Open access peer-reviewed chapter

Plasma Preheating Technology for Ablation Studies of Hypersonic Reentry Vehicles

Written By

Daniel Odion Iyinomen

Submitted: 08 August 2021 Reviewed: 25 August 2021 Published: 04 May 2022

DOI: 10.5772/intechopen.100129

From the Edited Volume

Hypersonic Vehicles - Applications, Recent Advances, and Perspectives

Edited by Giuseppe Pezzella and Antonio Viviani

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Measurements of mass ablation rates in hypersonic flows are used to calibrate computational models. A novel plasma technique for preheating axisymmetric samples of heatshield materials has been developed and applied to ablation of graphite sample through this work. The experimental probe was very similar to the European standard probe, normal to the flow. The Scanning Electron Microscope (SEM) was used to examine the surface characteristics; and the experiments show a significant spatial variation in thickness loss for the graphite test material over the disc radius though the spatial variation was still largely axisymmetric. The present work does support reasonable contributions to reentries and can further be developed to validate computational models under conditions that replicate characteristics of reentry flights.


  • Heatshield
  • Plasma
  • Preheating
  • Technology
  • Ablation
  • Hypersonic
  • Planetary
  • Re-entry vehicles
  • Surface temperatures

1. Introduction

To achieve practical reentries of flight vehicles, speeds in excess of the earth’s escape velocity (≥ 11 km/s) are needed [1]. On reentry, this hypersonic speed is slowed down by viscous drag: firstly, thermal energy is created through the deceleration and compression of the high velocity incoming flow to a high pressure and high temperature in the hypersonic boundary layer close to the surface (heatshield); and secondly the convective and radiative heating associated with high temperatures accelerates the wall reactions [2]. The blunt shape of heatshields consisting of carbon-based materials [3] are effective structures that enable most of the generated heat to be carried away from the vehicle [4]. Computational models of the heat loads [5, 6] experienced during atmospheric reentries are continually being updated [7, 8]. The validation of these models is critical for the safe and economic design of future flight vehicles [9]. Proper analysis of heat flux [10], gas/surface interactions [11], and properties of heatshield materials [12] are needed for ablation performance and evaluations [13, 14]. Various thermochemical processes during reentry [15], descent, and landing also support the ablation of heatshields [16]. Flow parameters [17] such as enthalpy, stagnation pressure, velocity and heat flux are identified in Figure 1. It explains how the velocity of a reentry Stardust is slowed from 12.8 km/s to almost 2 km/s. A significant reduction in velocity is experienced between 70 to 40 km altitude due to higher atmospheric density resulting in more effective drag. At the same time, the heat-flux rises to a peak of 11 MW/m2 at 62 km altitude. All space vehicles like Apollo, Huygens, Stardust, Hayabusa and Orion generate heat at reentry as a consequence of the drag used to reduce speed [19].

Figure 1.

Stardust trajectory parameter [18].


2. Overview

This work helps to explain quantitative measurements of graphite ablation and oxidation rates in a hypersonic boundary layer using a new hot-surface plasma preheating technique to raise the temperature of the model to the required surface temperature for reentry studies in a hypersonic impulse facility [20]. Hot-wall reentry tests have been carried out in hypersonic impulse facilities [21] with temperatures characteristic of ablators in hypervelocity reentry of approximately 2000–3000 K in 8.6 km/s for Earth reentry flow [22]. Experiments with electrically preheated graphite samples using uniform width profiles were conducted at the University of Queensland’s X2 expansion tunnel [23] for surface temperatures from 1770 to 2410 K used to target the carbon–nitrogen violet band [24]. All of these used resistive heating on a uniform width profile and not a plasma preheating technique. Recent experiments based on Orion reentry conditions used a surface temperature of about 2800 K [25] and Apollo 4 lunar return speed of 11 km/s reportedly experienced a surface temperature of about 2400 K [26], while typical reentry surface temperatures for Space Shuttles was about 1740 K [25]. This present work has achieved temperatures in excess of 2500 K on a preheated graphite surface used for the assessment of mass loss through ablation in a Mach 4.5 and Mach 6 flows. The use of a plasma as the means of heating the disc was proven to be effective for reaching temperatures of these magnitudes [27] and the process has potential to be more widely used for similar experiments [28]. The technique uses a plasma heating source with argon flow [29]. The work reports the first-time quantification of ablation and oxidation rates of a heated carbon disc up to 500 milliseconds duration in 500 Pa vacuum pressure in a hypersonic wind tunnel facility using a plasma preheating methodology.


3. Plasma preheating methodology

The model in the present work can be likened to the Euro-model of the European Space Agency (ESA) which is a flat-faced cylinder of 50 mm in diameter with rounded corners of 1 mm radius, positioned at a 0° angle of attack to the incoming flow. Preheating of the graphite disc was achieved with plasma [30, 31] generated by a DC current [32, 33] between a tungsten electrode and the back (downstream) side of the disc. Figure 2 is a sectional view of the model which illustrates the heat transfer processes from the hot plasma to the disc. The orientation of the tungsten inert gas (TIG) is centralised in the model to enable even thermal spread from the centre to the edges. This made it possible for the probe model to assume an axisymmetric orientation.

Figure 2.

The new plasma preheating technique illustrating the heat transfer processes from hot plasma to graphite disc [34].

The disc was set up and heated for 15 seconds. The flow was initiated after 15 seconds of heating, manually timed. The heating remained on during the flow and for about 0.5 seconds after the flow stopped. The present work adopts heatshields of a constant thickness of 2 mm and the thickness was considered suitable for ablation experiments [27]. Using graphite material and a current rating of 400 A, it took about 15 seconds for the plain disk-sample to attain a steady state. Using this current and invoking graphite material properties into FEA simulation with Ansys [29], the results from simulations were found to have a good agreement with experiments as shown in Figure 3. The first 15 seconds was for heating, followed by the flow which lasted for about 0.5 seconds, and then finally cooling due to power cut-off from supply. The Schlieren technique based on the principle of changing densities in the gas was used to identify the establishment of the bow shock and therefore the commencement of flow [35]. Obtaining the plenum pressure from pressure survey array, run-times were matched with Schlieren images to identify when flow starts [27]. The high-speed camera was set at a frame-rate of 2500 fps and frames at the specific points during the flow are shown in Figure 3.

Figure 3.

Transient behaviour of stagnation surface temperatures with heatshields of 2 mm thickness [28].

Figure 4 illustrates the flow-dynamics and the associated aerothermodynamic gradient along the surface, where the source of heat flux for the heatshield is the plasma inside the probe. The temperature is driven by the plasma at the backside of experimental sample, while the aerodynamic flow is driven by the forced convection from hypersonic impulse facility at the frontside of sample (heatshield specimen). The aerodynamic flow-velocity at the surface of the experimental sample increases from the stagnation point to the edges, while the surface temperature decreases from the stagnation point to the edges. While the aerothermodynamic flow (cooling of sample) occurs at the front, the plasmadynamic flow (heating of sample) occurs at the backside of the disk. The plasma zone describes the region occupied by the plasma [36]. At steady state conditions, the inert gas flows through the shroud, gain some heat energy from the centralised hot tungsten electrode, then experiences a drastic rise in enthalpy as it passes through the hot plasma towards the heatshield, before finally exiting via the vent as shown in Figure 5.

Figure 4.

Schematic illustration of surface aerothermodynamic flow properties [36].

Figure 5.

Test conducted at Mach 4.5 to validate plasma preheating technology [28].

The aerothermodynamic flow properties for the plain disk of radius 25 mm were simulated using Ansys Fluent CFD. The CFD results using an axisymmetric graphite sample, showed a good temperature distribution for ablation rate experiments in hypersonic impulse facilities. The temperature is driven by the plasma at the backside of experimental sample, while the flow-field is driven by the hypersonic impulse facility at the frontside of sample [36]. Figure 6 shows the density variations from Mach 0 to 6, using 2D axisymmetric simulations. The density gradually increased from no-flow conditions to Mach 6 hypersonic flow. Numerical and experimental analysis at Mach 4.5 have been presented extensively by the author in other publications [28, 29].

Figure 6.

Density variations from stagnation, via transonic, to supersonic and hypersonic flows. (a) Mach zero. (b) Mach 1 flow. (c) Mach 2 flow. (d) Mach 3 flow. (e) Mach 4 flow. (f) Mach 4.5 flow. (g) Mach 5 flow. (h) Mach 6 flow.

The reaction species were simulated by using 21% O2 and 79% N2 at the hypersonic inlet, to flow over heated graphite surface [29]. The pressure gradient in Figure 7a shows the bow shock standoff distance, while Figure 7b shows the contour result of Carbon II oxide species using 2D axisymmetric simulation in Mach 4.5 flow. The ablation rate gradually decreased from stagnation point to the edges of the disk. Along the stagnation line of Figure 7a, the upstream edge of the experimental bow shock was about 15 mm from the surface, very similar to the bow shock position described in Figure 5. The numerical and experimental results from Mach 4.5 were very similar to that of Mach 6.

Figure 7.

Flow and species parameters at Mach 4.5 flow from simulations. (a) Bow shock standoff distance. (b) CO mass fraction [29].

Flow properties associated with the present work were obtained from CFD simulations using ANSYS Fluent. Figure 8 shows the simulation results using a density-based solver for the physical parameters along the stagnation line for a plain disk sample [29]. Surface temperatures and boundary conditions in the CFD were set using measurement information from experiments. Figure 8 indicates that the temperature has not changed significantly until within 1 mm from the wall, so the concentration of products was not entirely driven by temperature alone. The Mach 4.5 flow parameters were compared with that of Mach 6. The values along the stagnation line were roughly the same. The origin of the horizontal-axes in these figures is at the surface of the disk. In the case of Mach 4.5, the pressure increased from 400 Pa to 18.6 kPa across the shock in the flow direction. The gas density increases from 0.045 kgm3 to 0.22 kgm3 across the shock, and continues to rise to a maximum value of 0.23 kgm3 at about 4 mm from the wall; and finally dropping to a minimum value of 0.033 kgm3 at the wall. The static temperature rises from 60 K to 272 K across the shock, maintaining this relatively constant value to about 4 mm from the hot surface, then rises to a maximum value of about 2500 K at the wall, which explains the density change [34]. The turbulence kinetic energy is greatest at the shock and falls to almost zero at the stagnation point, suggesting a laminar boundary layer. The gas velocity also decreased across the shock from about 700 m/s to about 51 m/s, and then experiencing a small rise immediately after the shock and gradually falling to zero at the stagnation point.

Figure 8.

Boundary layer flow parameters from simulations at Mach 4.5 and Mach 6.

No changes in the gas properties were evident until approximately 4 mm from the wall as shown in Figure 9. The mass fraction of N2 behaves in a similar way to that of O2 with a sharp drop near the surface as a result of the increase in mass fraction of reaction products entering the flow from the surface. The mass fraction of molecular oxygen dropped from 21% to about 8% at the wall while that of molecular nitrogen dropped from 79% to about 67% at the wall. Neither the molecular oxygen nor nitrogen concentration dropped to zero at the wall as the temperature reached by the gas was not sufficient for complete dissociation. This indicates that a mixture of N2 and O2 was still present in the reacting boundary layer. A contribution to the reduced concentration of molecular oxygen results from its consumption in combustion. The molecular oxygen was only dissociated in very close proximity to the surface as the gas temperature only reached the dissociation temperature within 0.0025 mm from the wall. Species transports are driven by flow properties. Thermally initiated chemical reactions are the formation process of all carbonaceous species in Figure 9 except for the carbon sublimation species, C, C2, C3. The sublimation species C, C2, and C3 were almost zero. The CO2 species are formed from further oxidation of CO species, thus making CO2 a secondary reaction. The result also shows that CO2 has a sharp drop from the peak value to almost zero at the surface within the experimental temperature limit of 2530 K. The CN species distribution also show a similar pattern to that of CO2. This also suggests that CN is not a product of direct surface reaction within the experimental temperature limit. This further supports the absence of dissociated nitrogen atoms which would otherwise aid direct formation of CN at the surface [29]. The simulations did not predict atomic nitrogen in the present work. This is because the temperature needed to cause N2 dissociation was not achieved in the present work. The CO formation was the major contributor to graphite mass loss and contributions from all other carbonaceous species were insignificant [29]. All carbonaceous species have their peak values around the stagnation region at the wall. The mass fraction of CO species rises to a maximum value at the wall. Mass fraction of CO2 increases to a maximum near the wall and then drops closer to the wall. Simulation results from CFD show that the peak mass fraction of CO at the surface was about 7.9%. The CO2 mass fraction was about six orders of magnitude lower than that of CO.

Figure 9.

Boundary layer species concentrations from simulations at Mach 4.5 and Mach 6.

The information contained in this material is a new experimental method for heating material samples using a high temperature plasma arc fixture. This new fixture can be integrated within cold hypersonic flow wind tunnels in order to measure material ablation and mass loss due to aerodynamic flow effects. The mass loss measurements and microscopic images can be obtained to quantitatively and qualitatively observe material mass loss and surface modifications. Finite rate ablation chemistry has been used in the present work along with volumetric and wall surface reactions [29]. The technique makes it possible to document mass loss from their specimens using dimensional values. This invention focuses on ground-based characterisation techniques for thermal protection material analysis, thus making it a valuable tool for aerothermodynamics of reentry studies.

Well prepared experiments with clear scope and structure will be of importance to the hypersonic research community, especially in the field of new ground-testing facilities for ablation analysis. The presented technology is useful for detailed material characterisation, for example, in the context of material response model validation. It is a new test facility, which can adopt a new measurement technique to generate data for model calibration and validation. The technique is able to accurately replicate the hypersonic flow characteristics and heatshield conditions at reentries. This technology is indispensable when it comes to presenting a new experimental technique for ablation analysis, highlighting new techniques for ablation measurements and providing experimental data for code validation. The key contributions to reentries include: (1) the development of the new test technique, (2) the quantifiable measurements of mass loss, and (3) connecting the measurements made to ablation theory or models.


4. Surface ablation using plasma preheating technology

4.1 Theoretical approach

To effectively couple the solid and the flow-field during ablation, the mass, momentum, energy, and species have to be conserved. The surface mass balance for each species is given in Eq. (1), where mṡ is the mass flux of species s per second determined from the surface thermochemistry, ρs is species density in kg/m3, v is the velocity vector representing the mass-averaged velocity leaving the surface, n̂ is the unit normal vector to the surface but away from the wall, and the last term is the diffusion of species to or from the surface, where D is the diffusion coefficient in m2/s, Cs is the species mass fraction [37].


The surface energy balance is expressed in Eq. (2), where qw contains both the heat conduction and the diffusive chemical heat flux, the first term on the right is the heat flux conducting energy away from the surface into the body, the second term is the radiation of heat from the surface into the flow, and the last term is the removal of energy from the surface due to mass removal [38].


Generally, Ksolid is the thermal conductivity of graphite sample in W/(m.K), Tsolid is the temperature of graphite at the edges, ε is graphite emissivity, σ is the Stefan-Boltzmann constant in W/(m2.K4), Tw is the surface temperature, Ta is the surrounding temperature, and ho,w is species enthalpy in J/kg [39]. Unlike the negative heat flux to the wall from the flow in high enthalpy facilities, the present work adopts a positive heat flux from the wall to the flow.

4.2 Surface ccharacteristics at Mach 4.5 flow using the micrometre gauge and SEM

The micrometre measurements also recorded losses in thickness over successive runs. Four points were used across the disc at 0 mm, 8 mm, 16 mm, and 24 mm from the centre as shown in Figure 10. The measured thickness across the four locations is shown in Figure 11. The anvil and spindle of the micrometre (where the micrometre contacted the disc) were 6.5 mm in diameter. The disc deformation and the ablating surface meant that the contact from the micrometre was unlikely to accurately measure the local thickness at each point, but more likely measured a general thickness at the location in the vicinity of each point. Figure 11 shows the general trend resulting from the ablation of the disc. The thickness from the micrometre measurements in Figure 11 shows a general agreement with that obtained from the measuring arm [34]. The laser sheet visualisation confidently identifies that the material loss is not at a consistent rate along a radius of the disc, but that the trend of the material loss rate is relatively consistent over the duration of the experiments. The experiments show a significant spatial variation in thickness loss for the graphite test material over the disc radius though the spatial variation was still largely axisymmetric.

Figure 10.

Scanning locations on graphite surface. (a) Graphite specimen (top view). (b) Graphite specimen (side view not to scale).

Tools such as the Scanning Electron Microscope (SEM) was used for surface characteristics and the results from microscopy are presented in Figures 1218. Five positions in total were chosen at 4.5 mm increments, shown as A, B, C, D, and E in Figure 10. This resulted in position E being 7 mm from the outside edge of the disc (6 mm from the chamfer used to retain the disc). During experiments, the actual points scanned in the SEM were not exactly the same; each scan was a representative area in close proximity to the points described by A, B, C, D and E.

Figure 11.

Variation in thickness across disc along paths perpendicular to the surface at various radial positions using the micrometre gauge after 1, 2, 4, 8, 12 and 16 heated-with-flow runs.

Figure 12.

SEM images after the second heated with flow run in proximity of points A, C and E. (a) In the proximity of point A. (b) In the proximity of point C. (c) In the proximity of point E.

Figure 13.

SEM images after the fourth heated with flow run in proximity of points A, C and E. (a) In the proximity of point A. (b) In the proximity of point C. (c) In the proximity of point E.

Figure 14.

SEM images after the eighth heated with flow run in proximity of points A, B, C, D and E. (a) In the proximity of point A. (b) In the proximity of point B. (c) In the proximity of point C. (d) In the proximity of point D. (e) In the proximity of point E.

Figure 15.

SEM images after the twelfth heated with flow run in proximity of points A, B, C, D and E. (a) In the proximity of point A. (b) In the proximity of point B. (c) In the proximity of point C. (d) In the proximity of point D. (e) In the proximity of point E.

Figure 16.

SEM images after the sixteenth heated with flow run in proximity of points A, B, C, D and E. Back of disc. (a) In the proximity of point A. (b) In the proximity of point B. (c) In the proximity of point C. (d) In the proximity of point D. (e) In the proximity of point E.

Figure 17.

SEM images after the twelfth heated with flow run in proximity of points A, B, C, D and E (back). (a) In the proximity of point A. (b) In the proximity of point B. (c) In the proximity of point C. (d) In the proximity of point D. (e) In the proximity of point E.

Figure 18.

SEM images after sixteenth heated-with-flow run in proximity of points A, B, C, D and E (back). (a) In the proximity of point A. (b) In the proximity of point B. (c) In the proximity of point C. (d) In the proximity of point D. (e) In the proximity of point E.


5. Future work

The future work has optimised the aerothermodynamic efficiency for future experimental tests and can now generate stagnation point temperatures in excess of 3000 K. The two major improvements that are associated with the Next Generation Experimental Model (NGEM) include: (a) reducing the heat losses by conduction on heatshield sample by incorporating a thermal barrier in-between the test-sample and the backshell; and (b) incorporating six degrees of freedom in order to account for the variable angle of attacks for better manoeuvrability during reentry, descent and landing [36]. These modifications enable the NGEM to be smarter and more practically replicates real-flight vehicles as shown in Figure 19. The future experimental model has been fully developed for series of ablation tests in any reliable aerospace laboratory. Future experimental tests will encourage the use of spectroscopic measurements of ablation species and spatial microstructural studies using X-ray Microtomography. Infrared pyrometers and thermo-cameras are also needed to adequately monitor the surface temperature profiles across experimental samples. The next generation experimental model is not only expected to generate a stagnation point temperature of about 3200 K under the same experimental conditions, but also enables variable angle of attack for better manoeuvrability during reentry, descent and landing. These inclusions have never been attempted anywhere else and will enable the next generation models to be smarter and more practically replicate real-flight vehicles.

Figure 19.

Sectional view showing improved operational capabilities for the NGEM [36]. (a) Thermophysics and heat transfer. (b) Full view of the NGEM.

Some of the significant improvements in the NGEM for the purpose of improving the aerothermal capabilities of the plasma preheating technology to reliably replicate planetary reentry surface temperatures with high degree of confidence are presented in Table 1. This novel invention is able to accurately replicate the planetary reentry surface temperatures and any associated hypersonic flow characteristics within the boundary layer.

PartsPresent workFuture workScientific relevance for future work
Backshell thickness5.0 mm2.5 mmReduced mass and heat-sink effects
GeometryConventionalAdvancedBetter thermodynamic performance
Thermal barrier systemAbsentPresentReduced conduction losses
Holding ringMild steelTiCHigher thermal capability over steel
BackshellMild steelMild steelHousing and support
Ceramic shroud96% AluminaZirconiaHigher thermal capability over 96% Alumina
GTAW torchSeparatedAttachedThermal optimisation over detached features
StandFixedAdjustableFlexibility in elevations
6DOFAbsentPresentBetter aerodynamic manoeuvrability

Table 1.

Materials selection for components and parts for hypersonic/ablation experiments [36].

Figure 20 shows the temperature profile from FEA simulation using Ansys Workbench. The contact regions in the future work have been designed to minimise conduction losses at heatshield edges (shoulder regions) using a thermal barrier material (Zirconia). The improvement of surface temperature profile for future work can be seen in Figure 20. The details from spatial temperature profiles for the present work have been extensively published by the author [27].

Figure 20.

Spatial temperature profiles showing heatshield thermal spread [36].


6. Conclusions

The major area of interest is using plasma to preheat probe models that can be used for ablation studies in cold flow hypersonic wind tunnels. This novel plasma preheating technology is capable of generating the needed heat flux for the surface temperature characterisation without using any Arc-Jet or plasmatron facilities. This technique aims to produce a newer and better method of aerothermodynamic tests for investigating ablation samples of reentry probes in expansion tubes. The probe used in the present work was very similar to the European Standard Probe with a stagnation temperature of about 2500 K. The advantages that are associated with this newly innovated plasma preheating technique include: (1) light weight and portability of model; (2) surface temperature control; (3) the ability to replicate entries for different planetary missions due to its capability to perform well in most types of reentry gases like O2 N2, Air, CO2, He, Ne, Ar, H2, CH4, NH3, etc.; (4) ability to be applied to different types of heatshield materials like PICA, SIRCA, Avecoats, C/C Composites, graphite, etc.; (5) ability to be used for different axisymmetric payload geometries like Stardust, Orion, Hayabusa, SpaceX-dragon, etc.; (6) ability to perform well in both short and long duration wind tunnels including shock tunnels; and (7) highly economical and less than 10% of the operational costs of running the high enthalpy plasmatron and NASA Ames Arc-Jet facilities. This work presents reasonable contributions to reentry studies under conditions that replicate characteristics of hypersonic reentry flights. Future applications for the technique are expected to be found in hypersonic impulse facilities that can simulate the true flow energy under reentry conditions.



The author would like to acknowledge his totally self-funded independent PhD student research work including the 100% payment of all tuition fees in addition to the purchasing of plasma preheating test device, accessories and consumables that were used for experimental tests during his doctoral candidature at the University of Southern Queensland (USQ). Evidence of financial self-sponsorship and all academic financial statements throughout the full-time doctoral candidature from 2 March 2015 to 13 September 2019 are held at the Graduate Research School of USQ. The author acknowledges the tremendous financial supports that he received from family and friends during difficult times. The author also acknowledges the six-times refusal of student scholarship applications during his doctoral candidature despite high distinctions; but perseverance led to earning a PhD in Mechanical/Aerospace Engineering along with a Doctoral Research Excellence Award for achieving the highest possible result for a Higher Degree by Research thesis examination 2019 at the Institute for Advanced Engineering and Space Sciences, USQ. The author is grateful for the opportunities to continuously exchange professional ideas with his entire research team and other aerospace research professionals around the world. The novel plasma preheating technology for ablation experiments reported herein is not elsewhere classified and the intellectual property is fully protected by the intellectual property (IP) right of Australia with Patent Number 2019205004.


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Written By

Daniel Odion Iyinomen

Submitted: 08 August 2021 Reviewed: 25 August 2021 Published: 04 May 2022