Fuel Ignition in HTP Hybrid Rockets at Very Low Mass Fluxes: Challenges and Pulsed Preheating Techniques Using Palladium-Coated Catalysts

Abstract

In a worldwide scenario which sees an increasing number of small satellite launches, novel mission concepts may be unlocked providing the spacecrafts with the very precise and rapid maneuvering capability that electric thrusters cannot guarantee. In this context, chemical thrusters appear to be a possible solution. This work aimed to experimentally study and solve the problem of ignition for 10 N hybrid rockets based on hydrogen peroxide. Firstly, the study analyzed the performance of a monopropellant engine capable of functioning as a hybrid injection system. In particular, the effects of the liquid mass injected, the initial temperature, and the supply pressure on the pulsed engine performance were experimentally investigated. The injected mass showed a greater impact on the performance with respect to the starting chamber temperature and injection pressure. This thruster also showed a good potential for space applications. In the second part of the work, the objective was to find an ignition procedure that reduced propellant consumption and eliminated the need for a glow plug. This is important because the electrical power consumption in real applications significantly affects other subsystems and is undesirable for chemical engines. Different ignition procedures were tested to emphasize their respective advantages and disadvantages, and the findings indicated that the concept of pulsed preheating is feasible with only a small amount of propellant consumption, while substantially decreasing the ignition duration from approximately 45 min to a maximum of just 3 min. Finally, similar ignition procedures were adopted using different fuels. The results showed that PVC and ABS, under the same operating conditions, ignite more easily than HDPE, which requires an oxidizer consumption approximately double that of the other two fuels. Considerations about the effect of chamber pressure and oxidizer mass flow rate on engine ignition were also included.

1. Introduction

A propulsion system is the main mobility device for a spacecraft and is used for attitude control, drag recovery, and orbit changes [1,2]. Recent years have witnessed a huge increase in interest in Cube Satellites (CubeSats). There are two main drivers: low access-to-space cost for demonstration technology purpose, science proof of concept validation, communication, and education; second, utilizing commercial off the shelf technologies in the design architecture. These two factors have provided a significant reduction in overall costs for a CubeSat mission [3].

Typically, a CubeSat unit has the dimensions of 10 × 10 × 10 cm³ and a mass of 1 kg. CubeSat platforms go from three units (3 U) up to 27 U. Thus, CubeSats are restricted in their operations due to their size that limits their on-board capabilities leading to a limited mission life and travel range. Nevertheless, these small satellites like nano- or micro-satellites have evolved from simple passive in-orbit missions to being capable of performing active orbital operations that would require propulsive capabilities [4].

Chemical thrusters are preferred when rapid maneuvers are required as the electrics are not as effective due to their low thrust. Among chemical thrusters, liquid monopropellants and hybrids present some features which make them appealing to comply with the stringent requirements of a CubeSat, including safety, compactness, reliability, simplicity, and re-ignition capability.

Thus, the present work was focused on the characterization of a pulsed hydrogen peroxide-based thruster which could have a dual role in space missions. Firstly, it could serve as a monopropellant for fine orbit maneuvers and secondly, it could be a in ignition system for a hybrid capable of performing fast maneuvers.

The choice of the propellant lies in the following reasoning: hydrazine is known as the most used propellant in monopropellant engines. The problem with this substance is its high level of toxicity, which results in increased costs. These cover the entire process from production to storage in tanks. Furthermore, there is also the environmental problem due to the toxicity of the vapors expelled in the atmosphere. All these factors contributed to a renewed interest in the utilization of hydrogen peroxide [5,6], that can be used also as an oxidizer for hybrids [7].

Over the last few years, tests regarding the performance of both monopropellant and hybrid thrusters based on hydrogen peroxide have been conducted by several groups of scientists. As for monopropellants, numerous tests examine the improvement of the decomposition efficiency by studying catalysts type, preheat temperature, or catalytic bed size [8,9,10,11,12,13,14]. Some works [15,16] have carried out static firing tests of a hybrid engine to demonstrate the possibility of improving these performances by modifying the injection of the liquid oxidizer. Additionally, multipulsed operation has been demonstrated with constant propulsive performance. Pasini et al. [17,18] tested a monopropellant thruster in pulse mode operation and analyzed how the duration of the valve open range time, during which the propellant is injected, affects the chamber pressure performance.

Thus, in the literature there are few works that study monopropellants in pulsed mode, and their analysis is mainly focused on the influence of catalysts on engine performance at the steady state. For this reason, the first part of this work was an attempt to obtain information on the decomposition efficiency trends as a function of the injected mass, injection pressure, and initial temperature (hot and cold start) for pulsed injection.

Considering hybrids, they are largely studied all over the word by several research groups; hydrogen peroxide finds application in numerous works about engines with a thrust class greater than 100 N. Some examples are shown by Mayer et al. [19] who developed and tested a multiport 200 N hydrogen peroxide-based hybrid rocket, or Behiang University, where researchers are studying various configurations of solid fuels [20,21]. Also, the Korean research group guided by Prof. Kwon has presented several works over the years [22,23].

However, 10 N scale hybrid rockets have not been studied extensively in the past, as there was little interest in them until a few years ago. The primary issue is that these thruster types are designed for use on CubeSats, which have extremely strict mass, volume, and available power limitations. Because of this, it is crucial to investigate the ignition process in order to reduce the amount of propellant and power needed. In addition, the mass fluxes involved, strongly influencing regression rate [24], are relatively low (ranging from 15 to 60 kg/m² s). It would be interesting to assess the impact of this factor on the ignition process. For this reason, the focus and the novelty of the present work are in the study of the ignition of a 10 N hybrid rocket through different procedures. Specifically, the data gathered from the initial tests conducted on monopropellant configuration were utilized to establish a potential pulsed preheating method without relying on a glow plug, aiming to eliminate the need for electrical power consumption. In the following paragraphs, the logical progression that guided the authors to acquire crucial information through numerous experimental tests will be elucidated.

2. Experimental Breadboard and Setup

Functional demonstration tests were carried out at the propulsion laboratory facility of the University of Naples “Federico II”, which is available at the Grazzanise (CE) military base.

A schematic of the experimental setup is displayed in Figure 1 [25]. The high pressurized nitrogen was stored in a cylindrical tank with a pressure of 200 bar. Then, a pressure regulator reduced the gas pressure to the desired value along the motor feed line. Hydrogen peroxide was stored in a tank with a volume of 2 L; downstream of the latter, the measurement of the mass flow rate was provided by means of an electronic Coriolis flow meter/regulator model (Bronkhorst Cori-Flow M55). Upstream of the injector was located the solenoidal electro-valve model Parker Miniature Calibrant Valve Series 9. It can carry out both continuous and pulse mode operations, with a response time of less than 5 ms. The facility was provided with pressure, thrust, and temperature sensors whose signals, sampled at 5 kHz, were recorded on the hard disk by a National Instruments PXI 1082e. Then, these signals were downsampled for a quick visualization on the computer, which was wired with PXI by means of optical fiber cables.

An on-ground breadboard was developed and equipped with housing for pressure measurements and thermocouples. The design and the concept are explained in a previous publication [26]. Figure 2 shows the main components; the supporting flange in Figure 2a represents the interface between the hydrogen peroxide feeding line and the thruster. This flange has the possibility of housing a glow plug which, as will be shown in the next sections, was used in some tests to preheat the catalytic compartment with a power consumption of the order of 10 W. The flange was connected to an injection plate characterized by three 0.3 mm diameter holes (Figure 2b). The hydrogen peroxide was injected into the decomposition chamber, Figure 2c, which was filled with catalysts and downstream it was separated from the nozzle by a pellet containment plate, Figure 2d.

The decomposition chamber was a 36 mm long and 25 mm diameter cylinder. The system was oversized with respect to the mass flow rate treated during the tests, in order to ensure a slightly higher residence time and thus improving the decomposition efficiency [26]. Finally, the nozzle of 2 mm throat diameter closed the entire system.

Since the pressure ratio between chamber and ambient pressure was relatively low, a simple convergent nozzle was employed; it is indicated with the red arrow in Figure 3.

A Tersid K-type thermocouple (IndiaTersid, Sesto San Giovanni, Italy) (accuracy of ±5 K) and Setra C206 (Setra Systems, Boxborough, MA, USA) gauge pressure transducer (reference value at rest 1 atm and accuracy of 0.7 × 10⁴ Pa) were set up in the centerline of the catalytic chamber for temperature and pressure measurements. Finally, the engine was mounted on the test bench with a mechanical support connected to a Tedea Huntleigh load cell (accuracy of ±0.05 N) to measure the thrust.

In the second part of the work, it was demonstrated that this system is useful for allowing the ignition and mass flow supply of a 10 N class hybrid engine (Figure 4). The engine was assembled in the following way: downstream of the simple convergent nozzle, the combustion chamber was mounted, equipped with a prechamber which housed the solid grain and the after chamber with a convergent/divergent graphite nozzle with an area ratio of 1.49 to avoid overexpansion and flow separation that would lead to a collapse of thrust coefficient [27].

3. Materials and Methods

3.1. Propellants and Catalysts

In this work, hydrogen peroxide concentrated at 87.5% by mass was used as the main propellant. It was obtained and stored at 60% by mass, then concentrated through a distillation process before the tests were conducted. The evaluation of the hydrogen peroxide concentration was performed using an online tool provided by Evonik Company (Essen, Germany) [29]. To perform this calculation, it was essential to ascertain both the density and temperature of the propellant. For this purpose, the mass was determined using a high-precision balance with an accuracy of ±10⁻⁴ g. The volume was measured using a graduated beaker, with a measurement error of less than 0.25 mL. Lastly, the temperature was assessed using a thermocouple, which was immersed in the propellant at various locations. Given that the temperature of the propellant was uniform, the measurement error was limited to the inherent precision of the instrument, which was ±5 K.

In the case of applications with the engine in a hybrid configuration, three different types of fuels were used, as indicated below:

(1)

Polyvinyl Chloride.

(2)

High-Density Polyethylene.

(3)

Acrylonitrile Butadiene Styrene.

For the decomposition of hydrogen peroxide, catalysts in the form of pellets with cylindrical geometry were used. They are commercially available from the Alfa Aesar Company (Ward Hill, MA, USA). In this specific case, these pellets are characterized by an alumina support, coated with a quantity of palladium corresponding to 5% of the mass of the pellet. These catalysts have been used in previous studies [26]. It is important to emphasize that the type of catalyst can have a significant effect on the performance of the catalytic chamber. In this study, only one type of catalyst was utilized to ensure the independence of the study from any potential related effects. The specific choice of palladium pellets was based on their availability and ease of procurement compared to other types of catalysts. As previously mentioned, they have been used in past studies; this allowed for the elimination of uncertainties that could arise from the authors’ use of a “novel” catalyst, whose behavior is not well understood.

3.2. Measurement Data Reduction Techniques

Several tests were performed and data were recorded to allow the post-processing and to analyze the results.

Each experiment included a sequence of impulses which allowed the injection of a small mass (order of grams) of hydrogen peroxide through the catalytic chamber to heat the system without the use of a glow plug.

Each test was conducted by setting the feeding/injection pressure and the duration of the impulse, which determined the range of time for the valve to open and close, allowing the passage of hydrogen peroxide through the line. The propellant input valve had an open/close lag of 0.005 s, according to the datasheet. To account for this, a lag of 0.01 s (including both open and close times) was added to the impulse duration. In this way, a possible overestimation of performance was avoided.

The data processing was based on elaborations of the mass flow rate, pressure, and thrust signals. For each, the moving average was calculated to obtain a smoother signal, as reported in Figure 5. The red line represents the average signal, while the black line is the high frequency signal.

Regarding pressure and thrust signals, the threshold for selecting the start and end time of the signal was set when pressure and thrust reached 5% of the maximum peak observed during the impulse.

Physical Parameters Computation

Because of the inherently unsteady nature of the pulses, the main performance parameters were computed integrating on time the signals for each pulse, with a time interval based on the criteria expressed in the previous paragraph.

To define the monopropellant performance, characteristic velocity, c*, and $\mathsf{\Delta }\mathrm{V}$ are estimated with the following equations:

$$c_i^{*}={\displaystyle \frac{\int p_{i}dt{A}_t}{\int {\dot{m}}_{i}dt}}$$

(1)

$$\mathsf{\Delta }v={\displaystyle \frac{\int F_{i}dt}{m_{tot}}}$$

(2)

where m_tot is the overall satellite mass. In the present work, the analyzed prototype thruster was considered for a low earth orbit mission of a 12 U CubeSat, thus having an overall mass of 16 kg. This choice was related to the projects in which the University of Naples is involved. “i” refers to the signal acquired at the i-th pulse.

The impulse was computed in two different ways:

-

Integrating the thrust signal over time;

-

Integrating the pressure signal over time.

The second method is also used for in-vacuum estimation of the impulse:

$$I_{bit}=\int p_{i}dt{A}_{t}cf$$

(3)

where c_f is the thrust coefficient computed by using the CEA code [30] and considering an expansion ratio of 20. The resulting value is c_f = 1.66. For the on-ground estimation $c_f=0.7$ is used, based on previous experimental campaigns [31].

For a clear representation of the obtained results, impulses are organized in sequences, thus, the average values, $x_{avg}$, of the performance parameters and their corresponding standard deviations, σ, were evaluated for each sequence of impulses:

$$c_{avg}^{*}={\displaystyle \frac{\sum _{i=1}^N{c}_i^{*}}{N}}={\displaystyle \frac{\sum _{i=1}^N\frac{\int p_{i}dt{A}_t}{\int {\dot{m}}_{i}dt}}{N}}$$

(4)

$$m_{avg}={\displaystyle \frac{m_{seq}}{N}}={\displaystyle \frac{\sum _{i=1}^N{m}_i}{N}}={\displaystyle \frac{\sum _{i=1}^N\int {\dot{m}}_{i}dt}{N}}$$

(5)

The standard deviation of these values was calculated in the following way [32]:

$$\sigma =\sqrt{{\displaystyle \frac{1}{N}}\sum _{i=1}^N{(x_i-x_{avg})}^2}$$

(6)

The errors related to the calculation of the consumed mass were determined via a conservative procedure. This involved considering the instrument error as independent of the read flow rate and fixed at 0.3 g/s, which was the maximum error specified in the catalogue. The mass error was calculated by integrating the flow rate error over the total time the propellant valve was open. This calculation assumed that the error was entirely positive or negative, aiming to provide an estimate of the maximum potential error derived from the consumed mass measurements.

Finally, the average flow rate over the entire sequence was calculated as follows:

$${\dot{m}}_{seq}={\displaystyle \frac{m_{seq}}{t_{valveON}}}={\displaystyle \frac{\sum _{i=1}^N{m}_i}{t_{valveON}}}$$

(7)

where $t_{valveON}$ is the total time that the valve is open during the sequence.

4. Results and Discussion

4.1. Test Cases Presentation

To make the discussion clearer, the tests are divided according to the engine configuration used: monopropellant or hybrid.

4.1.1. Monopropellant Test Campaign

Several pulse sequences were conducted to evaluate the performance of a monopropellant engine operating in pulsed mode. The most relevant sequences, which are discussed in this work, are presented in

Table 1. Tests of the experimental campaign.

Sequence Name	Pulse Duration	Δtoff	Fires Number	Injection Pressure (bar)	Hot/Cold Start-Up	Glow Plug
M1	0.3	2	10	4.10	Hot	No
M2	0.1	2	4	3.90	Cold	No
M3	0.3	2	20	3.90	Hot	No
M4	0.5	3	10	3.90	Hot	No
M5	0.5	3	10	3.55	Hot	Yes
M6	1.0	3	10	3.55	Hot	Yes
M7	0.5	3	10	4.00	Hot	Yes
M8	1.0	3	10	4.00	Hot	Yes
M9	1.0	2	10	4.00	Hot	Yes
M10	0.5	3	12	4.30	Hot	Yes
M11	0.5	3	8	4.30	Hot	Yes
M12	0.5	3	8	4.30	Hot	Yes
M13	0.1	3	8	14.60	Cold	No
M14	0.2	3	15	14.60	Hot	No
M15	0.3	3	10	14.60	Hot	No
M16	0.5	3	10	14.60	Hot	No
M17	0.1	3	15	14.30	Cold	No
M18	0.2	3	15	14.30	Hot	No
M19	0.3	3	10	14.30	Hot	No
M20	0.5	3	10	14.30	Hot	No

. Each test sequence was performed with a different feeding pressure and considering cold/hot start with or without the use of a glow plug. As indicated, the second column reports the single impulse duration, the third the time interval between two impulses, and the fourth the number of impulses for the given sequence.

The performance parameters were computed for each test sequence and for each single pulse by using a MATLAB R2022b code. In

Table 2. Sequence M20.

Start MFR [s]	End MFR [s]	fire n.	MFR [g/s]	Mass [g]	c* [m/s]	ΔV [cm/s]	Δpinj [bar]
118.06	118.57	1	5.72	2.92	731.10	7.25	13.30
121.56	122.07	2	5.38	2.74	844.76	8.21	13.30
125.05	125.56	3	5.52	2.82	805.96	7.68	13.30
128.55	129.06	4	5.58	2.85	849.22	8.08	13.30
132.06	132.57	5	5.74	2.93	796.42	7.90	13.30
135.55	136.06	6	5.54	2.83	888.83	8.57	13.30
139.05	139.56	7	5.55	2.83	824.71	8.71	13.30
142.53	143.04	8	5.52	2.82	821.73	8.10	13.30
146.04	146.55	9	5.67	2.89	843.67	8.01	13.30
149.54	150.05	10	5.67	2.89	843.16	8.29	13.30

, an example of sequence (composed of 10 pulses) is shown, in which MFR is the integral of mass flow rate signal divided by impulse duration and Δp_inj is the pressure drop across the injector. The results present a good repeatability and a good performance in terms of c* and ΔV.

Example signals of mass flow rate, pressure, and thrust are shown in Figure 6 and Figure 7. The green line of Figure 6 represents the mass flow rate of hydrogen peroxide, while the blue line is the pressure signal. It is possible to see that the pressure rose with a little time delay with respect to the mass flow rate signal. The physical meaning is that the decomposition started with a slight lag with respect to the liquid injection. Even if difficult to see in the figures, it was also observed that the thrust signal started to be significant with a slight lag with respect to the pressure signal; this delay was in the order of one hundredth of a second.

4.1.2. Hybrid Test Campaign

The monopropellant tests helped to understand the best sequence capable of efficiently heating the catalytic system. Building on the knowledge acquired about the catalytic system, tests and sequences conducted in a hybrid configuration aimed to determine the amount of propellant consumed before igniting the flame. Therefore, the focus of the discussion is related to the mass required to ignite the engine. Various tests were conducted using different propellants, and

Table 3. Hybrid tests.

Test Name	Ignition Procedure	Type of Regulation	Mass Flow Rate/Upstream Pressure	Glow Plug	Fuel
H1	Continuous Ignition	MFR ***	3 g/s	Yes	PVC
H2	Continuous Ignition	MFR	3 g/s	Yes	PVC
H3	Fixed Sequence *	PR ****	4.3 bar	Yes	PVC
H4	Fixed Sequence	PR	4.3 bar	Yes	HDPE
H5	Progressive Sequence **	PR	10.9 bar	No	PVC
H6	Progressive Sequence	PR	11 bar	No	PVC
H7	Progressive Sequence	PR	12 bar	No	ABS
H8	Progressive Sequence	PR	12 bar	No	HDPE
H9	Progressive Sequence	PR	16.7 bar	No	HDPE
H10 §	Progressive Sequence	PR	19.4 bar	No	HDPE
H11	Progressive Sequence	PR	23 bar	No	HDPE

* Fixed Sequence: 0.5-3-10; ** Progressive sequence: succession of 4 sequences 0.1-3-15 -> 0.2-3-15 -> 0.3-3-10 -> 0.5-3-10; *** MFR: mass flow regulated; **** PR: pressure regulated; ^§ test performed with a hybrid nozzle throat diameter of 2 mm, instead of 3 mm.

summarizes the operating conditions for each test.

It is noted that initially, two tests (H1 and H2) were conducted with hydrogen peroxide continuously injected after heating the catalytic compartment with a 60 W glow plug for more than half an hour. Subsequently, even after using the glow plug, two tests (H3 and H4) were carried out with a pulse procedure, established following the analyses conducted on the monopropellant engine. Finally, several tests (H5 to H11) were carried out using a preheating procedure based on the experience gained from tests conducted on the monopropellant. This method eliminated the need for a glow plug, thereby eliminating the power consumption associated with it.

In the next paragraphs, the main obtained results will be discussed as follows:

-

Effect of injected mass, heating, and injection pressure on the monopropellant performance (monopropellant).

-

Analysis and in-space estimation of the ΔV available for a low earth orbit satellite with an overall mass of 16 kg (monopropellant).

-

Progressive sequence analysis for hybrid ignition (monopropellant).

-

Analysis of propellant consumption during the ignition phase (hybrid).

4.2. Effect of Injected Mass, Heating, and Injection Pressure on the Pulsed Monopropellant Performance

With the aim of igniting a hybrid rocket minimizing mass and power consumption, the performance of the pulsed monopropellant engine was first analyzed via dedicated post-processing. In particular, the study was based on the analysis of the characteristic velocity, which represented the best tool that allowed for a quantitative estimation of the decomposition efficiency. In fact, it was indicative of the amount of heat released and therefore of the “quality” of the gas that was injected into the combustion chamber.

Firstly, it is noteworthy that the characteristic velocity decreased with an increasing value of the injected mass per pulse. The points in Figure 8 are related to sequences performed at a fixed feeding pressure in the range of 3.5 to 4.5 bar, to remove possible influence of the injection pressure on the performance. One possible reason for this phenomenon could be that the increase in injected mass caused a build-up of hydrogen peroxide in the liquid phase within the chamber, which did not immediately decompose and absorbed some of the energy released during decomposition [33]. This led to a decrease in decomposition efficiency.

Another explanation comes from considering the heat exchange that occurred within the thruster. The increase in injected mass reduced the net energy obtained from the decomposition reaction because it brought an increase in the energy lost through convective heat exchange within the thruster. In any case, this explanation requires further investigations and calculations on the heat exchange are required to assess the reason for this behavior.

In the end, another possible reason, already demonstrated, could be that increasing the injected mass resulted in insufficient active sites available on the catalyst used in these tests, which hindered complete decomposition [34].

All the sequences shown had a hot start with a chamber temperature above 100 °C, except for Sequence M2. As already demonstrated in previous works, the preheating is useful for starting the decomposition. Once the latter started, the preheating no longer had a decisive effect, also because the power released by the decomposition was about two orders of magnitude (about 9000 W) higher than that released to heat the thruster, which was around tens of Watts [26]. Therefore, there was a temperature limit beyond which decomposition began and performance was no longer affected by higher temperature values. However, sequence M2 seemed to be not affected by this phenomenon because the quantity of injected mass was very small, and the catalytic system decomposed it completely in a very short time.

Regarding the effect of preheating on pulsed operation, it is possible to draw several conclusions from the same graph for groups of pulses with equal injected mass, conducted under identical operating conditions except for the initial temperature of the pulses. The results indicated that the starting temperature does not significantly affect the performance of pulsed operation. Specifically, the following observations can be made:

-

Comparing the pulses of group M4, conducted without the use of a glow plug, with the pulses of groups M5, M7, M10, M11, and M12, which involved the use of a glow plug, they exhibit practically the same performance. This suggests that the preheating methodology had no impact on performance.

-

The same groups of tests mentioned above were carried out with starting temperatures ranging from 140 °C to 320 °C, but the performance did not appear to be affected by this parameter. It should be noted that all tests were performed with a starting temperature at least equal to the propellant vaporization temperature, so part of the heat released by the decomposition must not be “used” for vaporization.

For completeness, the average performance of the various pulse sequences and their corresponding starting temperature values are reported in

Table 4. Average performance and starting temperatures.

Sequence Name	Tstart [°C]	${\mathit{c}}_{\mathit{a}\mathit{v}\mathit{g}}^{\mathit{*}}$ [m/s]
M4	140 ± 5	590.83 ± 23.92
M5	150 ± 5	581.14 ± 39.00
M7	160 ± 5	597.88 ± 47.80
M10	148 ± 5	569.35 ± 50.59
M11	321 ± 5	604.65 ± 36.25
M12	206 ± 5	557.86 ± 39.70

.

As for the effect of injection pressure, no particular results were reported. From the numerous tests carried out, no defined trend emerged in the performance as the injection pressure varied. In a previous study [26], it was demonstrated that in continuous operation, an increase in flow rate and chamber pressure favors the decomposition process. In this case, given the very short injection intervals, it is possible to affirm that the decomposition occurs almost at ambient pressure. Therefore, the increase in injection pressure certainly impacts the propellant mass flow rate since the pressure difference at the injector will be linear with the feeding pressure. The increase in mass flow rate involves two contrasting phenomena: on the one hand, the increase in liquid jet velocity leads to an increase in kinetic energy, which is dependent on the square of the velocity. The impact of liquid jet on the catalyst active phase could be more forceful, resulting in better breakage and thus higher decomposition efficiency. On the other hand, the same amount of liquid mass is injected in a shorter time. This could be critical because the catalytic system may not be able to instantly decompose the amount of liquid injected. Since the phenomenon is strongly unsteady, it could occur that a quantity of heat developed by the decomposition is absorbed by the liquid phase, shifting the thermal equilibrium to lower temperature values, similar to what happens in the case of large quantities of injected mass. One of these phenomena can prevail over the other depending on the volumes involved and the quantity of injected mass. Probably for this reason, even with the attempt to obtain trend curves for fixed values of injected mass, a clear result was not reported.

4.3. ΔV Analysis and Theoretical In-Space Estimation

After analyzing the performance in terms of decomposition efficiency, the trend of ΔV and how injected mass affects it was investigated.

The results have been computed for all of the tests, and Figure 9 shows an example. The ΔV showed a linear trend with respect to the injected mass. This was an expected phenomenon, and the reason could be trivially found in the fact that increasing the amount of decomposed hydrogen peroxide in the chamber improved performance.

Furthermore, it should be noted that in a vacuum, performance is further enhanced. The mean value of ΔV was computed for each sequence, both for on-ground tests and the vacuum estimations. All the results are coherent with those in Figure 9, thus, good repeatability and propulsive performance for space applications were demonstrated.

4.4. Hybrid Ignitions via Pulsed Preheating

After evaluating the performance of the pulsed monopropellant engine, the main objective of the pulsed sequences was to heat the catalytic compartment starting from low temperatures. By implementing this method, it is possible to enhance the efficiency of decomposition, resulting in the production of a high-temperature gas that could ignite a hybrid system. Essentially, this system serves as an ignition mechanism for a thruster.

The hybrid engine used for this application was a 10 N scale thruster, operating at significantly lower oxidizer mass fluxes compared to hybrid rockets in the literature, which typically have thrusts of 100 N and above [35,36]. This difference could pose a challenge for solid fuel ignition, as the initiation of pyrolysis and subsequent ignition are likely influenced by the velocity of the hot gas on the fuel surface.

For this reason, the study progressed gradually, and the ignitions were carried out following various procedures as indicated in paragraph Section 4.1.2. Therefore, the results in this paragraph are reported in the following order:

-

Ignition was achieved by using a glow plug.

-

Ignition was achieved by using a glow plug, combined with a short series of pulses.

-

Ignition by preheating exclusively with pulses.

4.4.1. Ignitions with the Support of a Glow Plug

In this paragraph are the results obtained from the hybrid rocket ignitions using a glow plug. In particular, tests H1 and H2 were performed using only the glow plug, while in tests H3 and H4 a preheating was performed by combining the glow plug power with a sequence of pulses.

Decomposition was studied by measuring the pressure in the chamber and taking temperature measurements with a thermocouple on the centerline along the axial direction of the decomposition chamber and in a variable position from the axis to the periphery, considering the radial direction. It is important to say that, during the operational phase, the temperature measurement may not be entirely indicative of the decomposition efficiency as it is influenced by the thermocouple’s placement. In fact, the figures report different levels of temperature depending on the thermocouple position. In any case, the temperature measurements are useful for qualitative comments on the experiments. Another thermocouple was positioned downstream of the nozzle to have knowledge of the gas temperature that would be injected onto the fuel.

Returning to the ignition problem, since the continuous injection of hydrogen peroxide did not allow the catalytic system to function properly, it was necessary to preheat the environment with a glow plug, as demonstrated in a previous study on the same monopropellant [26]. Consequently, the initial tests in a hybrid configuration involved preheating the catalytic chamber with a glow plug for over half an hour, consuming 60 W of power. Subsequently, hydrogen peroxide was continuously injected, and fuel ignition was awaited. In this scenario, both power and some propellant were consumed; this is elaborated on shortly. The delay in ignition likely stemmed from the catalytic chamber’s inability to immediately process all the incoming hydrogen peroxide. To address this issue, a brief series of pulses was considered to guarantee an instantaneous decomposition and to initiate fuel heating before transitioning to continuous mode.

For this purpose, several attempts were made on the monopropellant configuration of the engine to verify the validity of the applied methodology. An example is shown in Figure 10 where the pressure in the catalytic chamber, the thrust, the temperature at the center of the chamber in the axial direction and at the periphery in the radial direction, and the temperature of the gas at the exit of the convergent nozzle are reported. After preheating with a glow plug up to a temperature of 433 K, 10 pulses of 0.5 s and 20 of 1 s were performed. The last train of pulses of 1 s was performed by spacing the pulses by 2 s instead of 3. It is observed that in this case the catalytic chamber was not able to empty completely (the pressure did not return to 1 bar) and a further quantity of propellant was already injected. However, the engine performances remain the same, and there did not seem to be a determining effect of the pause interval in this case. At the end of the pulse sequences, hydrogen peroxide was injected continuously to verify whether the catalytic system was able to “manage” the amount of propellant entering. The pressure signal was very stable and demonstrates that this methodology can work.

After a series of attempts with the engine in a hybrid configuration, it was noted that just 10 pulses of 0.5 s were enough to ignite the flame after heating with a glow plug up to 433 K. This revised approach proved successful and helped to reduce the quantity of hydrogen peroxide needed for ignition. Figure 11 illustrates the ignition sequence in Test H3, indicating that the flame ignited at the tenth pulse. The red curve represents the temperature acquired on the axis of the catalytic chamber.

Table 5. Mass consumption at ignition.

Test Name	Glow Plug Power [W]	Tstart [°C]	Mass Consumption [g]
H1	60	250 ± 5	40.45 ± 3.90
H2	60	250 ± 5	53.00 ± 4.50
H3	60	160 ± 5	26.60 ± 1.35
H4	60	320 ± 5	48.75 ± 4.47

shows the consumption of hydrogen peroxide before reaching the flame ignition in tests H1 to H4, whose operating conditions are reported in Table 3; it is evident that the mass consumption in the Test H3 was lower than that in the first two tests. In addition, to allow ignition, it was sufficient to reach a preheating temperature about 100 °C lower than in the first two cases. This implies that the preheating time was also reduced.

As for Test H4, since it involved a different fuel, it was evident that HDPE was more challenging to ignite than PVC. In fact, the amount of mass required for ignition was approximately double. This contrast is also evident in the following paragraph, where the same fuels are compared using the progressive sequence procedure.

4.4.2. Ignitions via Pulsed Preheating

It was observed from the monopropellant tests that injecting a large amount of hydrogen peroxide at low temperatures prevented the complete decomposition of the entire mass. Therefore, there was a need to preheat the system to raise its temperature and initiate the decomposition process, which can then continue on its own.

Therefore, the plan was to inject a small amount of mass in pulses. As the decomposition process began, the intention was to gradually increase the amount of mass to be injected by extending the duration of the valve opening range. For these reasons, several pulsed sequences were performed with the objective of finding an optimum procedure to guarantee the ignition of a hybrid rocket with the minimum possible mass consumption.

For these experiments, the choices of the opening and closing intervals of the valve were based on the observation of temperature drop between pulses, as can be seen in Figure 12. To prevent heat dissipation to the external environment when the valve was closed, it was necessary to increase the amount of injected mass and simultaneously reduce the time interval between pulses as the sequence progressed. The realized effect was that increasing the opening/closing ratio helped the temperature to rise (yellow line).

It can be also observed that the temperature in the catalytic chamber initially increased, but as the pulses continued, it remained constant. The problem was thought to be related to the decomposition phenomenon, thus, the concept of the so called “decomposition point” was introduced, as reported by [37]. Increasing the injected mass caused the motion of the decomposition point downstream, resulting in a thermal wave passing through the center of the chamber and moving downstream. As a result, the hottest point, at a higher temperature, reached upstream of the nozzle (nozzle inlet). This phenomenon can be explained by the trend of red and yellow curves in Figure 12. On the other hand, in cases where a glow plug was used for preheating, both temperatures increased rapidly, and the exit temperature quickly surpassed the one reached in the middle of the catalytic chamber. This is consistent with the earlier discussion, as the decomposition point had already shifted downstream due to the preheating of the engine.

Furthermore, Figure 12 shows that for a certain number of pulses, the pressure never managed to return to atmospheric pressure. This indicates that the valve opening/closing interval ratio was too high, and the growth of this factor should have occurred more gradually. The optimal ratio, in general, can depend on many factors. First and foremost, the feed pressure combined with the valve opening interval determined the amount of injected mass. Additionally, numerous factors complicate the modeling of the heat exchange phenomenon, including the type of catalyst (both the active component and the support), and the dimensions of the chamber, which affect heat transfer with the surrounding environment. The shape of the catalyst and the way the pellets are packed in the catalytic chamber can also influence conduction phenomena within the engine. In this study, these latter factors are clearly considered fixed, and the most determining parameter was the feed pressure. Thus, following the reasoning previously explained and essentially based on an experimental approach, the progressive sequence defined for the ignition of a hybrid was as follows (

Table 6. Progressive sequence for hybrid ignition.

Step	Pulse Interval [s]	Closing Interval [s]
1	0.1	3
2	0.2	3
3	0.3	3
4	0.5	3

):

After testing this sequence on the monopropellant, which is not reported as it adds nothing to the discussion, the same procedure was applied to the hybrid configuration and has proven to be suitable within the operational range of this engine class. Figure 13 illustrates the first ignition of the hybrid engine without the use of a glow plug.

In this test, there was no need to run the entire sequence since the flame ignition occurred at the fourth 0.5 s pulse. After that, the authors switched to continuous mode to complete the test.

For the purposes of example, the masses consumed during the sequence and the average performance of each step are reported in

Table 7. Mass consumption and performance during progressive sequence.

Test-Step	Mass Consumed [g]	${\mathit{c}}_{\mathit{a}\mathit{v}\mathit{g}}^{\mathit{*}}$ [m/s]
Monoprop–1	9.88 ± 0.45	//
Monoprop–2	19.76 ± 0.90	652.36 ± 49.11
Monoprop–3	18.82 ± 0.90	707.98 ± 76.42
Monoprop–4	28.54 ± 1.5	824.96 ± 39.67
H5–1	7.85 ± 0.45	//
H5–2	15.7 ± 0.90 s	605.92 ± 48.45
H5–3	14.79 ± 0.90	708.97 ± 34.66
H5–4	4.73 ± 0.45	782.13 ± 46.12

, comparing them with those obtained from the monopropellant configuration under similar operating conditions, to demonstrate the repeatability of the experiments. The test called “Monoprop” refers to the progressive sequence that was used for the M17 to M20 tests.

As can be seen, the results in terms of c* were very similar, and the experimental analysis starting from the monopropellant configuration was successful. The small difference in mass consumed at each step was simply related to the difference in upstream pressure, as the monopropellant test was conducted with slightly higher pressure.

It should be also noted that in the initial steps, it was not possible to calculate the characteristic velocity values because the pressure levels were very low, and the signal was comparable to the sensor measurement error. In other words, there was no actual increase in pressure, as can also be observed from the graphs.

Since then, many tests have been carried out in hybrid configuration.

Table 8. Mass consumptions in hybrid tests without the use of a glow plug.

Test Name	Mass Consumption [g]	Mox,Hybrid [g/s]	Combustion Chamber Pressure [bar]
H5	43.07 ± 2.70	2.65 ± 0.32	7.32 ± 0.07
H6	37.02 ± 3.60	1.99 ± 0.30	6.17 ± 0.07
H7	37.81 ± 2.85	2.58 ± 0.30	6.36 ± 0.07
H8	73.33 ± 4.35	2.48 ± 0.30	6.50 ± 0.07
H9	56.02 ± 1.35	4.25 ± 0.30	10.83 ± 0.07
H10	67.55 ± 2.40	2.71 ± 0.30	15.91 ± 0.07
H11	54.73 ± 1.23	5.56 ± 0.30	14.56 ± 0.07

is a summary containing the mass consumption, the oxidizer mass flow rate, and the combustion chamber pressure for all the tests that were conducted.

Comparing the tests for fixed fuel, without a doubt the combination of glow plug and pulses was the best solution in terms of mass consumption. In fact, tests H3 and H4 had the lowest consumption when compared to the other tests with the same fuel. However, tests H5 and H6 showed a consumption comparable to tests H1 and H2, in which preheating was performed only by a glow plug. It is therefore evident that it is advantageous to remove power consumption at the expense of a small amount of hydrogen peroxide, especially when considering the preheating times. In the case of the glow plug, preheating times were significantly longer. Pulsed preheating, for instance, took a maximum of about 3 min, while the glow plug required approximately 45 min.

Looking at Tests H6 to H8, they were performed under approximately the same operating conditions, except for the fuel used. The results indicate that PVC and ABS were easier to ignite than HDPE, as HDPE required approximately double the mass to ignite. This finding aligns with the observations made in the previous paragraph when discussing tests H3 and H4.

HDPE was tested in several operating conditions, as reported in Table 3, allowing for the evaluation of the potential impact of mass flow rate and combustion chamber pressure on ignition times. To give a better visualization of the analysis conducted, Figure 14 shows the quantity of mass consumed with the relative pressure values in the combustion chamber for the tests carried out by fixing HDPE as the fuel. Tests H8 and H10 were conducted with different throats and different supply pressures to achieve the same oxidizer flow rate in the combustion chamber. It is evident that, at an equivalent mass flow rate, the chamber pressure positively influenced ignition, albeit slightly. Conversely, tests H9 and H11 were carried out with a higher mass flow rate compared to the previous cases. In this scenario, the flow rate appeared to have a more significant impact on the ignition phase than pressure. This observation aligns with the theory suggesting that regression in hybrid rockets is largely unaffected by pressure but rather influenced by the oxidizer’s mass flux. This notion is further supported by comparing tests H10 and H11, which were conducted at different flow rates but similar combustion chamber pressures, indicating that the higher flow rate was more advantageous.

The ignition sequence for test H11, which had the shortest ignition time compared to all the other tests, is shown in Figure 15, although the fuel consumption was not the lowest.

In the end, Figure 16 reports a photo of the plume during a firing test.

5. Conclusions

In this work, the issue of igniting hybrid rockets operating at low mass fluxes was tackled. One potential method to aid in the ignition of such engines involves preheating the engine through impulsive injections. To this end, the monopropellant engine serving as an injection system for the hybrid engine was first examined, and numerous experiments were conducted, yielding the following results:

• The injected mass on the performance of a monopropellant pulsed thruster was a very impactful parameter. The characteristic velocity trend was analyzed by varying the mass while fixing the feeding pressure range. It was observed that when the mass increased, the hydrogen peroxide in liquid phase accumulated within the chamber without undergoing decomposition.
• The starting temperature was essential to allow the decomposition process to start but, after a given threshold, it did not have an important effect on the overall performance.
• The feeding pressure did not show a clear trend in performance.
• The propulsive performance of the thruster was analyzed by computing the ΔV. An increasing linear trend with respect to the injected mass was observed, since increasing the amount of decomposed propellant in the chamber improved performance. Moreover, the vacuum performance was computed and as expected, the ΔV values were higher than the on-ground ones. Results show that the system is reliable and has good potentialities for space propulsive applications.
• These analyses were used to define an optimal pulsed ignition sequence for a hybrid engine that utilizes this monopropellant engine as an oxidizer injection system. In this context, several experiments were conducted, resulting in the following outcomes:
• Ignition via glow plug preheating and continuous injection involved a power consumption of approximately 60 W for around 45 min with a considerable mass consumption.
• A pulsed injection of 10 shots followed by a continuous one allowed the engine to start with half the propellant consumption and with preheating that took approximately half the time compared to the previous case.
• Under the same conditions, HDPE was more difficult to ignite than PVC and ABS. Specifically, it required a propellant consumption approximately double that required by the latter, whether ignited by a glow plug or exclusively pulsed preheating.
• The pressure in the chamber seemed to have a positive effect on the ignition times, but the mass flow appeared to have a greater beneficial impact, as predicted by theory.
• By removing the consumption of electrical power and preheating only by pulses, it was possible to start the engine with a smaller mass than with continuous injection and in much shorter times than in the cases analyzed with the glow plug (about 3 min compared to 45).

In conclusion, the study focused on addressing the ignition issue in hybrid engines with a 10 N thrust capacity using catalysts composed of a palladium-coated alumina support. It proposed a solution to enhance ignition, which proved to be a favorable balance between electric power and propellant consumption.