[American Institute of Aeronautics and Astronautics AIAA/CIRA 13th International Space Planes and Hypersonics Systems and Technologies Conference - Capua, Italy ()] AIAA/CIRA 13th

American Institute of Aeronautics and Astronautics

1

PRORA-USV SHS: Ultra High Temperature Ceramic

Materials for Sharp Hot Structures

L. Scatteia*, A. Riccio†, G. Rufolo‡, F. De Filippis

§, A. Del Vecchio

**, G. Marino

††

CIRA – Italian Aerospace Research Centre –

Via Maiorise 81043 CAPUA (CE) – ITALY

The Italian Aerospace Research Centre (CIRA) is running the National Aerospace

Research Program (PRO.R.A.) funded by the Italian Ministry of Education and Research.

In this frame, the Space Program Office of CIRA is managing System and Technology

activities finalized to the development of Flying Test Beds (FTB) aimed at the in flight

experimentation to test new advanced technologies useful for the next generation of re-entry

vehicles.

This paper describes the work performed within the technology project named Sharp

Hot Structures (SHS), that was started four years ago in support to the system activities

related to the development of the orbital re-entry vehicle FTB-X whose first flight is

currently scheduled for the year 2010.

In order to provide new aerodynamic design criteria and improved manoeuvre capability

to the experimental platform, SHS project was focused on the assessment of the applicability

of Ultra High Temperature Ceramics (UHTC) to the fabrication of high performance and

slender shaped hot structures for reusable launch vehicles.

Fundamental technology advancements, progressively reached during the performed

activity, will be summarized and critically analyzed. In particular the results of the design

phase of two technology demonstrators (Nose_1 and Nose_2) of the nose cap of FTB-X will

be shown and compared. Furthermore the paper describes the manufacturing process of the

first concept of multi-material hot structure (Nose_1), that has been already built and is now

ready to be tested in the CIRA Arc-Jet facility, SCIROCCO.

Moreover, a report will be given concerning the on-ground plasma test of a preliminary

scaled demonstrator, dubbed Nose_0, that was fabricated to assess the manufacturing

technologies and tested to verify the thermal-oxidative stability of the interface between a

C/SiC frame and a ZrB2 coating, under consistent heat flux conditions.

Considering the specific typology of different materials investigated, up to date, an

extensive tests campaign at laboratory level has been performed and concluded in order to

create a complete materials data base.

The measured materials properties have been then used, together with the aero-thermal

loads associated with a reference re-entry mission, as input for the design phase.

Our major preliminary findings indicate that the structure is thermally fully compliant

with the environment requirements and shows local mechanical criticalities in specific areas

such as the materials interfaces and hot/cold joining parts.

* Researcher, Advanced materials and technology Lab., e-mail: [email protected]

† Researcher, Computational Mechanics Lab., e-mail: [email protected]

‡ Researcher, Aerothermodynamics Lab., e-mail: [email protected].

§ Researcher, Plasma Wind Tunnel Unit, e-mail: [email protected]

** Researcher, Plasma Wind Tunnel Unit, e-mail: [email protected]

†† Project Manager, Space Programs Office, e-mail: [email protected]

AIAA/CIRA 13th International Space Planes and Hypersonics Systems and Technologies AIAA 2005-3266

Copyright © 2005 by CIRA. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.


2

Nomenclature

x = spatial coordinate

t = time h = coefficient of proportionality between heat flux and enthalpy

H0 = total enthalpy

Cp = specific heat

CFDq& = heat flux calculated under the cold wall assumption

Twall = assumed wall temperature

Twall,FEM = FE-calculated wall temperature

I. Introduction

Thermal protection systems represent the key issue for the successful re-entry of a space vehicle1. Future

concepts for space launchers foresee sharp aerodynamic profiles as conventional aircrafts2. This kind of architecture

offers several advantages with respect to current blunt shapes: maneuverability improvement, decrease of

electromagnetic interferences and communication black-out and drag reduction during the ascent phase. As main

drawback, the aerodynamic heat fluxes increase dramatically over the vehicle profile, exceeding conventional values

of 650-800 kW/m2, and determining the non-applicability of state of art hot structures materials.

The Sharp Hot Structures Project (SHS) is focused on the development of modified diboride compounds as

potential candidate materials for the manufacturing of innovative high performance and slender shaped hot

structures. Zirconium and Hafnium diborides/Silicon carbide composites are under investigation: those compounds

are actually addressed as the sole materials that can be conveniently employed at temperatures above 2200K3.

Due to the ambitious target, SHS project activities are performed within a national research network, managed

by CIRA, and involving the University of Naples “Federico II”, Centro Sviluppo Materiali S.p.A. (CSM), the

University of Rome “La Sapienza”, the Institute of Science and Technology for Ceramics of the Italian National

Research Council (CNR-ISTEC), Fabbricazioni Nucleari (FN) and the University of Turin.

The main objective of this project is to provide technology products identified as critical parts of re-entry

vehicles such as nose cap and wing leading edges that will be first qualified on-ground, and then tested and validated

in flight conditions. The development of a technology demonstrator of the nose cap started four years ago funded by

PRO.R.A., while a feasibility study (Phase A) on the wing leading edges was launched last year with the support of

the Italian Space Agency (ASI). The design of these components meets the high level requirements of the re-entry

Flying Test Bed (FTB-X) and at the same time provides experimentation requirements to be met by the vehicle in

flight conditions.

II. Project structure and logic

The Sharp Hot Structure Project is articulated into the following phases: a) a basic research on selected UHTC

materials conducted in parallel with the related manufacturing processes assessment; b) hot structures thermo-

mechanical design, which drives the materials and process assessment; c) Hot structure scaled demonstrator

manufacturing; d) On-ground qualification test at CIRA Scirocco Plasma Wind Tunnel; e) In-flight validation test of

the component, in the re-entry mission of the USV Flying Test-Bed..

III. Selected Materials

The reference re-entry mission that has been considered as input for our studies, is characterized by very high

thermal loads that conventional CMCs such as C/C and C/SiC, although reliable and well tested, are not able to

sustain.

Aerospace research is moving towards ceramic systems based on hafnium, zirconium and titanium borides, in

account of their high configuration stability (ablation resistance) in the presence of high velocity dissociated air,

high thermal shock and thermal fatigue resistance4.

Those material are characterized by very high melting point and, if blended with a proper reinforcing phase, they

exhibit excellent oxidation resistance, thanks to the growth of a protective oxide layer on the surface of the

component in oxidizing atmosphere, that hinders further oxidation of the bulk.


3

In the SHS project, conventional C/SiC materials produced by polymer infiltration and pyrolisis have been

coupled to novel diboride materials in order to create a multi-material structure able to withstand the severe

condition associated with slender-shaped hot structures and non-conventional reentry mission profiles.

An extensive trade-off between different ZrB2-X composition blends was performed, in order to maximize

mechanical properties, oxidation resistance and emissivity, while minimizing the material’s density in order to save

weight in the final nose cap component. At the end of the material trade-off phase a ZrB2-SiC compound was

chosen.

ZrB2-SiC compound, when exposed to high temperature oxidizing environments, forms boro-silicate glass based

surface layer, which protects the bulk from further oxidation.

Massive diboride production by hot pressing sintering has been set up in order to realize a small massive

diboride conical tip that is intended to sustain the greatest thermal load in the final nose cap structure. Contextually,

the use of the diboride was further extended identifying the non-conventional Plasma Spray Deposition technique5 in

order to obtain thin protective ZrB2-SiC coating on the structural C/SiC long fiber composite frame.

In the following paragraph describe the architecture of the Nose_cap scaled demonstrator.

IV. Prototype Structural Concept

Figure 1 depicts a schematic of the component dubbed Nose cap 1. The nose is composed by: a) a bulk graphite

core; b) a truncated conical C/SiC frame manufactured by polymer infiltration and Pyrolisis process c) a ZrB2-SiC

coating applied on the C/SiC frame by plasma spray deposition technique; d) a ZrB2-SiC massive conical tip

produced by sintering technique. Each of the identified (material)/(manufacturing process) systems was subjected to

a complete characterization test campaign, in order to provide the thermo-mechanical design with the required

database of properties.

Figure 1. Schematic of the nose cap scaled demonstrator

Moreover, in order to test the adhesion between the C/SiC frame and the ZrB2 coating in operating conditions,

an intermediate step between the laboratory scale characterization and the on ground testing of the above described

nose cap 1 demonstrator was conducted, by producing a “Nose cap 0” demonstrator constituted of a graphite bulk, a

C/SiC frame, and a ZrB2 coating. This prototype 0 was tested in the Scirocco plasma wind tunnel. Test results are

presented in the last section of this paper.


4

V. Nose Cap 1 Prototype manufacturing

A. C/SiC Frame by Polymer Infiltration and Pyrolisis

Basically a preform of C fibers properly shaped is placed in a frame. This preform is the infiltrated with allil-

polycarbosilazane resin, the infiltrated form is placed in a oven and pyrolyzed. The same process is used for the

fabrication of the conical frame, the only difference being in the shape of the preform.

The infiltration step usually occurs at low pressure, allowing for the ceramic precursor to polymerize among the

fibers. The sample in then pyrolized in inert atmosphere. Several cycles can be necessary at temperatures as high as

1000°C. In this step low molecular weight oligomers and hydrocarbons are lost by the matrix. The final product can

have a high porosity percentage more than 30%, and other infiltration/pyrolysis steps can be necessary to increase its

density.

Matrix porosity makes carbon fibers scarcely protected from oxidation, which significantly lowers the

mechanical performances at high temperatures. This drawback is overcome by the presence in the SHS nose cap

structural concept, of the protective ZrB2 sprayed coating.

The chosen composite configuration is based on a bi-directional carbon mat, draped in a conical shape and then

processed by PIP.

B. ZrB2-SiC Coating by Plasma Spray Deposition

Due to the well known impossibility to directly spray materials such as SiC for the absence of a stable liquid

phase, the development of the new fabrication methodology included also a suitable ZrB2-SiC powder production

method. To this purpose a mixture of silicon carbide and zirconium diboride powders was selected as starting

material. The precursors had a flaked, angular shaped morphology and an average size of about 0.5 µm for SiC and

5 µm for ZrB2. A pre-consolidation treatment of the powders was performed to achieve the desired flowability and

to optimize the deposition efficiency. Spray drying method was selected for mixing, agglomerating and correctly

forming fine starting powders. Final results were composite powders with coarser ZrB2 positioned on the surface of

the dried particle, while the core consisted mainly of small SiC precursors.

The high-melting ZrB2 gives a contribution to protect the thermally unstable SiC during plasma-particle

interactions. Agglomerated powders were plasma sprayed in different pressure conditions (low pressure, LPPS, and

high pressure, HPPS) using a controlled atmosphere (CAPS) system operating in a close pressure-vessel (Figure 3).

Spraying operating parameters for the different pressure conditions were selected using statistical design of

experiments. Goal functions for the optimization of coating morphology were porosity, roughness and thickness, and

the variable process parameters were pressure in the spraying chamber, substrate to torch distance and plasma gas

composition.

Figure 2. a) plasma spray deposition facility; b) plasma spray torch

C. Massive Diboride Conical Tip Sintering

Isostatic Hot-Pressing has been chosen as the most appropriate technique. Slip casting has also been considered

as an alternative process but the results were not satisfactory.

To ensure that favourable properties in the sintered diboride are attained, control of densification and

microstructure is necessary, because strength and corrosion resistance, for instance, are adversely dependent on

b) a)


5

porosity in sintered bodies. Due to the high melting point and high vapour pressure of the constituents, the sintering

of ZrB2 powders needs very high temperatures. Relatively high densities are achieved only by pressure-assisted

sintering procedures at temperatures approaching or even higher than 2000 °C.

The addition of sintering aids is a strategy to overcome the intrinsic low sinterability of highly refractory

compounds. High density materials were obtained through liquid phase sintering at processing temperatures well

below those needed for the undoped ZrB2, and enhanced properties like hardness, toughness and strength were

successfully achieved as well. The introduction of sintering aids improves the final density and allows to lower

the densification temperature, increasing volume diffusion and retarding evaporation mechanisms.

The properties of the dense materials become then strictly dependent on the starting powders and processing

parameters as they determine micro-structural features such as grain size, volume and chemistry of the secondary

phases, etc.

The starting powder composition and the processing parameters were optimized and a billet was produced using

isostatic hot pressing. The billet was then electrical discharge-machined to obtain the 1th

prototype of the nose cap

conical tip, showed in figure 3.

Figure 3. Nose cap massive diboride conical tip


6

VI. Thermo-structural Analysis

On the basis of the review concept analysis of Nose_1, especially referred to the manufacturing criticalities, the

design of the second technology demonstrator (Nose_2) has been carried out.

In this paragraph, the thermo-structural design of the Nose_2 is described and critically analyzed in comparison

with previous results obtained on Nose_1 design.

The main assumptions introduced in the Nose_2 activities (Nose_2 final design and Nose_2 optimization) are the

following:

1. C/SiC is considered orthotropic and more precisely transversely isotropic since the material properties, such

as thermal conductivity and CTE, are characterized by different values in fabric plane direction and through

the thickness direction;

2. A difference between ZrB2-SiC coating and massive ZrB2-SiC with different thermal and mechanical

properties has been considered;

3. Temperature dependent properties for massive ZrB2-SiC and C/SiC (thermal conductivity, specific heat,

CTE, emissivity) have been implemented;

4. The verification of the integrity of the structure is carried out when the thermal flux is maximum along the

vehicle trajectory.

The design configuration has been defined by adopting three kind of complementary approaches:

1. BASE MODEL

2. LOCAL MODEL

3. GLOBAL MODEL

The base model is a simplified two-dimensional model able to take into account complex boundary conditions

such as temperature and displacement distributions; it is useful to understand qualitatively the thermo-structural

behaviour of two or three material in series with respect to variability of the material properties and of the thickness.

The local model is an axial symmetric two dimensional piece of the global model; it is used to determine the

quantitative changes in thermo-structural behaviour of the model with respect to thickness variations and then to

choose the optimum values of the thickness of the materials examined. Local model’s boundary conditions are

achieved by interpolating the global model temperature and displacement distributions in two fixed geometries

(thickness, curvature radius, etc). Then at least two global model analyses with different architectural configurations

are necessary.

The global model is an axial symmetric model useful to provide boundary conditions for the local model and it is

adopted to verify the optimum configuration obtained by the local model.

Starting from the Nose_1 base configuration, its residual critical areas (where no full satisfaction of the adopted

failure criteria was registered) have been analysed and alternative architectural configurations have been taken into

account in order to remove those criticalities. As a result the Final Design activity of the Nose_2 with the relevant

technical drawings has been carried out.

In figure 4 a schematic representation of the Nose_2 configuration showing the materials layout is presented.


7

Figure 4. SHS Nose_2 FEM model.

The FEM model has been adopted to perform non stationary thermal analyses together with geometrically non-linear

mechanical analyses. The time instant chosen for failure verifications is after 72 seconds along the trajectory since

the maximum temperature of some “hot key points” is reached at the same time instant.

The failure criteria, adopted in Nose_1 and Nose_2 final design, is “the Maximum Normal Stress Criterion”, also

known as the Normal Stress, Coulomb, or Rankine Criterion, which is often used to predict the failure of brittle

materials. Failure occurs when the maximum (normal) principal stress reaches either the uniaxial strength or the

uniaxial compression strength,

TC S

S

S

S

S <

<−

3

2

1

(1)

where S1, S2 e S3 are the point-wise principal stresses induced by thermo-mechanical loads, ST is the tensile

strength and SC is the compressive strength. In a 2-D stress state, graphically, the maximum stress criterion requires

that the two principal stresses lie within the green zone (Fig. 5):

Figure 5: Safety zone for a 2D-stress, according to the Maximum Stress Criterion.

The Maximum Normal Stress Criterion has been preferred to the Von Mises Stress criterion because it makes a

distinction between Tensile and Compressive behaviours of the material. Moreover, the Von Mises Stress criterion

considers an equivalent stress which is an average of the stresses and does not take account the difference between

Massive

ZrB2/SiC

Coating

ZrB2/SiC

C

C-SiC

Graphite

Cold-Interface


8

compressive strength and tensile strength. This distinction is fundamental for brittle materials like ceramics.

Moreover in these analyses, the strength has been considered as a function of the temperature.

In figure 6, the comparison between the Nose_ 1 and Nose_2 configurations in terms of percentage of failed areas

for each material are presented. From this figure it is clear that the modifications in configuration adopted for the

Nose_2 design have allowed to eliminate the tensile failure in ZrB2-SiC coating and in C-SiC and to reduce the

tensile failure in Graphite. Nevertheless, the compressive failure in ZrB2-SiC coating was not eliminated even if the

thickness of coating was strongly reduced.

Figure 6: Nose_1 vs Nose_2 prevision of the failed area percentage

In order to show the distributions of the principal stresses and the position of the critical areas in the Nose_2

configuration some pictures are shown hereafter.

Figure 7: Nose_2, ZrB2-Sic coating - S1 principal stress and temperature distribution

NOSE_2

Damaged Area Percentage Material

0%

0%

0%

100%

Compressive

696 K

0.3% Graphite

1987 K

0% Massive

ZrB2-SiC

1493 K

0% C/SiC

1519 K

0%

ZrB2-SiC

coating

Max

Temperature

Tensile

Damaged Area Percentage Material

0%

0%

0%

100%

Compressive

451 K 1.6% Graphite

1955 K 0% Massive

ZrB2-SiC

944 K 0.1% C/SiC

1403 K 1.5%

ZrB2-SiC

coating

Max

Temperature

Tensile

NOSE_1

Principal stress S1 for ZrB2SiC coating -

Temperature for ZrB2SiC coating -


9

Figure 8: Nose_2, graphite - S1 principal stress and temperature distribution

Figure 9: Nose_2, massive ZrB2-SiC - S1 principal stress and temperature distribution

Figure 10: NOSE_2, C/SiC - S1 principal stress and temperature distribution

Principal stress S1 for massive ZrB2SiC -

Temperature for massive ZrB2SiC

Damaged area for Graphite

Temperature for graphite

Principal stress S1 for C/SiC Temperature for C/SiC


10

In figure 7 the principal stresses and the temperature found for the ZrB2-SiC Coating are shown. The damaged

area caused by tensile load in the coating is located at the interface between nose and the Cold Structure. The

damaged area caused by compressive loading is distributed throughout the component for this reason it has not been

shown. In figure 8 the S1 principal stress and the temperature found for the graphite are shown and a quite extended

damaged area is visible at the interface between graphite and C-SiC near the axis of symmetry of the nose. From

figure 9, no damaged area can be noticed in the massive ZrB2-SiC cone; the maximum values of the principal

stresses lie below the strength limit.

Finally, in figure 10 the S1 principal stress and the temperature found for the C-SiC are shown. The maximum

temperature and the maximum stress are located at the interface with the massive cone, in this region (like in the rest

of the component) no damage area is visible.

The bonding between C/SiC and the massive cone has also been checked by evaluating the interface stresses

accounting for the stress induced by the mass body forces of the massive cone, and no criticality from this point of

view has been found.

In conclusion, with respect of nose_1, the nose_2 design activity has shown the feasibility of improvements in

thermo-mechanical behaviour of the technology demonstrator by changing some design parameters .


11

VII. Computational Fluid-dynamic analysis

As already mentioned in the abstract, an intermediate scaled demonstrator, dubbed nose_0, was fabricated in

order to assess the manufacturing capabilities of such shaped components. In addition to that, nose_0 was also

intended to be tested in CIRA PWT in order to study and improve the adhesion of the plasma coating over the

carbon/silicon carbide substrate.

In this light a dedicated test was then designed and set up. In the following, the computational fluid dynamic

(CFD) analysis performed prior to the wind tunnel test is presented.

The on-ground facilities for high-enthalpy tests do not always allow the simultaneous experimental reproduction

of all the thermo-fluid-dynamics conditions that characterize above all the low-earth orbit part of a typical space

vehicle re-entry path. Even in the case of the CIRA Plasma Wind Tunnel, SCIROCCO, the largest in the world, it

may be difficult to contemporary reproduce both stagnation point heat flux and pressure. Moreover, due to the

strong dissociation of the airflow occurring trough the arc-heater, the environment reproduced within the arc-jet

facilities can be quite different from the real one. Even if the energetic level of the flow within the arc jet is the same

of the flight one, in the former case a large amount of energy is frozen within the fluid as formation enthalpy of

dissociated atomic species. For this reason, if, for instance, the material has a partially catalytic behaviour it is

essential to be able to properly characterize the difference between the flight and ground environment in order to

better understand which mechanism of heat release to the wall surface prevails: conductive or chemical. Moreover,

when the article to be tested is scaled and/or of slightly different shape with respect to the real one, the correct

reproduction of heat flux and pressure at the stagnation point do not assure in general that we are simulating the

same environment downstream of the stagnation point. In the same way, the small radius of curvature of sharp

structures together with the low Reynolds flow obtainable with an arc-jet facility may give rise to undesired

rarefaction effects. For the above reason an extensive use of CFD is required both for the extrapolation from simulated flight condition to suitable operating condition of the plasma wind tunnel and for the extrapolation of the

test results to flight condition.

Within the framework of the PRORA-USV Program numerical activities have been carried out and others are

currently in progress in order to characterize the aero-thermal environment that the vehicle will experience during

the reference re-entry mission. CFD simulations of the flow field surrounding the forebody part of the vehicle have

been performed in correspondence of the maximum heat flux trajectory point that in the case of the SRT mission

take place at an altitude of about 20Km at a Mach number of about 7.5. At this low altitude and relatively low speed

Figure 11. Flow Field Around the nose_0. Iso-Contour of Mach

Number.


12

the high heat flux value over the sharp nose (blunted cone with 1cm radius of curvature) is mainly due to high

pressure effects rather than to high enthalpy ones. Unit Reynolds number for the above trajectory point is about 20

millions so that the boundary layer will be turbulent for the most part of the vehicle surface and transition will

probably occur immediately downstream the sphere-cone junction. For this reason CFD simulation have been

performed with fully turbulent assumptions in order to provide realistic and conservative heat flux distribution. In

figure 11 the flow field surrounding the first part of the vehicle is shown in terms of Mach Number iso-contours. It

is evident how, due to the low radius of curvature of the nose, the shock is very close to the body surface. Sharp

nosed vehicle are commonly characterized by flying at low angle of attack along the trajectory being this a crucial

factor to gain aerodynamic efficiency and than cross-range capabilities. In figure 12 the heat flux profile derived

from a full three dimensional computation (4deg of angle of attack) is compared with that obtained with a 2D

axysimmetric simulation (0deg angle of attack). It is clearly evident that, apart from the spread (that is emphasized

by the turbulent state) due to the angle of attack, the axysimmetric distribution is a good approximation of the real

situation.

In particular, over the sphere the distributions are identical. This result allowed to consider for the preliminary

nose concept design a sphere-cone geometry subject to axisymmetric heat loads.

The analysis of the flight environment allow to identify the most critical condition that the material has to

withstand and that has to be reproduced in wind tunnel testing.

Therefore, starting from the value of heat flux to be realized at the stagnation point, theoretical-numerical

activities have been conducted in order to aid the set up of the wind tunnel operating conditions. In order to properly

execute the test, it is necessary to know the value of heat flux and pressure to be realized on a calibration

hemispherical (10cm dia.) probe made of copper and cooled at a constant temperature of about 50°C. When the

desired conditions are obtained over the probe this is extracted and the model is injected into the plasma flow and

the test take place for the desired time. Therefore, aim of the numerical activities in this phase is the translation of

the heat flux requirements over the model to be tested into operating conditions for the calibration probe. This

process is influenced by several factors that cause differences between the probe and the test article: 1) different

shape; 2) different positioning within the test chamber. Due to the effects of the nozzle expansion the conditions

along the axial direction are not uniform; 3) different wall temperature condition. Wall temperature of the cooled

probe is constant in time and uniform in space, while that of the test article comes out from the balance of heat

convected from the fluid towards the surface, heat radiated from the surface towards the fluid and heat conducted

into the solid. By neglecting the latter contribution a radiative equilibrium assumption is made that allow to

decouple the external flow field simulation from the thermal computation inside the solid.; 4) different catalytic

behaviour of the copper probe (fully catalytic) and of the test article (finite rate catalysis).

In figure 13 and figure 14 the results of the CFD simulations of the SCIROCCO nozzle flow and of the flow

surrounding the test article are respectively shown. The above mentioned computations have been performed with

Figure 12. Heat Flux and temperature Profiles with Radiative Equilibrium Hypothesis. Turbulent.


13

the CIRA code H2NS that is capable of solving the Navier-Stokes equation in thermo-chemical non-equilibrium

conditions. Results of this analysis in terms of heat flux distribution over the scaled nose geometry have constituted

an input for the thermo-structural analysis that will be illustrated in a following paragraph.

VIII. On Ground Arc Jet testing

The nose_0 demonstrator was tested into the arc-jet Plasma Wind Tunnel that is available in CIRA, named

Scirocco6,7,8

. This plant represents one of the most powerful hypersonic test facility in the world (70 MW) and

produces a very uniform and large test jet (up to 2 m diameter). The process air is thermally energised into the

segmented constricted Arc Heater reaching temperature values between 2000 and 10000 K. This energy is

transformed in kinetic by the air passage through a convergent-divergent Conical Nozzle and an hypersonic test jet

is generated with velocity ranging between 2000 to 6000 m/s and Mach number between 6 to 12 depending on the

exit nozzle size.

Figure 13. CFD Simulation of SCIROCCO Nozzle Flow.

.

Figure 14. Nose_0 Test-Article. Iso-Contours Mach Numbers.


14

The test requirement is the achieving on the test article ( “nose_0”: stagnation radius of 6.5 mm measured on the real

test article) of a stagnation pressure of about 10 mbar and a stagnation heat flux of about 900 kW/m2. These data

have been transformed in the equivalent ones to be obtained on the calibration probe that is used in Scirocco to

verify that the requested flow conditions are reached. This probe is a cooled copper semi-sphere with a radius of 50

mm, considered fully catalytic, able to measure the stagnation pressure and the thermal load by means of a Gardon

gauge heat flux sensor. With the CFD support the flow requested conditions, specifically on the probe, have been

calculated: 1) Stagnation Heat Flux of 300 kW/m2; 2) Stagnation Pressure of 10 mbar.

These flow conditions in Scirocco wind tunnel implies a plasma total enthalpy of about 5 MJ/Kg, i.e. lower than 10

MJ/kg. This value is the minimum that the Arc Heater is able to produce, otherwise the electrical discharge cannot

be sustained. The test total enthalpy value of 5 MJ/kg will be obtained by injection of the process air in part into the

arc heater and in part downstream in a mixing chamber between the arc-heater and the nozzle inlet9,10

.

In Figure 15 pictures of the test model before the test mounted on the TA support CPA2 and during the test are

shown.

Figure 15. SHS Nose_0 before and during the test.

The test campaign has been successfully conducted. The test article has been tested three times consecutively by

increasing the plasma exposure of the model from 20 seconds to 50 seconds. The measured experimental main

parameters and quantities are reported in Table 1.

Parameter Numerical Values

Arc heater current (A) 1700±100

Arc heater voltage (V) 8000±150

Total mass flow rate (kg/s) 1.22±0,02

Arc Heater total enthalpy (MJ/kg) 6,4±0,5

Arc heater total pressure (bar) 4,3±0,1

Probe stagnation pressure (mbar) 12,3±1,1

Probe stagnation heat flux (kW/m2) 360±90

Table 1. Arc-Jet Experimental Conditions

In Figure 16 the temperature map from the IR thermography is shown in the case of the longest run at the end of the

plasma exposure. In the same figure it is also reported the temperature profile of the hottest point visible from the

side view by means of the IR thermographic investigations. This point do not corresponds to the stagnation point

but, approximately, to the conjunction point between the spherical stagnation part and the cone shape.


15

Figure 16. Scirocco PWT Performance Map.

IX. Conclusions

Thermo–structural design of two different concepts of PRO.R.A. FTB-X nose cap technology demonstrators has

been completed. The first prototype has been fabricated and its on ground qualification test will be performed in few

months. Taking advantages from the fabrication of this first prototype, and from the on-ground test campaign

conducted on an intermediate scaled demonstrator (nose_0), an improved design was finally executed over the

second demonstrator (nose_2).

All the proposed and investigated concepts are based on the application of an innovative typology of thermal

protection system (TPS) and on the extensive utilization of a manufacturing technology never applied before to the

fabrication of hot structure. The second prototype of the PRO.R.A. FTB-X nose cap technology demonstrator is

currently under production and will be on-ground tested into the arc-jet Plasma Wind Tunnel December 2005.

Acknowledgments

The authors wish to acknowledge all the research team members belonging to the organizations that are currently

co-operating with CIRA on this technology project.

References

1. Beherens, B., “Technologies for Thermal Protection Systems Applied to Reusable Launchers” in Proceedings of the 54th

International Astronautical Congress, Bremen, 2003. 2. McKenzie P., “Lockheed Martin Orbital Spaceplane Program” in Proceedings of the 54th International Astronautical

Congress, Bremen, 2003.


16

3. Monteverde F., Bellosi, A., “Advanced Diboride Ceramics”, Scripta Materalia, 46, 223-228, (2002).

4. Levine S. R., Opila E. J., Halbig M. C., Kiser J. D., Singh M., Salem J. A., “Evaluation of Ultra-High Temperature Ceramics

for Aeropropulsion Use”, Journal of the European Ceramic Society, 22, 2757-2767, (2002). 5. Valente T., Bartuli G., Visconti G., Tului M., “Plasma Spray Deposition” in Thermal Spray Surface Engineering via Applied

Research, edited by Berndt C., ASM International Material Park, OH, 2000, pp. 837-841. 6. Caristia, S., De Filippis, F., Del Vecchio, A., Purpura, C., “SCIROCCO final tests measured data: comparison between

theory and experiments,” in Proceedings of 4th European Symposium Aerothermodynamics for Space Applications, Capua,

2001. 7. Russo G. and Marino G., “The USV Program & UHTC Development” in Proceedings of 4th European Workshop on

Thermal Protection Systems for Space Vehicles, Palermo, 2002, pp.157-163. 8. Caristia, S., De Filippis, F., Del Vecchio, A., Graps, E., “SCIROCCO PWT Facility for High Temperature Material

Assembly Testing,”, in Proceedings of 54th International Astronautical Congress, Bremen, 2003. 9. De Filippis, F., Del Vecchio, A., Caristia, S., “SCIROCCO Plasma Wind Tunnel: Low Enthalpy by use of cold air

transverse injection”, in Proceeding of 4th European Workshop on Thermal Protection Systems for Space Vehicles,

Palermo, 2002. 10.

Del Vecchio, A., De Filippis, F., “SCIROCCO Plasma Wind Tunnel: low enthalpy by use of cold air transverse injection,” in

Proceedings of 12th AIAA International Space Planes and Hypersonic Systems and Technologies, Norfolk, 2003.

Documents

[American Institute of Aeronautics and Astronautics AIAA/CIRA 13th International Space Planes and Hypersonics Systems and Technologies Conference - Capua, Italy ()] AIAA/CIRA 13th