Deployable Heat Shield Research Articles

Mini-IRENE, a Successful Re-entry Flight of a Deployable Heatshield Capsule

Mini-IRENE, a Successful Re-entry Flight of a Deployable Heatshield Capsule

Incremental model predictive control for satellite de-orbiting based on drag modulation

This paper deals with the problem of controlled atmospheric re-entry, which has recently gained growing interest from the aerospace scientific and industrial communities for the economic and environmental benefits related to the development of reusable space vehicles. In this context, an original technique based on the incremental formulation of the Model Predictive Control methodology is proposed to tackle the trajectory tracking problem of a small satellite equipped with a deployable heat shield in the de-orbiting phase, focusing on a range of altitude from 300 km to 130 km, i.e., before the atmosphere is dense enough to begin a terminal re-entry phase. The proposed control logic allows following a given pre-determined guidance law, limiting position offsets due to unmodelled disturbances or uncertainties in the knowledge of spacecraft parameters (e.g., drag coefficient). An extensive numerical simulation campaign to assess the control performances in different perturbing conditions has demonstrated its ability to reach a km-error level in the position error at the final altitude of 130 km.

Precise re-entry and landing of propellantless spacecraft

• Propellant-free spacecraft ground impact point targeting algorithm. • Guidance and feedback control targeting strategy using exclusively aerodynamic drag. • Monte Carlo simulation results proving algorithm robustness and performance. Spacecraft returning scientific samples or humans from space must be capable of surviving re-entry and landing in a desired location. Traditionally, this has been accomplished via a propulsive de-orbit burn. However, it is not always possible to mount a propulsion system on board a satellite or a capsule. In the case of small satellites deployed from the International Space Station, for example, on-board propulsion systems are forbidden for safety reasons. Our work proposes a new technological solution for re-entering and landing a spacecraft in a desired location from a low Earth orbit using exclusively aerodynamic drag and eliminating the need for chemical propulsion. First, an iterative procedure is utilized to compute the desired state at the re-entry interface (100 km) such that a propagation of the vehicle dynamics in the nominal re-entry drag configuration from this initial state leads to a landing at a desired latitude and longitude on the surface of the Earth. Next, a re-entry point targeting algorithm is utilized to determine the on-orbit ballistic coefficient profile necessary to target the desired re-entry point. Finally, the ballistic coefficient profile during the final hours of the trajectory before the re-entry interface is iteratively modified to correct any remaining along-track error in the landing location. The proposed solution is applied to a small satellite system that is jettisoned from the ISS and is equipped with a deployable heat shield that also serves as a drag device.

Feasibility Study of Aerocapture at Mars with an Innovative Deployable Heat Shield

Performing orbital insertion around Mars using aerocapture instead of a propulsive orbit insertion maneuver allows the saving of resources and/or the increase of the payload mass fraction. Aerocapture has never been employed to date because of the high uncertainties in the parameters from which it depends, mainly related to atmospheric density modeling and navigation errors. The purpose of this work is to investigate the feasibility of aerocapture at Mars with an innovative deployable drag device, whose aperture can be modulated in flight, and to assess the effects of the main uncertainties on the success of the maneuver. This paper starts with the presentation of a parametric bidimensional analysis of the effectiveness of aerocapture, for which a wide range of uncertainty levels in the atmospheric density and the ballistic coefficient are considered. Then, an application to a real mission scenario is carried out, including the error of the targeting maneuver performed at the limit of the sphere of influence of the planet. The analyses show the strong influence of the uncertainties in the atmospheric density and the ballistic coefficient, which significantly narrow the solution space and limit its continuity. However, viable solutions for aerocapture can still be identified, even in the worst conditions.

Aerodynamic Control System for a Deployable Re-entry Capsule

Deployable aerobrakes for Earth re-entry capsules may offer many advantages in the near future, including the opportunity to recover on Earth payloads and samples from Space with reduced risks and costs with respect to conventional systems. Such capsules can be accommodated in the selected launcher in folded configuration, optimizing the available volume, and when foreseen by the mission profile the aerobrake can be deployed in order to increase the surface exposed to the hypersonic flow and therefore to reduce the ballistic parameter.The ballistic parameter reduction offers as main advantage the opportunity to perform an atmospheric re-entry with reduced aerothermal and mechanical loads. It makes also possible the use of relatively lightweight and cheap thermal protection materials.There are plenty of algorithms and control methods for classical re-entry capsules, such as the Orion and Soyuz re-entry modules. However, a lot of attention is now paid towards how to control low ballistic coefficient capsules with inflatable and mechanically deployable heat shield.The aim of this study is to prove the feasibility of an aerodynamic control system, applied to a mechanically deployable re-entry capsule, in order to increase landing precision at the targeted site.The deployable heat shield is equipped with an actuation system, consisting of eight aerodynamic surfaces, referred to as flaps. The assumed control strategy is to deflect the flaps independently, in order to trim the capsule and produce enough lift and side force to give downrange and cross range maneuverability during the re-entry phase.A control algorithm was developed, implemented and tested in a 3DOF simulation environment. Capsule performances both for uncontrolled ballistic re-entries and for controlled lift re-entries starting from a low Earth orbit have been assessed, verifying the capability of the controller in guiding the capsule toward the chosen target.Monte Carlo simulations were run assuming errors and uncertainties at the de-orbit burn, and the control system has been proved to obtain good results in reducing dispersions at the landing site.

Aerothermodynamics and thermal design for on-ground and in-flight testing of a deployable heat shield capsule

The article deals with the aerothermodynamic and thermal analysis of two technological demonstrators suitable for research activities on deployable heatshield entry capsules. Variable geometry bodies represent a formidable way to build low ballistic coefficient entry platforms able to limit convective heat fluxes, mechanical loads, and final descent velocity. These characteristics suggest several interesting applications for space missions such as to exploit those systems for human or large payloads landing on Mars. The main feature of the demonstrators is the umbrella-like deployable heatshield made of a flexible thermal protection system. In this framework, the work reports on the design analyses carried out for a flight and a ground demonstrator developed to prove, with suborbital flight and plasma wind tunnel test campaigns, several key technologies of the deployable heatshield. Several non-equilibrium three-dimensional flowfield simulations were performed to assess the aerothermal loading conditions expected for the demonstrators. Pressure and heat flux distributions on the deployable heatshield are provided and discussed. Finally, those surface loading conditions are set to feed a finite-element thermal analysis of the demonstrators, also detailed in the work.

Downrange manoeuvre and oscillation suppression of a self-regulating centrifugally deployed flexible heat shield using a controlled reaction wheel

A recent study has introduced a flexible deployable heat shield that passively deploys and stiffens due to centrifugal forces generated from a self-regulated autorotation. This paper demonstrates that the heat shield is similar to a PI controlled second order nonlinear system, which explains why the deployment is accompanied by a limit cycle structural oscillation that persists throughout a simulated re-entry. The heat shield design offers a unique capability to actively adjust the deployment using conventional attitude control devices. This operation is explored by simulating the re-entry of a CubeSat-sized vehicle equipped with an off-the-shelf reaction wheel controlled by a switching phase shift controller and gain-scheduled controllers. The effects of the control parameters are investigated, and successful oscillation suppression as well as an open-loop downrange manoeuvre of over 300 km is predicted for re-entry from low earth orbit.

TANDEM: Tension Adjustable Network for Deploying Entry Membrane

The Tension Adjustable Network for Deploying Entry Membrane (TANDEM) is a tensegrity rover designed to also be the frame of a deployable heat shield. TANDEM integrates the heatshield technology developed for the ADEPT entry system with an active infrastructure to combine entry, descent, and landing as well as surface locomotion into a single multifunctional system. Reusing the same active infrastructure for every phase of the mission makes TANDEM an efficient and versatile system for medium- to large-sized payloads. This efficiency translates directly to longer mission lives, increased scientific payload, or a reduction in mission cost and total mass of the whole mission. A conceptual mission to the surface of Venus is presented to illustrate the merit of the concept and is supported by proof-of-concept level calculations and simulations. The TANDEM vehicle designed for the conceptual mission is compared with two alternative designs from the literature, the VITAL mission and ADEPT-VITaL mission.

Aerodynamics and flight mechanics activities for a suborbital flight test of a deployable heat shield capsule

MINI-IRENE is the Flight Demonstrator of IRENE, a new-concept capsule with a variable geometry, originally conceived by ASI to widen the range of available platforms to retrieve payloads and/or data from low Earth orbit. The main characteristics of IRENE is the “umbrella-like" deployable front structure that reduces the capsule ballistic coefficient, leading to acceptable heat fluxes, mechanical loads, stability and final descent velocity. Following the feasibility studies carried out since 2011, with also preliminary Thermal Protection System materials tests in plasma wind tunnels, the objective is now to design and build a Flight Demonstrator and a Ground Demonstrator to prove, with a suborbital flight and with a Plasma Wind Tunnel (PWT) test campaign, the functionality of the deployable heat shield. The Flight Demonstrator shall be included as a secondary payload in the interstage adapter of a VSB-30 launcher from ESRANGE, then ejected during the ascent phase of the payload section, perform a 15-min ballistic flight, re-enter the atmosphere and hit the ground. The Ground Demonstrator, representative of the Thermal Protection System of the Flight Demonstrator, shall be instead exposed to a heat flux similar to that expected for an atmospheric re-entry from low Earth orbit inside the SCIROCCO Plasma Wind Tunnel at CIRA. The paper, after a short description of the mission profile both for orbital and suborbital flights, focuses on the aerodynamics and flight mechanics activities held for the suborbital flight and PWT test campaigns.

Deployable heatshields for future multistage missiles

On a multistage ballistic missile, extensive insulation is usually required on the aft side of powered upper stages to protect the components and structure from excessive heat generated by the upper main stage motor, The objective of this program is the development of a lightweight, inexpensive, reliable heatshield to protect the aft end components and structure of a powered upper stage from the thermal environment generated by direct plume impingement and radiation from the upper-stage motor. Presented in this paper are two deployable heatshield concepts—inflatable torus and flex-rib—which offer a potential range performance improvement and cost advantage when compared to a current approach of individually insulating each component and structure as is effected on the Trident I fleet ballistic missile.