7 FAST2
INDUCER
ROTORDYNAMIC
TESTS
In the following section a detailed description of the experiments carried out to validate the facility in the CPRTF (Cavitating Pump Rotordynamic Test Facility) configuration will be presented.
In particular, the experimental tests were performed on the FAST2 inducer to evaluate the pump performance in noncavitating and cavitating conditions, varying the shaft eccentricity through the dedicated eccentricity mechanism and the engines speed ratio. The results have demonstrated that the pump performance in noncavitating and cavitating conditions are independent from the eccentricity and speed ratio variations and an increased tip clearance has a detrimental effect on the pump performance.
7.1 Introduction
As previously mentioned, the CPRTF is an upgraded test facility configuration: it allows to impart a whirl motion to the pump shaft though an eccentricity mechanism and then to measure the rotordynamic forces through an equipped dynamometer mounted downstream the pump on the main shaft. The characterization of the main rotordynamic instabilities is intended to be performed in the next future. The experimental campaign, described in this section, was performed to validate the eccentricity system and the speed ratio regulation of the two engines.
The experimental tests were carried out on FAST2 inducer in noncavitating and cavitating conditions. In particular, the following parameters were varied simultaneously for both conditions:
− Eccentricity: it can be regulated by a mechanism for adjusting and rotating the eccentricity of the impeller axis in the range 0-2mm (Figure 7.2)
− Speed ratio(ω/Ω): it can be regulated by the electronic controlling system connected to the secondary engine. Once the main engine speed value is set by the dedicated engine program, the speed ratio is fixed by directly imposing the secondary engine speed value on the secondary engine electronic panel.
girante
telaio(fisso)
Albero motore cilindro interno
cilindro esterno
Figure 7.1 – Schematic of the eccentricity mechanism
As in the previous experiments, the following parameters were measured:
− The pump inlet pressure, by means of the Druck absolute transducer installed upstream of the Plexiglas inlet duct;
− The pressure rise given by the pump, by means of the Kulite differential transducer, installed between the Plexiglas inlet duct and the discharge pipe immediately downstream the inducer inside the test section;
− The mass flow rate, by means of the two electromagnetic flow meters; − The torque given by the main motor, obtained by the electronic drive itself; − The water temperature, by means of the probe installed in the main tank.
7.2 Experimental tests in noncavitating conditions
The first set of experimental tests have been carried out in order to analyze the performance of the FAST2 inducer in noncavitating conditions. The eccentricity was varied between 0-0.749 mm by the dedicated mechanism and the speed ratio (ω/Ω) was varied between -1.2-1.2 according to the Table 7.1. As previously mentioned in the past sections, the maximum nominal eccentricity value allowable by the eccentricity mechanism is 2 mm; in these experimental campaign the maximum eccentricity value was fixed 0.749 mm due to the diameter of the Plexiglas conduct (85 mm), which was available at the time. In fact, it was necessary to prevent any contact between the inducer and the conduct during the whirl motion. The maximum tip clearance allowed resulted 1.4 mm. In order to perform the experimental tests at the maximum eccentricity value (2 mm) a Plexiglas conduct of 87 mm has to be mounted.
Table 7.1 – Experimental tests matrix of the FAST2 inducer according to different speed ratio and eccentricity values in noncavitating conditions.
Figure 7.2 presents a comparison between the FAST2 inducer noncavitating performance for two different tip clearances (0.4 mm, same tip clearance value of the experimental tests presented in the previous section, and 1.4 mm, according to the dimension of the new Plexiglas conduct) between the inducer and the Plexiglas conduct. The speed ratio and the efficiency were maintained zero. As expected, an increased tip clearance has a detrimental effect on the pump performance, which decreases of about the 20%, due to a rise of the fluid dynamic losses.
Figure 7.2 – Comparison between the noncavitating performance of the FAST2 inducer at Ω =2000 rpm, two different tip clearances and ω/Ω =0
The next Figure shows a comparison of the inducer performance obtained at a fixed speed ratio and a variable eccentricity, according to Table 7.1. The inducer performance result completely independent from the eccentricity values at constant speed ratio, showing a total matching apart from the zones at low flow coefficient where some flow instabilities can be detected.
Figure 7.3 – Comparison between the noncavitating performance of the FAST2 inducer at a variable eccentricity and constant speed ratio ω/Ω
Another set of experimental tests were carried out at a fixed eccentricity and a variable speed ratio to verify that the inducer performance were still invariant. The results are proposed in Figure 7.4. Also in this case, at low flow coefficient the inducer performance appear slightly different due to the flow instabilities. The results demonstrate that, the inducer performance in noncavitating conditions are
Figure 7.4 – Comparison between the noncavitating performance of the FAST2 inducer at a constant eccentricity and variable speed ratio ω/Ω
7.3 Experimental tests in cavitating conditions
The second set of experimental tests have been carried out in order to analyze the performance of the FAST2 inducer in cavitating conditions. The eccentricity was varied between 0-0.501 mm by the dedicated mechanism and the speed ratio (ω/Ω) was varied between -0.5-0.5 according to the next Table. It was necessary to fix the maximum speed ratio and eccentricity values to 0.5 and 0.501, respectively, in order to prevent any contact between the inducer and the conduct during the whirl motion because cavitation phenomenon can trigger rotordynamic instabilities on the main shaft. The main engine rotational speed was set 3000 rpm in order to test the inducer at higher flow coefficients.
Figure 7.5 presents a comparison between the FAST2 inducer cavitating performance at several flow coefficients for two different tip clearances between the inducer and the Plexiglas conduct, 0.4 mm and 1.4 mm, as in the previous noncavitating tests. The speed ratio and the eccentricity were maintained zero. As expected, an increased tip clearance has a detrimental effect on the inducer performance in cavitating conditions: it decreases of about the 35%, due to a rise of the fluid dynamic losses. The cavitation breakdown moves towards higher cavitation number, causing the performance breakdown at higher inlet pressures.
Table 7.2 – Experimental tests matrix of the FAST2 inducer in respect to different speed ratio and eccentricity values in cavitating conditions.
Figure 7.5 – Comparison between the cavitating performance of the FAST2 inducer at several tip clearances and flow coefficients for ω/Ω=0 and e=0
Figure 7.6 shows a comparison of the inducer performance in cavitating conditions obtained at several flow coefficients, at a fixed eccentricity and a variable speed ratio, according to the above table. As in noncavitating conditions, the inducer performance in cavitating conditions result completely independent from the eccentricity values if the speed ratio is maintained at a constant value, showing a complete matching at the various flow coefficients.
An other set of experimental tests were carried out at a fixed speed ratio and a variable eccentricity to verify that the inducer performance were still invariant. The results are proposed in Figure 7.7. The results demonstrate that the inducer performance in cavitating conditions are independent from the eccentricity value, ε, as well as from the engine speed ratio, ω/Ω.
Figure 7.6 – Comparison between the cavitating performance of the FAST2 inducer inducer at a constant eccentricity and variable speed ratio ω/Ω
Figure 7.7 – Comparison between the cavitating performance of the FAST2 inducer at a variable eccentricity and constant speed ratio ω/Ω
7.4 Whirling eccentricity instabilities
Two whirling eccentricity experiments were carried out on the FAST2 inducer at 3000 rpm and nominal flow rate (φ φ =/ ref 0.7). The eccentricity of the rotating whirl motion was set equal to 0.244
Figure 7.8 and Figure 7.9 show the waterfall plots of the pressure fluctuation spectra measured at the inducer inlet section in the two experimental conditions. The most interesting results of these experiments are:
• whirl frequency ω was clearly detected in both cases, together with the inducer rotating speed Ω. Optical visualization of cavitation on the inducer showed that, for the case ω / Ω = 0.02 (corresponding to a whirl frequency of 1 Hz), the whirl motion acted as an exciting factor for the development of a violent surge-mode instability having the same frequency. On the other hand, this surge-mode instability is not observed for ω / Ω = 0.2 because the exciting whirl frequency (10 Hz) is too distant from surge frequency and does not couple with the flow oscillations.
• at ω / Ω = 0.2 a new instability was detected, whose frequency is denoted by f6. Phase
and coherence analysis of the cross-correlation of pressure signals at the inducer inlet section showed that it is an axial instability, probably a subsynchronous “cavitation surge” with a frequency of about 0.8 Ω.
• at ω / Ω = 0.02 the high-order cavitation surge instability f2 is still observed at a
frequency 4.4Ω., thus indicating that the relatively slow whirl motion imposed to the inducer has little influence on the development of this kind of high-frequency instability.
Figure 7.8 – Waterfall plot of the power spectrum of the inlet pressure fluctuations in the FAST2 inducer under forced vibration conditions at φ φ =/ ref 0.7, Ω = 3000 rpm, ω / Ω = 0.02 and room water
Figure 7.9 – Waterfall plot of the power spectrum of the inlet pressure fluctuations in the FAST2 inducer under forced vibration conditions at /φ φ =ref 0.7, Ω = 3000 rpm, ω / Ω = 0.2 and room water temperature.
8 NACA0015
HYDROFOIL
EXPERIMENTS
In the following section a detailed description of the TCT (Thermal Cavitating Tunnel), an upgraded configuration of the CPTF (Cavitating Pump Test Facility), will be presented. The Thermal Cavitating Tunnel is an hydrodynamic tunnel designed to investigate the fluid-dynamic over test bodies. The experimental campaign performed to characterize the fluid-dynamic behaviour of a NACA0015 hydrofoil in noncavitating and cavitating conditions will be described. Furthermore a detailed analysis of the cavity length at different flow conditions and of the instabilities triggered by the cavitation phenomenon will be presented. The Thermal Cavitating Tunnel was designed and procured under the sponsorship of ASI, in order to validate a numerical code developed by CIRA.
8.1 Introduction
The first step for understanding cavitation instabilities and thermodynamic scaling effects is typically represented by experimentation on test bodies in hydrodynamic tunnels. A rough initial approximation of the cavitating behaviour of a rotating machine can be related, in fact, to that of a static cascade of hydrofoils. Cavitation instabilities on hydrofoils are generated by fluctuations of the cavity length caused by the inherent unsteady nature of the flow and the interaction with the pertinent boundary conditions. Franc (2001) distinguishes two classes of instabilities: system instabilities, in which the unsteadiness comes from the interaction between the cavitating flow and the rest of the system (inlet and outlet lines, tanks, valves), and intrinsic instabilities, whose features, such as frequency content, are independent from the rest of the system. The most known example of system instability is represented by cavitation surge, sometimes observed in supercavitating hydrofoils (Wade & Acosta, 1966) as a result of the extreme sensitivity of long cavities to external pressure fluctuations generated by other components of the circuit. When a partial cavity is formed on a two-dimensional
hydrofoil, a short sheet cavity appears near the leading edge of the hydrofoil and grows to its limitation (Figure 8.1). Then, vortex cavities accompanied by small cavity bubbles are shed from the termination of the fully developed sheet cavity. These bubble filled vortices are responsible of the cloud cavitation which produces harmful effects such as strong vibration of fluid machineries, radiated noise and erosion damage on solid surfaces of the machineries. The cloud cavitation phenomenon was originally found by Knapp [1955] and has been investigated by many researchers Acosta et al.[1958], Kubota et al [1989], Kato et al [1998], Brennen [1995], Franc [2001].
Figure 8.1 – Cavitation number on the suction side of a NACA 16-012 hydrofoil as function of incidence angles α in different flow conditions (Franc).
Kubota et al. (1989), using laser Doppler anemometry with a conditional sampling technique, showed that the shed cloud consists of a large-scale vortex containing a cluster of many small bubbles. The intrinsic nature of this form of instability has been largely elucidated: experiments carried out in various facilities of different characteristics and hydraulic impedances, or in adjustable configurations of the same facility (Tsujimoto, Watanabe & Horiguchi, 1998), lead to very similar Strouhal numbers for cloud oscillations (Figure 8.2).
Figure 8.2 – Strouhal number as function of the cavitation number on a NACA0015 hydrofoil (Kato, 1998).
Yamaguchi, Tagaya & Tanimura, 1997; Sakoda, Yakushiji, Maeda & Yamaguchi, 2001). The development of the re-entrant jet at the back of the cavity and the subsequent oscillations of the cavity length is the most common type of intrinsic instability. The principle of the re-entrant jet instability has been described by many researchers (Furness and Hutton 1975, De Lange 1996, Boehm et al. 1997, Kawanami et al. 1997). The closure of the cavity is the region where the external flow re-attaches to the wall. The flow which originally moves along the cavity has locally the structure of a jet which hits the wall. It separates in two parts. One is the re-entrant jet which moves upstream towards the cavity detachment. The other one makes the flow re-attach to the wall (Figure 8.3). According to Franc, a succession exists between periods of development of the re-entrant jet, which tends to fill the cavity, and periods of emptying of the two phase mixture. The Strouhal number of this instability was proved to be around 0.3 (Kawanami et al, 1998).
Figure 8.3 – Development of the cavity on the suction side of the hydrofoil and of the re-entrant jet (left). Vorticity component at different instants at s=1.2 and a=6.2° (right) (Franc).
Figure 8.4 – Development of the cavity on the suction side of a NACA0012 hydrofoil as function of the time at two angles of attack and cavitation numbers (de Lange, 1996).
The next Figures shows several CFD simulations performed by Song or a NACA 0015 hydrofoil at 8° angle of attack at different flow conditions (Re= 105). The pressure field, vorticity field and the lift coefficient as function of the time are presented for each cavitation number. In the Figures, where the cavitation is present, the related Strouhal number is introduced. It is defined as:
( 8.1 )
where c is the chord of the hydrofoil, f the frequency of the oscillations triggered by the cavitation phenomenon and V is the flow velocity. The cavitation region develops when the cavitation number decreases till it fully covers all the hydrofoil surface (“supercavitation”).
• σ=2.5 (noncavitating flow)
Figure 8.5 – NACA 0015 pressure profile (a) and vorticity (b) for a noncavitating flow (σ=2.5) and α=8° (left); the lift coefficient as function of the time (right).
• σ=2 (“bubble cavitation”), ST= 0.1 for the principal frequencies ST= 0.3 for the secondary
frequencies
Figure 8.6 – NACA 0015 pressure profile (a, b, c) and vorticity (d, e, f) for a cavitating flow (σ=2) and c f
St V
⋅ =
• σ=1.5 (“partial cloud cavitation”), ST= 1.3 for the principal frequencies
Figure 8.7 – NACA 0015 pressure profile (a, b, c) and vorticity (d, e, f) for a cavitating flow (σ=1.5) and
α=8° (left); the lift coefficient as function of the time (right).
• σ=1 (“completed cavitation”), ST= 0.5 for the principal frequencies
Figure 8.8 – NACA 0015 pressure profile (a, b, c) and vorticity (d, e, f) for a cavitating flow (σ=1) and
α=8° (left); the lift coefficient as function of the time (right).
• σ=0.05 (“supercavitation”)
Figure 8.9 – NACA 0015 pressure profile (a, b, c) and vorticity (d, e, f) for the supercavitating condition
Thermal cavitation effects on hydrofoils have been extensively investigated in the past. Kato et al. (1996) observed a temperature depression along the cavity, which becomes more significant when freestream temperature increases. At different freestream temperatures, cavities with comparable lengths were found to have comparable thickness, but could be observed at different cavitation numbers (bigger for higher temperatures). The conclusion drawn by Kato and his collaborators was that the reference length scale of thermodynamic cavitation effect must be the separated layer thickness at the leading edge of the cavity. Tani & Nagashima (2002) compared the cavitation behaviour of a NACA 0015 hydrofoil in water and in cryogenic fluids, showing the important role of the cavity Mach number, different in the two cases because of the different sound speeds. Yet, a number of aspects of unsteady flow phenomena in cavitating turbopumps and hydrofoils, including their connection with thermal cavitation effects, are still partially understood and imperfectly predicted by theoretical means alone. Technology progress in this field must therefore heavily rely on detailed experimentation on scaled models. To this purpose Centrospazio has developed a low-cost instrumented facility, the CPRTF, Cavitating Pump Rotordynamic Test Facility. The inlet section of the facility has been recently reconfigured allowing for the installation of a Thermal Cavitation Tunnel, TCT, where current experiments have been conducted.
8.2 The Thermal Cavitating Tunnel (TCT configuration)
The CPRTF has been designed for general experimentation on noncavitating/cavitating turbopumps and test bodies in water under fluid dynamic and thermal cavitation similarity. An alternative CPRTF configuration used the Thermal Cavitation Tunnel (TCT). It was specifically designed for analyzing 2D or 3D cavitating flows over test bodies. In this configuration the pump is simply used to generate the required mass flow. The TCT was designed in order to accomplish the following requirements:
• adaptability to 2D and 3D test bodies
• to allow variable angle of attack of the test body • to allow the visualization of the phenomenon
• possibility to characterize the pressure field (Cp ) on test body surface
• possibility to carry out experiments in cavitating (σ < =Cpmin ) or non-cavitating (σ > =Cpmin )conditions
The TCT was designed and positioned in the suction line of the water loop according to the next Figure. The total length of the hydrodynamic tunnel is 1709 mm (304 mm contraction + 500 test section + 905 mm expansion). The rectangular test section (120x80x500mm), made of AISI 316 stainless steel, is constituted by four machined rectangular plates welded to the two interface flanges. Optical access is allowed through three large Plexiglas windows located on the lateral and top sides of the test section. O-ring seals are places between the Plexiglas windows and the stainless structure in order to avoid unwanted leakages.
The contraction starts from a square section inscribed into the circular one of the 6” preceding pipe and finishes with a rectangular shape identical to that of the test section. The pipe has been designed following a cubic polynomial geometrical law, whose characteristics have been carefully chosen in order to avoid cavitation along the contraction (cavitation troubles are widely reported by the open literature as a serious problem in designing the contraction pipes). The sudden passage from
not considered a serious trouble, because it is located exactly downstream of the flow straightener. Generation of large-scale turbulence in the sudden section change is so significantly limited. The expansion pipe, on the other hand, starts from a rectangular section equal to that of the test section and finishes with a square one designed in such a way to have the same frontal area of the successive Plexiglas circular pipe, in order to limit the pressure losses in the sudden section change. The pressure at the inlet of the test pump, in fact, has to be kept as high as possible, because cavitation on it has to be avoided in this configuration of the facility.
Figure 8.10 – Schematic of the position of the Thermal Cavitating Tunnel in the CPRT (left). Three-dimensional sketch of the Thermal Cavitating Tunnel (right).
The main characteristics of the Thermal Cavitating Tunnel are highlighted in the next Table. The tunnel maximum velocity is 8 m/s, the maximum pressure is 6 atm and the Reynolds number (Re= ⋅c V/υ) is maintained higher than 5⋅105in the tests performed to assure the completely turbulent
conditions. The minimum free-stream pressure attainable at the inlet of the test body is 0.1 atm.
Working Fluid water
Test section dimensions 120x80 mm
Total length of the hydrodynamic tunnel 1709 mm
Test body NACA 0015
Cord length 115 mm
Total n° pressure taps 12
n° pressure taps on the suction side 10
n° pressure taps on the pressure side 2
Maximum operative temperature 100°C
Upstream minimum pressure 0.13 atm
Maximum pressure in the tunnel 6 atm
Upstream maximum velocity 6 m/sec
Table 8.1 –Main characteristics of Thermal Cavitating Tunnel (TCT).
The test body, a NACA 0015 hydrofoil with 115mm chord and 80mm span length, is mounted on a blind panel on the bottom of the rectangular test section. The hydrofoil is made of marine bronze in
order assure the resistance to corrosion. The NACA0015 hydrofoil exploits the stall for an incidence angle of 15.4°. The experimental tests were performed at a lower angle of attack.
Figure 8.11 – Geometrical characteristics of the NACA0015 hydrofoil in term of percent of chord (right).
Figure 8.12 shows a schematic of the test section, with the hydrofoil instrumented with 12 pressure taps, 10 on the suction side and 2 on the pressure side in order to assure the symmetry. Three taps are located on the bottom panel upstream and 3 downstream of the test section to monitor the inlet/outlet pressure.
The incidence angle can be manually adjusted as necessary for the specific experiment. Each pressure tap on the hydrofoil was made by perforating vertically (from the bottom) the hydrofoil with holes of 2.5 mm till the middle height of the hydrofoil (90 mm) and intercepting them with horizontal holes of 1 mm. On the bottom of the hydrofoil the vertical holes are intercepted by the relative screw connections (1/8”) which are connected to the pressure transducer through Teflon tubes (Figure 8.15 and Figure 8.14). The pressure signal is read by an absolute pressure transducer.
Figure 8.12 – Schematic of the test section with the NACA 0015 hydrofoil and the locations of the pressure taps on the hydrofoil surface (x), at the section inlet (o) and outlet (o).
Figure 8.13 – Two-dimensional schematic of the test section and detail of the pressure taps.
Figure 8.14 – Schematic of the NACA0015 hydrofoil and detail of the pressure taps.
The next Figure shows assembly and a cut-off drawings of the test section with a NACA 0015 test body installed in it. The following Figures show some pictures of the Thermal Cavitation Tunnel assembled in the suction line of the facility and a detail of the test section with a NACA 0015 test body installed in it.
8.2.1 The effect of lateral constrains
The Thermal Cavitating Tunnel presents a rectangular test section in which the hydrofoil is mounted. The presence of the lateral boundaries can promote:
- an increase of the dynamic pressure, increasing all forces and moments at a given angle of attack
- an alteration to the local angle of attack along the span
- an alteration of the normal curvature of the flow so that the moment and lift coefficients and angle of attack are increased in a closed hydrodynamic tunnel
Figure 8.15 – Several views of the NACA0015 hydrofoil test section
8.3 Experimental results
8.3.1 Pressure coefficient
The present experimental tests have been conducted to analyze the pressure coefficient in noncavitating and inertial/thermal cavitating conditions on the suction side of the NACA 0015 hydrofoil. During these experiments pressure was measured at each pressure tap for constant values of the tunnel velocity, water temperature, incidence angle and cavitation number. The cavitation number, the pressure coefficient and the Strouhal number are defined as follows:
(8.2)
(8.3)
In particular, Figure 8.17 shows the pressure coefficient profile in noncavitating conditions for different incidence angles at room water temperature. Experimental uncertainty in the evaluation of the pressure coefficient is about 1.5% of the nominal value.
As expected, the experimental results are different from the surface pressure distribution in unconstrained flow. Figure 8.19 presents the comparison between the experimental noncavitating pressure coefficient on the suction side of the NACA 0015 hydrofoil and the theoretical one in unconstrained flow, for three incidence angles. The lateral constraints to the flow pattern promote “solid blockage” with the increase of the dynamic pressure, the hydrofoil forces and moments at given incidence angle (Kubota, Kato & Yamaguchi, 1992).
The absolute value of the pressure coefficient on the suction side of the test body is higher in constrained flows with respect to the unconstrained case: this increase is more significant in test sections where the “solid blockage” effect is more relevant. As a result, different pressure profiles can be measured on the same test body in different test facilities, depending on the position of the lateral constraints (compare, for instance, the data of Figure 8.17 with the results obtained by Kjeldsen et al, 1998). In the same way, when the “solid blockage” is more effective, the cavitation inception begins at higher cavitation numbers.
The results have been also compared with a two different CFD simulations in order to validate the numerical code developed for noncavitating/cavitating flows around 2D or 3D test bodies. The data presented in the Figures refer to simulations at 8° (Figure 8.17) and 4° Figure 8.18 incidence angles under constrained conditions. Figure 8.17 and Figure 8.18 are related to CFD simulations performed by two different teams, in particular by a group at CIRA (Centro Italiano Ricerca Aerospaziale) and another at Pisa University, respectively. Experimental results are in good agreement with the numerical reconstructions. 2 2 1 V p p L V in ⋅ − =
ρ
σ
2 2 1 V p p C L in p ⋅ − =ρ
Figure 8.17 – Pressure coefficient on the suction side of the NACA 0015 hydrofoil in noncavitating conditions for various incidence angles α at room water temperature. CFD simulation at 8° incidence
angle and room water temperature (solid line).
Figure 8.18 – Comparison of the pressure coefficient distribution obtained by the experimental results and four numerical solver methods. Initial conditions: ambient temperature and angle of attack 4° [Beux et Al,
2005; Bramanti, 2002].
Figure 8.19 – Comparison between the experimental noncavitating pressure coefficient on the suction side of the NACA 0015 hydrofoil and the theoretical one in unconstrained flow, for three incidence angles.
The next Figures present a comparison of the pressure coefficient profiles in noncavitating and cavitating conditions at several angles of attack and fixed cavitation number. The related pictures highlights the cavitation phenomenon over the suction side of the hydrofoil and confirm the experimental results. The next Figures are all referred to flow conditions in which cloud cavitation oscillations are clearly detected. The first part of the pressure profile, where the value of the pressure coefficient is practically constant and the condition Cp = -σ is almost verified, is related to a region in which cavitation is steady and no cavity oscillations are observed. The second region, where the pressure coefficient is no more constant but still different from the noncavitating value, is referred to the zone in which cloud cavitation oscillations extend. The mean pressure coefficient over the 1-second acquisition time is therefore an intermediate value between the full-cavitating and the noncavitating ones. Finally, the last part of the pressure profile, where the cavitating and noncavitating values of the pressure coefficient exactly match, is related to the part of the foil surface on which no cavitation, steady or unsteady, is observed. It has to point out that at higher incidence angles (α = 8°) the noncavitating region is not present, and therefore cloud cavitation oscillations extend through the whole test body.
Figure 8.20 – Pressure coefficient on the suction side of the NACA 0015 hydrofoil in noncavitating and cavitating conditions for 5° (top), 6° (left) and 8° (right) incidence angles and room water temperature.
Figure 8.21 compares the pressure profiles in cavitating conditions for three different freestream water temperatures at the same cavitation number and incidence angle. At higher temperatures the absorption of the latent heat at the cavity interface increases, reducing the vapor pressure under the unperturbed saturation value. This trend is well reflected in the figure: at 70 °C, due to pressure decrement under saturation value, the pressure recovery occurs more upstream than at room temperature.
The experimental uncertainty indicated in the Figures has been evaluated taking into account the random oscillations of the acquired data and the precision of the pressure transducers used for the experimentation. Experimental uncertainty in the evaluation of the pressure coefficient is about 2% of the nominal value.
Figure 8.21 – Influence of thermal cavitation effects on surface pressure distribution on the NACA 0015 hydrofoil at constant angle of attack α and cavitation number
σ
for several water temperatures T .8.3.2 Cavity oscillations
A number of experiments have been carried out in order to determine the characteristics of the cavity length and oscillations at different incidence angles, cavitation numbers and freestream temperatures. Cavity length for each nominal condition was calculated by taking pictures of the cavitating hydrofoil at a frame rate of 30 fps, during a period of 1 second (
Figure 8.22).
The mean cavity length along the span was determined for each picture with a maximum estimated error of 4% of the chord length (Figure 8.23). As a final result of this process, maximum, minimum and mean value of the 30 cavity lengths were obtained (Figure 8.24). At the same time, frequency spectrum of the upstream pressure was measured at each flow condition.
Figure 8.25 shows the maximum and minimum cavity lengths for various incidence angles at room water temperature.
The cavity length and the frequency spectrum of the upstream pressure are shown in Figure 8.26 and Figure 8.27 for the case of 8° incidence angle and room water temperature.
Figure 8.22 – Pictures of the cavity length development within 1 sec (30 pictures).
Figure 8.23 – Optical identification of cavitating region (left) and evaluation of mean cavity length along the span (right).
Figure 8.24 – Optical identification of cavitating region and evaluation of minimum (left) and maximum (right) cavity length along the span.
Figure 8.25 – Normalized maximum and minimum lengths of the cavity as function of the cavitation number σ for various incidence angles α at room water temperature.
Figure 8.26 – Characteristics of cavity length at 8° incidence angle and room water temperature. Experimental uncertainty in the evaluation of the cavity length is about 4% of the chord length.
Figure 8.27 – Frequency spectrum of the upstream pressure at 8° incidence angle and room water temperature.
Analysis of the Figures shows that the maximum and minimum cavity lengths provide a good qualitative indication of cavitation behavior on the hydrofoil at different cavitation numbers. One can recognize three different regimes of cavitation, corresponding to different ranges of values of σ :
- Supercavitation (σ < 1.3): both minimum and maximum cavity lengths are larger than the chord length. There are practically no cavity oscillations and therefore the frequency spectrum is almost flat (Figure 8.28).
Figure 8.28 – Typical cavitation appearance in “Supercavitation” case α=8°, σ=1.1, T=25°C.
- Bubble+Cloud cavitation (1.3<σ <2): the flow pictures show the occurrence of an initial zone of bubbly cavitation, followed by a second zone where the bubbles coalesce and strong cloud cavitation oscillations are observed. The frequency of these oscillations is almost constant at different cavitation numbers with a Strouhal number (
St
= ⋅
f c V
/
) of about 0.2, similar to those obtained by Tsujimoto et al. (1998) and Kjeldsen et al. (1998). Other frequencies, multiple of the first, are present in the spectrum. One can speculate that the second frequency might be related to the second, faster re-entrant jet sometimes observed by Sakoda et al. (2001). More likely, since higher frequencies areinteger multiples of the fundamental one, they may be the result of the frequency spreading caused by nonlinear effects in the flow field (Figure 8.29).
Figure 8.29 – Typical cavitation appearance in “Bubble+Cloud” case α=8°, σ=1.3, T=25°C, St=0.2.
- Bubble cavitation (σ >2.1): after a short transition zone, cloud cavitation disappears. Only the traveling bubble cavitation zone remains, with drastically reduced pressure oscillations (flat frequency spectrum) (Figure 8.30).
Thermal cavitation tests were carried out with a similar procedure for 8° incidence angle at two different freestream temperatures (50 °C and 70 °C). Results are shown in Figure 8.31, Figure 8.32 and Figure 8.33. The Figures show that, for higher freestream temperatures, the “Bubble+Cloud cavitation” zone tends to spread over a wider range of cavitation numbers and to begin at higher values of σ . Similarly, supercavitation also begins at higher cavitation numbers. These findings seem to be in accordance with the results obtained by Kato et al. (1996), who compared the temperature depressions in the cavity in water tests at 120 °C and 140 °C. At higher freestream temperatures and constant cavitation number, the cavity tends to become thicker and longer, even when there are no oscillations (“Bubble cavitation” zone), as shown in the Figure 8.34.
Figure 8.31 – L.E., maximum and minimum lengths of the cavity for three different water temperatures T at 8° incidence angle.
Figure 8.32 – Frequency spectrum of the upstream pressure at 8° incidence angle and 50 °C water temperature.
Figure 8.33 – Frequency spectrum of the upstream pressure at 8° incidence angle and 70 °C water temperature.
At higher freestream temperatures and constant cavitation number, the cavity tends to become thicker and longer, even when there are no oscillations (“Bubble cavitation” zone), as shown in the next Figure.
Figure 8.34 – Cavity thickness for three different water temperatures T at the same incidence angle α and cavitation number σ (α= °8 ,σ =2.5).
The above results appear significantly different from the typical data presented in the open literature, where at higher temperatures the cavitation breakdown tends to be shifted towards lower cavitation numbers and the cavity is shorter. A suitable explanation for this apparent disagreement can be given taking into account the influence of the lateral constraints: the observed increase of the cavitation thickness at higher temperatures, when the lateral wall is sufficiently near to the suction side, can promote cavity spreading along the longitudinal direction. This effect is more significant at higher incidence angles, for which the flow is forced to pass through a reduced cross section. The above statement is confirmed by analysis of Figure 8.21 and Figure 8.31: at a lower incidence angle (α
outlined opposite behaviour. Cavitation at higher freestream temperatures looks quite different: bubbles are smaller and tend to coalesce more easily, resulting in a narrower and less defined “bubble zone” compared to the “cloud zone” (Figure 8.35).
Figure 8.35 – Cavitation appearance at higher freestream temperature (α = °8 ,σ =2,T=70°C).
8.3.3 Thermal effects on the pressure drop
Another set of experimental tests was conducted in order to determine the pressure drop caused by the hydrofoil at various incidence angles, freestream temperatures and cavitation numbers. Tests were performed using a differential pressure transducer mounted between one of the pressure taps upstream of the test body and the corresponding pressure tap downstream. The pressure drop obtained using this procedure represents therefore a “punctual” value that cannot be directly related to the drag, but should nevertheless exhibit a similar general behaviour. Figure 8.36 and Figure 8.37 show, respectively, the results obtained for different incidence angles and for the same incidence angle at different freestream temperatures. The experimental uncertainty indicated in the Figures has been evaluated taking into account the random oscillations of the acquired data and the precision of the pressure transducers used for the experimentation. Experimental uncertainty in the evaluation of the pressure drop is about 5% of the nominal value.
Figure 8.36 – Normalized pressure drop caused by the hydrofoil for various incidence angles α at room water temperature.
Figure 8.37 – Normalized pressure drop caused by the hydrofoil for three different water temperatures T at 8° incidence angle.
The Figures show a sudden rise of the pressure drop for cavitation numbers slightly smaller than those corresponding to the onset of supercavitation; at higher freestream temperatures this “breakdown” effect occurs at higher cavitation numbers, just like supercavitation and cloud cavitation oscillation. This behaviour, strictly related to the increase of cavity thickness (Figure 8.34), is different from that observed on hydrofoils in free flows and in cascades, probably as a result of the more significant “solid blockage” effect (in a free flow there is no solid blockage, in a staggered cascade solid blockage is less effective because the blades overlap only partially).
9 O
THER RESEARCH
ACTIVITIES
The following section is dedicated to the description of some the other activities performed in parallel with the main research activity illustrated in the previous sections.
A brief introduction on the properties and possible applications in space of hydrogen peroxide and will be performed. The activity carried out at ALTA on the design and testing of a hydrogen peroxide monopropellant rocket and on an assessment study on an innovative concept on a bipropellant rocket using hydrogen peroxide and ethane as propellants will be illustrated. Finally, the experimental characterisation of different advanced materials for the catalysis of hydrogen peroxide will be also presented.
9.1 Hydrogen peroxide in space applications
As space missions become more ambitious, the need for reducing the costs and increasing the capabilities of rocket systems through the enhancement of their propulsion performance, safety and reliability represents a major aspect in the development of competitive space engines. A variety of factors have resulted in an increasing interest in the exploration of alternatives to widely employed cryogenic and hypergolic propellant combinations. These factors include heightened sensitivity to cost, environmental concerns and personnel protection from the hazards associated with the use of present highly toxic propellants.
The market of LEO satellites is the most promising one in this respect, due to their lower cost and the consequent reduction of the risk associated with the use of innovative technologies and propulsion concepts. The main producers of small and microsatellites for LEO applications show strong interest for low toxicity, or green, storable liquid propellants as possible substitutes for hydrazine and nitrogen tetroxide. The main advantage of “green” propellants is represented by the significant cost savings associated with the drastic simplification of the health and safety protection procedures necessary
during the production, storage and handling of the propellant [Anderson et al (1998), Austin et al (2002), Bombelli et al (2003), Humble et al (1995) and (2000)].
Because of their superior propulsive performance the current standard in high-performing, storable bipropellants is the combination of nitrogen tetroxide (NTO) and hydrazines (N2H4, MMH and UDMH). Typical values of the specific impulse and combustion temperature developed by NTO-hydrazines systems are shown in Figure 9.1. However, NTO-hydrazines are highly toxic and carcinogenic and bear the risk of unwanted detonation if exposed to high rates of change of pressure and temperature, depending also on the containment conditions. Hydrazine-propelled satellite systems must therefore be designed to incorporate preventive measures against all identified hazards. On-ground operations as well as transportation and handling of hydrazines are subjected to very restrictive safety procedures and have to be carried out with costly dedicated special precautions and specific infrastructure provisions (Figure 9.2).
Figure 9.1 – Flame temperature (left) and specific impulse (right) versus oxidizer/fuel mass ratio for equilibrium adiabatic reaction at 3.45 MPa and frozen flow expansion to 13.8 kPa of nitrogen tetroxide, N2O4, and several hydrazine fuels (hydrazine, N2H4, monomethyl hydrazine, MMH, and unsymmetrical
dimethyl hydrazine, UDMH).
Figure 9.2 – Historical evaluation of fueling costs versus payload hardware costs for space missions
The use of non-toxic “green” propellants would greatly contribute to reduce these drawbacks and significantly lower the life-cycle cost of small- and medium-size spacecrafts. Besides, overcoming
contingency operations. Finally, the elimination of safety hazards associated with the use of dangerous propellants contributes to drastically decrease the environmental impact and clean-up costs in the case of inadvertent spills or satellite launch failures.
The most promising high-energy green propellants (like for instance ADN and HAN) are based on complex organic molecules and compensate the large molecular weight of their decomposition products with high operational temperatures of the exhaust gases. As a result, it is necessary to use extremely expensive materials and manufacturing processes for the thrust chamber and, at the same time, the operational life of the catalytic beds is drastically reduced.
Hydrogen Peroxide, on the other hand, does not suffer from these disadvantages and has therefore been reconsidered, despite its lower specific impulse, as a promising green propellant for low and medium thrust applications. It can be used as a monopropellant, as an oxidizer for bipropellant engines, or in conjunction with a solid propellant in hybrid rockets.
The most significant technology challenge for the realization of hydrogen peroxide monopropellant and bipropellant thrusters is represented by the development of effective, reliable, long-lived catalytic beds, giving fast and repeatable performance, not sensitive to poisoning by the stabilizers and impurities contained in the propellant, capable of sustaining the large number of thermal cycles imposed by typical mission profiles and not requiring (if possible) preheating for efficient operation. To this purpose new catalytic beds have to be developed, integrated in a small thruster and tested.
Hydrogen peroxide has been manufactured commercially since the end of the 19th century, but the first to recognize its potential as a propellant was Hellmuth Walter in Germany. Its first aerospace application was for take-off assistance of the Heinkel He 176 aircraft, flown in 1938, which used a 5783 N thrust unit where HP was decomposed by the simultaneous injection of a liquid catalyst. The German V2 rocket also used HP decomposition for powering the propellant feed turbopumps. After the war, tests on hydrogen peroxide rockets were continued in the UK and USA, first using 80% German HP and, during a second phase, with purer 85% HTP (High Test Peroxide) decomposed by silver plated gauze catalyst packs. In the UK these activities led to the use of hydrogen peroxide as an oxidizer for bipropellant rocket engines like, for instance, the Gamma 201 and 301 (in the Black Knight and Black Arrow vehicles), the Spectre and the Stentor, while in the USA HP was the propellant of the AR (Aircraft Rocket) series of engines used by the USAF for aircraft thrust augmentation. In the 1950’s and 1960’s monopropellant rocket engines based on the decomposition of hydrogen peroxide have also been extensively used as Reaction Control Systems for aircrafts (1, X-15) or space vehicles (Mercury, Syncom, Comsat).
Hydrogen Peroxide (HP) is a high density liquid having the characteristic of being able to decompose exothermically into water (steam) and oxygen according to the reaction:
H2O2(l)→ H2O(g)+
1
2O2(g)
At room temperature HP has a low vapor pressure, it remains in the liquid state at ambient pressure in a wide range of temperatures, and is relatively easy to handle with respect to other
common rocket propellant liquid oxidizers like nitrogen tetroxide, nitric acid and liquid oxygen. The physical properties of HP are close to those of water, with two notable differences: hydrogen peroxide has a significantly higher density and a much lower vapor pressure, as shown in the following Tables.
50% HP 70% HP 85% HP
Density (kg/m3) 1198 1294 1366
Freezing point (°C) -52 -40 -17
Boiling point (°C) 114 125 137
Table 9.1 – Physical properties of hydrogen peroxide at various concentrations
Hydrogen peroxide is a storable, low-toxicity, non-hypergolic, relatively inexpensive liquid oxidizer. It does not react with the elements and compounds in the atmosphere, unlike hydrazine, which in contact with carbon dioxide forms compounds that can seriously attack the structural materials of the thrust chambers. HP’s very low vapor pressure allows pumping machinery, if present, to operate without cavitation at lower inlet pressures.
Characteristic Hydrazine Hydrogen Peroxide (90%)
Melting Point, °C 1.5 -11.5
Boiling Point, °C 113.5 141.7
Specific Gravity at 20°C, gm/ml 1 1.4
Explosion Temperature, °C 232 149
Vapor Pressure at 20°C, kPa 1.4 0.3
Long-Term Storage Stability Excellent if kept blanketed with inert gas
Slowly decomposes to form oxygen and water Other Precautions Corrosive, Flammable,
toxic
Need to be stabilized, corrosive
Table 9.2 – Comparison of hydrazine, hydrogen peroxide and ethane main characteristics
Features Benefits
Non-toxic and storable Commercial, easier propellant packaging and no insulation, simplified ground operations
Favorable thermo-chemistry
High density impulse, simpler thermal management
Oxidizer/monopropellant No separate ignition systems required, high range of thrust variation, smoother starts and shutdowns, low cost pump feed systems
Gas-liquid injection Increased stability margin, high combustion efficiency, simple injection system
Table 9.3 – Features and benefits of hydrogen peroxide as propellant
When it was developed as a monopropellant for rockets, hydrogen peroxide was one of the most powerful known chemicals that could be safely used to this purpose. The situation changed in the late 1950’s when technology advancements led to the industrial production of ultra-high purity hydrazine and high-performance Iridium-based hydrazine decomposition catalysts. These advancements opened
HP in almost all RCS thruster applications. The next Figures show the specific impulse of HP as a monopropellant and a comparison with the propulsive performance of hydrazine. Although the propulsive performance of HP is about 20% lower than for hydrazine, the significant savings allowed by the use of HP still make it an attractive alternative in small and medium-size propulsion systems for cost-driven rather than performance-driven missions.
Figure 9.3 – Vacuum specific impulse of hydrogen peroxide at various concentrations (Ventura and Muellens, 1999)
Figure 9.4 – Vacuum specific impulse of hydrazine and hydrogen peroxide at various concentrations, as a function of the nozzle expansion ratio (Ventura and Muellens, 1999)
With respect to bipropellant and hybrid rocket engines, hydrogen peroxide yields a specific impulse comparable to other liquid oxidizers like nitrogen tetroxide, nitric acid and even liquid oxygen. The next two Figures illustrate the performance of HP as a bipropellant and a hybrid propellant in comparison to other commonly used oxidizers. In particular, the specific impulse of HP/hydrocarbons combinations is similar to that of mono-methyl-hydrazine (MMH) and nitrogen tetroxide (NTO).
Figure 9.5 – Vacuum specific impulse of hydrogen peroxide at different concentrations, compared to other oxidizers, with various fuels in bipropellant rockets (Ventura and Muellens, 1999)
Figure 9.6 – Vacuum specific impulse of hydrogen peroxide at different concentrations, compared to other oxidizers, in hybrid rockets (Ventura and Muellens, 1999)
9.2 Hydrogen Peroxide as propellant for monopropellant rocket
The present section is dedicated to a brief description of the activities conducted in the framework of a LET-SME program funded by the European Space Agency. The focus of the activity is the design and realisation of two prototype monopropellant thrusters (5N and 25N), and the experimental characterisation of different catalytic beds made of “advanced materials”. The next Table presents the design requirements and specifications of the prototype, which were driven from a detailed literature review on the thrust chambers and catalytic beds for hydrogen peroxide. The propellant for which the thrusters have been specifically designed a high-concentration (87.5%) hydrogen peroxide solution, stable and relatively easy-handling, with a significantly low content of impurities and stabilizers. With regard to operating pressures, a nominal operating pressure of approximately 10 bar has been chosen for the combustion chamber, taking into account the typical values used in similar applications in the
and effective decomposition of the propellant in the catalytic bed. The target value of the catalytic bed loading has been chosen near the lowest limit of the range of values typically used in the past. Finally, the nozzle geometry (in particular the convergent/divergent angles and the fillet radius) has been defined following the typical values used for conical nozzles.
Table 9.4 – Main preliminary requirements and specifications for the prototype thrusters.
Starting from the specifications highlighted, preliminary dimensioning of the prototype thrusters has been carried out using simplified isentropic mono-dimensional relations, which will be briefly illustrated in the following. The combustion chamber temperature can be evaluated using the classical reaction enthalpy balance equations or, as an alternative, one of the several available software tools for chemical equilibrium calculations. The specific heat ratio and the molecular weight of the exhaust gases are evaluated in a similar way. Then, the characteristic velocity of the propellant is calculated by means of the classical equation:
( 1) 2 1 1 2 c RT c γ γ γ γ + − ∗= ⎛ + ⎞ ⎜ ⎟ ⎝ ⎠ (1)
where Tc is the combustion chamber temperature, γ is the specific heat ratio of the exhaust gases
and R is their gas constant (which, in turn, is a function of the molecular weight). According to the frozen-flow approximation, the throat temperature (Tt), pressure (pt), density (ρt) and velocity (ut) are
given by the following relations:
1 2 2 1 1 t t c t c t t t t p T T p p u RT RT γ γ ρ γ γ γ − ⎛ ⎞ ⎛ ⎞ = ⎜ ⎟ = ⎜ ⎟ = = + + ⎝ ⎠ ⎝ ⎠ (2)
where pc is the combustion chamber pressure. Exhaust Mach number (Me), temperature (Te) and
( 1) 2 2 1 1 1 2 1 c e e e c e e e e p p M T T M p RT γ γ γ ρ γ − ⎡⎛ ⎞ ⎤ ⎛ − ⎞ ⎢ ⎥ = − ⎢⎜ ⎟ − ⎥ = ⎜⎝ + ⎟⎠ = ⎝ ⎠ ⎣ ⎦ (3)
where pe is the exhaust pressure. Then, isentropic nozzle equations give the possibility of
calculating the exhaust/throat area ratio and the thrust coefficient:
( 1 2) ( 1) 2 1 2 1 1 1 2 e e t e A M A M γ γ γ γ + − ⎡ ⎛ − ⎞⎤ = ⎢ ⎜ + ⎟⎥ + ⎝ ⎠ ⎣ ⎦ (4) 1 1 1 2 2 1 1 1 e e a e F c c t p p p A C p p A γ γ γ γ γ γ γ + − − ⎡ ⎛ ⎞ ⎤ ⎛ − ⎞ ⎛ ⎞ ⎢ ⎥ = ⎜ + ⎟ − ⎢ −⎜ ⎟ ⎥+⎜ ⎟ ⎝ ⎠ ⎢⎣ ⎝ ⎠ ⎥⎦ ⎝ ⎠ (5)
where pa is the ambient pressure. Finally the exhaust velocity, throat area and propellant mass
flow rate are given by the following equations:
c t F t F c p A F c C c A m C p c ∗ ∗ = = = (6)
where F is the required thrust. The main performance characteristics of the two prototype thrusters, evaluated by means of the above simplified equations, are provided in next Table, where the “corrected” thrust (evaluated taking into account the effects of the displacement thickness at the throat) is also indicated, together with an estimation of the pressure drop across the catalytic bed, carried out using a simplified model based on the equations for the pressure drop across a porous medium.
Table 9.5 – Main performance characteristics of the prototype thrusters, evaluated by means of simplified isentropic relations.
Three-dimensional models of the two prototype thrusters, based on the design obtained by the above procedure, are presented in next Figures, on the other hand, shows a cut-off drawing of the 25 N thruster, highlighting its main components.
The main guidelines in the design of the prototypes have been the following:
• The thrusters have been designed in a modular manner, i.e. the nozzle, the catalytic bed and the connecting flange are separate components and each one of them can be substituted without changing the others.
• Two static pressure taps are present in each thruster, one at the catalytic bed inlet and the other at its outlet. One temperature and one dynamic pressure tap are designed at the catalytic bed outlet.
• Sealing has been assured by means of copper-coated, inconel C-rings.
• Because of the moderately high value of the combustion chamber temperature, the engine thrust chamber and nozzle do not require the use of rare materials like Iridium-coated Rhenium or Niobium. For this reason, the material of all the non-catalytic components is the relatively inexpensive and easily machinable AISI 316L stainless steel.
Figure 9.7 –Three-dimensional drawings of the 5 N (left) and the 25 N thruster (right).
The main elements which characterize the thruster are:
• the injection plate, for the supply and distribution of the liquid hydrogen peroxide to the catalytic reactor;
• the catalyst and the catalytic bed, which effects the decomposition of HP in steam and gaseous oxygen;
• the convergent/divergent nozzle, for the acceleration of the exhaust gases and the generation of the propulsive thrust.
The catalytic bed is installed as a unique “cartridge” and joined to the other components of the prototype thruster by means of flanges. Different catalytic bed cartridges, for example with different lengths or catalyst geometry, can be installed in the prototype using the same nozzle and connecting flange. This conceptual solution has the interesting advantage of being adaptable to different catalyst geometries, like gauzes or pellets. The choice of the materials for the catalytic bed derived from the experimental activity (which will be briefly described at the end of this section) conducted in order to evaluate the catalytic behaviour of different substances with the hydrogen peroxide. The bed configurations to be tested are the following:
• Pure silver grids, obtained by a 80 mesh gauze woven from a 0.115 mm diameter wire. In this case, some gauzes with greater wire diameter (made of silver or nickel) are alternated with the more active ones, in order to provide sufficient mechanical strength to the bed. • Alternation of different metallic gauzes.
• Commercially available manganese oxide-covered pellets.
• Self-impregnated manganese oxide pellets (on ceramic substrate). A technique is under development by Alta S.p.A., in collaboration with the Department of Chemistry of the University of Pisa, for the impregnation and calcination of high contact surface (up to 300 m2/g) spheres made of Alumina-γ, using particular organic compounds as promoters. The diameter of the spheres is between 0.5 and 1 mm and, as a consequence, retaining grids are needed at the beginning and the end of the catalytic bed to avoid loss of the spheres in the flow through the distribution plate. The same calcination technique can also be applied to platinum-covered spheres, which are being used as another possible catalyst material for the tests.
More complex bed casing configurations are currently under design, including the possibility of measuring temperature at various stations along the catalytic bed by means of suitably designed ports and the insertion of interstitial elements of suitable form between the screen or pellets catalysts, in order to prevent possible problems due to channeling effects.
The injection and distribution plates have been chosen with an open area ratio of about 0.5 for both the prototypes. Considering 1.5 mm diameter holes, this leads to an 18-holes plate for the 5 N prototype and a 92-holes plate for the 25 N prototype. The plate thickness has been chosen to be 1.5 mm. Considering the really low estimated values of the pressure drop across the plates, the distribution plate has been designed equal to the injection plate (differently from what is typically observed in larger-size thrusters, where the distribution plate has normally a wider open area ratio than the injection plate).
The connecting flange (pink element in the above Figures) is designed in order to match the interface dimensions of the catalytic bed (on one side) and the thruster valve (on the other side). The thruster valve presently used for the experimentation is a commercial solenoid valve able to provide a response time of no more than 20 milliseconds.
9.3 Hydrogen peroxide-ethane propellants for bi-propellant rocket engines
9.3.1 Introduction
The high operational complexity associated with hydrazines leads to renewed consideration of alternatives aiming at reducing both the involved costs and risks. To this purpose, in fact, low toxicity, or “green”, storable liquid propellants have recently become considerably more attractive as possible substitutes for NTO and hydrazines. In a number of respects hydrogen peroxide and ethane represent an example combination of high-performance, “green” bipropellant propulsion systems. As previously mentioned, hydrogen peroxide is non-hypergolic with ethane, but in contact with suitable catalysts decomposes exothermically into hot oxygen and steam, capable to ignite and sustain ethane combustion. In particular, ethane is a stable, low-toxicity, inexpensive hydrocarbon with critical state at 5.01 MPa and 305.9 K. At room temperature (280 to 300 K) its vapor pressure is ideally suited for propellant pressurization, given the desired values of the thrust chamber pressure (1 to 3 MPa) and the expected losses in the feed system (20 to 40% of the chamber pressure) (see next Figure).
Figure 9.9 – Saturation pressure (left) and density (right) of ethane as function of the temperature
The next Table summarizes the main characteristics of ethane.
Characteristic Ethane
Melting Point, °C -183
Boiling Point, °C 88.6
Specific Gravity at 20°C, gm/ml 1.049
Explosion Temperature, °C 472
Vapor Pressure at 20°C, kPa 3850
Long-Term Storage Stability Excellent if prevent leakage Other Precautions Flammable, explosive with air.
Table 9.6 – Comparison of hydrazine, hydrogen peroxide and ethane main characteristics
Typical values of the specific impulse and combustion temperature of H2O2-C2H6 systems are shown in Figure 9.10. Computations have assumed adiabatic equilibrium reaction at constant pressure
(3.45 MPa) in the combustion chamber and frozen flow expansion to 13.8 kPa in the nozzle. The maximum specific impulse (298 s at 98% H2O2 concentration and O/F = 7) is less than 7% lower than for NTO/MMH systems (320 s) and relatively insensitive to mixture ratio changes. For optimum specific impulse NTO-hydrazines systems operate fuel-rich at dangerously high temperatures (3050 K) lower than the maximum, and are therefore easily exposed to the risk of local overheating due to uneven propellant injection and mixing. Conversely, H2O2-C2H6 systems operate oxidizer-rich near their maximum flame temperature (and therefore with no risk of soothing and dangerous hot spots) at 2872 K, which is fully compatible with current radiation-cooled coated-niobium thrust chambers.
Figure 9.10 – Combustion temperature (left) and specific impulse (right) versus oxidizer/fuel mass ratio for equilibrium adiabatic reaction at 3.45 MPa and frozen flow expansion to 13.8 kPa of hydrogen
peroxide and ethane for different H2O2 mass concentration (0.70; 0.85; 0.98).
The density of hydrogen peroxide at high concentrations (above 70%) is comparable to that of nitric acid and nitrogen tetroxide and significantly higher than that of liquid oxygen. As a consequence, the volume and dry mass of the oxidizer tank can be reduced significantly and, as shown in the next Figure, the volume specific impulse of 90% HP is higher than of most other propellants. This is particularly useful for systems with significant aerodynamic drag losses and/or stringent volumetric constraints.
Figure 9.11 - Ideal volume specific impulse of several bipropellants, as a function of the oxidizer/fuel mixture ratio
An innovative green bipropellant thrusters with Fuel Vapor Pressurization (FVP) of hydrogen peroxide and ethane (C2H6), where the catalytic reactor provides the oxidizing stream for C2H6 combustion, has been studied and its propulsive and operational performance has been analyzed and assessed. In the US fuel vapor pressurization has been actively studied since 1994 [Moser (1994), (1995), (2001)] and is now undergoing development tests. However, no similar experience is presently available in Europe, nor FVP has ever attained flight readiness anywhere in the world.
9.3.2 Fuel Vapor Pressurization (FVP) principle of operation
Fuel vapor pressurization (Figure 9.12) systems exploit the high vapor pressure of light hydrocarbons to transfer both the fuel and the oxidizer into the combustion chamber. Exploratory analyses of propellants thermodynamic and propulsive properties have indicated that hydrogen peroxide and ethane represent the most promising propellant combination that best exploits the potential advantages of FVP for the realization of simple, safe, reliable, inexpensive and high-performance rocket propulsion systems.
Figure 9.12 – Schematic of H2O2-C2H6 rocket engine with fuel pressurization The main elements which characterize the configuration of FVP systems are:
• the dual tank for the pressurization of both fuel and oxidizer;
• the injection plate, for supply and distribution of liquid hydrogen peroxide to the catalytic reactor;
• the catalytic bed, which effects the decomposition of HP in gaseous oxygen and steam;
• the convergent/divergent nozzle, for accelerating the exhaust gases and generating the propulsive thrust.
When the oxidizer valve is opened, allowing the oxidizer flow from the tank to reach the injection plate and the catalytic bed, hydrogen peroxide undergoes an energetic exothermic decomposition, creating a hot oxidizing gas downstream of the catalytic reactor. After a short delay, the fuel valve is also opened and gaseous ethane is tapped from the upper end of the propellant tank and delivered to the injector. As it expands through the injection orifices, almost the 95% of the ethane stays in its vapor state, while the remainder condenses into liquid droplets, which are immediately vaporized again in the combustion chamber. The fuel spray impinges on the hot, high velocity oxidizer stream, resulting in spontaneous ignition and rapid, efficient gas-gas combustion.
Since the same pressure drives both the oxidizer and fuel flows, only passive devices (cavitating venturis for the liquid oxidizer and sonic injectors for the gaseous fuel) can be used to accurately control the oxidizer/fuel ratio regardless of the value of the driving pressure. No pressurization tanks, pressure regulators and flow control valves are needed, realizing a very simple and reliable propellant management system where the only active components are the propellant shut-off valves.
Because of the moderately high value of hydrogen peroxide/ethane combustion temperature (about 3000 K for 90% H2O2), the radiation-cooled thrust chambers and nozzles do not require the use of rare and extremely expensive materials like Iridium-coated Rhenium, but can be realized with the more traditional coated-Niobium technology.
An important aspect of fuel vapor pressurization systems consists in the use of just one tank for both propellants. Significant savings of system mass and complexity are gained by storing the two propellants in the same tank, using a flexible diaphragm or a bladder to separate the fuel and the oxidizer. In this configuration ethane and hydrogen peroxide are in thermal contact with each other and the heat capacity of the oxidizer can be exploited to minimize the tank temperature and pressure drifts due to fuel evaporation during propellant extraction.
As the propellants are expelled from the tank, the fuel temperature and pressure will tend to decrease as a consequence of the evaporation required to fill the larger available volume. Counteracting this effect is the transfer of heat from the oxidizer to the fuel through the thickness of the separating bladder. Since the fuel is initially close to its critical temperature, the latent heat of vaporization is low. Besides, since the O/F mass ratio is relatively large (between 7 to 9), the thermal inertia of the oxidizer effectively reduces the drift of the tank temperature and pressure during propellant extraction.
Preliminary calculations indicate that the high value of the optimum oxidizer/fuel ratio (O/F) is indeed quite beneficial in stabilizing the tank temperature and pressure, whose drifts for adiabatic propellant extraction do not exceed 20 K and 1.5 MPa, respectively. These calculations are based on a simplified thermodynamic model of the propellants in the tank where the following assumptions have been introduced:
• adiabatic conditions; • incompressible oxidizer;
• uniform temperature and pressure; • constant mixture ratio φ;
• constant tank volume V ;
• thermodynamic equilibrium between liquid fuel and its vapor. From the continuity equation:
