Investigation of Performance Stability of a Nytrox Hybrid Rocket Propulsion System

Abstract

Nitrous oxide is a highly suitable oxidizer for hybrid rockets due to its self-pressurizing properties, moderate cost, and high accessibility. However, its vapor pressure and density are highly dependent on ambient temperature, requiring careful consideration of temperature variations in real applications. To mitigate this issue, an oxidizer called Nytrox was produced by adding a small fraction of oxygen to bulk nitrous oxide. This modification enables the hybrid rocket propulsion system to maintain a nearly constant average thrust and total impulse across a wide range of ambient temperatures. A series of 7 s hot-fire tests of a small Nytrox/polypropylene hybrid rocket engine operating at ~60 barA of running tank pressure demonstrated a consistent average thrust of 45.3 ± 0.7 kgf and a total impulse of 307.6 ± 3.9 kgf·s within a N₂O temperature range of 5.9–22.6 °C, compared to highly varying values of the N₂O/polypropylene one within a N₂O temperature range of 10.8–29.8 °C. Furthermore, the specific impulse of the Nytrox hybrid rocket engine increases mildly with decreasing temperature because of the increasing amount of added oxygen that benefits the combustion for generating the thrust.

1. Introduction

The hybrid rocket propulsion system combines features of solid and liquid propulsion systems. Figure 1 shows the conceptual design of a hybrid rocket engine (hereafter HRE) in which the liquid or gaseous oxidizer flows into the combustion chamber containing a solid fuel grain through a control valve and an oxidizer injector for combustion and generates the designed thrust. Recently, due to their throttle ability, low-cost effectiveness, simplicity, reliability, and safety, hybrid rocket propulsion systems have the potential to serve the future rocket propulsion system. Therefore, many commercial companies such as Gilmour in Australia, Innospace in South Korea, HyImpulse in Germany, and ATspace in Australia, to name a few, are currently developing satellite launch vehicles using HREs as their primary propulsion system [1].

Among all the possible oxidizer choices for hybrid rockets, nitrous oxide may become the preferred choice for relatively small rockets due to its self-pressurizing properties because of its high vapor pressure at normal temperature. A blowdown type with a self-pressurizing system could achieve a high thrust in a simple and cost-effective way because it reduces the additional weight, complexity, and cost of the pressurization system. For example, Virgin Galactic’s SpaceShip One was launched to an altitude of 90 km with a crew of three which was powered by an HRE using N₂O and HTPB [2,3]. The Hybrid Engine Development (HyEnD) hybrid sounding rocket project at the University of Stuttgart has also presented several hybrid rockets using N₂O as the oxidizer [4], and the N₂ORTH sounding rocket reached an altitude of 64.4 km in 2023 [5]. The other important advantages of N₂O include stable combustion and easy access commercially with a modest cost. However, the main disadvantage of N₂O is that the vapor pressure and density highly depend on the ambient temperature (assuming it matches the temperature of N₂O), as shown in

Table 1. N₂O properties with respect to temperature [7].

Temperature (°C)	Vapor Pressure (barA)	Liquid Density (kg/m3)	Vapor Density (kg/m3)
−20	18.01	995.4	46.82
−15	20.83	975.2	54.47
−10	23.97	953.9	63.21
−5	27.44	931.4	73.26
0	31.27	907.4	84.86
5	35.47	881.6	98.41
10	40.07	853.5	114.5
15	45.10	822.2	133.9
20	50.60	786.6	158.1
25	56.60	743.9	190.0
30	63.15	688.0	236.7
35	70.33	589.4	300.4
Tcrit (36.42 °C)	72.51	452.0	452.0

and Figure 2; this leads to an immense impact on the control of the injected mass flow rate and thrust in a blowdown system, hence impairing the stability of the propulsion performance, which is otherwise important for a real application [6].

To enhance the use of N₂O as an oxidizer in hybrid rocket propulsion, Karabeyoglu et al. [8,9] introduced Nytrox—a nitrous oxide/oxygen mixture, which is designed for improved safety and easier preparation. In Nytrox, oxygen serves as the pressurizing agent, while N₂O acts as the densifying component. Compared to its pure components, Nytrox offers a balanced combination of specific impulse performance, density, self-pressurization characteristics, chemical stability, storability, and ease of handling. Additional advantages include the ability to adjust the composition ratio based on mission requirements and enhanced gas-phase combustion efficiency. Moreover, Nytrox provides greater overall safety, as O₂ acts as a diluent in gaseous N₂O, increasing the activation energy required for self-decomposition.

Regarding the properties and practical applications of Nytrox, Whitmore et al. [10,11,12,13] explored the methods and conditions for its preparation and storage and conducted a feasibility analysis on replacing gaseous O₂ with Nytrox in hybrid motors for small spacecraft. Their research established a method for calculating the properties of Nytrox under various temperatures and pressures. In addition, the data (diagram format) of O₂ dissolved mass fraction and density for saturated gas and liquid Nytrox at different pressures and temperatures were provided. Nytrox was produced by simply bubbling gaseous oxygen under high pressure into liquid nitrous oxide until the solution reached saturation. The study showed that while Nytrox has a slightly lower specific impulse and regression rate compared to pure gaseous oxygen, it offers significantly higher volumetric efficiency.

Compared to conventional oxidizers, Nytrox retains the excellent self-pressurization characteristics of nitrous oxide while offering enhanced safety and cost benefits without sacrificing performance. Since the mixture composition is determined by two control variables—temperature and pressure—different Nytrox formulations have been explored in several studies and can be easily fine-tuned for specific applications. However, in the past, the application of Nytrox still necessitated maintaining a fixed oxidizer temperature, as the solubility of oxygen is highly dependent on the temperature of liquid N₂O and the applied pressure.

An interesting concept is that during pressure variations in a running tank of a blowdown system, additional oxygen is released from liquid nitrous oxide due to changes in oxygen solubility. This process generates a self-pressurization effect like N₂O gasification. Since a lower temperature and a higher applied pressure of the nitrous oxide allow for larger oxygen dissolution, more oxygen is released as tank pressure fluctuates, enabling Nytrox to exhibit a similar blowdown pressure change curve across different temperature conditions.

This research aims to utilize Nytrox as an oxidizer to address the primary shortcomings of N₂O-based hybrid rocket propulsion, specifically the sensitivity of thrust and total impulse performance under various ambient temperatures. A series of ground static hot-fire tests with the running tank operating at approximately 60 barA were performed using a swirling-injection small hybrid rocket engine with a single-port fuel grain. The test results of the thrust and total impulse were analyzed under N₂O temperatures ranging from 5.9 °C to 29.8 °C, comparing using both pure N₂O (N₂O temperatures: 10.8 °C to 29.8 °C) and Nytrox (N₂O temperatures: 5.9 °C to 22.6 °C) as oxidizers. The goal is to achieve a stable thrust, a total impulse, and a specific impulse in this temperature range, simplifying future implementation in real applications of propulsion technology.

2. Research Methods

This research is based on a 45 kgf thrust-level hybrid rocket propulsion system with a swirling-injection oxidizer, and designed to utilize Nytrox oxidizer by incorporating a N₂O and O₂ filling system. Additionally, a theoretical model was developed to determine the liquid/gas mass ratio and oxygen solubility in nitrous oxide at known N₂O temperatures and vapor pressures. This model was used to characterize the properties of the Nytrox, which then guided the assessment of the propulsion system’s available burn time and combustion performance. Subsequently, a series of horizontal ground static hot-fire tests were conducted using the propulsion system with different oxidizers (nitrous oxide and Nytrox) under various N₂O temperature conditions. The differences in engine performance between the two oxidizers were then analyzed and compared. The detailed methods are described in the following.

2.1. Combustion Chamber Design and Specification

Figure 3 illustrates the cross-sectional view and overview of the combustion chamber used in this study. The oxidizer is injected into the combustion chamber through a swirling injector. The polypropylene (PP) solid fuel grain features a single-port design for simplicity of manufacturing. It has a cylindrical shape with an outer diameter of 80 mm, an initial inner diameter of 26 mm, and a length of 240 mm, and this size provides sufficient fuel supply for the 8 s burn time. Both ends of the grain have a tapered configuration, and a step is designed upstream of the fuel to create a recirculation zone, ensuring smooth engine ignition by a pyrograin igniter. The nozzle has a simple conical design, which is made of graphite, with a throat diameter of 10 mm and an expansion ratio of 6.25. This hybrid rocket engine is designed to produce a thrust of 45 kgf, a chamber pressure of 40 barA, an average oxidizer/fuel (O/F) ratio of 7 (based on pure N₂O), and a specific impulse of 220 s.

This study utilizes a swirling oxidizer injector to generate a swirling flow field within the combustion chamber, enhancing the mixing of the oxidizer and fuel while improving the diffusion flame characteristics of hybrid combustion. To quantify the intensity of the swirling flow field produced by the oxidizer passing through the injector, Beer et al. [14] defined the geometric swirl number (SNg) based on the injector’s geometric dimensions as follows:

$${SN}_g={\displaystyle \frac{r_o(r_o-r_{\varphi })}{N{r}_{\varphi }^{ 2}}}$$

(1)

where $r_o$ is the injector outlet radius, $r_{\varphi }$ is the injector tangential inlet port radius, and N is the number of tangential inlet ports, as shown in Figure 4.

In this research, the geometric swirl number of the oxidizer injector is designed to be 58.5 to enhance the swirling effect. This value is higher than the typical design due to its liquid state when injected into the chamber. The detailed geometric dimensions of the oxidizer injector are shown in Figure 5, and the geometric swirl number parameters are summarized in

Table 2. The design parameters of the swirling injector.

Parameter	Value	Unit
N	16	-
$r_{\varphi }$	0.5	mm
$r$	15	mm
${\theta }_r$	70.8	°
${\theta }_a$	60	°
$r_o$	18	mm
$r_{in}$	15.9	mm
SNg	58.5	-

. Considering the injection angles of the tangential inlet ports in both the radial and axial directions, the geometric swirl number is further derived from Beer et al. [14] and defined as follows:

$${SN}_g=sin{\theta }_a\times {\displaystyle \frac{(r_o-r_{\varphi })r_o}{N{r}_{\varphi }^{ 2}}}=sin{\theta }_a\times {\displaystyle \frac{{(r}_{in}\times sin{\theta }_r)r_o}{N{r_{\varphi }}^2}}=sin{\theta }_a\times {\displaystyle \frac{r\times r_o}{N{r_{\varphi }}^2}}$$

(2)

where $r_{in}$ is the distance between the center of the tangential inlet port outlet and the axis, r is the distance from the center of the tangential inlet port inlet to the injector axis ($r=r_o-r_{\varphi })$, ${\theta }_a$ is the angle between the direction of the tangential inlet port and the axial direction, and ${\theta }_r$ is the angle between the direction of the tangential inlet port and the radial direction.

2.2. Theoretical Models for the Mixture of Liquid Nitrous Oxide and Oxygen

2.2.1. Determination of the Filling Method of Nitrous Oxide

When nitrous oxide is used as a rocket oxidizer, it is stored in a liquid/vapor coexisting state. Liquid nitrous oxide serves as the primary oxidizer entering the combustion chamber to participate in the combustion reaction, while the gaseous portion is utilized for pressurization. As shown in Figure 2, nitrous oxide has a specific saturated vapor pressure at different temperatures. By measuring the pressure and temperature inside the storage tank and comparing it with the available N₂O properties, the corresponding densities of N₂O in the gas and liquid phases in the storage tank can be determined. Taking 25 °C as an example, the gas and liquid densities can be determined as 190 kg/m³ and 744 kg/m³, respectively, as illustrated in Figure 6.

Given the storage tank volume $\left(V_{tank}\right)$ and the total mass of N₂O filled into the storage tank ${(m}_{N_{2}O,total})$, the filling volume ratio of liquid N₂O $\left(X_{N_{2}O,liquid}\right)$ and its corresponding weight ${(m}_{N_{2}O,liquid},m_{N_{2}O,gas})$ can be calculated as follows:

$$\begin{gathered} m_{N_{2}O,total}={\left(\right(X}_{N_{2}O,liquid}) {\mathsf{\rho }}_{liquid}+(1-X_{N_{2}O,liquid}){\mathsf{\rho }}_{gas})V_{tank}\to X_{N_{2}O,liquid} \\ ={\displaystyle \frac{\left(\frac{m_{N_{2}O,total}}{V_{tank}}\right)-\left({\mathsf{\rho }}_{gas}\right)}{\left({{\mathsf{\rho }}_{liquid}-\mathsf{\rho }}_{gas}\right)}} \end{gathered}$$

(3)

$$m_{N_{2}O,liquid}=X_{N_{2}O,liquid}\left(m_{N_{2}O,total}\right)$$

(4)

$$m_{N_{2}O,gas}=\left({1-X}_{N_{2}O,liquid}\right)\left(m_{N_{2}O,total}\right)$$

(5)

Based on the target thrust $\left(F\right)$, the specific impulse $\left(I_{sp}\right)$, and the O/F ratio, the required oxidizer mass flow rate ${(\dot{m}}_{oxidizer})$ can be estimated as follows:

$$\begin{gathered} I_{sp} \left(s\right)={\displaystyle \frac{F \left(\mathrm{k}\mathrm{g}\mathrm{f}\right)}{\dot{m} (\mathrm{k}\mathrm{g}/\mathrm{s})}}={\displaystyle \frac{F (\mathrm{k}\mathrm{g}\mathrm{f})}{{\dot{m}}_{oxidizer} \left(\frac{O/F ratio+1}{O/F ratio}\right)(\mathrm{k}\mathrm{g}/\mathrm{s})}}\to {\dot{m}}_{oxidizer}(\mathrm{k}\mathrm{g}/\mathrm{s}) \\ ={\displaystyle \frac{F \left(\mathrm{k}\mathrm{g}\mathrm{f}\right)}{I_{SP} \left(\frac{O/F ratio+1}{O/F ratio}\right)(\mathrm{s})}} \end{gathered}$$

(6)

Dividing the available liquid oxidizer mass by the mass flow rate yields the maximal burn time ${(t}_{b,max})$ of the storage tank at that filling mass as follows:

$$t_{b,max}(\mathrm{s})={\displaystyle \frac{m_{N_{2}O, available liquid}\left(\mathrm{k}\mathrm{g}\right)}{{\dot{m}}_{oxidizer} (\mathrm{k}\mathrm{g}/\mathrm{s})}}$$

(7)

Since nitrous oxide will vaporize during the operation of the blowdown process, the amount of usable liquid N₂O ($m_{N_{2}O, available liquid})$ can be estimated by subtracting the mass of the fully vaporized N₂O gas at the same temperature from the total filled N₂O. It is important to note that Equation (7) is a simplified estimation only for convenience. In reality, as N₂O vaporizes, its temperature decreases, leading to a lower gas density, which means that a higher amount of liquid nitrous oxide can be utilized. The error in combustion time depends on the tank volume and the temperature drop. The difference in combustion time can be calculated by multiplying the change in N₂O gas density by the tank volume and dividing by the oxidizer mass flow rate. For a 2.8 L storage tank, a 10 °C temperature change before and after the test will result in an error of approximately 0.52 s (initial N₂O vapor pressure 40 barA) to 0.89 s (initial N₂O vapor pressure 60 barA). Additionally, because the density of N₂O increases significantly with decreasing N₂O temperature, the maximum amount that can be filled into the fixed-volume storage tank at lower temperatures is much higher than at higher temperatures. This also leads to an increase in maximal burn time.

To evaluate the impact of the N₂O filling mass on the experimental parameters at different temperatures, this research employs the calculation method described above as a basis. Under the condition of a fixed storage tank volume of 2.8 L and varying temperature conditions (expressed by N₂O vapor pressure, with 60 barA as the benchmark and intervals of 5 bar downward to 40 barA), three filling methods are analyzed: the total filling mass of N₂O (1.922 kg), the liquid N₂O percentage (95%), and the maximum burn time (7.4 s). The results, as shown in Figure 7, indicate that the method of the total filling mass of N₂O provides a tradeoff solution among the three varying parameters. Compared to the other methods, it has a relatively small impact on the overall change in parameters. Additionally, the fixed filling mass method allows for easier measurement of the storage tank mass as a benchmark, simplifying the operational process and making it more intuitive. Therefore, this research adopts the method of fixed total filling mass of nitrous oxide as the filling approach for the formal tests.

2.2.2. Estimation of the Solubility of O₂ in Liquid/Gaseous N₂O

For the theoretical model of Nytrox, Karabeyoglu [8] employed the Peng–Robinson method to model the properties of Nytrox. Whitmore and Stoddard [10] also applied the same model to establish a method for calculating the properties of Nytrox at various temperatures and pressures. They developed mass fraction and density diagrams for O₂ dissolved in both gaseous and liquid N₂O at different pressures and temperatures. In this study, the O₂ solubility data in the temperature range of 0 °C to 30 °C were extracted and replotted in Figure 8 in which the curves are presented as isotherms in the vapor phase (Figure 8a) and liquid phase (Figure 8b), respectively. For example, by measuring the actual temperature and pressure in the running tank and referring to the corresponding temperature and pressure solubility plot in Figure 8b, the mass fraction of O₂ in the liquid-phase mixture can be readily obtained. A typical procedure for estimating the oxygen mass fraction in the liquid is illustrated in Figure 9. For example, the N₂O temperature is 11.3 °C, and gaseous O₂ is bubbled into the liquid N₂O to reach a pressure of 60.4 barA. Then, the mass fraction of O₂ in the liquid N₂O can be determined as 4.4%. This is obtained by interpolating data between two isotherms (10 °C and 20 °C).

2.3. Experimental Facility and Instrumentation

Figure 10 shows the oxidizer filling and feeding system for the ground static hot-fire tests of the nitrous oxide and Nitrox hybrid rocket engine. It consists of a running tank with an integrated mass load cell, a N₂O and O₂ filling system, a nitrogen cooling pipeline, multiple pressure transducers, many K-type thermocouples for monitoring the running tank and pipeline, a load cell for measuring instantaneous thrust, and an open-loop control valve capable of delivering both low and full flow rates of oxidizer, respectively, for ignition and hot-fire burning. All the test procedures were automated using a DAQ system (Model No. cRIO-9074, National Instruments, Austin, TX, USA).

2.4. Experimental Configuration and Procedure

To establish suitable oxidizer configurations and effectively control the experimental parameters, this study maintained a consistent N₂O filling mass. For the Nytrox tests, the gaseous O₂ was bubbled under high pressure into liquid N₂O to form the Nytrox until a predetermined target pressure of 60 barA was reached. The varying parameters of the hot-fire test included the oxidizer type (N₂O/Nytrox) and the N₂O temperature in the range of 5.9–29.8 °C. It is important to note that the venting process influenced the N₂O temperature during the filling procedure. The N₂O filling mass was fixed as 1.922 kg, corresponding to 95% of the N₂O liquid-phase volume within the fixed tank volume (2.8 L) at a N₂O temperature of 28 °C. This filling mass provided an approximate combustion duration of 7.4 s at a constant temperature and oxidizer mass flow rate. The target tank pressure for the Nytrox tests was set to 60 barA, closely matching the vapor pressure of pure N₂O at 28 °C. This approach served two purposes: first, it facilitated a performance comparison with pure N₂O at a N₂O temperature of 28 °C; second, it simplified implementation by providing a clear benchmark.

Figure 11 illustrates the automated procedural sequence of the ignition voltage for the ignition of the pyrograin and the valve angle for controlling the flow rate of the oxidizer, with the corresponding flame images shown in Figure 12. Upon initiating the automated procedure, the ignition signal was activated for one second, igniting the pyro grain igniter via the electronic ignition system. Subsequently, the control valve was opened to the preheat angle, delivering an approximate oxidizer flow rate of 10 g/s to ignite the solid fuel and preheat the combustion chamber. This allowed the combustion chamber to fully heat up. Following preheating, the control valve was gradually opened to avoid combustion instability, and then fully opened and maintained for four seconds. Finally, the control valve was rapidly closed within approximately 0.2 s, and the nitrogen gas cooling valve was opened, injecting cold nitrogen gas into the combustion chamber to extinguish any residual flames.

3. Results and Discussion

3.1. Analysis of Pure Nitrous Oxide HRE Hot-Fire Test Results

We followed the engine automated procedure in Figure 11 to carry out the subsequent hot-fire tests. Figure 13 illustrates the measured temporal variations in tank pressure, chamber pressure, and thrust in two ground hot-fire tests using pure N₂O at the highest (29.8 °C, vapor pressure 63.1 barA) and the lowest (10.8 °C, vapor pressure 40.9 barA) N₂O temperatures, respectively. This result shows the typical pressure variation characteristics of the N₂O blowdown system. Tank pressure begins to drop continuously during preheating, with a more significant pressure drop after the valve is officially opened. This occurs because, during the N₂O blowdown process, the temperature of N₂O decreases due to liquid vaporization, leading to a continuous decrease in N₂O vapor pressure. Therefore, even if the valve opening angle remains unchanged, the N₂O flow rate tends to decrease gradually during the hot-fire test. After the test, the combustion time and total impulse are calculated for the period when the combustion chamber pressure is above 20 barA. The fuel and oxidizer mass flow rates are obtained by dividing the weight difference between the fuel and oxidizer before and after the test (minus the preheating amount) by the combustion time. Finally, Isp is obtained by dividing the total impulse by the weight difference in the total propellant before and after the test (minus the preheating amount).

Table 3. Comparison of pure N₂O HRE hot-fire tests at different N₂O temperatures.

N2Ovapor pressure (barA)	40.9	45.5	48.5	52.4	58.0	63.1
Tank pressure before test (barA)	40.9	45.5	48.5	52.4	58.0	63.1
Avg. chamber pressure (barA)	26.2	31.7	34.1	35.5	38.3	41.6
Total impulse (kgf·s)	188.1	228.4	249.8	259.5	284.8	314.3
Avg. thrust (kgf)	26.7	33.8	35.9	38.5	43.6	45.0
Avg. Isp at sea level (s)	193.3	205.3	211.7	203.2	215.6	216.5
Liquid N2O temperature (°C)	10.8	14.9	18.1	21.5	26.3	29.8

compares the propulsion performance of these pure N₂O hot-fire tests. The data clearly demonstrate the significant influence of ambient temperature on the performance of pure N₂O hybrid rocket propulsion systems. A decrease in N₂O vapor pressure with decreasing temperature leads to an obvious reduction in the self-pressurized vapor pressure, thus resulting in reduced combustion pressure. It naturally leads to a reduced thrust, a total impulse, and an average sea-level Isp at the same time. These findings underscore the importance of taking into account the variation in ambient temperature in the design and operation of N₂O hybrid rocket systems, necessitating the design of an auxiliary pressurization system or regulating the temperature of N₂O across different ambient conditions to maintain upstream pressure stability without too much engineering work required.

3.2. Analysis of Nytrox HRE Hot-Fire Test Results

Figure 14 illustrates the temporal variations in tank pressure, chamber pressure, and thrust in two ground static hot-fire tests using the Nytrox oxidizer at the highest (22.6 °C, vapor pressure 53.6 barA) and the lowest (5.9 °C, vapor pressure 36.3 barA) N₂O temperatures, respectively. Similarly to the results in Figure 13, this result shows the typical pressure variation characteristics of the blowdown system. The key difference is that similar pressure variations occur even with different initial N₂O temperatures. This is because lower tank N₂O temperatures lead to lower N₂O vapor pressures, which in turn increase O₂ solubility, as shown in Figure 8. Higher oxygen solubility allows for increased oxygen release during the pressure drop, compensating for variations in N₂O vaporization at lower N₂O vapor pressures. It is expected that the primary process before the tank pressure reaches the N₂O vapor pressure will be the release of dissolved oxygen from the Nytrox, resulting in a gradual decrease in the oxygen content of the Nytrox injected into the combustion chamber.

Table 4. Comparison of Nytrox HRE hot-fire tests at different N₂O temperatures.

N2Ovapor pressure (barA)	36.3	41.3	46.0	48.0	53.6
Tank pressure before test (barA)	60.6	60.4	60.0	60.8	60.6
Avg. chamber pressure (barA)	37.3	38.4	37.8	37.1	38.8
Total impulse (kgf·s)	310.7	306.2	305.9	311.7	303.7
Avg. thrust (kgf)	45.3	45.1	45.7	45.6	44.6
Avg. Isp at sea level (s)	229.1	222.6	227.1	216.7	219.2
Liquid N2O temperature (°C)	5.9	11.3	15.9	18.3	22.6
Dissolved O2 mass fraction of Nytrox (%)	2.9	2.7	2.6	2.2	1.2

summarizes the comparison of the propulsion performance of the Nytrox hot-fire tests. The Nytrox test results indicate that the addition of O₂ to N₂O and pressurization to a consistent tank pressure of 60 barA greatly improved the stability and consistency of the hybrid rocket propulsion system. The minimal variation in tank pressure produced a similar engine performance variation. In contrast with the pure N₂O tests, the combustion pressure, the thrust, and the total impulse exhibited greatly reduced sensitivity to the variation in N₂O vapor pressure due to different ambient temperatures. The coefficients of variation for the combustion chamber pressure, the thrust, and the total impulse were calculated as 1.9%, 1.0%, and 1.1%, respectively. In addition, the coefficient of variation for the average sea-level Isp was only 2.3%, indicating similar performance was observed.

Figure 15 and Figure 16 illustrate the average thrust and total impulse data, respectively, using two different oxidizers with varying N₂O vapor pressures due to different N₂O temperatures during the hot-fire tests. The results indicate that the thrust and total impulse of the pure N₂O tests exhibit an almost linear regression with the N₂O vapor pressure (pure N₂O in this case), whereas the Nytrox tests maintained a nearly constant thrust (45.3 ± 0.7 kgf) and total impulse (307.6 ± 3.9 kgf·s) in the range of N₂O vapor pressures (N₂O with a small fraction of O₂ in this case) tested. The trending lines of both oxidizers (N₂O and Nytrox) converge near the target O₂ pressurization pressure (~60 barA), suggesting that deviations in O₂ pressurization pressure during Nytrox preparation would result in the total impulse and thrust values approaching those of the pure N₂O result at the same tank pressure. It should be noted that the initial tank pressure of the Nytrox system was 60.4 ± 0.4 barA, while the initial tank pressure of the pure N₂O system matches its vapor pressure.

Furthermore, Figure 17 compares the average sea-level specific impulse of the pure N₂O and the Nytrox hot-fire tests with varying N₂O vapor pressures. The results indicate a positive correlation for the pure N₂O tests. Conversely, a negative correlation was observed for the specific impulse of the Nytrox tests in the range of N₂O vapor pressures tested. As shown in the dissolved O₂ mass fraction data in Figure 8 and Table 4, this discrepancy arises because lower tank N₂O temperatures result in lower N₂O vapor pressures, which in turn lead to higher O₂ solubility. In this case, as described in the study by Whitmore and Stoddard [10], increased dissolved O₂ in the liquid oxidizer enhances combustion performance. Additionally, the O₂ released from the Nytrox liquid helps maintain tank supply pressure automatically. This allows the combustion chamber pressure to remain at the optimal value for the nozzle design, thereby improving the overall theoretical Isp. These findings confirm that the Nytrox approach effectively improves and stabilizes the performance of the N₂O-based hybrid rocket propulsion system. Moreover, an experimentally observed trend indicates that the lower the N₂O temperature is, the better the Nytrox hybrid rocket engine performance is.

4. Conclusions

In this study, we investigate the effect of ambient temperature on the performance of a N₂O hybrid rocket propulsion system and explore the use of Nytrox as an oxidizer to enhance and stabilize its propulsion performance. To achieve this goal, a propellant feed system capable of preparing and utilizing both pure N₂O and Nytrox was developed, and multiple static hot-fire tests were conducted at various N₂O temperatures using both oxidizers. The major findings are summarized as follows:

1. For a consistent engine burn time, the thrust, total impulse, and specific impulse of a pure N₂O hybrid rocket propulsion system exhibit an almost linear relation with N₂O vapor pressure, which is highly dependent on the N₂O temperature.

2. By bubbling gaseous O₂ under high pressure into liquid N₂O to prepare the Nytrox as an oxidizer in a hybrid rocket propulsion system while maintaining a consistent running tank pressure, a nearly constant average thrust and total impulse can be achieved with various N₂O temperatures.

3. As the temperature of N₂O decreases and the O₂ fill ratio increases, the HRE’s specific impulse increases accordingly.

In summary, under consistent oxidizer running tank pressure conditions, a simple bubbling of gaseous oxygen into the liquid nitrous oxide allows for adjusting the target engine thrust level by modifying the target oxygen pressurization pressure. In addition, this approach enables the HRE to maintain a stable average thrust and total impulse across a wide range of N₂O temperatures (5.9–22.6 °C). These characteristics significantly contribute to system simplification and enhance the applicability of the Nytrox in real applications, e.g., sounding rockets, with different mission requirements.