crew while avoiding any risk to population, i.e. descent on land is forbidden except in specially prepared areas, and shall minimize the number of rescue teams. Different strategies are analyzed, abort to Atlantic, abort to orbit and abort to Woomera (Australia). The launch and ascent trajectory of a launcher similar to Ariane5 EPS-V (two burns) is optimized (maximum payload into desired orbit) to obtain the initial conditions at failure. Other sub-optimal launch and ascent trajectories are computed to fulfill additional constraints of the CTV trajectories not applicable in conventional Ariane5 launches. The emergency trajectories computation will imply the optimization of the CTV propelled trajectory, the exo-atmospheric flight, and the capsule descent trajectory. Different lift-to-drag ratios are analyzed corresponding to different capsule designs (bi-conic and Viking-type). The CTV thrust level, capsule control strategy and launcher ascent trajectory constraints will be derived in order to assure the overlapping of the strategies, the avoidance of landing on Europe and the safety of the crew during the re-entry. [Copyright &y& Elsevier]

In the framework of Research Task Group 043 of the NATO Research and Technology Organization, an analysis of the capabilities in prediction of aerothermal loads acting on a crew exploration vehicle at the higher altitudes of its reentry trajectory has been performed. In particular, the focus of this investigation is to provide information where overlap between the continuum approach (i.e., computational fluid dynamics) and particle approach (i.e., direct simulation Monte Carlo) occurs and improves understanding of relevant physics in transitional regime. Computational fluid dynamics calculations with slip flow boundary conditions had shown good predicting capabilities of slip velocity, slip temperature, and pressure, but not the surface heat flux. The analysis of the contributions to the total surface heat flux had exhibited the lack of continuum model in evaluating the convective heat flux in the transitional zone of the Orion reentry trajectory and an underestimation of shock wave thickness. [ABSTRACT FROM AUTHOR]

This paper presents a numerical investigation of the hypersonic reacting flow around the FIRE II reentry capsule. At the chosen freestream conditions, the forebody boundary layer and the separated flow on the afterbody are turbulent. The Reynolds-averaged Navier-Stokes method along with two commonly used turbulence models are used to compute the flowfield. Accurate prediction of turbulent separated flow at hypersonic conditions is challenging due to the limitations of the underlying turbulence models. The presence of turbulent eddy viscosity in the flow simulation results in a smaller separation bubble than the laminar solution at identical conditions. Also, the two turbulence models predict different levels of eddy viscosity in the neck region. This has a dominant effect on the separation bubble size and the surface pressure. On the other hand, the eddy viscosity values in the near-wall region determine the heat transfer rate to the body. The two models predict comparable heating rates on the conical frustum, and the results match in-flight measurement well. By comparison, surface pressure predictions are appreciably higher than the data. [ABSTRACT FROM AUTHOR]

Abstract: This paper highlights the design, qualification and mission performance of the tether deployer system on the second Young Engineers’ Satellite (YES2), that featured a tethered momentum transfer. The deployer is designed with a broad range of near-term tether applications in mind. The system contains the tether, including features to enhance safety and wound up in controlled manner onto a spool core, optical deployment sensors, a “barberpole” friction brake controlled by a stepper motor and a triple tether cutter system. To initiate the deployment a spring-based ejection system was developed, and to apply accurate momentum transfer a timer and release system is present on the subsatellite side. A small, 6kg re-entry capsule was developed as subsatellite. On September 25th, 2007, YES2 deployed a 32km tether in orbit and gathered a wealth of data. Confidence is gained from the mission results for use of the deployer in future missions. [Copyright &y& Elsevier]

Abstract: The history of the deployment of nuclear reactors in Earth orbits is reviewed with emphases on lessons learned and the operation and safety experiences. The former Soviet Union''s “BUK” power systems, with SiGe thermoelectric conversion and fast neutron energy spectrum reactors, powered a total of 31 Radar Ocean Reconnaissance Satellites (RORSATs) from 1970 to 1988 in 260km orbit. Two of the former Soviet Union''s TOPAZ reactors, with in-core thermionic conversion and epithermal neutron energy spectrum, powered two Cosmos missions launched in 1987 in ∼800km orbit. The US’ SNAP-10A system, with SiGe energy conversion and a thermal neutron energy spectrum reactor, was launched in 1965 in 1300km orbit. The three reactor systems used liquid NaK-78 coolant, stainless steel structure and highly enriched uranium fuel (90–96wt%) and operated at a reactor exit temperature of 833–973K. The BUK reactors used U-Mo fuel rods, TOPAZ used UO2 fuel rods and four ZrH moderator disks, and the SNAP-10A used moderated U-ZrH fuel rods. These low power space reactor systems were designed for short missions (∼0.5kWe and ∼1 year for SNAP-10A, <3.0kWe and <6 months for BUK, and ∼5.5kWe and up to 1 year for TOPAZ). The deactivated BUK reactors at the end of mission, which varied in duration from a few hours to ∼4.5 months, were boosted into ∼800km storage orbit with a decay life of more than 600 year. The ejection of the last 16 BUK reactor fuel cores caused significant contamination of Earth orbits with NaK droplets that varied in sizes from a few microns to 5cm. Power systems to enhance or enable future interplanetary exploration, in-situ resources utilization on Mars and the Moon, and civilian missions in 1000–3000km orbits would generate significantly more power of 10''s to 100''s kWe for 5–10 years, or even longer. A number of design options to enhance the operation reliability and safety of these high power space reactor power systems are presented and discussed. [Copyright &y& Elsevier]

Abstract: The paper concerns the general problem of a bounded final state control of non-linear dynamic systems with reference to near-optimal predictive guidance for low lift-to-drag ratio re-entry vehicles. More specifically, it addresses deriving guidance strategies capable to provide a maximal downrange maneuvrability for a maximal remaining flight time. Such robust, “guaranteed”, or assured, guidance keeps the remaining range-to-go to be coincident with the center of instant attainability domain. The paper discusses the existing guaranteed guidance strategy, and presents more general approach that provides an on-board planning of the entry trajectory, thus giving future state and control profiles. As a consequence the proposed guidance law is able to satisfy not only specified terminal conditions but also typical inequality constraints such as the maximal load factor and heat load. Computer simulations show that the algorithm can generate the feasible trajectories with equal downrange margins, using simple two-parametric families of control functions. [Copyright &y& Elsevier]

Abstract: In early 2008, the need arose to predict the orbital decay of the American spacecraft USA-193, whose characteristics, function and orbit were classified information. With no orbit data and independent Italian tracking capability available, we turned our attention on the orbits determined by a worldwide network of about 20 visual satellite observers. The orbits of USA-193 obtained from their visual observations were therefore used as the sole source of orbit information. Contrary to our expectations, this exercise was extremely successful and we learned a lot in the process. The orbits provided by the visual observers were very accurate for such a low satellite (although the minimum and very stable level of solar activity helped considerably); however, data gaps of a few days were sometimes possible, due to unfavorable pass geometry or weather and light conditions. In any case, the orbital period and the semimajor axis were so accurate that it was possible for us to obtain very good decay fits using special perturbation software, including various atmospheric density models together with all the other relevant perturbing accelerations. We were therefore able to estimate accurate values of the ballistic parameter and the resulting decay and reentry predictions were extremely stable. Amateur optical observations and images of USA-193 had also led to a rough estimation of the shape and sizes of the satellite, revealing that the solar arrays had never been deployed. With this information, and taking into account our estimates of the ballistic parameter, we obtained reasonable and consistent values of the spacecraft mass. Based on previous reentry fragmentation analyses, we were then able to guess the expected USA-193 casualty area, casualty expectancy, debris ground footprint and probability of impact in Italy. Lastly, after the decision by the US Government to destroy the satellite, we independently predicted the interception time windows and the post-event ground tracks. Following the successful spacecraft breakup, we analyzed the evolution of the resulting debris cloud and assessed its (very limited) adverse impact on the circumterrestrial environment. [Copyright &y& Elsevier]

Abstract: At IRS the RESPECT sensor system for in-flight emission spectroscopic measurements of re-entry body post-shock regimes is currently under development. Design solutions for this sensor system based on the basic conditions of ESA''s EXPERT program are presented. The sensor head design is described with emphasis on the optical access to the plasma in the stagnation area. To justify the presented design the thermal and mechanical tests already performed are described. The present tests include thermal tests of the TPS window assembly and transmittance investigations of the sapphire window at high temperatures using plasma wind tunnel facilities at DLR and IRS. Both test campaigns have been concluded successfully, showing that the window assembly is withstanding the thermal loads of the EXPERT re-entry and that no substantial changes in the sapphire transmittance properties could be detected up to temperatures of . Mechanical testing has been limited so far to the spectrometer unit and has been performed using shaker facilities. The tests demonstrate that the chosen OceanOptics S2000 spectrometer can withstand the EXPERT-Volna vibrational launch environment and the EXPERT ballistic re-entry. In addition, the functionality of the sensor system in the harsh environment imposed by plasma wind tunnel tests has been approved. Summarizing all the tests performed to support the development of the sensor system it can be concluded that the presented sensor design can be fully applied for atmospheric (re-)entry mission such as EXPERT. [Copyright &y& Elsevier]

Exoatmospheric nuclear effects, such as those that would be encounter by reentry bodies, provide instantaneous (near zero-duration), impulsive loading of structures. Endoatmospheric nuclear effects possess an impulse that is finite in duration, but whose rise time is still instantaneous. The commonality of these loadings is that they initiate waves propagating through structures, resulting in extremely short duration accelerations to free surfaces where accelerometers are mounted. Over the years, attempts have been made to measure free surface accelerations using ceramic, quartz, and piezoresistive accelerometers. This paper describes the lessons learned, and looks to the future. It also provides a history of shock accelerometer development. [ABSTRACT FROM AUTHOR]

Abstract: The unmanned space vehicle (USV) program has been undertaken by the Italian Center for Aerospace Research with the aim of developing flying test beds of next generation reentry launch vehicles. In this framework, the development of small demonstrators is also foreseen to validate technological and operational aspects of full-scale vehicles and missions. In this paper, a small-scale demonstrator of the sub-orbital re-entry test mission of the USV program is described. Both mission profile and objectives are very challenging in terms of demonstrator guidance, navigation and control. After a short description of the mission and demonstrator architectures, particular emphasis is given to the guidance and navigation analysis. To this end, mission objectives and reduced-scale system constaints are integrated and translated into innovative guidance solutions relying on optimization techniques. Then, performance of a commercial-off-the-shelf GPS-aided, miniature inertial navigation system over the proposed trajectories is evaluated by Monte Carlo analysis. Standalone inertial and GPS-aided inertial navigation performance is also compared considering GPS loss conditions due to antenna plasma effects. [Copyright &y& Elsevier]